Flight trajectory control device and trajectory control method

Abstract

The objective is to propose a thrust control algorithm for maintaining a constant argument of perigee in an elliptical orbit where the perigee is in an extremely low orbit, and to provide an orbit control device and orbit control method for an aircraft that enables efficient orbit maintenance and high-resolution observation. [Solution] We propose a control method that keeps the variations of each orbital parameter—major axis, eccentricity, and argument of perigee—within a threshold. Based on the differences in thrust sensitivity of each orbital parameter, priority is given to correcting the eccentricity and argument of perigee, which have higher sensitivity, and then the major axis is corrected. By appropriately adjusting the thrust application region at an orbital position where the correction efficiency of the perigee and argument of perigee is maximized, both maintaining the perigee and minimizing propellant consumption are achieved.

Description

Technical Field

[0001] The present invention relates to an orbit control device and an orbit control method for a flying object. Specifically, in an elliptical orbit where the perigee is in a very low orbit, a thrust control algorithm for keeping the argument of perigee constant is proposed, and the present invention relates to an orbit control device and an orbit control method for a flying object that achieve efficient orbit maintenance and high-resolution observation.

Background Art

[0002] In recent years, an orbit at an altitude of less than 400 km, called the very low Earth orbit (VLEO), has attracted attention as an orbit that enables high-resolution data acquisition and low-latency communication in the fields of Earth observation and communication. However, in this altitude range, atmospheric drag becomes a dominant perturbation factor, and continuous thrust addition is essential to maintain the satellite's orbit. In conventional orbit control methods, it is common to apply thrust when the orbital semi-major axis falls below a certain threshold, and since it mainly targets circular orbits, the movement of the perigee has rarely been regarded as a problem.

[0003] On the other hand, recent research has proposed the use of an elliptical orbit with the perigee placed in a very low orbit in order to suppress propellant consumption. For example, Non-Patent Document 1 shows the possibility of designing an elliptical orbit in VLEO by utilizing air-breathing electric propulsion technology. This research presents a new approach to achieve continuous orbit maintenance by using air as a propellant in outer space where propellant replenishment is difficult.

[0004] Also, Non-Patent Document 2 details the theoretical background regarding changes in orbit parameters, and in particular, it shows that the time variation of the argument of perigee depends on the oblateness of the Earth and the orbit inclination angle. This theoretically supports the need to set the orbit inclination angle to a critical inclination angle (about 63° or 117°) in order to keep the argument of perigee constant.

Prior Art Documents

Non-Patent Documents

[0005] [Non-Patent Document 1] Yuxian Yue et al., Elliptical Orbit Design Based on Air-Breathing Electric Propulsion Technology in Very-Low Earth Orbit Space, Aerospace, 2023, 10, 899. [Non-Patent Document 2] Minoru Han'yō, "Fundamentals of Mission Analysis and Trajectory Design," Gendai Sugakusha (Modern Sugaku Publishing). [Overview of the project] [Problems that the invention aims to solve]

[0006] However, the aforementioned non-patent literature does not provide specific control methods for the problem of the perigee continuously shifting across the Earth's surface due to fluctuations in the argument of perigee. As a result, it has been difficult to maintain the perigee above the desired observation point, leaving the challenge of reduced observation frequency.

[0007] This invention was conceived in view of the above points, and proposes a thrust control algorithm for maintaining a constant argument of perigee in an elliptical orbit where the perigee is in an extremely low orbit, and proposes an orbit control device and orbit control method for an aircraft that realize efficient orbit maintenance and high-resolution observation. [Means for solving the problem]

[0008] To achieve the above objective, the present invention provides a control device for an aircraft that orbits on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in an altitude region of 400 km or less where atmospheric resistance is dominant, and comprises: an orbital parameter acquisition unit that acquires orbital parameters consisting of the semi-major axis, eccentricity, and perigee argument of the aircraft while it is flying on the elliptical orbit; a determination unit that calculates the difference between the orbital parameters and a target orbital parameter which is a target value, and determines whether the difference exceeds a preset threshold; a thrust addition calculation unit that, based on the result of the determination unit, calculates the position on the elliptical orbit in which thrust should be added to the aircraft, and the thrust to be added at that position, so that the semi-major axis, eccentricity, and perigee argument are maintained within a predetermined range; and a thrust control unit that controls the output of the aircraft based on the calculation result of the thrust addition calculation unit.

[0009] Here, by providing an orbital parameter acquisition unit that acquires orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee of an aircraft flying in an elliptical orbit, it is possible to accurately grasp in real time the state of orbital parameters (semi-major axis, eccentricity, argument of perigee) that indicate the size, shape, and orientation of an elliptical orbit, which are constantly changing under atmospheric drag, and based on this, precise thrust addition calculations can be performed to maintain each orbital parameter within a predetermined range.

