Power and Propulsion Element (PPE) Spacecraft Reference Trajectory Document

This document captures example reference trajectories for the PPE including a reference delivery orbit and orbit maintenance, an example cislunar orbit transfer and end-of-mission (EOM) disposal trajectory. The flexibility of electric propulsion offers, by its low thrust nature, multiple different trajectory options to transfer from one orbit to another. The trajectories captured in this document are representative examples of a low thrust transfer from the NRHO and to multiple cislunar orbits. This document provides a consistent set of data from mission design to be used in the design of the vehicle capable of flying the trajectory described. The data in this document will be used to create conference papers. In order to do so, we are ending this document through for external release

Implicit Formulations of Bounded-Impulse Trajectory Models for Preliminary Interplanetary Low-Thrust Analysis

The bounded-impulse approach to low-thrust interplanetary trajectory optimization is widely used. In an effort to efficiently implement this approach using NASAs OpenMDAO optimization software, the authors have implemented implicit formulations of the forward shooting/backwards-shooting methods commonly used in bounded-impulse models. These implicit approaches allow for vectorization of the underlying calculations which can significantly reduce runtime in interpreted languages. An implicit approach may be either converged by using an underlying nonlinear solver to converge the state propagation, or as a constraint in an optimizer-driven multiple-shooting approach. Significant computational efficiency gains are realized through the utilization of the modular approach to unified derivatives. Further computational efficiency is achieved by capitalizing on the sparsity of the constraint Jacobian matrix. This work demonstrates that a vectorized multiple-shooting approach for propagating a state-time history is superior in terms of computational efficiency as the number of segments in the state-propagation is increased

Analysis of Cislunar Transfers from a Near Rectilinear Halo Orbit with High Power Solar Electric Propulsion

As government and commercial interest in the exploration of the Moon and cislu- nar space has grown, Near Rectilinear Halo Orbits (NRHOs) have shown to be of particular interest as staging orbits for human exploration of the Moon. Once in such staging orbits, low thrust solar electric propulsion (SEP) can enable efficient transfer to other orbits in cislunar space. This paper captures ongoing analysis to design efficient transfers of a massive spacecraft from a L2 Southern NRHO to a Distant Retrograde Orbit, L1 Northern NRHO, and Flat L2 Halo Orbit using low thrust SEP. For each transfer type, reference transfer is designed for an assumed 39 t spacecraft with 26.6 kW SEP system. For each reference transfer, analysis is completed to understand the sensitivity of the transfer to changes in initial mass and SEP power and identify the optimal number of thrusters to use for a given combination of mass and power

Comparison of Propulsion Options for Human Exploration of Mars

NASA continues to advance plans to extend human presence beyond low-Earth orbit leading to human exploration of Mars. The plans being laid out follow an incremental path, beginning with initial flight tests followed by deployment of a Deep Space Gateway (DSG) in cislunar space. This Gateway, will serve as the initial transportation node for departing and returning Mars spacecraft. Human exploration of Mars represents the next leap for humankind because it will require leaving Earth on a long mission with very limited return, rescue, or resupply capabilities. Although Mars missions are long, approaches and technologies are desired which can reduce the time that the crew is away from Earth. This paper builds off past analyses of NASA's exploration strategy by providing more detail on the performance of alternative in-space transportation options with an emphasis on reducing total mission duration. Key options discussed include advanced chemical, nuclear thermal, nuclear electric, solar electric, as well as an emerging hybrid propulsion system which utilizes a combination of both solar electric and chemical propulsion

Mission Design for the Exploration of Ice Giants, Kuiper Belt Objects and Their Moons Using Kilopower Electric Propulsion

The exploration of Ice Giants, Kuiper Belt Objects (KBOs) and their moons pose unique challenges from a mission design standpoint. NASA is currently developing a scalable 1-10 kilowatt electric class in-space fission reactor, known as Kilopower. The focus of this paper is to investigate the applicability of Kilopower Electric Propulsion to orbiting missions to Uranus, Neptune, and Pluto. This effort is broken into two pieces for each destination. First, a broad search of interplanetary trajectories with multiple gravity assists is completed to identify a range of mission opportunities from 2025 to 2045. Second, preliminary analysis is completed to understand the accessibility of various destination orbits, including elliptical orbits around the primary body and circular orbits around the largest moons. Results suggest that orbital missions to Uranus and Neptune are feasible with reasonable time of flight. Further work is necessary to achieve similar success with Pluto missions, but preliminary results are promising

Analysis of Near Rectilinear Halo Orbit Insertion with a 40-kW Solar Electric Propulsion System

