ALBERT

Description: Solar electric propulsion (SEP) technology is truly at the "intersection of commercial and military space" as well as the intersection of NASA robotic and human space missions. Building on the use of SEP for geosynchronous spacecraft station keeping, there are numerous potential commercial and military mission applications for SEP stages operating in Earth orbit. At NASA, there is a resurgence of interest in robotic SEP missions for Earth orbit raising applications, 1-AU class heliocentric missions to near Earth objects (NEOs) and SEP spacecraft technology demonstrations. Beyond these nearer term robotic missions, potential future human space flight missions to NEOs with high-power SEP stages are being considered. To enhance or enable this broad class of commercial, military and NASA missions, advancements in the power level and performance of SEP technologies are needed. This presentation will focus on design considerations for the solar photovoltaic array (PVA) and electric power system (EPS) vital to the design and operation of an SEP stage. The engineering and programmatic pros and cons of various PVA and EPS technologies and architectures will be discussed in the context of operating voltage and power levels. The impacts of PVA and EPS design options on the remaining SEP stage subsystem designs, as well as spacecraft operations, will also be discussed.

Description: The NASA Space Technology Mission Directorate's (STMD) Green Propellant Infusion Mission (GPIM) Technology Demonstration Mission (TDM) is comprised of a cross-cutting team of domestic spacecraft propulsion and storable green propellant technology experts. This TDM is led by Ball Aerospace & Technologies Corp. (BATC), who will use their BCP- 100 spacecraft to carry a propulsion system payload consisting of one 22 N thruster for primary divert (DeltaV) maneuvers and four 1 N thrusters for attitude control, in a flight demonstration of the AF-M315E technology. The GPIM project has technology infusion team members from all three major market sectors: Industry, NASA, and the Department of Defense (DoD). The GPIM project team includes BATC, includes Aerojet Rocketdyne (AR), Air Force Research Laboratory, Aerospace Systems Directorate, Edwards AFB (AFRL), NASA Glenn Research Center (GRC), NASA Kennedy Space Center (KSC), and NASA Goddard Space Flight Center (GSFC). STMD programmatic and technology oversight is provided by NASA Marshall Space Flight Center. The GPIM project shall fly an operational AF-M315E green propulsion subsystem on a Ball-built BCP-100 spacecraft.

Description: NASA's goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Description: The COncurrent Multidisciplinary Preliminary Assessment of Space Systems (COMPASS) Team partnered with the Applied Research Laboratory to perform a NASA Innovative Advanced Concepts (NIAC) Program study to evaluate chemical based power systems for keeping a Venus lander alive (power and cooling) and functional for a period of days. The mission class targeted was either a Discovery ($500M) or New Frontiers ($750M to $780M) class mission. Historic Soviet Venus landers have only lasted on the order of 2 hours in the extreme Venus environment:temperatures of 460 degrees Centigrade and pressures of 93 bar. Longer duration missions have been studied using plutonium powered systems to operate and cool landers for up to a year. However, the plutonium load is very large. This NIAC study sought to still provide power and cooling but without the plutonium. Batteries are far too heavy but a system which uses the atmosphere (primarily carbon dioxide) and on on-board fuel to power a power generation and cooling system was sought. The resuling design was the Advanced Long-Life Lander Investigating the Venus Environment (ALIVE) Spacecraft (S/C) which burns lithium (Li) with the CO2 atmosphere to heat a Duplex Stirling to power and cool the lander for a 5-day duration (until the Li is exhausted). While it does not last years a chemical powered system surviving days eliminates the cost associated with utilizing a flyby relay S/C and allows a continuous low data rate direct to earth (DTE) link in this instance from the Ovda Regio of Venus. The five-day collection time provided by the chemical power systems also enables science personnel on earth to interact and retarget science - something not possible with an approximately 2-hour spacecraft lifetime. It also allows for contingency operations directed by the ground (reduced risk). The science package was based on that envisioned by the Venus Intrepid Tessera Lander (VITaL) Decadal Survey Study. The Li Burner within the long duration power system creates approximately 14000 W of heat. This 1300 degree Centigrade heat using Li in the bottom "ballast" tank is melted to liquid by the Venus temperature, drawn into a furnace by a wick and burned with atmospheric CO2. The Li carbonate exhaust is liquid at 1300 degrees Centigrade and being denser than Li drains into the the Li tank and solidifies. Since the exhaust product is a dense liquid no "chimney" is required which conserves the heat for the stirling power convertor. The Duplex Stirling provides about 300 W of power and removes about 300 W of heat from the avionics and heat that leaks into the 1-bar-insulated payload pressure vessel kept at 25 degrees Centigrade. The Na K radiator is run to the top of the drag flap.The ALIVE vehicle is carried to Venus via an Atlas 411 launch vehicle (LV) with a C3 of 7 km2/s2. An Aeroshell, derived from the Genesis mission, enables a direct entry into the atmosphere of Venus (-10 degrees Centigrade, 40 g max) and 6 m/s for landing (44 g) using a drag ring. For surface science and communication, a 100 WRF (WebEx Recording Format), X-Band 0.6-meter pointable DTE (Direct-to-Earth) antenna provides 2 kbps (kilobits per second) to DSN (Deep-Space Network) 34-meter antenna clusters.Table 1.1 summarizes the top-level details of each subsystem that was incorporated into the design. Cost estimates of the ALIVE mission show it at approximately $760M which puts it into the New Frontiers class.The ALIVE landed duration is only limited by the amount of Li which can be carried by the lander. Further studies are needed to investigate how additional mass can be carried, perhaps by a larger launcher and larger aeroshell.