[0010] Furthermore, by incorporating a determination unit that calculates the difference between the orbital parameters and the target orbital parameters, and determines whether the difference exceeds a preset threshold, it is possible to accurately detect the point at which fluctuations in orbital parameters due to atmospheric resistance exceed the acceptable range and control becomes essential. This avoids excessive thrust application in response to minute fluctuations, and by performing minimal corrections only when necessary, it is possible to reduce propellant (fuel) consumption and contribute to stable orbit maintenance and mission extension.

[0011] Furthermore, by incorporating a thrust addition calculation unit that calculates the position on the elliptical orbit where thrust should be added to the aircraft, and the thrust to be added at that position, based on the results of the determination unit, the semi-major axis, eccentricity, and argument of perigee should be maintained within a predetermined range. This allows for the instantaneous determination of the optimal thrust addition position and thrust while minimizing the impact on other orbital parameters, in response to requirements that necessitate correction of any of the semi-major axis, eccentricity, or argument of perigee. As a result, based on precise control logic, propellant can be used efficiently, and the control targets for orbital inclination and argument of perigee can be simultaneously achieved over a long period of time.

[0012] Furthermore, by including a thrust control unit that controls the output of the aircraft based on the calculation results of the thrust addition calculation unit, the output of the aircraft, such as electric propulsion systems, can be controlled in real time based on the position and thrust determined by the thrust addition calculation unit. This makes it possible to offset the influence of atmospheric drag perturbations on orbital parameters and maintain the orbit within a certain range.

[0013] Furthermore, when the thrust addition calculation unit corrects the semi-major axis simultaneously with the correction of eccentricity or the argument of periphery, it first calculates the increase in the semi-major axis due to the correction of eccentricity, and / or first calculates the increase in the semi-major axis due to the correction of the argument of periphery. In this case, it takes into account the increase in the semi-major axis due to the correction of eccentricity and the argument of periphery prior to the calculation of the main semi-major axis correction. This allows for accurate calculation of the amount of thrust addition necessary to achieve the target value of the semi-major axis, enabling efficient thrust addition without excess or deficiency. This minimizes interference between orbital parameters in simultaneous multi-element control, enabling highly accurate orbit maintenance.

[0014] Furthermore, if the thrust addition calculation unit, based on the determination by the determination unit, finds that the difference between the target eccentricity (the target value for eccentricity among the target values) and the actual eccentricity exceeds a preset threshold, it calculates the position where thrust should be added to maintain the eccentricity within a predetermined range, such that the position is near the perigee where the true anomaly is approximately 0°. In this case, when a deviation in eccentricity is detected, the unit identifies a position near the perigee where the true anomaly is approximately 0°, which allows for the most efficient correction of the eccentricity with minimal impact on the semi-major axis. This minimizes unnecessary interference with other orbital parameters, reduces propellant consumption, and allows for the accurate and continuous maintenance of the elliptical orbit shape.

[0015] Furthermore, the thrust addition calculation unit, based on the determination by the determination unit, calculates the position where thrust should be added so that the true angle of separation, where the cosine coincides with the negative value of the eccentricity, is in the range of 0° to 180°, in order to increase the eccentricity argument and maintain it within a predetermined range, if the difference between the target eccentricity argument and the eccentricity argument is a positive value. If the difference between the target eccentricity argument and the eccentricity argument is a negative value, the thrust addition calculation unit calculates the position where thrust should be added so that the true angle of separation, where the cosine coincides with the negative value of the eccentricity, is in the range of 180° to 360°, in order to decrease the eccentricity argument and maintain it within a predetermined range. In this way, the thrust addition calculation unit accurately determines the direction of deviation (insufficient / excessive) of the eccentricity argument and identifies a position in the range of 0° to 180° or 180° to 360° that minimizes the impact on other elements. This makes it possible to cancel out perturbations of the eccentricity argument that occur in orbits near the critical inclination angle. Furthermore, if the average value of the orbital inclination over time closely matches the critical inclination, it may be possible to maintain the variation in the argument of perimes within an acceptable range without performing any correction operations on the argument of perimes.

[0016] Furthermore, if the thrust addition calculation unit determines, based on the determination by the determination unit, that the eccentricity and argument of periphery are maintained within the range of target values, and the difference between the target semi-major axis and the semi-major axis exceeds a preset threshold, the unit calculates multiple symmetrical positions on the elliptical orbit where thrust should be added to maintain the semi-major axis within a predetermined range. By adding thrust at these multiple positions, the unit increases the semi-major axis while calculating a thrust that reduces the total change in the eccentricity and argument of periphery of the elliptical orbit to approximately zero. If other orbital parameters (eccentricity and argument of periphery) are within acceptable limits, the thrust addition can be performed at multiple symmetrical positions on the orbit to efficiently increase only the semi-major axis. This offsets the decrease in the semi-major axis due to atmospheric resistance, prevents unnecessary fluctuations from being added to an already stable orbital shape, and enables the long-term maintenance of a highly accurate recurring orbit.

