JP2015168315A - Guide device, and space machine mounting guide device - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a guide device capable of easily calculating motion of a machine body at real time by a small calculator mounted on the machine body, without depending on communication with ground, for reducing guide error in re-entry.SOLUTION: An angle speed generator 2 such as a thruster and reaction wheel is used for applying fine torque to the machine body for acquiring information related to mass characteristic of the machine body before start of guide, and a centroid position of the machine body is identified by a calculator 5. A reference orbit to a target point is created by using the identified centroid position, so that the machine body can be guided and controlled with simple calculation.

Description

  The present invention relates to a guidance device for guiding a spacecraft having a re-entry function to a target point.

  In recent years, research and development of spacecraft having a re-entry function has been actively conducted, and there are not a few situations where the ability to return to the ground after performing missions in orbit is required. At that time, in order to facilitate the recovery of the aircraft, it is necessary to reduce the error (hereinafter referred to as "guidance error") between the target point set on the ground surface or in the air and the arrival point of the spacecraft with respect to the target point Is natural. In addition, in order to realize future high-frequency space transportation, there are issues such as improving the responsiveness at the time of reentry and improving the autonomous guidance control capability of spacecraft that does not rely on communication with ground base stations. .

Depending on the type of mission in orbit, the fuselage mass characteristics at the time of re-entry, such as the fuselage center of gravity, mass distribution, and moment of inertia, differ from mission to mission, resulting in guidance errors due to differences from the standard aircraft mass characteristics. There is a possibility that. For example, in the mission of transporting goods from the orbit to the ground, if the material, shape, location, etc. of the goods stored in the aircraft are different for each mission, a guidance error occurs due to the difference in the materials.
In addition, the mass characteristics of the aircraft suddenly change during the motion of the aircraft, which may lead to induction errors. For example, if the materials in the aircraft are not fixed and the aircraft is tilted, the arrangement of the materials in the aircraft will change, resulting in changes in the mass characteristics of the aircraft, or the consumption of fuel in the aircraft fuel tank. When the fuselage mass characteristics change due to, there is a possibility of causing an induction error.

Conventionally, as a guidance method when a spacecraft re-enters, orbit calculation is performed based on the standard position, speed, fuselage mass characteristics, aerodynamic characteristics, atmospheric conditions, etc. (hereinafter referred to as "motion model") (1) A reference trajectory method, a state quantity integrating in real time, and (2) a prediction guidance method are used.
As described above, when the mass characteristics and atmospheric conditions of the aircraft are different for each mission, the aircraft can be controlled in real time by sequentially integrating the equations of motion by the predictive guidance method, and the guidance error can be reduced. However, this method places a great load on the computer mounted on the machine.
On the other hand, as a recent trend, spacecrafts are becoming smaller and more power-saving. In a small and power-saving spacecraft, it is not possible to apply a great load to the on-board computer such as the sequential integration of the equation of motion by the predictive guidance method.
For this reason, establishment of the technique which performs a highly accurate guidance by simple calculation by a standard orbit system without performing sequential calculation like a prediction guidance system is calculated | required. In order to realize a means for reducing guidance errors based on simple calculations using this reference trajectory method, it is extremely important for the spacecraft to acquire accurate information on the motion model necessary to calculate the reference trajectory. It is.
Conventionally, a technique has been disclosed in which observation data relating to the motion of the aircraft is transmitted to the base station on the ground, the mass characteristics of the aircraft are calculated by the base station computer, and the calculation results are transmitted to the spacecraft ( For example, Patent Document 1).

JP 2000-272599 A I. Bogner. “Description of apollo entry guidance,” Technical Report NASA-CR-110924. Bellcomm, Inc., August 1966. Gerald L. Hunt. “Apollo reentry guidance and navigation equations and flow logic,” Technical Report MSC 65-FM-157, NASA Manned Spacecarft Center, December 1965.

In the conventional technology, observation data related to the movement of the aircraft was transmitted to the base station on the ground, and the results of calculating the mass characteristics of the aircraft were transmitted to the spacecraft by the base station computer. The load and power of the computer can be made relatively small.
However, since it takes time to transmit and receive data between the spacecraft and the base station, it lacks real-time characteristics. For example, even if the information obtained at the start of guidance is accurate, the calculation results from the base station are transmitted. There was a problem that the wind direction, wind speed, and the internal state of the fuselage suddenly changed before it arrived, and when the calculation result from the ground arrived at the spacecraft, it was not already effective.

