High-fidelity trajectory design to flyby near-Earth asteroids using CubeSats

Abstract

Fast development of CubeSat technology now enables the first interplanetary missions. The potential application of CubeSats to flyby near-Earth asteroids is explored in this paper in consideration of CubeSats' limited propulsive capabilities and systems constraints. Low-energy asteroid flyby trajectories are designed assuming a CubeSat is initially parked around to the Sun-Earth Lagrange points. High-impulse and low-thrust trajectories with realistic thrusting models are computed first in the Circular Restricted Three-Body Problem (CR3BP), and then in a high-fidelity ephemeris model. Analysis in the ephemeris model is used to confirm that trajectories computed in the CR3BP model also exist in a more realistic dynamical model, and to verify the validity of the results obtained in CR3BP analysis. A catalogue of asteroid flyby opportunities between years 2019 and 2030 is provided, with 80 m/s of available ΔV and departure from halo orbits around the first and second Sun-Earth Lagrange points (of similar size to those typically used by scientific missions). Results show that the CR3BP model can serve as an effective tool to identify reachable asteroids and can provide an initial estimation of the ΔV cost in the ephemeris model (with ±15 m/s accuracy). An impulsive maneuver model can also provide an accurate estimation of the ΔV requirement for a CubeSat equipped with a high-impulse thruster (with 4 m/s accuracy), even if its thrust magnitude is small and requires duty cycling; low-thrust ΔV requirements, however, may differ significantly from the impulsive results (±15 m/s).

CubeSat Autonomous Navigation and Guidance for Low-Cost Asteroid Flyby Missions

“…The baseline mission scenario assumes a CubeSat is initially parked in a halo orbit around the first or the second Sun-Earth Lagrange points. Short transfer trajectories to near-Earth asteroids (<150 days) at ΔV costs that are compatible even with 3U platforms (<80 m/s) appear from the Sun-Earth L1/L2 points [32,33]. As such, future piggyback opportunities to L1/L2 present an interesting opportunity for the design of low-cost asteroid exploration missions.…”

“…This avoids numerical integration of equations of motion and complex optimization routines to time and compute TCMs. The high-impulse asteroid flyby trajectories designed in [32] are used here as reference, nominal trajectories (Fig. 1).…”

“…1). Flyby trajectories to 36 different asteroids from halo orbits around L1/L2 are presented in [32] (listed in Table 1); such asteroids are reached through up to two propulsive maneuvers and with less than 80 m/s in total (note that the same ΔV capability is considered here for the baseline scenario-Table 2). Targeted asteroids are all larger than ~100 meters in diameter (i.e., absolute magnitude H≤22.5), and their flyby trajectories are designed from 9 different departure points along the halo orbit: equispaced along one orbital period (e.g., Fig.…”

“…The baseline mission scenario leverages asteroid flyby trajectories previously computed in [32], which are compatible with current CubeSat propulsive capabilities. These flyby trajectories were designed from halo orbits around the first and second Sun-Earth Lagrange points, and a total of 36 asteroids can be reached with 80 m/s of ΔV between years 2019 and 2030 (suitable even for 3U CubeSats).…”

Novel 3U Stand-Alone CubeSat Architecture for Autonomous Near Earth Asteroid Fly-By

“…In Reference [47], a solar electric propulsion system, with Isp = 1200 s and thrust level at 1 AU equal to 0.006 N, was considered, showing small ∆V for more than 10 NEAs, and comparing them with rendezvous missions. In Reference [48], an electrospray thruster characterized by Isp = 2300 s and T = 100 µN was considered, and the mission feasibility with low ∆V (less than 85 m/s) was demonstrated for more than 40 targets (using as initial condition Sun-Earth Mean Barycenter L1 and L2). However, given the ∆V, which can be retrieved from literature, the rocket equation cannot be used to compute the required propellant mass, since it is based on a constant ejection velocity assumption.…”