[0017] To achieve the above objective, the present invention provides a method for controlling the trajectory of an aircraft that orbits on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in an altitude region of 400 km or less where atmospheric resistance is dominant, comprising the steps of: acquiring orbital parameters consisting of the semi-major axis, eccentricity, and perigee argument of the aircraft while it is flying on the elliptical orbit; calculating the difference between the acquired orbital parameters and target orbital parameters which are target values, and determining whether the difference exceeds a preset threshold; calculating, based on the result of the determination step, the position on the elliptical orbit in which thrust should be added to the aircraft, and the thrust to be added at that position, so that the semi-major axis, eccentricity, and perigee argument are maintained within a predetermined range; and controlling the output of the aircraft based on the calculation result of the calculation step.

[0018] Through the above process, in an elliptical orbit affected by atmospheric drag, orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee are constantly monitored, and precise thrust application position and thrust are calculated and controlled. This allows for high-precision orbit maintenance and extended mission duration while minimizing propellant consumption. [Effects of the Invention]

[0019] The trajectory control device and trajectory control method for an aircraft according to the present invention propose a thrust control algorithm for maintaining a constant argument of perigee in an elliptical orbit where the perigee is in an extremely low orbit, thereby enabling efficient trajectory maintenance and high-resolution observation. [Brief explanation of the drawing]

[0020] [Figure 1] This is a block diagram of a track control device according to an embodiment of the present invention. [Figure 2] This graph shows the characteristics of the time rate of change (normalized value) of each orbital parameter when thrust is applied in the tangential direction of the orbit at the true anterior angle v. [Figure 3] This figure shows the control flow in the thrust addition calculation unit according to an embodiment of the present invention. [Figure 4] This diagram shows the orbital position where thrust is applied in a way that maximizes the rate of increase in eccentricity. [Figure 5] This diagram shows the orbital position where thrust is applied in a way that maximizes the rate of increase in the perigee argument. [Figure 6] This diagram shows the orbital position where thrust is applied in such a way that the rate of decrease in the perigee argument is maximized. [Figure 7] This diagram shows the orbital position where thrust is applied in such a way that the rate of increase in the semi-major axis is maximized while keeping the total change in eccentricity and the argument of perigee approximately zero. [Figure 8] This diagram shows the orbital position where thrust is applied in such a way that the rate of increase in eccentricity is maximized and the semi-major axis also increases, while keeping the total change in the argument of perigee approximately zero. [Figure 9] This diagram shows the orbital position where thrust is applied in a way that maximizes the rate of increase in the argument of periphery and also increases the semi-major axis, while keeping the total change in eccentricity approximately zero. [Figure 10] This diagram shows the position on the orbit where thrust is applied in such a way that the rate of decrease in the argument of periphery is maximized while increasing the semi-major axis. [Figure 11]This diagram shows the position on the orbit where thrust is applied to increase the semi-major axis while maximizing the rate of increase in eccentricity and the argument of perpendicularity. [Figure 12] This diagram shows the orbital position where thrust is applied in such a way that the rate of decrease in the argument of periphery and the rate of increase in the eccentricity are both maximized, while increasing the semi-major axis. [Modes for carrying out the invention]

[0021] Hereinafter, an embodiment of the aircraft trajectory control device and trajectory control method of the present invention will be described in detail with reference to drawings and other drawings to facilitate understanding of the present invention.

[0022] The orbit control device for an aircraft according to this embodiment (hereinafter referred to as the "orbit control device") is applied to aircraft, including satellites, that orbit in an elliptical orbit where the orbital inclination is maintained near the critical inclination and the perigee is located in the region below 400 km altitude (hereinafter referred to as "very low Earth orbit") where atmospheric drag is dominant. While such elliptical orbits passing through very low Earth orbit enable high-resolution observations, the influence of atmospheric drag is very large, and it is difficult for the argument of perigee to remain strictly constant due to the precession caused by the flattening of the Earth, making precise orbit maintenance over long periods of time a challenge. To solve this problem, this device efficiently controls orbital parameters while minimizing interference between orbital parameters.

[0023] Figure 1 is a schematic diagram (block diagram) of a track control device 1 according to an embodiment of the present invention. The track control device 1 according to the present invention mainly consists of a track parameter acquisition unit 10, a determination unit 20, a thrust addition calculation unit 30, and a thrust control unit 40. Each component will be described in detail below.

[0024] [Orbital parameter acquisition unit] The orbital parameter acquisition unit acquires in real time the current values ​​of orbital parameters consisting of the semi-major axis a, eccentricity e, and argument of perigee ω of an object orbiting in an elliptical orbit.

[0025] These parameters are derived based on data from external sensor 2. Specifically, in addition to the real-time position and velocity vector of the aircraft obtained mainly from the GNSS receiver, non-gravity acceleration measurements are acquired from a highly sensitive accelerometer to accurately grasp the effect of atmospheric drag, which is dominant in the ultra-low Earth orbit environment. The orbital parameter acquisition unit 10 takes these measurements as input values ​​and applies an arbitrary algorithm integrated with physical models such as the Earth's gravity field model to estimate highly accurate orbital parameters that eliminate the effects of measurement noise and disturbances, and outputs them to the subsequent computing unit.

