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This paper discusses the processes which should be taken to establish debris indices that assess the orbital environmental impact of spacecraft or missions in orbit. The first process is establishing a methodology and criteria for evaluating an impact on the orbital environment. This study proposed an assessment method to evaluate the cumulative environmental changes during an evaluation period. It evaluated the impact of an uncontrolled LC mission abandoned in orbit. This study used Near-Earth Orbital Debris Environment Evolutionary Model (NEODEEM), an orbital environmental evolutionary model that Kyushu University and JAXA developed. The second process is the debris index formulation. This study considered four steps to establish the debris index. To reveal the establishing process of a debris index, this study used debris indexes as candidates which consider expected number of fragments generated from an object. The calculated debris index values were compared to the impact on the environment assessed by the evolutionary model to see if there was any correlation. Further, this study conducted numerical simulations of un-controlled large-constellation satellites abandoned in low earth orbit in different scenarios, and of active debris removals of large intact objects by NEODEEM. This paper discusses the applicability of proposed environmental assessment methods and the debris index. • Two processes to establish a debris index are proposed. • A methodology and criteria for evaluating environmental impacts is discussed. • Four steps to formulate a debris index is discussed. • Correlations between debris indexes and environmental impacts are compared. [ABSTRACT FROM AUTHOR]

Abstract: Over 9000 satellites and other trackable objects are currently in orbit around the Earth, along with many smaller particles. As the low Earth orbit is not a limitless resource, some sort of debris mitigation measures are needed to solve the problem of unusable satellites and spent upper stages. De-orbiting devices based on the use of conducting tethers have been recently proposed as innovative solutions to mitigate the growth of orbital debris. However, electrodynamic tethers introduce unusual problems when viewed from the space debris perspective. In particular, because of their small diameter, tethers of normal design may have a high probability of being severed by impacts with relatively small meteoroids and orbital debris. This paper compares the results obtained at ISTI/CNR, the Kyushu University and NASA/JSC concerning the vulnerability to debris impacts on a specific conducting tether able to de-orbit spacecraft in inclinations up to and initial altitude less than 1400km. A double line tether design has been analyzed, in addition to the single wire solution, in order to reduce the tether vulnerability. The results confirm that the survivability concern is fully justified for a single line tether and no de-orbit mission, from the altitudes and inclinations considered, is possible if the tether diameter is smaller than a few millimeters. The survival probability is shown to grow for a double line configuration with a sufficiently high number of knots and loops. The results are strongly dependent on the environment model adopted and the MASTER-2001 orbital debris and meteoroids fluxes result in survival probabilities appreciably higher than those of ORDEM2000 coupled with the Grün meteoroids model. [Copyright &y& Elsevier]

This paper examines the effect of shortening the post-mission disposal (PMD) orbit lifetime (remaining orbit lifetime after PMD) from 25 years to 5 or 1 year. Using NEODEEM, the debris evolutionary model developed by Kyushu University and JAXA, the change in the orbital environment at various altitude bands for long-term stability and short-term safety are discussed. It was confirmed that the PMD compliance rate is more important than shortening the orbit lifetime for improving the long-term stability of the orbital environment. In the short term, the 5- or 1-year rule reduced the collision rate and collision avoidance frequency at an altitude below 700 km, more than the 25-year rule, thus shortening the PMD orbit lifetime improves short-term safety, especially at low altitudes. The paper also evaluates the effect of large constellations (LCs) and it was shown that shortening the PMD orbit lifetime of LCs improves short-term safety, but the long-term impact is dominated by the number of LCs that remain at high operational altitudes due to PMD failure. It shows that the post-mission disposal lifetime and required compliance rate should be set according to the number and size of objects expected to be launched in the future. • The effect of shortening post-mission-disposal lifetime from 25 years are shown. • Short-term safety can be improved by 5-year rule or 1-year rule. • PMD compliance rate is more important for the long-term stability. • Difference between an elliptical and a circular orbit for PMD is shown. • PMD requirement changed according to the characteristics of the large constellation. [ABSTRACT FROM AUTHOR]

