Practical guidelines for electro-dynamic tethers to survive from orbital debris impacts

Japan Aerospace Exploration Agency (JAXA) has proposed an active debris removal using electro-dynamic tether to reduce large space debris in the low-Earth orbit. However, a tether strand is thin but long enough to have a large area so that it is vulnerable to small particles. This vulnerability might be the weakest point of a tether system against orbital debris. In order to overcome this weakest point, a double tether system, in which two tether strands are tied together at even intervals to form equally spaced loops, has been suggested as one of the promising candidates. This paper provides a mathematical approach to estimate the survival probability of a double tether system and then apply the approach to evaluate the mission success rate of the active debris removal using electro-dynamic tether that JAXA has proposed. It can be concluded the countermeasure to get enough success rate can be obtained. The result is simulated for Advanced Earth Observing Satellite II (ADEOS-II) re-entry from 800 km sun synchronized orbit to atmosphere. The simulation shows that mission success rate over 90% can be obtained with number of loops over 1000 and 10 mm clearance between two strands.     

A deorbiter CubeSat for active orbital debris removal

This paper introduces a mission concept for active removal of orbital debris based on the utilization of the CubeSat form factor. The CubeSat is deployed from a carrier spacecraft, known as a mothership, and is equipped with orbital and attitude control actuators to attach to the target debris, stabilize its attitude, and subsequently move the debris to a lower orbit where atmospheric drag is high enough for the bodies to burn up. The mass and orbit altitude of debris objects that are within the realms of the CubeSat’s propulsion capabilities are identified. The attitude control schemes for the detumbling and deorbiting phases of the mission are specified. The objective of the deorbiting maneuver is to decrease the semi-major axis of the debris orbit, at the fastest rate, from its initial value to a final value of about 6471?km (i.e., 100?km above Earth considering a circular orbit) via a continuous low-thrust orbital transfer. Two case studies are investigated to verify the performance of the deorbiter CubeSat during the detumbling and deorbiting phases of the mission. The baseline target debris used in the study are the decommissioned KOMPSAT-1 satellite and the Pegasus rocket body. The results show that the deorbiting times for the target debris are reduced significantly, from several decades to one or two years.     

A flash on the moon caused by orbital debris

Sun glinting from a satellite in earth orbit is suggested as the probable cause of a reported lunar flash observed on May 23, 1985. This conflicts with the suggestion that a lunar ionization phenomenon is the source and points out concern that flashes from space debris must always be considered in the investigation of sky — flash discoveries.     

Analysis of tape tether survival in LEO against orbital debris

The low earth orbit (LEO) environment contains a large number of artificial debris, of which a significant portion is due to dead satellites and fragments of satellites resulted from explosions and in-orbit collisions. Deorbiting defunct satellites at the end of their life can be achieved by a successful operation of an Electrodynamic Tether (EDT) system. The effectiveness of an EDT greatly depends on the survivability of the tether, which can become debris itself if cut by debris particles; a tether can be completely cut by debris having some minimal diameter. The objective of this paper is to develop an accurate model using power laws for debris-size ranges, in both ORDEM2000 and MASTER2009 debris flux models, to calculate tape tether survivability. The analytical model, which depends on tape dimensions (width, thickness) and orbital parameters (inclinations, altitudes) is then verified with fully numerical results to compare for different orbit inclinations, altitudes and tape width for both ORDEM2000 and MASTER2009 flux data.     

Detection of small radar cross-section orbital debris with the Haystack radar

NASA has been making statistical measurements of the orbital debris environment for more than a decade using the MIT Lincoln Laboratory Haystack radar. The goal has been to characterize the environment for debris sizes as small as possible. Like all sensors which operate in the presence of noise, the Haystack radar has limited sensitivity. As the returned energy from small targets begins to approach the sensitivity limit of the radar, the probability-of-detection decreases, eventually approaching zero. The slope of the cumulative size distribution of debris begins to flatten out. This paper explores the possibility of extending the cumulative size distribution to smaller sizes by adjusting the distribution for probability-of-detection.     

Estimation of orbital parameters of broken-up objects from in-situ debris measurements

Even sub-millimeter-size debris could cause a fatal damage on a spacecraft. Such tiny debris cannot be followed up or tracked from the ground. Therefore, Kyushu University has initiated IDEA the project for In-situ Debris Environmental Awareness, which conducts in-situ measurements of sub-millimeter-size debris. One of the objectives is to estimate the location of on-orbit satellite fragmentations from in-situ measurements. The previous studies revealed that it is important to find out the right nodal precession rate to estimate the orbital parameters of a broken-up object properly. Therefore, this study derives a constraint equation that applies to the nodal precession rate of the broken-up object. This study also establishes an effective procedure to estimate properly the orbital parameters of a broken-up object with the constraint equation.     

