Dynamical cartography of Earth satellite orbits

We have carried out a numerical investigation of the coupled gravitational and non-gravitational perturbations acting on Earth satellite orbits in an extensive grid, covering the whole circumterrestrial space, using an appropriately modified version of the SWIFT symplectic integrator, which is suitable for long-term ($～$120?years) integrations of the non-averaged equations of motion. Hence, we characterize the long-term dynamics and the phase-space structure of the Earth-orbiter environment, starting from low altitudes ($～$400?km) and going up to the GEO region and beyond. This investigation was done in the framework of the EC-funded “ReDSHIFT” project, with the purpose of enabling the definition of passive debris removal strategies, based on the use of physical mechanisms inherent in the complex dynamics of the problem (i.e., resonances). Accordingly, the complicated interactions among resonances, generated by different perturbing forces (i.e., lunisolar gravity, solar radiation pressure, tesseral harmonics in the geopotential) are accurately depicted in our results, where we can identify the regions of phase space where the motion is regular and long-term stable and regions for which eccentricity growth and even instability due to chaotic behavior can emerge. The results are presented in an “atlas” of dynamical stability maps for different orbital zones, with a particular focus on the (drag-free) range of semimajor axes, where the perturbing effects of the Earth’s oblateness and lunisolar gravity are of comparable order. In some regions, the overlapping of the predominant lunisolar secular and semi-secular resonances furnish a number of interesting disposal hatches at moderate to low eccentricity orbits. All computations were repeated for an increased area-to-mass ratio, simulating the case of a satellite equipped with an on-board, area-augmenting device. We find that this would generally promote the deorbiting process, particularly at the transition region between LEO and MEO. Although direct reentry from very low eccentricities is very unlikely in most cases of interest, we find that a modest “delta-v” ($\mathrm{\Delta }V$) budget would be enough for satellites to be steered into a relatively short-lived resonance and achieve reentry into the Earth’s atmosphere within reasonable timescales ($～$50?years).     

Low-thrust transfer to Backflip orbits

The aim of the work is to design a low-thrust transfer from a Low Earth Orbit to a “useful” periodic orbit in the Earth–Moon Circular Restricted Three Body Model (CR3BP). A useful periodic orbit is here intended as one that moves both in the Earth–Moon plane and out of this plane without any requirements of propellant mass. This is achieved by exploiting a particular class of periodic orbits named Backflip orbits, enabled by the CR3BP. The unique characteristics of this class of periodic solutions allow the design of an almost planar transfer from a geocentric orbit and the use of the Backflip intrinsic characteristics to explore the geospace out of the Earth–Moon plane. The main advantage of this approach is that periodic plane changes can be obtained by performing an almost planar transfer. In order to save propellant mass, so as to increase the scientific payload of the mission, a low-powered transfer is considered. This foresees a thrusting phase to gain energy from a departing circular geocentric orbit and a second thrusting phase to match the state of the target Backflip orbit, separated by an intermediate ballistic phase. This results in a combined application of a low-thrust manoeuvre and of a periodical solution in the CR3BP to realize a new class of missions to explore the Earth–Moon neighbourhoods in a quite inexpensive way. In addition, a low-thrust transit between two different Backflip orbits is analyzed and considered as a possible extension of the proposed mission. Thus, also a Backflip-to-Backflip transfer is addressed where a low-powered probe is able to experience periodic excursions above and below the Earth–Moon plane only performing almost planar and very short transfers.     

Penetration testing for the Optical Probe for Regolith Analysis (OPRA)

The Optical Probe for Regolith Analysis (OPRA) is a spike-shaped subsurface analytical probe that will be delivered to a planet, asteroid, or cometary body by a lander and/or rover. OPRA will be pushed down into the subsurface to record near infrared spectra as a function of depth down to maximum of 50 cm. Therefore, knowledge of the required penetration force to specific depths can be helpful in estimating the length of the probe. Test probes covering the anticipated diameter (2.5, 1.9, 1.2 and 0.9 cm diameter) and tip angle (T.A. = 30°, 60°, 90° and 120°) of OPRA were inserted mechanically into dry playground sand. The results showed that tip angle does not have a major effect, while probe diameter and density of the regolith are the most important parameters. Increasing probe diameter from 0.9 to 1.9 cm (i.e. a factor of 2) leads to an increase in penetration force from 200 to 1000 N (i.e. a factor of 5) at 20 cm depth. An increase in bulk density (B.D.) from 1550 to 1700 kg m⁻³ leads to an increase in penetration force from 10 to 200 N at 20 cm depth. Bearing capacity theory was used to explain the downward movement of the penetrometer through regolith and showed good agreement with the experimental results. This model was then used to take into account the effect of gravity on other planetary bodies. We observed a good agreement between the theoretical model and results from penetration testings on the Moon by the Apollo missions. Since the maximum allowed force for penetration is the weight of the lander/rover on their targeted planetary surface, our results put a strong constraint on the maximum reachable depth without endangering the whole mission.     

一种空间协同探测系统的星座设计方法

研究了由多颗合成孔径雷达卫星和一颗可见光卫星构成的空间协同探测系统的工作模式.根据太阳同步轨道和冻结轨道的特点,结合近地轨道遥感卫星的应用需求,选择了轨道半长轴、偏心率、倾角、近心点幅角.考虑到卫星偏航控制对覆盖性能的影响,在不动的地心坐标系中推导了卫星观测方向与地表交点的表达式,提出了确定系统中各颗卫星的升交点赤经和过近心点时刻的算法.给出了包含两颗合成孔径雷达卫星与一颗可见光卫星的协同探测系统的星座设计结果,并利用国际公认的卫星软件工具包Satellite ToolKit进行了验证,表明该设计方法是正确的.     