[0026] Furthermore, the method for obtaining orbital parameters in this invention is not limited to the estimation algorithm described above. All technical means capable of deriving the position, velocity, and rate of change of the aircraft based on observational data and identifying orbital parameters are available. This includes orbit determination methods using batch processing such as the least squares method, other continuous state estimation methods, or ground control methods that determine orbital elements from observational values ​​obtained by Doppler tracking using ground antennas or satellite laser ranging.

[0027] In the very low Earth orbit where an object orbits, atmospheric drag is the dominant factor, and the orbital parameters of the object—the semi-major axis a, eccentricity e, and argument of perigee ω—are constantly fluctuating due to atmospheric drag. Here, the time rate of change of the semi-major axis a due to atmospheric drag is shown in equation (1), the time rate of change of the eccentricity e in equation (2), and the time rate of change of the argument of perigee ω in equation (3). Note that D is atmospheric drag, m is the mass of the object, μ is the geocentric gravity constant, and v is the true angle of perigee.

[0028]

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[0029]

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[0030]

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[0031] From equations (1) to (3), the semi-major axis a is constantly decreasing, the eccentricity e decreases with the mean orbital period, and the argument of perigee ω also changes according to the true angle of perigee v. Therefore, the orbital parameter acquisition unit 10 is essential for accurately grasping these fluctuations. The acquired orbital parameters are immediately sent to the determination unit 20, which serves as the basis for determining whether the difference with the target orbital parameters exceeds a threshold.

[0032] [Judgment section] The determination unit 20 uses the current values ​​of the semi-major axis a, eccentricity e, and perigee argument ω sent from the orbital parameter acquisition unit 10, and the preset target orbital parameters (target semi-major axis a tar , target eccentricity e tar , target near point argument ω tar The difference between this value and the predetermined threshold (Δa) for each parameter is calculated. thres Δe thres , Δω thres Determine whether it exceeds ).

[0033] Here, the function of the determination unit 20 is to accurately detect the point at which fluctuations in orbital parameters due to atmospheric resistance exceed the acceptable range, making orbital correction by thrust addition essential. This avoids instability caused by excessive thrust addition in response to minute fluctuations and optimizes propellant consumption. Each threshold used for determination is set considering the following control characteristics.

[0034] First, the threshold for the semi-major axis a (Δa thres As shown in equation (1) above, this value is constantly decreasing due to atmospheric drag and directly leads to a decrease in orbital altitude. This threshold is set based on the mission duration of the aircraft, the precision required to achieve the desired returnability, and the minimum altitude at which the orbit can be maintained, taking into account the atmospheric drag sphere.

[0035] Threshold of eccentricity e (Δe thres) is an important parameter that determines the shape (flattening ratio) of the elliptical orbit and particularly defines the altitude difference between the apogee and perigee. In the high-resolution observation mission by the flying object targeted by the present invention, the perigee altitude closest to the earth's surface directly affects the resolution of the observation and the magnitude of the atmospheric drag. Therefore, since the variation of the eccentricity e is directly related to the accuracy of high-resolution observation, it is necessary to maintain the eccentricity e within a strict range.

[0036] The threshold value of the argument of perigee ω (Δω thres ) is such that when the orbital inclination angle is maintained near the critical inclination angle, the rate of change of the argument of perigee ω with time can be suppressed. Therefore, theoretically, the control frequency of the argument of perigee ω can be reduced. Also, if corrections are made even for very small variations, the additional thrust may become excessive, and there is a risk that the orbit will diverge. Therefore, in this embodiment, in order to avoid this risk, by setting Δω thres to a larger threshold value compared to the threshold values of other parameters, the control frequency of the argument of perigee ω is intentionally suppressed to achieve stable control.

[0037] [Thrust addition calculation unit] Based on the result of the determination by the determination unit 20, the thrust addition calculation unit 30 calculates the position on the elliptical orbit where thrust should be added and the thrust to be added at that position so that the semi-major axis a, eccentricity e, and argument of perigee ω are maintained within a predetermined range.

[0038] The thrust addition calculation unit 30 is based on the calculation of the rates of change with time of the semi-major axis a, eccentricity e, and argument of perigee ω when thrust is added in the orbital tangent direction. Here, the rate of change with time of the semi-major axis a when thrust is added in the orbital tangent direction is shown in Equation (4), the rate of change with time of the eccentricity e is shown in Equation (5), and the rate of change with time of the argument of perigee ω is shown in Equation (6). Note that F is the thrust, m is the mass of the flying object, μ is the geocentric gravitational constant, and v is the true anomaly. From these equations, by adjusting the thrust addition position and thrust addition time, the optimal correction amount for converging each orbital parameter to the target value is calculated.