Faster algorithm of debris cloud orbital character from spacecraft collision breakup

Space debris is polluting the space environment. Collision fragment is its important source. NASA standard breakup model, including size distributions, area-to-mass distributions, and delta velocity distributions, is a statistic experimental model used widely. The general algorithm based on the model is introduced. But this algorithm is difficult when debris quantity is more than hundreds or thousands. So a new faster algorithm for calculating debris cloud orbital lifetime and character from spacecraft collision breakup is presented first. For validating the faster algorithm, USA 193 satellite breakup event is simulated and compared with general algorithm. Contrast result indicates that calculation speed and efficiency of faster algorithm is very good. When debris size is in 0.01–0.05 m, the faster algorithm is almost a hundred times faster than general algorithm. And at the same time, its calculation precision is held well. The difference between corresponding orbital debris ratios from two algorithms is less than 1% generally.     

Cosmic dust and orbital debris: Collection on MIR space station

Upon the last joint Soviet-French mission on the MIR Space Station, on December 1988, an experiment devoted to the collection and detection of cosmic dust and space debris has been deployed in space during 13 months.

A variety of sensors and collecting devices has make possible the study of effects and distribution of cosmic particles after recovery of exposed material. Remnants of particles, suitable for chemical identification are expected to be found within the stacked foil detectors. Discrimination between true cosmic particles and man-made orbital debris is expected.

Some preliminary results are presented here.     

Re-entry of orbital debris observed by Chung-Li VHF radar with MEDAC

MEDAC (Meteor Echo Detection and Collection) system, a product of University of Colorado, has become part of Chung-Li VHF facilities since July 12, 1989. MEDAC is installed to observe the mesospheric winds from Doppler echos returned by meteor trails in the upper atmosphere. However, the time variations in the in-phase and quadrature components of the signals can be used to derive the time history of the meteor trail formation. The meteor flight speed in the atmosphere is hence deduced. Preliminary analysis of some data taken from July 12 to July 17 of 1989 indicates that there are some “meteor” trails that could have been produced by the reentry of orbital debris into the atmosphere. The criteria of the flight speed and the ionization height are used for selecting an orbital debris trail from pools of “meteor” trails. The relative flux intensities between the reentry orbital debris flux as tentatively identified in this paper and the meteor flux is about 1 to 100.     

The impact of alternative mission models on the future orbital debris environment

In reviewing discussions of future directions for space activity, it becomes obvious that there are a large number of groups formulating a wide diversity of plans for the future use of space. These plan alternatives are being made to account for user needs, technology development constraints, economic constraints, and launch support, and each of the plans will have direct or indirect effects on the orbital debris environment in terms of mass to orbit, deposition of operational debris, and control of accidental breakups. Thus it is important to develop the ability to project future debris states for a range of possible space traffic scenarios. The impact that these possible traffic environments would have on space operations forms the basis for studies of alternative options for the usage of space. In this paper, the effects on the orbital debris environment of a base-line mission model and two alternatives are investigated, using a numerical debris environment simulation code under development at JSC.     

Semi-analytical theory of mean orbital motion for geosynchronous space debris under gravitational influence

This paper provides a hamiltonian formulation of the equations of motion of an artificial satellite or space debris orbiting the geostationary ring. This theory of order 1 has been formulated using canonical and non-singular elements for eccentricity and inclination. The analysis is based on an expansion in powers of the eccentricity and of the inclination. The theory accounts for the influence of the Earth gravity field expanded in spherical harmonics, paying a particular attention to the resonance occurring for geosynchronous objects. The luni-solar perturbations are also taken into account. We present the resonant motion and its main characteristics: equilibria, stability, fundamental frequencies and width of the resonant area by comparison with a basic analytical model. Finally, we show some results concerning the long term dynamics of a typical space debris under the influence of the gravitational field of the Earth and the luni-solar interactions.     

Stabilization of an injected conducting layer for artificially enhancing drag on orbital debris

The evolution of a magnetized conducting medium suspended in magnetic and gravitational fields is examined. In this paper some effects of the influence of velocity fields on the linear stability properties of such layers are investigated. A fully compressible, three-dimensional analysis of the layer is described. The relevant equations are derived and then solved by the MagnetoHydroDynamic SPEctral Compressible Linear Stability (MHDSPECLS) algorithm, a Chebyshev collocation code. The code allows for the computation of magnetic and thermal effects. A complete stabilization of the system is found above a critical velocity of approximately 2500 m/s.     

