Physical Characterization of Main-belt Comet (248370) 2005 QN173

We report results from new and archival observations of the newly discovered active asteroid (248370) 2005 QN173 (also now designated Comet 433P), which has been determined to be a likely main-belt comet based on a subsequent discovery that it is recurrently active near perihelion. From archival data analysis, we estimate g ′ -, r ′ -, i ′ -, and z ′ -band absolute magnitudes for the nucleus of H g = 16.62 ± 0.13, H r = 16.12 ± 0.10, H i = 16.05 ± 0.11, and H z = 15.93 ± 0.08, corresponding to nucleus colors of g ′ − r ′ = 0.50 ± 0.16 , r ′ − i ′ = 0.07 ± 0.15 , and i ′ − z ′ = 0.12 ± 0.14 ; an equivalent V-band absolute magnitude of H V = 16.32 ± 0.08; and a nucleus radius of r n = 1.6 ± 0.2 km (using a V-band albedo of p V = 0.054 ± 0.012). Meanwhile, we find mean near-nucleus coma colors when 248370 is active of g ′ − r ′ = 0.47 ± 0.03 , r ′ − i ′ = 0.10 ± 0.04 , and i ′ − z ′ = 0.05 ± 0.05 and similar mean dust tail colors, suggesting that no significant gas coma is present. We find approximate ratios between the scattering cross sections of near-nucleus dust (within 5000 km of the nucleus) and the nucleus of A d /A n = 0.7 ± 0.3 on 2016 July 22 and 1.8 < A d /A n < 2.9 in 2021 July and August. During the 2021 observation period, the coma declined in intrinsic brightness by ∼0.35 mag (or ∼25%) in 37 days, while the surface brightness of the dust tail remained effectively constant over the same period. Constraints derived from the sunward extent of the coma and width of the tail as measured perpendicular to the orbit plane suggest that the terminal velocities of ejected dust grains are extremely slow (∼1 m s−1 for 1 μm particles), suggesting that the observed dust emission may be aided by rapid rotation of the nucleus lowering the effective escape velocity.

Research on Task Satellite Selection Method for Space Object Detection LEO Constellation Based on Observation Window Projection Analysis

Aiming at the task planning and scheduling problem of space object detection LEO constellation (SODLC) for detecting space objects in deep space background, a method of SODLC task satellite selection based on observation window projection analysis is proposed. This method projects the spatial relative relationships of the SODLC observation blind zone, observation range, and the initial spatial position of the objects onto the surface of the earth for detectable analysis of satellites and targets and binds the dynamic observation conditions to the satellite trajectory after projection calculation of the visible relationship between target changes. On this basis, combined with the features of SODLC with high orbital symmetry, the task satellite selection is divided into two steps: orbit plane selection and task satellite selection. The orbit planes are selected based on the longitude range of the ascending node with the geographic location of the targets, and the task satellites are selected according to the relative motion relationship between the satellites and the targets together with the constraints of observable conditions. The selection method simplifies the calculation process of scheduling and selecting task satellites. Simulation analysis prove the method has better task satellite selection efficiency. The method has high practical value for task planning and scheduling for event-driven SODLC.

Effect of the Altitudes and Eccentricity of the Initial Orbit on Satellite Transition Efficiency

This research dealt with choosing the best satellite parking orbit and then the transition of the satellite from the low Earth orbit to the geosynchronous orbit (GEO). The aim of this research is to achieve this transition with the highest possible efficiency (lowest possible energy, time, and fuel consumption with highest accuracy) in the case of two different inclination orbits. This requires choosing a suitable primary parking orbit. All of the methods discussed in previous studies are based on two orbits at the same plane, mostly applying the circular orbit as an initial orbit. This transition required the use of the advanced technique of the Hohmann transfer method for the elliptical orbits, as we did in an earlier research, namely the transition from the perigee of the initial orbit to the final orbit and then conducting the rotation of the orbit plane to match the plane for the desired final orbit. The effect of the perigee altitude of the initial orbit on the transition efficiency calculated for the values ​​between 300 to 3000 km. It was found that increasing the altitude reduces the energy and fuel needed for transportation, but the time required for transportation increases, into account that the increased height of the initial or parking orbit also implies the requirement of higher energy to reach it. The effects of eccentricity (e) values of the initial orbit between 0.01 to 0.2 on the transition efficiency were calculated. It was found that the increase in (e) reduces the energy and fuel, but does not affect the time, required for transportation.