This study introduces a new method to estimate the orbital parameters (inclination, nodal precession rate, and right ascension of the ascending node) of the broken-up object using in-situ debris measurements. To reduce collision probability with tiny debris generated from breakups of satellites, it is required to identify the origin of breakups and predict the orbit of debris. In general, small debris cannot be tracked from the ground, thus an in-situ measurement is an effective method to detect sub-millimeter-sized debris. Previous studies developed a method to identify the location of breakup origin from an in-situ measurement, however, it can be applied to an ideal situation where the detection occurs at the intersection of the orbit planes. This study introduces a new method to estimate the orbital parameters by fitting the model of geocentric declination to the measurement data. This method can be applied to practical situations that cannot be validated in previous studies. This paper also adopts iteratively reweighted least squares to improve estimation accuracy and mitigate initial value dependencies. [Display omitted] • A New method to estimate the orbital parameters of the broken-up object is derived. • Origin identification of breakup events in practical situations is simulated and validated. • Estimation accuracy is improved by adopting iteratively reweighted least squares. [ABSTRACT FROM AUTHOR]

Abstract To date, there has been a lot of on-orbit satellite fragmentations since a Thor-Able upper stage exploded in space in 1961. Considering that even sub-millimeter-size debris are a big threat to operational satellites, then it can be said that detailed debris modelling is becoming an urgent task for sustainable space development. For accurate modelling, it is essential to know the accurate time of breakup. However, there is no effective method which can estimate the time precisely. Therefore, this study aims at constructing a new method for estimating the time of breakup precisely. This study focuses on orbital plane as a parameter and derives a new equation. This equation constraints orbital planes of fragments at breakup by using orbital elements of the broken-up object. The advantage of this equation is that the fragments can be evaluated by only inclination and right ascension of the ascending node which can guarantee relatively high accuracy even in a long-term propagation. For the estimation, this study proposes a hybrid method which combines the advantages of the constraint equation and the close approach analysis. As a result, by using officially reported data, it was confirmed that this hybrid method can estimate the time of breakup with an error of 1 min or less. It was also suggested that the difference in the estimated time between the hybrid method and the close approach analysis causes a large difference in predicted fragments. Highlights • A new constraint equation which can evaluate orbital planes of fragments is derived. • A hybrid method which combines the equation with a close approach analysis is proposed. • The hybrid method could estimate the time of breakup with an error of 1 min or less. [ABSTRACT FROM AUTHOR]

The in-situ debris environment awareness system has been developed. The objective of the system is to measure small debris (between 100 μm and several cm) in orbit. The orbital distribution and the size distribution of the debris are not well understood. The size distribution is difficult to measure from the ground, although the size distribution is very important for the risk evaluation of the impact of debris on spacecraft. The in-situ measurement of the size distribution is useful for: (1) verification of meteoroid and debris environment models, (2) verification of meteoroid and debris environment evolution models, (3) real time detection of unexpected events, such as explosions and/or collisions on an orbit. This paper reports the development study of the in-situ debris measurement system and shows demonstration experiments and their results to describe the performance of the micro-debris sensor system. The sensor system for monitoring micro-debris with sizes ranging from 100 μm to a few mm must have a large detection area, while the constraints of space deployment require that these systems be low in mass, low in power, robust and have low telemetry requirements. For this reason, we have been developing a simple trans-film sensor. Thin and conductive stripes (copper) are formed with fine pitch (100 μm) on a thin film of nonconductive material (12.5-μm thick polyimide). A hypervelocity micro-particle impact is detected when one or more stripes are severed by perforation of the film. We designed a debris detector specialized for measuring the micro-debris size and collision rate. We then manufactured and calibrated the detector. [ABSTRACT FROM AUTHOR]