基于非开普勒轨道的高超声速临近空间飞行器自主天文导航研究

针对航天器自主导航方法不适合高超声速临近空间飞行器的问题, 研究了基于非开普勒轨道的高超声速临近空间飞行器自主天文导航方案. 论述了基于非开普勒轨道的自主天文导航机理, 通过对高超声速临近空间飞行器受力分析, 建立了动力学方程; 利用矢量倒数法则推导出空间运动方程; 设计了基于非开普勒轨道的状态模型和基于星光折射间接敏感地平的观测模型, 采用卡尔曼滤波进行了仿真验证. 仿真结果表明, 基于非开普勒轨道的高超声速临近空间飞行器自主天文导航可达到较高的位置和速度精度.      

X-ray pulsar-based navigation system with the errors in the planetary ephemerides for Earth-orbiting satellite

The objective of this paper is to investigate and reduce the impact of the errors in the planetary ephemerides on X-ray pulsar-based navigation system for Earth-orbiting satellite. Expressions of the system biases caused by the errors in the planetary ephemerides are derived. The result of investigation has shown that the impact of the error in Earth’s ephemeris is must greater than the errors in the other ephemerides and would greatly degrade the performance of X-ray pulsar-based navigation system. Moreover, the system bias is modeled as a slowly time-varying process, and is handled by including it as a part of navigation state vector. It has been demonstrated that the proposed navigation system is completely observable, and some simulations are performed to verify its feasibility.     

屏蔽材料对木星轨道卫星内部介质充电效应的影响研究

在木星轨道的空间辐射环境中,占主导地位的粒子是能量大于1 MeV(甚至高于100 MeV)的高能电子,这可能会产生卫星内部介质充电效应.在卫星的防辐射设计中,通常需要一定厚度的材料来屏蔽这些电子,使得进入卫星内部的电子通量达到安全的水平.利用所建立的GEANT4-RIC(radiation induced conduc...     

Inter-satellite link augmented BeiDou-3 orbit determination for precise point positioning

Precise Point Positioning(PPP) requires precise products, including high-accuracy satellite orbit and clock parameters. It is impossible to obtain an orbit solution that is sufficiently accurate for PPP services with a regional tracking network; therefore, satellite orbits are usually estimated by a global tracking network with a large number of ground stations. However, it is expensive to build globally distributed stations. Fortunately, BeiDou-3 satellites carry an InterSatellite Link(ISL) pay...     

Close relative motion on distant retrograde orbits

Distant Retrograde Orbits (DROs) in the Earth-Moon system have great potential to support varieties of missions due to the favorable stability and orbital positions. Thus, the close relative motion on DROs should be analyzed to design formations to assist or extend the DRO missions. However, as the reference DROs are obtained through numerical methods, the close relative motions on DROs are non-analytical, which severely limits the design of relative trajectories. In this paper, a novel approach is proposed to construct the analytical solution of bounded close relative motion on DROs. The linear dynamics of relative motion on DRO is established at first. The preliminary forms of the general solutions are obtained based on the Floquet theory. And the general solutions are classified as different modes depending on their periodic components. A new parameterization is applied to each mode, which allows us to explore the geometries of quasi-periodic modes in detail. In each mode, the solutions are integrated as a uniform expression and their periodic components are expanded as truncated Fourier series. In this way, the analytical bounded relative motion on DRO is obtained. Based on the analytical expression, the characteristics of different modes are comprehensively analyzed. The natural periodic mode is always located on the single side of the target spacecraft on DRO and is appropriate to be the parking orbits of the rendezvous and docking. On the basis of quasi-periodic modes, quasi-elliptical fly-around relative trajectories are designed with the assistance of only two impulses per period. The fly-around formation can support observations to targets on DRO from multiple viewing angles. And the fly-around formation is validated in a more practical ephemeris model.     

Asteroid interception and disruption for terminal planetary defense

Modern techniques for planetary defense from comets and asteroids involve the deflection of the bolide via kinetic, gravitational, ablative, or radiative means. While potentially effective, none of these methods are capable of operating in a terminal interdiction mode wherethe threat is discovered with little time prior to impact. We present a practical and effective method for planetary defense which enables extremely short interdiction time scales, but can also operate within longer time scales and can be effective for extremely large threats. Called PI (“Pulverize It”), the method makes use of an array of hypervelocity penetrators which uses the kinetic energy of the asteroid or comet to disrupt it. In the terminal interdiction mode, the fragments of maximum $～10$ m diameter disperse laterally as they continue towards the Earth, and then enter the Earth’s atmosphere where they burn up as a series of airburst events which spatially and temporally de-correlate the energy of the original parent bolide for any arbitrary observer on the ground in the form of acoustical shockwaves and optical pulses. We show that terminal interdiction modes ranging from 2 minutes prior to impact for 20-meter class bolides (such as the Chelyabinsk asteroid), 1 day prior to impact for 100 m-class asteroids, 10 days prior to impact for Apophis-class asteroids ($～370$ m), and even 60 days prior to impact for 1 km-class threats are all possible, though longer warning times are always preferred. Using only technologies readily available today, the PI method allows for a cost-effective and practical roadmap towards robust planetary defense capability.