Gravity Field Estimation from GRACE Follow-On Data: Results from CSR Analyses

&lt;p&gt;At the time of presentation, nearly two years of flight data from the joint NASA/GFZ GRACE Folllow-On mission will have been collected. In this time, gravity field models have been produced using two independent inter-satellite tracking systems - the MWI and the LRI using radio and optical interferometry, respectively. The data have been analyzed over more than two complete cycles of the sun relative to the orbit plane, allowing a characterization of the environmental impacts on the flight data. Extended duration of analyses have also permitted an assessment of the GRACE-FO data relative to the corresponding GRACE data.&lt;/p&gt;&lt;p&gt;This poster presents the status and lessons learned from two years of estimation of Earth gravity field models from the GRACE-FO data at the science data system component at the University of Texas Center for Space Research.&lt;/p&gt;

Astrodynamical middle-frequency interferometric gravitational wave observatory AMIGO: Mission concept and orbit design

AMIGO is a first-generation Astrodynamical Middle-frequency Interferometric Gravitational Wave (GW) Observatory. The scientific goals of AMIGO are to bridge the spectra gap between first-generation high-frequency and low-frequency GW sensitivities: to detect intermediate mass BH coalescence; to detect inspiral phase and predict time of binary black hole coalescences together with binary neutron star & black hole-neutron star coalescences for ground interferometers; to detect compact binary inspirals for studying stellar evolution and galactic population. The mission concept is to use time delay interferometry (TDI) for a nearly triangular formation of three drag-free spacecraft with nominal arm length 10,000 km, emitting laser power 2–10 W and telescope diameter 300–500 mm. The design GW sensitivity in the middle frequency band is [Formula: see text] Hz[Formula: see text]. Both geocentric and heliocentric orbit formations are considered. All options have LISA-like formations, that is, the triangular formation is [Formula: see text] inclined to the orbit plane. For all solar orbit options of AMIGO, the first-generation TDI satisfies the laser frequency-noise suppression requirement. We also investigate for each option of orbits under study, whether constant equal-arm implementation is feasible. For the solar-orbit options, the acceleration to maintain the formation can be designed to be less than 15 nm/s2 with the thruster requirement in the 15 [Formula: see text]N range. AMIGO would be a good place to test the feasibility of the constant equal-arm option. Fuel requirement, thruster noise requirement and test mass acceleration actuation requirement are briefly considered. From the orbit study, the solar orbit option is the mission orbit preference. We study the deployment for this orbit option. After a last-stage launch from 300 km Low Earth Orbit (LEO), each S/C’s maneuver to an appropriate 2-degree-behind-the-Earth AMIGO formation in 95 days requires only a [Formula: see text]v of about 80 m/s.

An Analysis of Flight Dynamics of a Space Debris Collector Transferring from its Orbital Plane to the Orbital Plane of a Debris Fragment

A space debris collector and a debris fragment move along random noncoplanar orbits ranging from 400 to 2000 km in height. The space collector leaves the base station, transfers into the orbital plane of a debris fragment, aligns itself with and approaches the fragment, grabs it and returns to the base station. The execution time of the flight mission is 24 hours. This paper examines only the stage when the debris collector transfers from its orbital plane to the fragment’s orbital plane. Dampening is provided by repeated activation of a cruise propulsion unit with the thrust of no less than 20000 N and the fuel specific impulse of no less than 15000 m/s. An analysis of dynamics of the orbital flight is performed by numerically integrating the equations of orbital movement of the debris collector and the debris fragment using the fourth order Runge-Kutta methods. The change of the scalar product sign of the vector of the orbit area integral of the debris fragment and the radius-vector of the debris collector is the criterion for intersecting the final orbit plane. Fuel depletion and the nonsphericity of the Earth’s gravitational field in the second zonal harmonic are taken into account, and an example of the calculations is given. Convergence estimates for the integration procedure with regard to the final orbit inclination relative to the orbit’s eccentricity are provided.

Probabilistic Analysis of the Requirements for an Elliptical Orbit Displacement Goniometric Detection System

The article addresses the requirements for a goniometric system intended to detect displacements of the space body orbit, which measures angles at discrete moments of time with a random error. Two statistical hypotheses about the zero and specified nonzero displacement values are considered. The likelihood ratio statistics is obtained, and the information distance for it is determined, which characterizes the discrimination confidence level, and which depends on the measurement device accuracy s and discreteness Dt, and also on the orbit parameters. By specifying the orbits and discrimination confidence level, the accuracy s and discreteness Dt are determined through information discrepancy. The information discrepancy is calculated on a family of orbits unfavorable for the observer, the plane of which touches the orbit of the moving observer, and the touching point is the intersection point of the corresponding trajectories. To refine the information distance, a simple flat approximate motion-measurement model is substantiated, in which the true motion of the observer (outside the space body orbit plane) is replaced, without changing its speed, by motion in the space body orbit plane over the earth's sphere circumference. The justification is corroborated by calculating the errors. The obtained simple model is used for estimating the potential possibility of detecting the specified displacement. Tables and graphs for practically determining the discreteness and accuracy of the system are given. The appendix contains all formulas necessary for carrying out similar calculations. Examples show that from angular measurements it is possible to obtain quite reliable detection for the range of orbit parameters that are of practical interest.