An autonomous guidance method for a returning aircraft carried on an ultra-low orbit satellite platform

Abstract

The present invention discloses an autonomous guidance method for a returning aircraft carried on an ultra-low orbit satellite platform. The returning aircraft is arranged to be carried on the ultra-low orbit satellite platform, and the ultra-low orbit satellite platform is utilized for long-term stable operation. In the process of entering the braking orbit of the returning aircraft from the operating orbit, the ultra-low orbit satellite measures the atmospheric density in real time and calculates the estimated value of the atmospheric resistance of the returning aircraft. Two circles after the scheduled return time of the returning aircraft, the ultra-low orbit satellite releases the returning aircraft and transmits the estimated value of the atmospheric resistance to the returning aircraft. After separation, the returning aircraft adopts a combined method of inertial navigation orbit extrapolation and GNSS measurement correction for navigation, and autonomously corrects the ascending node and inclination errors in the return circle. The typical algorithm of the Lambert problem is used to solve the velocity increment at the first braking point and the second pulse point of the given return circle, and uses this as the initial value to output the velocity vector of the returning aircraft before and after the braking moment.

Description

Technical Field

[0001] The present invention belongs to the field of aerospace technology, and relates to the design and control of a recoverable aircraft, and specifically to an autonomous guidance method for a recoverable aircraft carried on an ultra-low-orbit satellite platform. Background Art

[0002] With the recent advancement of electric propulsion technology, the realization of ultra-low earth orbit (ULEO) satellites has become possible. An ULEO satellite generally refers to a satellite operating at an altitude between 300 km and 150 km. Deploying satellites in this orbit significantly reduces the distance between payloads and the ground, thereby improving payload efficiency and reducing satellite development and launch costs. If a traditional recoverable vehicle is integrated with an ULEO satellite, the low orbit of the ULEO satellite, close to the return orbit of the recoverable vehicle, can reduce the chemical fuel consumption required for braking, thus lowering costs. Furthermore, releasing a recoverable vehicle from an ULEO satellite platform offers greater flexibility and unpredictability, making its return process more abrupt and covert. Furthermore, ULEO satellites are significantly affected by atmospheric drag and frequently undergo low-thrust orbit maneuvers with electric propulsion. This results in significant fluctuations in orbital characteristics and a degree of stealth.

[0003] Prior Art 1 (Transition Phase Trajectory Design and Guidance for Orbital Bombardment Vehicles, Hu Zhengdong et al., Solid Rocket Technology, Vol. 32, No. 5, 2009) uses a Lambert guidance method in a reentry vehicle, compensating for J2 perturbations but failing to address atmospheric drag. Prior Art 2 (An Improved Algorithm for Exoatmospheric Trajectory Planning Under the Influence of J2 Perturbations, Wei Qian et al., Control Theory and Applications, Vol. 33, No. 9, 2016) analyzes the exoatmospheric trajectory planning problem under the influence of J2 perturbations and proposes an improved BP network prediction method based on the classic Lambert guidance theory. However, this method also fails to account for non-negligible issues such as atmospheric drag for ultra-low orbit vehicles. Consequently, the accuracy of reentry guidance for vehicles using prior art is limited. Summary of the Invention

[0004] In response to the shortcomings of the existing technology, the present invention proposes an autonomous guidance method for a returning aircraft carried on an ultra-low-orbit satellite platform. By utilizing the ability of ultra-low-orbit satellites to stay in low orbit for a long time, the method measures or estimates relatively accurate atmospheric density parameters, providing a reference for the autonomous guidance process of the returning aircraft, thereby improving its return guidance accuracy.

[0005] A method for autonomous guidance of a returning aircraft carried on an ultra-low-orbit satellite platform is described as follows:

[0006] Step 1: As the ultra-low-orbit satellite enters the reusable vehicle's braking orbit from its operational orbit, it measures atmospheric density in real time and calculates an estimated atmospheric drag value for the reusable vehicle. Two orbits before the reusable vehicle's scheduled return time, the ultra-low-orbit satellite releases the reusable vehicle and simultaneously transmits the estimated atmospheric drag value to the reusable vehicle.