MULPEX: A compact multi-layered polymer foil collector for micrometeoroids and orbital debris

Detailed studies of preserved hypervelocity impact residues on spacecraft multi-layer insulation foils have yielded important information about the flux of small particles from different sources in low-Earth orbit (LEO). We have extended our earlier research on impacts occurring in LEO to design and testing of a compact capture device. MUlti-Layer Polymer EXperiment (MULPEX) is simple, cheap to build, lightweight, of no power demand, easy to deploy, and optimised for the efficient collection of impact residue for analysis on return to Earth. The capture medium is a stack of very thin (8 and 40 μm) polyimide foils, supported on poly-tetrafluoroethylene sheet frames, surrounded by a protective aluminium casing. The uppermost foil has a very thin metallic coating for thermal protection and resistance to atomic oxygen and ultra-violet exposure. The casing provides a simple detachable interface for deployment on the spacecraft, facing into the desired direction for particle collection. On return to the laboratory, the stacked foils are separated for examination in a variable pressure scanning electron microscope, without need for surface coating. Analysis of impact residue is performed using energy dispersive X-ray spectrometers. Our laboratory experiments, utilising buck-shot firings of analogues to micrometeoroids (35–38 μm olivine) and space debris (4 μm alumina and 1 mm stainless steel) in a light gas gun, have shown that impact residue is abundant within the foil layers, and preserves a record of the impacting particle, whether of micrometer or millimetre dimensions. Penetrations of the top foil are easily recognised, and act as a proxy for dimensions of the penetrating particle. Impact may cause disruption and melting, but some residue retains sufficient crystallographic structure to show clear Raman lines, diagnostic of the original mineral.     

Identification of a debris cloud from the nuclear powered SNAPSHOT satellite with haystack radar measurements

Data from the Massachusetts Institute of Technology Lincoln Laboratory Long Range Imaging Radar (known as the Haystack radar) have been used in the past to examine families of objects from individual satellite breakups or families of orbiting objects that can be isolated in altitude and inclination. This is possible because, for some time after a breakup, the debris cloud of particles can remain grouped together in similar orbit planes. This cloud will be visible to the radar, in fixed staring mode, for a short time twice each day, as the orbit plane moves through the field of view. There should be a unique three-dimensional pattern in observation time, range, and range rate which can identify the cloud. Eventually, through slightly differing precession rates of the right ascension of ascending node of the debris cloud, the observation time becomes distributed so that event identification becomes much more difficult.     

The influence of orbital maneuver on autonomous orbit determination of an extended satellite navigation constellation

The right ascension of the ascending node is unobservable if only the inter-satellite ranging is used for autonomous orbit determination (AOD) of an Earth navigation constellation. However, if an Earth-Moon libration point satellite is added to the Earth navigation constellation to construct an extended navigation constellation, all the orbital elements can be determined with only the inter-satellite ranging. Furthermore, the extended navigation constellation can provide navigation information for interplanetary probes. For such an extended navigation constellation, orbital control needs to be considered due to the instability of the libration-point satellite orbit. This study concerns the influence of satellite orbital maneuver on the AOD of the extended navigation constellation. An AOD method under orbital maneuver is proposed. A low thrust controller is designed to achieve libration point satellite autonomous orbit maintenance by using AOD results. A navigation constellation consisting of three GPS satellites and one libration point satellite are designed for simulation. The simulation results show that libration point satellites can achieve autonomous navigation and autonomous orbit maintenance by only using inter-satellite ranging information. The rotation drift error of the Earth navigation constellation is also suppressed.     

Two-dimensional phase plane structure and the stability of the orbital motion for space debris in the geosynchronous ring

Based on the orbital resonance model, we study the two-dimensional phase plane structure of the motion of space debris orbiting the geosynchronous ring under the combined effects of the tesseral harmonics J₂₂, J₃₁ and J₃₃ of the Earth’s gravitational field. We present the main characteristic parameters of the two-dimensional phase plane structure. We also analyze the stability of the two-dimensional phase plane structure with numerical method. Our main findings indicate that the combined effects of the tesseral harmonics J₂₂, J₃₁ and J₃₃ fully determine the two-dimensional phase plane structure of the space debris, and it remains robust under the effect of the Earth’s actual gravitational field, the luni-solar perturbations and the solar radiation pressure with the normal area-to-mass ratios.     