[0039] [Equation]

[0040]

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[0041]

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[0042] Figure 2 shows the characteristics of the time rate of change (normalized value) of each orbital parameter with respect to the true anomaly v shown by equations (4) to (6), corresponding to the case where the eccentricity e = 0.039, the periapsis is 250 km, and the apoapsis is 800 km. As this figure shows, the rate of change of the eccentricity e and the argument of peripate ω have a point where they are zero at a specific true anomaly v, and the range of variation in the rate of change due to the change in true anomaly v (the difference between the maximum and minimum values) is very large. In contrast, the rate of change of the semi-major axis a has the characteristic that the range of variation in its rate of change is relatively small compared to other orbital parameters over the entire range of true anomaly v.

[0043] Figure 2 shows that the most efficient way to increase the semi-major axis a is to add thrust in the tangential direction at the perigee, but when the eccentricity e is small to begin with, the difference in efficiency due to the true anterior angle v is very small. While the eccentricity e increases with the addition of thrust at the perigee, the cosine of the true anterior angle v (cosv) coincides with the negative value of the eccentricity e (cosv = -e) v = cos ―1 The rate of increase is 0 at the point where (-e) is true.

[0044] Such angles exist in two cases when the eccentricity e is greater than 0 and less than 1, i.e., as long as the trajectory shape of the flying object is elliptical, and v = cos ―1 Let v be the angle such that (-e) + Let v = cos ―1 Let v be the angle such that (-e) - Let's assume that if the eccentricity e is small, then v + and v -These are approximately 90° and 270°, respectively. Regarding the argument of perigee ω, the rate of increase is 0 even when thrust is applied in the tangential direction between the perigee and apogee, and v + In this case, the rate of increase is greatest, v - The rate of decrease is greatest in this case.

[0045] This means that, in the correction of the semi-major axis a, the sensitivity of the rate of change of orbital parameters due to the selection of the thrust application position is lower compared to the eccentricity e and the argument of perpendicularity ω. Therefore, it can be seen that the correction of the semi-major axis a has the characteristic of not causing a significant decrease in efficiency per unit thrust, regardless of where the thrust is applied in the orbit.

[0046] Based on these characteristics, the control algorithm of the thrust addition calculation unit 30 employs the control flow shown in Figure 3, prioritizing the correction of the eccentricity e and the perigee argument ω, which are highly sensitive to position and suppress interference with other orbital parameters, over the correction of the semi-major axis a.

[0047] In this control flow, the thrust addition calculation unit 30 first determines whether or not correction of the eccentricity e or the argument of periphery ω is necessary (S1). Furthermore, if the thrust addition calculation unit 30 is instructed to perform composite control in which the correction of the semi-major axis a is performed simultaneously with the correction of the eccentricity e or the argument of periphery ω (when "Yes" is selected in S1, S2, and S3 in Figure 3), it first calculates the increase in the semi-major axis a due to the correction of the eccentricity e or the argument of periphery ω. This logic takes into account that when thrust is added to correct the eccentricity e or the argument of periphery ω, the semi-major axis a will inevitably increase due to orbital dynamics interference.

[0048] Therefore, the thrust addition calculation unit 30 first calculates the increase in the semi-major axis a that is incidentally caused by the correction of other parameters, and adds it to the current semi-major axis a to obtain the updated semi-major axis a. upd Calculate this a. upd Using the target semi-major axis a tarBy evaluating the difference, it is determined whether an additional correction for the semi-major axis a is necessary. This calculation priority avoids excessive thrust application to the semi-major axis a, suppresses propellant consumption, and enables highly accurate orbit maintenance while maintaining a balance of multiple orbital parameters.

[0049] The above describes the main control logic in the thrust addition calculation unit 30. Next, we will explain the detailed correction logic for each trajectory parameter.

[0050] (Correction for eccentricity) First, regarding the eccentricity e, in elliptical orbits passing through very low orbits, the eccentricity e tends to decrease due to the effect of atmospheric drag, therefore, the target eccentricity e tar and the actual eccentricity e sat The difference Δe is the threshold Δe thres The criterion for judgment is whether or not it exceeds [a certain value].

[0051] And the target eccentricity e tar and the actual eccentricity e sat The difference Δe = e tar -e sat The threshold Δe thres If it exceeds (if "Yes" is selected in S1 or S3 in Figure 3), the actual eccentricity e sat To maintain the thrust within a predetermined range, the position where thrust should be applied is calculated to be near the perigee where the true angle of separation v is approximately 0°.

[0052] The thrust addition calculation unit 30 sets the range of true anomaly separation for thrust addition to v = -θ in order to effectively increase the eccentricity e without changing the argument of periphery ω. e / 2~θ e Let's define it as / 2, and the angle θ e We numerically determine this such that it satisfies the following equation (7) (S11, S31 in Figure 3).

[0053]

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[0054] Here, de / dv in equation (7) is obtained from the time rate of change de / dt using thrust F and atmospheric drag D, and the time rate of change dt / dv with respect to the true anterior separation, by setting de / dv = de / dt·dt / dv. Figure 4 schematically shows the range when only thrust is added. η e This is a coefficient applied to the correction, and is selected from numbers greater than 0 and less than or equal to 1. For example, η e When it is 1, it corresponds to the case where thrust is added to return to the target trajectory, η e A value of 0.5 has the effect of adding thrust to the satellite even in orbits between the target orbit and an orbit below the threshold. Such correction factors are appropriately selected depending on the actual propulsion system used and the mission requirements.