The present invention has been made to solve such a problem, and it is a case where a sudden wind direction, wind speed, and state change inside the aircraft occur without applying a large calculation load or power to the computer mounted on the aircraft. Another object of the present invention is to provide a guidance device that can accurately guide a spacecraft to a target point.

  A guidance device according to the present invention is a guidance device that guides a spacecraft having a re-entry function to a target point, angular velocity generation means for applying torque to the spacecraft, and angular velocity detection for detecting an angular velocity generated by the torque. A body center of gravity position identification that identifies the center of gravity position of the spacecraft using the means, the acceleration detection means for detecting the acceleration caused by the torque, the angular velocity detected by the angular velocity detection means, and the acceleration detected by the acceleration detection means And calculating the trajectory of the spacecraft to be re-entered by using the airframe center of gravity position identifying means comprising the means and the position of the center of gravity of the airframe output from the position identifying means, and guiding the spacecraft to the destination based on the trajectory A control calculation unit.

  According to the present invention, it is possible to identify the state quantity of the airframe required for the calculation calculation of the reference trajectory at the time of re-entry of the spacecraft without using communication with the ground station. Furthermore, the induction error can be reduced by simple calculation without increasing the calculation load due to sequential calculation. In addition, it is possible to reduce induction errors caused by sudden changes in the motion model.

It is a lineblock diagram of the guidance device concerning Embodiment 1 of the present invention. It is a block diagram of the spacecraft concerning Embodiment 2 of this invention. It is an example of the data table of the prediction induction | guidance | derivation error memorize | stored in the memory area of the arithmetic unit which concerns on Embodiment 2 of this invention.

Embodiment 1.
Hereinafter, the guidance device according to the first embodiment will be described with reference to the drawings. FIG. 1 is a diagram illustrating a configuration of a guidance device 50 according to the first embodiment.
The guidance device 50 obtains a reference trajectory from the aircraft center-of-gravity position identification means 1, the RF receiver 6 for receiving information on the atmospheric state over the target point transmitted from the ground facility, and the given motion model 7. A guidance calculation device 8 for calculating, a GPS receiver 9 including a GPS antenna, an inertial measurement device IMU 10, a barometric altimeter 11, and a control input to be performed by the aircraft to move along a reference trajectory are calculated. The control calculation device 12 is provided.

  The control amount calculated by the control calculation device 12 is sent as a control command 13 to a device such as a thruster or a steering device.

Here, the body center-of-gravity position identifying means 1 detects an angular velocity generating means 2 for applying torque to the aircraft, an angular velocity detecting means 3 for detecting the generated angular velocity, and an acceleration generated by the torque applied to the aircraft. Acceleration detection means 4 and a body gravity center position calculation device 5 for identifying the body gravity center position from the detected angular velocity and acceleration.
The aircraft center-of-gravity position calculation device 5, the guidance calculation device 8, and the control calculation device 12 are preferably implemented by the same arithmetic unit in order to simplify the aircraft design, but are separately provided for convenience of design. There is no problem even if it is installed.

Next, the operation of the guidance device according to the present embodiment will be described.
First, in the vicinity of the earth orbit before the start of reentry, the angular velocity generating means 2 such as a thruster or reaction wheel is operated for a short time to apply a minute torque to the aircraft.

After the generation of the angular velocity due to the minute torque, the angular velocity and acceleration generated in the aircraft by applying torque are acquired using the angular velocity detection device 3 such as a rate gyro mounted on the aircraft and the acceleration detection device 4 such as an accelerometer. .

The body centroid position identification device 5 identifies the body centroid position using the detected angular velocity, acceleration, and arithmetic unit.
An example of the process performed by the body center of gravity position identification is shown in the following formula (1). However, although it is desirable to use numerical values after noise removal using calculation means such as a smoothing filter, the output results of the rate gyro 15 and the accelerometer 16 are omitted here.