This paper examines low thrust trajectories for delivery of a 40-kW solar electric propulsion spacecraft and potential additional payload to a desired NRHO. One option considered is a trans-lunar injection launch as a co-manifested payload on the Space Launch System. For this option, a reference trajectory is designed and a scan of launch dates is completed to understand the propellant mass sensitivity. A 15-day period cyclical variation in required propellant is observed that is attributed to solar gravity effects. A second option considered is to launch on a smaller commercial launch vehicle to a less energetic elliptical orbit and use SEP to spiral out to NRHO. For this option, analysis is completed to understand the trades between delivered mass to NRHO, total propellant required, time of flight, and solar array degradation. Results show that, while launching to lower altitudes can deliver greater payload mass to NRHO, significant solar array degradation can be observed

Parallel Monotonic Basin Hopping for Low Thrust Trajectory Optimization

Monotonic Basin Hopping has been shown to be an effective method of solving low thrust trajectory optimization problems. This paper outlines an extension to the common serial implementation by parallelizing it over any number of available compute cores. The Parallel Monotonic Basin Hopping algorithm described herein is shown to be an effective way to more quickly locate feasible solutions, and improve locally optimal solutions in an automated way without requiring a feasible initial guess. The increased speed achieved through parallelization enables the algorithm to be applied to more complex problems that would otherwise be impractical for a serial implementation. Low thrust cislunar transfers and a hybrid Mars example case demonstrate the effectiveness of the algorithm. Finally, a preliminary scaling study quantifies the expected decrease in solve time compared to a serial implementation

Use Of EOS-AURA Observation In The MERRA-2 Reanalysis

Meteorological reanalyses provide multi-year gridded datasets that describe the evolution of the atmosphere. Such products use a data assimilation system, comprising of an atmospheric model, a broad suite of observations, and an analysis system that optimally combines the model forecast with the observations, using an algorithm that includes information about model and data accuracy. The mixture of observations is of central importance to the quality of the assimilated datasets. The Modern-era Retrospective Analysis for Research and Applications (MERRA) included constraints on the thermal structure of the middle atmosphere from nadir sounders on the NOAA polar-orbiting platforms (Stratospheric Sounding Units and Advanced Microwave Sounding Units). These instruments have peak sensitivities that occur well below the stratopause. As such, the radiance measurements do not provide strong constraints on stratopause temperature. The new MERRA-2 reanalysis is using EOS-MLS temperature retrievals after they are available: it will be demonstrated that these data lead to a more realistic stratopause structure in MERRA-2 than in MERRA. Similarly, the work demonstrates the improvements in lower stratospheric ozone in MERRA-2 than in MERRA, for the period when EOS-MLS ozone data are assimilated. This improvement occurs because of the ozone profile information offered by MLS in the low stratosphere, in contrast to the SBUV/2 data used for the rest of MERRA-2. The impacts of choosing to use the EOS-MLS datasets are discussed in context of the continuity of the data record in MERRA- 2

Low Thrust Cis-Lunar Transfers Using a 40 kW-Class Solar Electric Propulsion Spacecraft

This paper captures trajectory analysis of a representative low thrust, high power Solar Electric Propulsion (SEP) vehicle to move a mass around cis-lunar space in the range of 20 to 40 kW power to the Electric Propulsion (EP) system. These cis-lunar transfers depart from a selected Near Rectilinear Halo Orbit (NRHO) and target other cis-lunar orbits. The NRHO cannot be characterized in the classical two-body dynamics more familiar in the human spaceflight community, and the use of low thrust orbit transfers provides unique analysis challenges. Among the target orbit destinations documented in this paper are transfers between a Southern and Northern NRHO, transfers between the NRHO and a Distant Retrograde Orbit (DRO) and a transfer between the NRHO and two different Earth Moon Lagrange Point 2 (EML2) Halo orbits. Because many different NRHOs and EML2 halo orbits exist, simplifying assumptions rely on previous analysis of orbits that meet current abort and communication requirements for human mission planning. Investigation is done into the sensitivities of these low thrust transfers to EP system power. Additionally, the impact of the Thrust to Weight ratio of these low thrust SEP systems and the ability to transit between these unique orbits are investigated

Overview of the Mission Design Reference Trajectory for NASA's Asteroid Redirect Robotic Mission

The National Aeronautics and Space Administration's (NASA's) recently cancelled Asteroid Redirect Mission was proposed to rendezvous with and characterize a 100 m plus class near-Earth asteroid and provide the capability to capture and retrieve a boulder off of the surface of the asteroid and bring the asteroidal material back to cislunar space. Leveraging the best of NASA's science, technology, and human exploration efforts, this mission was originally conceived to support observation campaigns, advanced solar electric propulsion, and NASA's Space Launch System heavy-lift rocket and Orion crew vehicle. The asteroid characterization and capture portion of ARM was referred to as the Asteroid Redirect Robotic Mission (ARRM) and was focused on the robotic capture and then redirection of an asteroidal boulder mass from the reference target, asteroid 2008 EV5, into an orbit near the Moon, referred to as a Near Rectilinear Halo Orbit where astronauts would visit and study it. The purpose of this paper is to document the final reference trajectory of ARRM and the challenges and unique methods employed in the trajectory design of the mission