Daniel J. H. Levack’s research while affiliated with Aerojet Rocketdyne and other places

Nuclear thermal propulsion (NTP) has been extensively researched as a potential main propulsion option for human Mars missions. NTP's combination of high thrust and high fuel efficiency makes it an ideal main propulsion candidate for these types of missions, providing architectural benefits including smaller transportation system masses, reduced trip times, increased abort capabilities, and the potential for transportation infrastructure reuse. Since 2016, AR has been working with NASA and members of industry as part of the NASA Space Technology Mission Directorate Game Changing Development Nuclear Thermal Propulsion Project. The overall goal of this project is to determine the feasibility and affordability of a low enriched uranium (LEU)-based NTP engine with solid cost and schedule confidence. Having shown feasibility and affordability, program planning has been underway for follow-on activities to continue to mature the LEU NTP engine technology. These activities include program planning for reactor fuels testing, reactor component design, engine component technology development, test facility design and demonstration, and a demonstration engine available for ground test and potentially flight test. These follow-on activities would set the stage for full scale development of a human rated NTP flight engine for use in human exploration missions. This paper presents details of a potential LEU NTP prototype flight test and corresponding first flight vehicle along with potential applications of an evolved vehicle for subsequent operational missions.

The future of human exploration missions to Mars is dependent on solutions to the technology challenges being worked on by the National Aeronautics and Space Administration (NASA) and industry. One of the key architecture technologies involves propulsion that can transport the human crew from Earth orbit to other planets and back to Earth with the lowest risk to crew and the mission. Nuclear thermal propulsion (NTP) is a proven technology that provides the performance required to enable benefits in greater payload mass, shorter transit time, wider launch windows, and rapid mission aborts due to its high specific impulse and high thrust. Aerojet Rocketdyne (AR) has stayed engaged for several decades in working NTP engine systems and has worked with NASA recently to perform an extensive study on using low-enriched uranium NTP engine systems for a Mars campaign involving crewed missions from the 2030s through the 2050s. Aerojet Rocketdyne has used a consistent set of NASA ground rules and they are constantly updated as NASA adjusts its sights on obtaining a path to Mars, now via the Lunar Operations Platform-Gateway. Building on NASA’s work, AR has assessed NTP as the high-thrust propulsion option to transport the crew by looking at how it can provide more mission capability than chemical or other propulsion systems. The impacts of the NTP engine system on the Mars transfer vehicle configuration have been assessed via several trade studies since 2016, including thrust size, number of engine systems, liquid hydrogen stage size, reaction control system sizing, propellant losses, NASA Space Launch System (SLS) payload fairing size impact, and aggregation orbit. An AR study activity in 2018 included examining NTP stages derived from Mars crew mission elements to deliver extremely large cargo via multiple launches or directly off the NASA SLS. This paper provides an update on the results of the ongoing engine system and mission trade studies.

Nuclear Thermal Propulsion (NTP) is a proven technology that provides the performance required to enable benefits in greater payload mass, shorter transit time, wider launch windows, and rapid mission aborts due to its high specific impulse (Isp) and high thrust. Aerojet Rocketdyne (AR) has stayed engaged for several decades in working NTP engine systems and has worked with NASA recently to perform an extensive study on using Low Enriched Uranium (LEU) NTP engine systems for a Mars campaign involving crewed missions from the 2030s through the 2050’s. AR has used a consistent set of NASA ground rules and they are constantly updated as NASA adjusts its sights on obtaining a path to Mars, now via the Lunar “Gateway.” Building on NASA work, AR has assessed NTP as the high-thrust propulsion option to transport the crew looking at how it can provide more mission capability than chemical or other propulsion systems. The impacts of the NTP engine system on the Mars transfer vehicle (MTV) configuration have been assessed via several trade studies since 2016, including thrust size, number of engine systems, liquid hydrogen stage size, reaction control system sizing, propellant losses, NASA Space Launch System (SLS) payload fairing (PLF) size impact, and aggregation orbit. AR study activity in 2018 included examining NTP stages derived from Mars crew mission elements to deliver extremely large cargo via multiple launches or directly off the NASA Space Launch System (SLS). This paper provides an update on the results of the on-going engine system and mission trade studies.

Future human exploration missions to Mars are being studied by NASA and industry. Several approaches to the Mars mission are being examined that use various types of propulsion for the different phases of the mission. The choice and implementation of certain propulsion systems can significantly impact mission performance in terms of trip time, spacecraft mass, and especially mission abort capability. Understanding the trajectory requirements relative to the round-trip Earth to Mars mission opportunities in the 2030’s and beyond is important in order to determine the impact of trajectory abort capability. Additionally, some propulsion choices for the crew vehicle can enable mission abort trajectories while others will most likely provide less flexibility and increase mission risk. This paper focuses on recent modeling of Earth to Mars abort scenarios for human missions to determine the capability to provide fast returns to Earth. The modeling assumed that the abort would occur after the Mars crew vehicle has been injected along the path to Mars (i.e., after the Trans Mars Injection (TMI) burn). These aborts have been defined as well as the timing of fly-by aborts to quickly return crew to Earth. These abort trajectory studies are based on missions NASA defined during the Evolvable Mars Campaign (EMC) with crew going to Mars in 2033, 2039, 2043 and 2048. Detailed trajectory analysis was performed with the NASA Copernicus program for the several crew missions that were in the EMC as well as other new missions being considered using finite-burn low thrust electric propulsion. The goal was to determine how the heliocentric trajectory elements change and the “abort trajectory” impulse requirements. Abort scenarios that were studied included fast returns N-days after TMI as well as fly-by aborts and multiple revolution cases, using all available propellants (e.g., main propulsion system and reaction control system (RCS)) to provide the required abort velocity change. Trajectories were investigated for impulsive maneuvers and for finite burn cases and the abort timelines for each are examined and compared. This paper and presentation will focus on the Copernicus trajectory analysis results that were performed to determine the abort trajectories that altered the primary mission to return to Earth as soon as possible.

Future human exploration missions to Mars are being studied by NASA, industry and academia. Many approaches to the Mars mission are being examined that use various types of propulsion for the different phases of the mission. The choice and implementation of propulsion system options can be optimally determined based on specific mission criteria such as: launch system, trip time, cargo or payload requirement at Mars, departure and arrival orbits, and mission campaign schedule. One way to reduce the size for a particular Mars mission architecture and possible cost of the mission is to split the cargo from the crew and preposition the cargo or payloads required over several trajectories or missions. This can reduce the size of the most expensive item, the Mars crew vehicle. This paper will discuss the current Solar Electric Propulsion (SEP) cargo stage sizing for Aerojet Rocketdyne (AR) Mars architecture work performed in 2017 that is examining the impact of SEP system design choices. Trades are presented varying SEP power and voltage, spanning various Space Launch System (SLS) departure orbits while taking into account the mission requirements across several Mars mission opportunities beginning in the late 2030's.