Description: Imagine sailing across the hot plains of Venus! A design for a craft to do just this was completed by the COncurrent Multidisciplinary Preliminary Assessment of Space Systems (COMPASS) Team for the NASA Innovative Advanced Concepts (NIAC) project. The robotic craft could explore over 30 kilometers of the surface of Venus, driven by the power of the wind. The Zephyr Venus Landsailer is a science mission concept for exploring the surface of Venus with a mobility and science capability roughly comparable to the Mars Exploration Rovers (MER) mission, but using the winds of the thick atmosphere of Venus for propulsion. It would explore the plains of Venus in the year 2025, near the Venera 10 landing site, where wind velocities in the range of 80 to 120 centimeters per second (cm/s) were measured by earlier Soviet landing missions. These winds are harnessed by a large wing/sail which would also carry the solar cells to generate power. At around 250 kilograms (kg), Zephyr would carry an 8 meter tall airfoil sail (12 square meters area), 25 kg of science equipment (mineralogy, grinder, and weather instruments) and return 2 gigabytes of science over a 30 day mission. Due to the extreme temperatures (450 degrees Centigrade) and pressures (90 bar) on Venus, Zephyr would have only basic control systems (based on high temperature silicon carbide (SiC)electronics) and actuators. Control would come from an orbiter which is in turn controlled from Earth. Due to the time delay from the Earth a robust control system would need to exist on the orbiter to keep Zephyr on course. Data return and control would be made using a 250 megahertz link with the orbiter with a maximum data rate of 2 kilobits per second. At the minimal wind speed required for mobility of 35 cm/s, the vehicle move at a slow but steady 4 cm/s by positioning the airfoil and use of one wheel that is steered for pointing control. Navigation commands from the orbiter will be based upon navigation cameras, simple accelerometers and stability sensors; Zephyr's stability is robust, using a wide wheel base along with controls to "feather" or "luff" the airfoil and apply brakes to stop the vehicle in the case of unexpected conditions. This would be the science gathering configuration. The vehicle itself would need to be made from titanium (Ti) as the structural material, with a corrosion-barrier overcoating due to extreme temperatures on the surface.

Description: Near Earth Objects (NEOs) and other primitive bodies are exciting targets for exploration. Not only do they provide clues to the early formation of the universe, but they also are potential resources for manned exploration as well as provide information about potential Earth hazards. As a step toward exploration outside Earth's sphere of influence, NASA is considering manned exploration to Near Earth Asteroids (NEAs), however hazard characterization of a target is important before embarking on such an undertaking. A small Solar Electric Propulsion (SEP) spacecraft would be ideally suited for this type of mission due to the high delta-V requirements, variety of potential targets and locations, and the solar energy available in the inner solar system.Spacecraft and mission trades have been performed to develop a robust spacecraft design that utilizes low cost, off-the-shelf components that could accommodate a suite of different scientific payloads for NEO characterization. Mission concepts such as multiple spacecraft each rendezvousing with different NEOs, single spacecraft rendezvousing with separate NEOs, NEO landers, as well as other inner solar system applications (Mars telecom orbiter) have been evaluated. Secondary launch opportunities using the Expendable Secondary Payload Adapter (ESPA) Grande launch adapter with unconstrained launch dates have also been examined.

Description: NASA s goal for human spaceflight is to expand permanent human presence beyond low Earth orbit (LEO). NASA is identifying potential missions and technologies needed to achieve this goal. Mission options include crewed destinations to LEO and the International Space Station; high Earth orbit and geosynchronous orbit; cis-lunar space, lunar orbit, and the surface of the Moon; near-Earth objects; and the moons of Mars, Mars orbit, and the surface of Mars. NASA generated a series of design reference missions to drive out required functions and capabilities for these destinations, focusing first on a piloted mission to a near-Earth asteroid. One conclusion from this exercise was that a solar electric propulsion stage could reduce mission cost by reducing the required number of heavy lift launches and could increase mission reliability by providing a robust architecture for the long-duration crewed mission. Similarly, solar electric vehicles were identified as critical for missions to Mars, including orbiting Mars, landing on its surface, and visiting its moons. This paper describes the parameterized assessment of power and propulsion technologies for a piloted solar electric vehicle to a near-Earth asteroid. The objective of the assessment was to determine technology drivers to advance the state of the art of electric propulsion systems for human exploration. Sensitivity analyses on the performance characteristics of the propulsion and power systems were done to determine potential system-level impacts of improved technology. Starting with a "reasonable vehicle configuration" bounded by an assumed launch date, we introduced technology improvements to determine the system-level benefits (if any) that those technologies might provide. The results of this assessment are discussed and recommendations for future work are described.

Description: NASA's new Orion Crew Exploration Vehicle has geometry that orients the reaction control system (RCS) thrusters such that they can impinge upon the surface of Orion's solar array wings (SAW). Plume impingement can cause Paschen discharge, chemical contamination, thermal loading, erosion, and force loading on the SAW surface, especially when the SAWs are in a worst-case orientation (pointed 45 towards the aft end of the vehicle). Preliminary plume impingement assessment methods were needed to determine whether in-depth, timeconsuming calculations were required to assess power loss. Simple methods for assessing power loss as a result of these anomalies were developed to determine whether plume impingement induced power losses were below the assumed contamination loss budget of 2 percent. This paper details the methods that were developed and applies them to Orion's worst-case orientation.

Description: Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.