Timothy S. Kokan

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

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.

Low Cost Small LOX/HC Launch Vehicle Enabled by Affordable Propulsion

The commercialization of space depends on having affordable access to space. This paper presents an affordable launch vehicle for delivering a small payload to low earth orbit, specifically a launch vehicle to place up to 1,000 lbm in low earth orbit (100 NM, circular orbit, due east, 28.5°). The enabling element for the launch vehicle is affordable propulsion, both in development and in production.

Commercial and Civil In-Space Applications of the Peacekeeper Stage IV (RS-34)

The Peacekeeper Stage IV propulsion and structure hardware will enable low-risk, cost effective in-space stages to provide the basic spacecraft bus propulsion and structure capability to support rapid hardware development schedules for commercial and civil in-space applications. The Peacekeeper was a four-stage vehicle consisting of three solid rocket motors and the liquid-fueled Stage IV. The fourth stage utilized the Stage IV propulsion, guidance and control systems to deploy vehicles in sequence with very precise state vectors and attitudes. Stage IV used Nitrogentetroxide and Monomethylhydrazine storable hypergolic propellants. The system is a simple, reliable, regulated pressure-fed system, with a high pressure helium tank, two propellant tanks with integrated propellant management devices, one axial thruster for orbit transfer maneuvers, and eight vernier thrusters for precise attitude control of the stage for inorbit operations. The propellant tanks were developed with surface tension propellant management devices to enable infinite operational and maneuvering flexibility. A total of 141 Stage IV propulsion systems were integrated, tested and delivered by Aerojet Rocketdyne to support development, qualification, fielding and sustainment operations. The Peacekeeper missile system was decommissioned and the program closed-out in 2006. The remaining stages were separated from their missiles and placed in controlled storage at Hill Air Force Base. About twenty-five stages are currently available. Aerojet Rocketdyne was the prime contractor responsible for development, production, and support of the Peacekeeper Stage IV. The Peacekeeper Stage IV is also referred to as the RS-34. Aerojet Rocketdyne recently successfully concluded support of the Ares 1-X flight test program where two Peacekeeper Stage IV propulsion systems were reconfigured into the Roll Control System (RoCS) modules for the successful Ares 1-X flight test vehicle for NASA MSFC. This successful program demonstrates the feasibility of reconfiguring Stage IV propulsion systems for civil space applications. This paper presents study results showing the capabilities of the Peacekeeper Stage IV for the orbital debris removal, resupply to the ISS, and general space bus applications.

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.