[0055] (Correction of the argument of perpendicularity) The periphrase argument ω can fluctuate in either direction, increasing or decreasing, even in regions where atmospheric drag is dominant, therefore, the target periphrase argument ω tar and current value ω sat The difference Δω = ω tar -ω sat The absolute value of the set threshold Δω thres The criterion for judgment is whether or not it exceeds [a certain value].

[0056] And Δω>Δω thres In this case, the thrust addition is v = v + This is done before and after this point (in the range where the true anomaly is 0° to 180°). Before and after this point, the rate of increase of the anomaly argument ω is maximum, while the rate of change of the eccentricity e reverses sign. Therefore, thrust is applied as v = v + -θ ω / 2~v + +θ ω Perform this within the range where it is equal to / 2. In this case, the angle θ ω η can be calculated based on equation (8). ω This is a coefficient applied to the correction, and is selected from a number greater than 0 and less than or equal to 1. Figure 5 schematically shows the range when this thrust is applied.

[0057]

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[0058] On the other hand, Δf < -Δf thres In this case, the thrust addition is v = v - This is done before and after this point (in the range where the true anomaly is between 180° and 360°). Before and after this point, the rate of decrease of the anomaly argument ω is maximum, while the rate of change of the eccentricity e reverses sign. Therefore, thrust is applied as v=v - -θ ω / 2~v - +θ ω Perform this within the range where it is equal to / 2. In this case, the angle θ ω η can be calculated based on equation (9). ω This is a coefficient applied to the correction, and is selected from a number greater than 0 and less than or equal to 1. Figure 6 schematically shows the range when this thrust is applied.

[0059]

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[0060] (Correction for semi-major axis) After the thrust addition calculation unit 30 has completed the correction logic for the eccentricity e and the perpendicular argument ω, if correction of the semi-major axis a is necessary, it applies either single correction or combined control logic.

[0061] First, let's explain the logic for the independent correction of the semi-major axis a. In the control flow of Figure 3, it is determined that the eccentricity e and the peripheral argument ω are maintained within the range of the target values, and the difference Δa = a tar -a sat The threshold Δa thres If it exceeds (if "Yes" is selected in S4 of Figure 3), the thrust required to increase the semi-major axis a is calculated (S41).

[0062] The thrust addition calculation unit 30 calculates the positions where thrust should be added so that they are multiple symmetrical positions on the elliptical orbit, in order to efficiently increase only the semi-major axis a while keeping the total change in the eccentricity e and the argument of periphery ω approximately zero. Specifically, the symmetrical range v=v is calculated before and after the periphery and aponeurosis where de / dv and dω / dv are zero, or the position where the sign of dw / dv is reversed. + -θ a / 4~v + +θ a / 4, and v=v - -θ a / 4~v - +θ a The correction is applied by / 4. In this case, the angle θ a η can be calculated based on equation (10). a This is a coefficient applied to the correction, and is selected from a number greater than 0 and less than or equal to 1. Figure 7 schematically shows the range when this thrust is applied.

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[0064] Next, we will explain the control logic for composite control, which corrects the semi-major axis a simultaneously with the correction of the eccentricity e or the peripheral argument ω. When the thrust addition calculation unit 30 is instructed to perform composite control, it applies a calculation priority that first calculates the increase in the semi-major axis due to the correction of other elements.

[0065] First, when correcting the semi-major axis a simultaneously with correcting the eccentricity e (when "Yes" is selected in S33 of Figure 3), the thrust range v required for correcting the eccentricity e is -θ e / 2~θ e The semi-major axis a also increases in / 2. The updated semi-major axis a takes this increase into account. upd This is calculated using the following formula (11).

[0066]

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[0067] This aupd Using this, the semi-major axis a of the target orbit tar The difference between this and Δa upd =a upd -a tar Defined as Δa upd >Δa thres In this case, an additional semi-major axis correction is required, and the range θ is θ a Using the left side of equation (10), Δa upd It is calculated to obtain [the desired result]. Figure 8 is a schematic diagram showing the range when this thrust is applied.

[0068] Next, when the correction of the semi-major axis a is performed simultaneously with the correction of the argument of periphery ω (when "Yes" is selected in S23 of Figure 3), the semi-major axis a also increases within the thrust addition range required for the correction of the argument of periphery ω (the range obtained by equation (8) or equation (9)). The updated semi-major axis a takes this increase into account. upd This is calculated using the following equations (12) and (13) (equation (12) and (13) are selected depending on the increase or decrease of the perpendicular argument).