[0007] Step 2: After separation, the returning vehicle uses a combination of inertial navigation orbit extrapolation and GNSS measurement correction for navigation.

[0008] Step 3: Autonomous correction of ascending node and inclination errors

[0009] S3.1. When the difference between the longitude of the ascending node of the return loop and the nominal orbit longitude is greater than the set threshold, the ascending node is automatically corrected when the vehicle first reaches the apogee. The control time point of the orbit correction is t l The ascending node of the return loop is at time t r , then the longitude deviation of the ascending node of the circle is returned ΔL(t r )for:

[0010]

[0011] Among them, ΔL(t l ) is t l The longitude deviation of the ascending node at the moment includes the effects of the initial orbit deviation, orbit control deviation and atmospheric parameter deviation on the longitude deviation of the ascending node. a0 is the semi-major axis of the nominal orbit, is the orbital semi-major axis decay rate, and Δa=a-a0, represents the semi-major axis deviation, a is the actual orbit semi-major axis, ω E is the Earth's rotational speed, R E is the equatorial radius of the Earth. l The semi-major axis deviation at the control time point for orbit correction.

[0012] Since the goal of orbit correction is to eliminate the longitude deviation of the ascending node of the return circle and ensure that the sub-satellite point passes through the predetermined landing site accurately in the return circle, let ΔL(t r )=0, then:

[0013]

[0014] By Δa l The orbit change increment Δv that eliminates the longitude deviation of the ascending node can be obtained l for:

[0015]

[0016] Where v0 is the flight speed of the nominal orbit.

[0017] s2.3. When the difference between the longitude of the ascending node of the return loop and the nominal orbit longitude is less than the set threshold, a combined correction of the ascending node residual error and the inclination error is performed. Specifically, the normal relative distance deviation ΔN and the normal relative velocity deviation ΔV are achieved by jointly correcting the orbit inclination deviation Δi and the orbit ascending node right ascension deviation ΔΩ. N Corrections:

[0018] Δi=i-i0

[0019] ΔΩ=Ω-Ω0

[0020] Where i and i0 are the current orbital inclination and the nominal orbital inclination, Ω and Ω0 are the current right ascension of the ascending node and the nominal right ascension of the ascending node, respectively. The velocity increment Δv of the joint correction of orbital inclination and orbital right ascension of the ascending node is h for:

[0021]

[0022]

[0023] Where h is the orbital moment of momentum, r is the distance from the center of the Earth, μ is the Earth's gravitational constant, and e is the actual orbital eccentricity. The latitude arguments of the two nodes are:

[0024] u1=arctan(ΔΩsin i0 / Δi)

[0025] u2=π+arctan(ΔΩsini0 / Δi)

[0026] Therefore, the relative distance deviation ΔN and relative speed deviation ΔV are represented by the “+-” sign in the normal direction of the orbital plane. N The control quantities are:

[0027]

[0028]

[0029] Preferably, the threshold value for the difference between the longitude of the ascending node of the return loop and the longitude of the nominal orbit of the returning aircraft is set to 0.08°.

[0030] Step 4: Calculation of return braking increment of returnable aircraft

[0031] Based on the given distances r1 and r2 between the first braking point and the second pulse point and the center of the earth in the return circle, as well as the real-time measured flight time between the first braking point and the second pulse point and the direction of movement of the initial point, the typical algorithm of the Lambert problem solves the velocity increments at these two positions.

