Claude R. Joyner

Propulsion System Choices and Their Implications

In defining a space vehicle architecture, the propulsion system and related subsystem choices will have a major influence on achieving the goals and objectives desired. There are many alternatives and the choices made must produce a system that meets the performance requirements, but at the same time also provide the greatest opportunity of reaching all of the required objectives. Recognizing the above, the SPST Functional Requirements subteam has drawn on the knowledge, expertise, and experience of its members to develop insight that will effectively aid the architectural concept developer in making the appropriate choices consistent with the architecture goals. This data not only identifies many selected choices, but also, more importantly, presents the collective assessment of this subteam on the “pros” and the “cons” of these choices. The propulsion system choices with their pros and cons are presented in five major groups. A. System Integration Approach. Focused on the requirement for safety, reliability, dependability, maintainability, and low cost.

Determining Mars Mission NTP Thrust Size and Architecture Impact for Mission Options

Future exploration missions across the Solar System need technologies that reduce the time of flight, provide efficient payload capability, and reduce the size and the number of launch systems in order to reduce mission risk and cost. NTP (Nuclear Thermal Propulsion) is the proven, high Technology Readiness Level technology, which provides the performance to enable rapid transit and can minimize the number of SLS launches due to the higher ISP (specific impulse). The future of human Mars exploration will see substantial benefit in terms of lower mission mass and faster trip times when NTP is employed. NTP has been proven scientifically and many of the engineering challenges have been addressed in past ground testing of reactor cores for a range of power levels needed for NTP systems. The challenge today is to create an affordable, highly capable in-space propulsion system. Aerojet Rocketdyne (AR) believes that this could be achieved based on using smaller reactors in the NTP designs (e.g., < 500 MWt) that spring-board off the knowledge gained from past research and development and applies new technologies to improve the life and provide eventually reusability. Also, mission architectures that have the local planetary exploration elements pre-deployed ahead of the human crew can have a significant impact on the design of the human NTP spacecraft, NTP power level (thrust size) and eventual NTP system reusability. The desire would be to have the human crewed vehicle as small as technically feasible, optimize the thrust size and optimize the number of engines based on mission need. This could drive the NTP design to have a smaller reactor, present a more affordable development plan, and lower cost operational footprint for future human Mars exploration transportation systems. Approaches to use a small NTP reactor core could provide a development cost benefit with less uranium content and be more easily tested with a smaller facility foot-print. The smaller facility and lower exhaust flow rate provides for less effluent to clean and manage, which, in turn, reduces the development cost due to environmental safety and nuclear material security concerns. Fundamentally a lower power NTP reactor core (< 500 MWt) can reduce the development, procurement, and operational costs making it a more affordable NTP system for a nuclear cryogenic propulsion stage. AR has been working on several NTP system designs that have a wide range of thrust (core) sizes and the scalability for any exploration mission. AR has performed various architecture and design studies from 2011 through 2015 that have identified NTP approaches for using the capability of smaller size NTP systems for robotic and human Solar System missions. AR has used our multidisciplinary design and architecture analysis capability to analyze a split cargo and crew approach where low to medium power (i.e., 100-150 KWe) Solar Electric Propulsion (SEP) pre-position mission cargo (e.g., long duration habitats, transfer stages, and lander systems) and the crew vehicle uses NTP for rapid transfer to Mars. This paper will focus on results of the study for the NTP stage Mars mission architecture element and the examinations of the thrust size and number of NTP engine systems and how they impact a human Mars mission. 1 Fellow, Mission Architecture, PO Box 109680, M/S 712-67, and AIAA Associate Fellow. 2 Program Manager, Advanced Programs, P.O. Box 7922 / MS RFB19, AIAA Member. 3 Sr. Engineer, Mission Architecture, 555 Discovery Dr.

Vehicle Architecture Study Using the Aerojet Rocketdyne Bantam Rocket Engine Family

fi = Inert Mass Fraction FLOX = Fluorinated Liquid Oxygen GPSM = General Purpose Stage Model GTO = Geosynchronous Transfer Orbit Isp = Specific Impulse, sec LEO = Low Earth Orbit LOX = Liquid Oxygen m = Mass, lbm MMH = Monomethyhdrazine NTO = Nitrogen Tetroxide O/F = Engine overall Oxidizer to Fuel Mixture Ratio POST = Program to Optimize Simulated Trajectories RP-1 = Rocket Grade Kerosene ρ = Density

Pros, Cons, and Alternatives to Weight Based Cost Estimating

Summary/Conclusions During architecture studies and conceptual design studies, details of any new items to be traded arerarely known: Part counts, features of individual parts, the manufacturing processes to be used, and eventhe weight are not easily obtainable. WBS details are not available without large expenditures of effort.What is available are the major choices of parameters and the choices of how to produce and how toprocure. Consequently, development and production cost models that do not need weight nor details aboutthe hardware are especially useful for studies at these levels. An ideal model would include factors thatallow examination not only ofthe change in cost due to design choices, but also the change in cost due tochanges in approaches to manufacturing, testing, and oversight.RECM and LLEGO are examples of such models. RECM has been incorporated by NASA intoNAFCOM. It has also been successfully used by PWR for many contractual and internal efforts. LLEGOhas been used by KSC.