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[0071] This updated semi-major axis a upd If additional semi-major axis correction is required using (Δa upd >Δa thres In this case, the range over which thrust is applied is calculated using equation (14). η a This is a coefficient applied to the correction, and is selected from a number greater than 0 and less than or equal to 1. Figures 9 and 10 schematically show the range when this thrust is applied. In this way, the thrust application calculation unit 30 can increase the semi-major axis a while suppressing interference with the eccentricity e and the near-point argument ω by applying thrust symmetrically to the four regions around the correction region of the near-point argument ω.

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[0073] Furthermore, when the correction of the semi-major axis a is performed simultaneously with the correction of the eccentricity e and the argument of perpendicularity ω (when "Yes" is selected in S1 in Figure 3), the range of thrust added for the semi-major axis correction is calculated by the following equations (15) and (16), and the range in which thrust is added is calculated by the aforementioned equation (14). Figures 11 and 12 schematically show the range when this thrust is added.

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[0076] [Thrust Control Unit] The thrust control unit 40 receives the final thrust addition command from the thrust calculation unit 30 and is responsible for controlling the generation of thrust in the tangential direction of the aircraft's orbit. Its main function is to accurately execute the thrust addition range and thrust generation amount necessary for correcting the semi-major axis a, eccentricity e, and argument of perigee ω. Specifically, it monitors the current true angle of separation in real time, and when it detects that the commanded angular range has been reached, it drives the onboard propulsion system 3 and adds thrust precisely for the calculated time. This realizes a control logic that maintains multiple orbital elements independently and efficiently.

[0077] As is clear from the above description, the orbit control device according to this embodiment employs a control logic that precisely considers the fluctuation characteristics of each orbital element and the sensitivity to thrust application position, thereby enabling the maintenance of a desired elliptical orbit with high precision over a long period of time while minimizing propellant consumption and maintaining a balance of multiple orbital parameters, even in ultra-low Earth orbit environments where atmospheric drag is dominant.

[0078] In summary, the trajectory control device and trajectory control method for an aircraft according to the present invention propose a thrust control algorithm for maintaining a constant argument of perigee in an elliptical orbit where the perigee is in an ultra-low orbit, thereby enabling efficient trajectory maintenance and high-resolution observation. [Explanation of Symbols]

[0079] 1. Orbital control device 10. Orbital parameter acquisition unit 20 Judgment section 30 Thrust addition calculation unit 40 Thrust Control Unit 2. External sensors 3 Propulsion system

Claims

1. An orbital control device for an aircraft orbiting on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in the region below 400 km altitude where atmospheric drag is dominant, An orbital parameter acquisition unit that acquires orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee of the aircraft flying in the aforementioned elliptical orbit, A determination unit calculates the difference between the orbital parameter and the target orbital parameter, which is the target value, and determines whether the difference exceeds a preset threshold. Based on the results of the determination unit, a thrust addition calculation unit calculates the position on the elliptical orbit where thrust should be added to the aircraft, and the thrust to be added at that position, such that the semi-major axis, the eccentricity, and the argument of perpendicularity are maintained within a predetermined range. The system includes a thrust control unit that controls the output of the aircraft based on the calculation results of the thrust addition calculation unit, The thrust addition calculation unit is, If, as a result of the determination by the determination unit, the difference between the target eccentricity (which is the target value of the eccentricity among the orbital parameters) and the actual eccentricity exceeds a preset threshold, the unit calculates the position where thrust should be applied to maintain the eccentricity within a predetermined range, such that the position is near the perigee where the true angle of separation is approximately 0°. An aircraft trajectory control device.

2. An orbit control device for an aircraft orbiting on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in a region below 400 km altitude where atmospheric drag is dominant, An orbital parameter acquisition unit that acquires orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee of the aircraft flying in the aforementioned elliptical orbit, A determination unit calculates the difference between the orbital parameter and the target orbital parameter, which is the target value, and determines whether the difference exceeds a preset threshold. Based on the results of the determination unit, a thrust addition calculation unit calculates the position on the elliptical orbit where thrust should be added to the aircraft, and the thrust to be added at that position, such that the semi-major axis, the eccentricity, and the argument of perpendicularity are maintained within a predetermined range. The system includes a thrust control unit that controls the output of the aircraft based on the calculation results of the thrust addition calculation unit, The thrust addition calculation unit is, If, as a result of the determination by the determination unit, the absolute value of the difference between the target near-point argument, which is the target value of the near-point argument among the orbital parameters, and the near-point argument exceeds a preset threshold, If the difference between the target near-point argument and the near-point argument is a positive value, the position where thrust should be applied is calculated to increase the near-point argument and maintain it within a predetermined range, such that the position is near the point where the cosine coincides with the negative value of the eccentricity, within the range of true near-point separation from 0° to 180°. If the difference between the target near-point argument and the near-point argument is a negative value, the position where thrust should be applied is calculated to decrease the near-point argument and maintain it within a predetermined range, such that the position is near the point where the cosine coincides with the negative value of the eccentricity, within the range of true near-point separation from 180° to 360°. An aircraft trajectory control device.