Protecting Earth-orbiting spacecraft against micro-meteoroid/orbital debris impact damage using composite structural systems and materials: An overview

Spacecraft that are launched to operate in Earth orbit are susceptible to impacts by meteoroids and pieces of orbital debris (MMOD). The effect of a MMOD particle impact on a spacecraft depends on where the impact occurs, the size, composition, and speed of the impacting object, the function of the impacted system. In order to perform a risk analysis for a particular spacecraft under a specific mission profile, it is important to know whether or not the impacting particle (or its remnants) will exit the rear of an impacted spacecraft wall. A variety of different ballistic limit equations (BLEs) have been developed for many different types of structural wall configurations. BLEs can be used to optimize the design of spacecraft wall parameters so that the resulting configuration is able to withstand the anticipated variety of on-orbit high-speed impact scenarios. While the level of effort exerted in studying the response of metallic multi-wall systems to high speed particle impact is quite substantial, the extent of the effort to study composite material and composite structural systems under similar impact conditions has been much more limited. This paper presents an overview of the activities performed to assess the resiliency of composite structures and materials under high speed projectile impact. The activities reviewed will be those that have been aimed at increasing the level of protection afforded to spacecraft operating in the MMOD environment, and more specifically, on those activities performed to mitigate the mechanical and structural effects of an MMOD impact.     

Characterizing slope correction methods applied to satellite radar altimetry data: A case study around Dome Argus in East Antarctica

Slope correction is important to improve the accuracy of satellite radar elevation measurements by mitigating the slope-induced error (SE), especially over uneven ground surfaces. Although several slope correction methods have been proposed, guidance in the form of stepwise algorithm on how to implement these methods in processing radar altimetric data at the coding level, and the differences among these methods need to be presented and discussed systematically. In this paper, three existing types of slope correction methods—the direct method (DM), intermediate method (IM), and relocation method (RM, further divided into RM1 and RM2)—are described in detail. In addition, their main differences and features for various scientific applications are analyzed. We conduct a systematic experiment with CryoSat-2 Low Resolution Mode (LRM) data in a physically stable area around Dome Argus in East Antarctica, where in-situ measurements were available for comparison. The slope correction is implemented separately using the three methods, with the latest high-accuracy Reference Elevation Model of Antarctica (REMA) as the a-priori topography model. The bias and precision of the slope-corrected CryoSat-2 data results from the RM2 is ?0.18 ± 0.86 m based on the comparison with the field Global Navigation Satellite System (GNSS) data. The results from the RM2 indicate higher precision compared to those from the RM1. According to the correlation analysis of the slope-corrected CryoSat-2 data results (RM1 and RM2), the bias enlarges and the precision becomes worse when the surface slope increases from 0 to 0.85°. After a comprehensively comparative analysis, we find that the results from the RM1 and RM2 are superior in precision (0.93 m and 0.86 m) with respect to the GNSS data. The relatively low precision (1.22 m) from the IM is due to the potential error from the a-priori digital elevation model (DEM). The DM has the lowest precision (2.66 m). Another experiment over rough topography in West Antarctica is carried out for comparison, especially between the RM1 (precision of 15.27 m) and RM2 (precision of 16.25 m). In general, the RM is recommended for the SE elimination among the three methods. Moreover, the RM2 is firstly considered over smooth topography due to the superior performance in bias and precision, while the RM1 is more suggested over the rough topography because of the slightly smaller bias and better precision. The IM relies much on the accuracy of the a-prior DEM and is not usually recommended, because of the strict requirement in the sampling time between the radar altimetry data and the a-priori DEM to avoid any surface change over time.     

Non-Gaussian clutter characterization applied to OTHR using a mixture of two Rayleigh probability density functions

A practical technique for characterizing non-Gaussian radar clutter is specified and demonstrated using Over The Horizon Radar (OTHR) data, as an example. The technique employs maximum likelihood to fit the probability density of the clutter amplitude returns to a mixture of two Rayleigh probability densities instead of the single Rayleigh density typically used for Gaussian clutter. This model for non-Gaussian clutter is fully specified for any set of clutter amplitudes by a log likelihood, two Rayleigh parameters, and a mixing coefficient. A 3D plot of these values yields an easily-visualized clutter characterization, as is illustrated using OTHR data. This technique is a demonstration of clutter characterization using OTHR data, but the method can be applied to characterize other types of clutter data.