[0032] The velocity increment obtained by using the typical algorithm of Lambert problem is used as the initial value, and the velocity increment of the two braking operations is corrected with t1, r1, V 01 is the initial state, t2, r2, V 02 To set the nominal target state of the return loop; use the orbit prediction calculation program to predict the state t2 at time t2, r 2pre 、V 02pre . Among them, t1, r1, V 01 are the time before braking, the position vector and the velocity vector of the returning aircraft respectively. 02 is the position vector and velocity vector of the returning aircraft at the moment after braking. It is required that the state after the ignition ends at time t2 reaches the nominal target state, so the optimization problem is described as:

[0033]

[0034] The returnable vehicle executes the first pulse at the apogee of the orbit, which can be measured and controlled. The second pulse is difficult to measure and control. In order to reduce the risk, the first pulse speed increment is designed to account for a large proportion, and the second pulse is a small increment mainly used for correction. Therefore, |V is set 01 |>0.7*(|V 01 |+|V 02 |), use Newton iteration to solve the above optimization problem, and get V 01 and V 02 The optimal solution of .

[0035] As a preferred method, the atmospheric density is estimated by using the ultra-low-orbit satellite flight data and sent to the returnable aircraft. The ultra-low-orbit satellite propulsion start-up data is used to estimate the theoretical value of orbital lift Δa within one orbit. F , and the actual orbit change value Δa during one orbit Real Find the difference, the difference is the orbital attenuation of the ultra-low-orbit satellite caused by atmospheric drag during one orbit, The estimated value of atmospheric drag F on ultra-low orbit satellites can be calculated air ,Depend on The atmospheric density can be estimated as ρ. Where T is the orbital period of the ultra-low orbit satellite, a is the semi-major axis of the ultra-low orbit satellite orbit, v is the flight speed of the ultra-low orbit satellite, m is the mass of the ultra-low orbit satellite, C is the mass of the ultra-low orbit satellite, and d is the drag coefficient, S is the frontal area of the ultra-low orbit satellite, v air is the speed of the ultra-low-orbit satellite relative to the airflow.

[0036] Preferably, the returning vehicle performs a one-step extrapolation using the orbital dynamics equation based on the velocity increment measured by the sensor, and the extrapolation formula is as follows.

[0037]

[0038] where x k-1 、y k-1 、z k-1 、 are the three-axis position and three-axis velocity of the returning aircraft in the inertial system at time k-1, x k 、y k 、z k 、 are the three-axis positions and three-axis velocities of the returning aircraft in the inertial system at the current moment obtained by extrapolation.

[0039] Preprocess the GNSS measurement data of the returnable aircraft to the current moment, and obtain the three-axis position and three-axis velocity measurement values x in the inertial system at the current moment. G 、y G 、z G 、 The navigation data obtained by extrapolating the orbital dynamics equation is corrected by constant gain filtering:

[0040]

[0041]

[0042] where x k 、y k 、z k 、 are the three-axis position and three-axis velocity of the returning vehicle in the inertial system after filtering and correction at time k.

[0043] Preferably, when the reusable vehicle is operating in a near-circular orbit, after obtaining the absolute position and velocity information, the instantaneous orbital elements and short-period terms are further calculated to obtain the orbital quasi-flat elements. After the reusable vehicle completes the return braking pulse, the orbital eccentricity increases, and its orbital dynamics equations are numerically integrated.

[0044] As a preference, orbit prediction calculations are performed in an inertial system, and the center-of-mass dynamics equation is:

[0045]

[0046]

[0047] Where (x I ,y I , z I ) is the position of the returning aircraft in the inertial system, (v xI , v yI , v zI ) is the speed of the returning aircraft in the inertial system, (a x , ay , a z ) is the acceleration of the returning aircraft in the inertial system:

[0048]

[0049]

[0050] (a px , a py , a pz ) is the perturbation acceleration of the returning vehicle, (a ex , a ey , a ez ) is the non-spherical gravitational acceleration of the Earth, (a sx , a sy , a sz ) is the gravitational acceleration of the sun, (a mx , a my , a mz ) is the gravitational acceleration of the moon, (a jetx , a jety , a jetz ) is the acceleration generated by the jet thrust expressed in the inertial system, (a slx , a sly , a slz ) is the solar pressure perturbation acceleration, (a airx , a airy , a airz ) is the atmospheric drag perturbation acceleration, and the subscripts x, y, and z represent the components along the x-axis, y-axis, and z-axis in the inertial system.