3. The thrust addition calculation unit is, When the correction of the semi-major axis is performed simultaneously with the correction of the eccentricity or the correction of the argument of the nearest point, the increase in the semi-major axis due to the correction of the eccentricity is calculated first, and / or the increase in the semi-major axis due to the correction of the argument of the nearest point is calculated first. A trajectory control device for an aircraft according to claim 1 or claim 2.

4. The thrust addition calculation unit is, If the determination by the determination unit determines that the eccentricity and the near-point argument are maintained within the range of the target values, and the difference between the target semi-major axis (which is the target value of the semi-major axis) and the semi-major axis exceeds a preset threshold, In order to maintain the semi-major axis within a predetermined range, a plurality of symmetrical positions on the elliptical orbit to which thrust should be applied is calculated, and by applying thrust at these plurality of positions, the semi-major axis is increased while calculating a thrust that results in approximately zero change in the total change in the eccentricity and the argument of perpendicularity of the elliptical orbit. A trajectory control device for an aircraft according to claim 1 or claim 2.

5. A method for controlling the trajectory of an aircraft orbiting on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in the region below 400 km altitude where atmospheric drag is dominant, The steps include obtaining orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee of the flying object while it is in flight along the elliptical orbit, The steps include: calculating the difference between the acquired orbital parameter and the target orbital parameter, which is the target value, and determining whether the difference exceeds a predetermined threshold; Based on the result of the determination step, the steps include calculating the position on the elliptical orbit where thrust should be added to the aircraft, and the thrust to be added at that position, such that the semi-major axis, the eccentricity, and the perpendicularity argument are maintained within a predetermined range; The process includes a step of controlling the output of the aircraft based on the calculation result of the calculation step, The step of calculating the thrust to be applied to the aforementioned aircraft is: In the step of determining whether the difference exceeds a preset threshold, if the result of the determination shows that the difference between the target eccentricity (which is the target value of the eccentricity among the orbital parameters) and the actual eccentricity exceeds a preset threshold, the step of calculating the position where thrust should be applied in order to maintain the eccentricity within a predetermined range is a position near the peripole where the true angle of separation is approximately 0°. A method for controlling the trajectory of an aircraft.

6. A method for controlling the trajectory of an aircraft orbiting on an elliptical orbit in which the orbital inclination is maintained near the critical inclination and the peripoint is located in a region below 400 km altitude where atmospheric drag is dominant, The steps include obtaining orbital parameters consisting of the semi-major axis, eccentricity, and argument of perigee of the flying object while it is in flight along the elliptical orbit, The steps include: calculating the difference between the acquired orbital parameter and the target orbital parameter, which is the target value, and determining whether the difference exceeds a predetermined threshold; Based on the result of the determination step, the steps include calculating the position on the elliptical orbit where thrust should be added to the aircraft, and the thrust to be added at that position, such that the semi-major axis, the eccentricity, and the perpendicularity argument are maintained within a predetermined range; The process includes a step of controlling the output of the aircraft based on the calculation result of the calculation step, The step of calculating the thrust to be applied to the aforementioned aircraft is: In the step of determining whether the difference exceeds a preset threshold, if the result of the determination is that the absolute value of the difference between the target near-point argument, which is the target value of the near-point argument among the orbital parameters, and the near-point argument exceeds a preset threshold, If the difference between the target near-point argument and the near-point argument is a positive value, the step of calculating the position where thrust should be applied in order to increase the near-point argument and maintain it within a predetermined range is to be near the point where the cosine coincides with the negative value of the eccentricity, within the range of true near-point separation from 0° to 180°. If the difference between the target near-point argument and the near-point argument is a negative value, the method includes the step of calculating the position where thrust should be applied in order to reduce the near-point argument and maintain it within a predetermined range, such that the position is near the point where the cosine coincides with the negative value of the eccentricity, within the range of true near-point separation from 180° to 360°. A method for controlling the trajectory of an aircraft.

7. The step of calculating the thrust to be applied to the aforementioned aircraft is: When the correction of the semi-major axis is performed simultaneously with the correction of the eccentricity or the correction of the argument of the nearest point, the increase in the semi-major axis due to the correction of the eccentricity is calculated first, and / or the increase in the semi-major axis due to the correction of the argument of the nearest point is calculated first. A method for controlling the trajectory of an aircraft according to claim 5 or claim 6.

8. The step of calculating the thrust to be applied to the aforementioned aircraft is: If it is determined that the eccentricity and the argument of the near point are maintained within a predetermined range, and the difference between the target semi-major axis, which is the target value of the semi-major axis, and the semi-major axis exceeds a preset threshold, The steps include: calculating a plurality of symmetrical positions on the elliptical orbit to which thrust should be applied in order to maintain the semi-major axis within a predetermined range; and calculating a thrust that increases the semi-major axis while reducing the total change in the eccentricity and the argument of perpendicularity of the elliptical orbit to approximately zero by applying thrust at the plurality of positions. A method for controlling the trajectory of an aircraft according to claim 5 or claim 6.

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