Here, explanation of each symbol is as follows.
r _os : Position vector from an arbitrary coordinate system origin fixed to the aircraft to the accelerometer.
r _og : A position vector from the origin of any coordinate system fixed to the aircraft to the center of gravity of the aircraft.
a _s . : Acceleration detected by accelerometer.
ω _s . : Angular velocity detected by the rate gyro.
ω _s . : Angular acceleration calculated from rate gyro output.

The motion model 7 used for the calculation for generating the reference trajectory is updated using the position of the center of gravity of the airframe obtained by the above calculation. The main thing of the motion model 7 that changes depending on the position of the center of gravity of the aircraft is the trim characteristics of the aircraft.
Expression (2) is a relational expression between the position of the center of gravity of the aircraft, the pitching moment coefficient of the aircraft, and the like.
By applying the position of the center of gravity of the airframe obtained by Expression (1) to Expression (2), the trim characteristics of the airframe can be obtained.

Here, description of each symbol is as follows.
Cm: Pitching moment coefficient of the aircraft.
CA: Axial force coefficient of the aircraft.
CN: Vertical force coefficient of the aircraft.
xcg: the center of gravity position in the x direction of the aircraft.
zcg: the position of the center of gravity of the aircraft in the z direction.
xmrc: Center of aerodynamic moment of the aircraft.
d: Aerodynamic reference length of the aircraft.
Cm0: Pitching moment coefficient when the center of gravity of the aircraft coincides with the center of aerodynamic moment.

Next, the guidance calculation device 8 calculates a reference trajectory from the guidance start point to the target point based on the position information obtained by the updated motion model 7 and the GPS receiver 9. The reference trajectory using the motion model 7 and the position information can be calculated by using the calculation methods shown in Non-Patent Documents 1 and 2.

After starting the guidance, the control calculation device 12 calculates a control command for performing a motion along the reference trajectory calculated by the guidance calculation device 8, and guides the aircraft to the target point.

Note that during guidance, atmospheric conditions and orbital errors are appropriately acquired by the RF receiver 6 and GPS receiver 9, and a control command for correcting the error is given to the spacecraft.

As described above, the guidance device according to the present embodiment includes the fuselage center-of-gravity position identification unit 1, and the angular velocity detection unit 3 and the acceleration detection unit 4 represent the angular velocity and acceleration generated by the minute torque applied to the fuselage by the angular velocity generation unit 2. The position of the center of gravity of the aircraft was identified by acquiring using.
According to this, the state quantity of the airframe required for the calculation of the reference trajectory at the time of re-entry of the spacecraft can be communicated with the ground station without applying a large computational load and power to the airborne computer. It can be identified without intervention. Furthermore, it is possible to reduce induction errors caused by sudden changes in the motion model.

Embodiment 2. FIG.
In the first embodiment, the description has been given of the mode in which the aircraft center-of-gravity position identifying unit identifies the center-of-gravity position of the aircraft, calculates the reference trajectory by using the identified aircraft center-of-gravity position, and guides the aircraft to the target point. In Section 2, a landing method for safely landing the aircraft guided to the target point will be described.

FIG. 2 is a diagram illustrating a configuration of the guidance device according to the second embodiment. The spacecraft in the present embodiment uses a parachute deployment mechanism when landing at a target point.
FIG. 2 is a diagram showing the configuration of the spacecraft 100 according to the first embodiment. The spacecraft 100 includes a fuselage center-of-gravity position identification unit 1, a calculation device 14, an RF receiver 6, a target point offset amount setting unit 17, a guidance calculation device 8, a motion model 7, a GPS receiver 9, A control calculation device 12, an IMU 10, a barometric altimeter 11, and a parachute deployment mechanism 15 are provided. In addition, about the structure which has the same function as Embodiment 1, the same number is attached | subjected and the description is abbreviate | omitted.
The spacecraft 100 according to the present embodiment has a parachute deployment mechanism 15 for descending to a target point by a parachute.
The parachute unfolding mechanism 15 unfolds the parachute according to the unfolding instruction issued by the control calculation device 12 at the timing when the guidance of the airframe ends just before reaching the ground surface, and lands on the target point.