[0051]

[0052]

[0053] Among them, (v xa , v ya , v za ) is the headwind velocity of the returning aircraft, k sma is the equivalent aerodynamic surface to mass ratio.

[0054] The present invention has the following beneficial effects:

[0055] 1. The recoverable aircraft is mounted on an ultra-low-orbit satellite platform, which can utilize the ultra-low-orbit satellite platform to operate stably in an ultra-low-orbit orbit for a long time. The proximity of the ultra-low-orbit orbit to the return orbit makes the return mission sudden and covert. The guidance process of the recoverable aircraft adopts a fully autonomous method, which has high military value.

[0056] 2. The returnable vehicle is carried on a very low-orbit satellite platform. The very low-orbit satellite can use its electric thrusters to control its orbit. Before the returnable vehicle returns, it can change its orbit to the return orbit (180-200km altitude). Before returning (the default is two orbits, and the minimum is 1.5 orbits in advance), the returnable vehicle is released, and the directly measured or estimated atmospheric density parameters are simultaneously injected into the returnable vehicle. This not only greatly reduces the fuel consumption for braking, but also achieves higher trajectory change accuracy. BRIEF DESCRIPTION OF THE DRAWINGS

[0057] Figure 1 Schematic diagram of ultra-low-orbit satellite and recoverable spacecraft;

[0058] Figure 2 Simplified diagram of the Lambert guidance principle. DETAILED DESCRIPTION

[0059] The present invention will be further explained below with reference to the accompanying drawings;

[0060] like Figure 1 As shown, the returnable vehicle is carried on a very low earth orbit (ULEO) satellite platform and together they enter a low earth orbit (LEO) of 500 km. The on-orbit experimental module and the returnable vehicle conduct in-orbit testing. The ULEO satellite platform then uses electric propulsion to enter an ULEO of 250 km along with the returnable vehicle. The on-orbit experimental module and the returnable vehicle conduct in-orbit testing. After the returnable vehicle completes its testing, the ULEO satellite platform, using electric propulsion, enters the returnable vehicle's braking orbit together with the returnable vehicle.

[0061] As the ultra-low-orbit satellite enters the return vehicle's braking orbit from its operational orbit, it measures atmospheric density in real time and calculates an estimated atmospheric drag for the return vehicle. Two orbits before the return vehicle's scheduled return time, the ultra-low-orbit satellite releases the return vehicle and simultaneously transmits the estimated atmospheric drag to it.

[0062] After separation, the return vehicle uses a combination of inertial navigation orbit extrapolation and GNSS measurement correction for navigation. In order to avoid long-term attitude control fuel consumption and ultra-low orbit altitude attenuation, the return vehicle needs to start its autonomous guidance process as soon as possible. First, it performs autonomous corrections to the orbital plane and the right ascension of the ascending node to eliminate the orbital plane errors of the return loop, and then performs Lambert guidance, such as Figure 2 As shown, two guidance velocity increments are calculated and implemented in the guidance algorithm using an orbit prediction model that includes J2 perturbation and atmospheric drag.