Design Options for a Versatile Nuclear Thermal Propulsion (NTP) Stage

A study was initiated to investigate propulsion stage and mission architecture options potentially enabled by fission energy. One initial concept is a versatile Nuclear Thermal Propulsion (NTP) system with a maximum specific impulse of 900 s and a maximum thrust (per engine) of 15 klbf. The system assumes a monopropellant stage (hydrogen), and is designed to also provide 300 lbf of thrust (potentially split between multiple thrusters) at an Isp > 500 s. Boost pumps are used to assist with engine decay heat removal and low thrust engine burns, and to compensate for partial tank depressurization during full thrust engine burns. Potential stage assembly orbits that take full advantage of launch vehicle payload mass and volume capabilities are being assessed. The potential for using NTP engines to also generate a small to moderate amount of electrical power is also being evaluated.

Solar Electric Propulsion Architecture for Mars Cargo for Affordable Exploration and Sustained Permanence

Future exploration missions throughout the Solar System will require high efficiency propulsion technologies using moderate spacecraft power that are extensible to multiple mission areas and customers. Solar Electric Propulsion (SEP) has been proven at various power levels and propulsion unit design types and should be part of such a technology portfolio. Many of the SEP thrusters used for recent missions have established the high Technology Readiness Level for this technology. SEP’s high specific impulse (ISP) has a dramatic impact on reducing the required spacecraft propellant while increasing launch window flexibility, making the propulsion technology an ideal choice for delivering large unmanned cargo payloads for sustaining human Mars exploration. The future of human Mars exploration will see substantial benefit in terms of lower mission mass when SEP systems are employed in the architecture. It must be remembered that key design parameters (e.g., power, ISP, specific mass) need to be optimized in order to effectively deliver the highest payload mass possible within a given program. Architectures that have the local planetary exploration elements pre-deployed ahead of the human crew can have a significant impact on the design of the human exploration mission in terms of total mass available at Mars, the size of the crew spacecraft, and the number of total systems employed to create low risk transportation. This paper will discuss a recent Aerojet Rocketdyne (AR) study that examines the impact of SEP power levels, ISP, and stage size required for large (20+mt) cargo prepositioning missions to Mars orbit. The study also examines the impact on the crew vehicle size when different combinations of architecture elements are pre-positioned. The AR analysis has focused on finding the combination of right size launch capability and SEP vehicle power level that delivers large cargo but maintains reasonable trip times for pre-positioning. For near-term missions, we show that 50kWe SEP modules can be used either singly or in combination, enabling a path to lower cost human exploration missions. An approach that uses 40-50kWe SEP modules creates mission extensibility where these modules can also be used for cis-lunar missions and large science missions such as the Asteroid Redirect mission and geosynchronous Earth orbit commercial and military satellite delivery missions launched on existing expendable launch vehicles ensuring that the SEP modules have multiple applications. We show that this modular approach enables a dramatic reduction in total development and production costs for high power SEP and provides for a more gradual phasing of system development to fit within available budgets. 1 Fellow, Space Systems and Mission Analysis, PO Box 109680, M/S 712-67, AIAA Associate Fellow. 2 Executive Director, Space, 1300 Wilson Blvd., AIAA Associate Fellow. 3 Sr. Engineer, Mission Architecture, 555 Discovery Dr., AIAA Member. 4 Sr. Manager, Advanced Space and Launch, P.O. Box 7922 / MS RFB19, AIAA Member. 5 Executive Director, Advanced Space and Launch, 11411 139 Place, AIAA Fellow.

Mission Comparison of Hall Effect and Gridded Ion Thrusters Utilizing Various Propellant Options

A mission comparison of solar electric propulsion vehicles with Hall effect and gridded ion thrusters utilizing bismuth, xenon, and krypton propellant options is presented. Three example missions are examined: (1) geosynchronous transfer orbit to geostationary orbit, (2) high Earth orbit to the near Earth asteroid 2008 EA9, and (3) high Earth orbit to Mars orbit at the Mars moon Phobos. Trades of power level and specific impulse are performed for each mission, recommended power level and specific impulse ranges are provided, and recommendations for potential future work are provided.