In the spacecraft 100 according to the second embodiment, identification of the center of gravity position of the aircraft and acquisition of information regarding the atmospheric state can be realized by the same method as described in the first embodiment.
However, the motion model used for the calculation in the present embodiment is not the numerical value calculated onboard as in the first embodiment, but the standard design value of the aircraft and the average numerical value assumed in operation. It is assumed that the fixed trajectory is generated by using the fixed motion model.
In this case, since the center of gravity position of the actual aircraft is different from the model used for the calculation, a guidance error occurs.

Therefore, in the present embodiment, the induction error caused by the motion model 7 is calculated in advance and stored in the storage area of the arithmetic device 14 mounted on the aircraft.

FIG. 3 is an example of a data table of the predicted guidance error table 16 stored in the storage area of the arithmetic device 14.
As shown in FIG. 3, the guidance error finally generated due to the difference between the center of gravity position of the body obtained by calculation and the center of gravity position of the body based on the specified motion model is stored in the storage area of the computing device 14 as a data table.

The target point offset amount setting unit 17 sets a target point obtained by offsetting the guidance error predicted in advance by the predicted guidance error 16 from the actual target point.
That is, if the guidance error predicted by the prediction guidance error 16 is r _err , the target point set in the aircraft is offset from the actual target point by −r _err .
Thereby, the induction | guidance | derivation error of the body by the control calculation apparatus 12 can be reduced.

As described above, according to the spacecraft according to the present embodiment, the predicted guidance error table 16 and the parachute deployment mechanism 15 are provided, and the target point is offset from the actual target point according to the guidance error predicted in advance. In the landing, the parachute deployment mechanism 15 deploys the parachute according to the deployment instruction issued by the control calculation device 12 at the timing when the guidance of the aircraft is completed immediately before reaching the ground surface, and landed at the target point. I did it.

  According to this, it is possible to further reduce the calculation load and power consumption of the computer mounted on the airframe, and it is possible to land safely at the target point by the parachute.

The target point in the first and second embodiments may be set on the parachute deployment altitude or may be set on the ground surface.

DESCRIPTION OF SYMBOLS 1 Aircraft center-of-gravity position identification means, 2 Angular velocity generation means, 3 Angular velocity detection means, 4 Acceleration detection means, 5 Airframe gravity center position identification apparatus, 6 RF receiver, 7 Motion model, 8 Guidance calculation apparatus, 9 GPS receiver, 10 IMU , 11 barometric altimeter, 12 control calculation device, 13 control command, 14 arithmetic device, 15 parachute expansion mechanism, 16 prediction guidance error table, 17 target point offset amount setting unit, 50 guidance device, 100 spacecraft.

Claims (4)

A guidance device for guiding a spacecraft having a re-entry function to a destination point,
Angular velocity generating means for applying torque to the spacecraft;
Angular velocity detection means for detecting the angular velocity generated by the torque;
Acceleration detecting means for detecting acceleration caused by the torque;
An aircraft center-of-gravity position identification unit comprising an angular velocity detected by the angular velocity detection unit and an aircraft center-of-gravity position identification unit that identifies the center of gravity position of the spacecraft using the acceleration detected by the acceleration detection unit;
Using the position of the center of gravity of the aircraft output by the position identification means, calculating the trajectory of the spacecraft to re-enter, and a control calculation section for guiding the spacecraft to a destination point based on the trajectory,
A guidance device comprising: The angular velocity generating means is a thruster or reaction wheel mounted on the spacecraft,
The angular velocity detection means is a rate gyro mounted on the spacecraft,
The guidance device according to claim 1, wherein the acceleration detecting means is an accelerometer mounted on the spacecraft. The guidance device according to any one of claims 1 and 2,
Equipped with a parachute deployment mechanism,
The guidance apparatus-equipped spacecraft, wherein the parachute deployment mechanism deploys the parachute in accordance with a deployment instruction issued at a timing when the guidance of the spacecraft is completed before the control calculation unit reaches the ground surface. The control calculation unit
Predicted guidance error table that tabulates the guidance error caused by the difference between the center of gravity of the aircraft obtained by calculation and the center of gravity of the aircraft based on the specified motion model, and offsets the guidance error predicted in advance by the forecast guidance error table 4. The spacecraft equipped with a guidance device according to claim 3, wherein the set destination point is newly set and guided to the set target point.

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