Claims

1. A method for autonomously guiding a recoverable vehicle mounted on a very low-orbit satellite platform, wherein the recoverable vehicle is mounted on the very low-orbit satellite platform and the electric thrusters of the very low-orbit satellite platform are used to perform very low-orbit orbit operation and orbit change, characterized in that: The return guidance method of the returning aircraft is as follows: Step 1: During the process of entering the braking orbit of the reusable vehicle, the ultra-low-orbit satellite measures the atmospheric density in real time and calculates an estimated atmospheric drag value for the reusable vehicle. At the scheduled return time of the reusable vehicle, the ultra-low-orbit satellite releases the reusable vehicle and simultaneously transmits the estimated atmospheric drag value to the reusable vehicle. The return time is two orbits later. Step 2: After separation, the returning vehicle uses a combination of inertial navigation orbit extrapolation and GNSS measurement correction for navigation; Step 3: When the difference between the longitude of the ascending node of the return circle and the nominal orbit longitude is greater than the set threshold, the returning vehicle performs an autonomous correction of the ascending node when it first reaches the apogee; the control time point of the orbit correction is set to t l , the ascending node of the return loop is at time t r , then the longitude deviation of the ascending node of the circle is returned ΔL(t r )for: Among them, ΔL(t l ) is t l The longitude deviation of the ascending node at time t includes the effects of initial orbit deviation, orbit control deviation and atmospheric parameter deviation on the longitude deviation of the ascending node; a0 is the semi-major axis of the nominal orbit, is the orbital semi-major axis decay rate, and Δa=a-a0, represents the semi-major axis deviation, a is the actual orbit semi-major axis, ω E is the Earth's rotational speed, R E is the equatorial radius of the Earth; Δa l Semi-major axis deviation at the control time point for orbit correction; Since the goal of orbit correction is to eliminate the longitude deviation of the ascending node of the return circle and ensure that the sub-satellite point passes through the predetermined landing site accurately in the return circle, let ΔL(t r )=0, then: By Δa l The orbit change increment Δv that eliminates the longitude deviation of the ascending node can be obtained l for: Where v0 is the flight speed of the nominal orbit; When the difference between the longitude of the ascending node of the return circle and the longitude of the nominal orbit of the returning vehicle is less than the set threshold, a combined correction of the ascending node residual error and the inclination error is performed; Step 4: Calculate the return braking increment of the returning aircraft based on the typical algorithm of the Lambert problem.

2. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform according to claim 1, characterized in that: When the difference between the longitude of the ascending node of the return circle and the nominal orbit longitude is less than the set threshold, the combined correction of the ascending node residual error and the inclination error is performed. Specifically, the normal relative distance deviation ΔN and the normal relative velocity deviation ΔV are achieved by jointly correcting the orbit inclination deviation Δi and the orbit ascending node right ascension deviation ΔΩ. N Corrections: Δi=i-i0 ΔΩ=Ω-Ω0 Where i and i0 are the current orbital inclination and the nominal orbital inclination, Ω and Ω0 are the current right ascension of the ascending node and the nominal right ascension of the ascending node, respectively; the velocity increment Δv of the joint correction of the orbital inclination and the right ascension of the ascending node is h for: Where h is the orbital moment of momentum, r is the distance from the center of the Earth, μ is the Earth's gravitational constant, and e is the actual orbital eccentricity; the latitude arguments of the two nodes are: u1=arctan(ΔΩsin i0 / Δi) u2=π+arctan(ΔΩsini0 / Δi) Therefore, the normal direction of the orbital plane is "+-" and the relative distance deviation ΔN and relative speed deviation ΔV are N The control quantities are:

3. The autonomous guidance method for a recoverable aircraft carried on an ultra-low-orbit satellite platform according to any one of claims 1 or 2, characterized in that: The set threshold value of the difference between the longitude of the ascending node of the return loop and the nominal orbit longitude of the returning vehicle is 0.08°.

4. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 1, characterized in that: According to the given distances r1 and r2 between the first braking point and the second pulse point and the center of the earth in the return circle, as well as the real-time measured flight time between the first braking point and the second pulse point and the movement direction of the initial point, the typical algorithm of the Lambert problem is used to solve the velocity increments at these two positions, and use them as the initial values to correct the velocity increments of the two brakings.

5. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 4, characterized in that: With t1, r1, V 01 is the initial state, t2, r2, V 02 To set the nominal target state of the return loop; use the orbit prediction calculation program to predict the state t2 at time t2, r 2pre 、V 02pre ; Among them, t1, r1, V 01 are the moment before braking, the position vector and velocity vector of the returning aircraft; t2, r2, V 02 is the moment after braking, the position vector and velocity vector of the returning aircraft; it is required that the state after the ignition ends at time t2 reaches the nominal target state. The optimization problem is described as: Set constraints |V 01 |>0.7*(|V 01 |+|V 02 |), use Newton iteration to solve the above optimization problem, and get V 01 and V 02 The optimal solution of .

6. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 1, characterized in that: Theoretical orbital lift value △a calculated from the electric propulsion startup data of ultra-low-orbit satellites within one orbit F , and the actual orbit change value △a during one orbit Real Find the difference, the difference is the orbital attenuation of the ultra-low-orbit satellite caused by atmospheric drag during one orbit, The estimated value of atmospheric drag F on ultra-low orbit satellites can be calculated air ,Depend on The atmospheric density estimate ρ is calculated; where T is the orbital period of the ultra-low orbit satellite, a is the semi-major axis of the ultra-low orbit satellite orbit, v is the flight speed of the ultra-low orbit satellite, m is the mass of the ultra-low orbit satellite, and C is the mass of the ultra-low orbit satellite. d is the drag coefficient, S is the frontal area of the ultra-low orbit satellite, v air is the speed of the ultra-low-orbit satellite relative to the airflow.

7. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 1, characterized in that: The returning vehicle uses the orbital dynamics equation to perform a one-step extrapolation based on the velocity increment measured by the sensor: where x k-1 、y k-1 、z k-1 、 are the three-axis position and three-axis velocity of the returning aircraft in the inertial system at time k-1, x k 、y k 、z k 、 are the three-axis position and three-axis velocity of the returning aircraft in the inertial system at the current moment obtained by extrapolation; Preprocess the GNSS measurement data of the returnable aircraft to the current moment, and obtain the three-axis position and three-axis velocity measurement values x in the inertial system at the current moment. G 、y G 、z G 、 The navigation data obtained by extrapolating the orbital dynamics equation is corrected by constant gain filtering: where x k 、y k 、z k 、 are the three-axis position and three-axis velocity of the returning vehicle in the inertial system after filtering and correction at time k.

8. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 1, characterized in that: When the returning vehicle is operating in a near-circular orbit, after obtaining the absolute position and velocity information, the instantaneous orbital elements and short-period terms are further calculated to obtain the quasi-flat orbital elements. After the returning vehicle completes the return braking pulse, the orbital eccentricity increases, and its orbital dynamics equation adopts the numerical integration method.

9. The autonomous guidance method for a recoverable aircraft mounted on an ultra-low-orbit satellite platform as claimed in claim 5, characterized in that: When performing orbit prediction calculation in an inertial system, the center-of-mass dynamic equation is: Where (x I ,y I ,z I ) is the position of the returning aircraft in the inertial system, (v xI ,v yI ,v zI ) is the speed of the returning aircraft in the inertial system, (a x ,a y ,a z ) is the acceleration of the returning aircraft in the inertial system: (a px ,a py ,a pz ) is the perturbation acceleration of the returning vehicle, (a ex ,a ey ,a ez ) is the non-spherical gravitational acceleration of the Earth, (a sx ,a sy ,a sz ) is the gravitational acceleration of the sun, (a mx ,a my ,a mz ) is the gravitational acceleration of the moon, (a jetx ,a jety ,a jetz ) is the acceleration generated by the jet thrust expressed in the inertial system, (a slx ,a sly ,a slz ) is the solar pressure perturbation acceleration, (a airx ,a airy ,a airz ) is the atmospheric drag perturbation acceleration, and the subscripts x, y, and z represent the components in the x-axis, y-axis, and z-axis directions in the inertial system; Among them, (v xa ,v ya ,v za ) is the headwind velocity of the returning aircraft, k sma is the equivalent aerodynamic surface-to-mass ratio.

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