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Ball Aerospace & Technologies Corp. (Ball) has conceptually developed a modular, reusable Solar Electric Propulsion (SEP) Tug that leverages heritage, commercial capabilities in spacecraft buses and electric propulsion (EP) to minimize the time and cost to market. The SEP Tug is a stand-alone, scalable, modular vehicle which combines a Mission Module based on commercially available spacecraft bus element from Ball, a SEP Module using high power electric propulsion, an Reaction Control System (RCS) Module based on monopropellant hydrazine (a green monopropellant option is under consideration) and a Payload Module/payload accommodation platform. SEP Module designs have examined implementation of both ion engine and Hall thruster systems and utilizes flexible blanket arrays. An optional Xenon Tank Module is available to accommodate larger propellant requirements. This modular architecture can satisfy the needs of the Asteroid Redirect Vehicle (ARV) and a broad array of customers and markets including: servicing, resupply, replenishment, payload delivery, operational orbit change and debris removal operations. The SEP Tug provides a reusable platform for industry to leverage for multiple mission needs while DOD benefits can include constellation resiliency.

Ball Aerospace & Technologies Corp participated in a Space Act Agreement with NASA GRC to determine the feasibility of accommodating enough Solar Electric Propulsion (SEP) on the Ball ESPA-class bus to result in a mission of interest to Ball customers. The baseline for the study was the ESPA-class BCP-100 bus. Since the BCP-100 bus has flight heritage on USAF programs, the approach for the study was to use the existing bus design and minimize changes to only those necessary to accommodate the SEP system. This approach maintains high heritage and minimizes the amount of Non-Recurring Engineering required for the bus. High heritage components were also selected for the SEP system when available, including an off-the-shelf Xenon tank, existing cathode, HET thruster and Xenon feed control, allowing future development funding to be focused on a PPU compatible with the existing BCP-100 28 V power bus. The results of the study show that while meeting the ESPA envelope and mass requirements, the BCP-100 can accommodate enough SEP capability to allow the orbit to be raised or lowered anywhere within LEO or change the inclination up to 10° from a LEO starting point. From a GTO starting point, an elliptical orbit with apogee at GEO is also possible.

The Solar Electric Propulsion (SEP) Technology Demonstration Mission (TDM) is a stepping stone leading to a reusable electric propulsion stage by demonstrating high power (10s kW) transfers from LEO to GEO and back to LEO. A scaled SEP TDM concept is described which can be launched on a Minotaur IV launch vehicle (LV). The mission concept is formulated in response to NASA’s BAA, meets all BAA requirements and demonstrates a modular and extensible solar electric propulsion system. It launches in early 2018 and flies multiple LEO-GEO transits over the 1-2 year-long operations period. The SEP TDM Space Vehicle (SV) is based on integrated SEP and Bus Modules allowing parallel development and efficient integration. The SEP Module includes two 5 kW Hall thruster strings (2 + 0) which can be operated singly or as a pair, or two NEXT strings which can be operated singly or as a pair. Advanced, light-weight, blanket solar array technology is employed for the SEP TDM instead of typically used, rigid panel technology. MegaFlex technology, using two 9 m-diameter wings, is baselined. The power and propulsion systems are at sufficient specific power to demonstrate the movement of large payloads from LEO to higher energy orbits at performance values consistent with future higher power electric propulsion capabilities (Isp, thrust-to-power, power-to-mass). The SEP TDM, and its SEP Module concept, represents a key infusion point to a reusable electric propulsion stage by executing a set of high ΔV trajectories to demonstrate long-term SEP operations, and fly the SEP TDM Space Vehicle through the radiation belts, sustained plasma environments, diverse distributed inertia spacecraft control environments and repeated spacecraft occultations.

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 green technology experts. The GPIM program has technology infusion-team members from all three major market sectors: Industry, NASA, and the Department of Defense (DoD). This team is led and managed by Ball Aerospace & Technologies Corp. (Ball), and includes Aerojet Rocketdyne (AR), Air Force Research Laboratory, Aerospace Systems Directorate, Edwards AFB (AFRL), NASA Glenn Research Center (GRC), and NASA Kennedy Space Center (KSC). STMD programmatic and technology oversight is provided by NASA Marshall Space Flight Center. The GPIM program shall fly an operational AF-M315E green propulsion subsystem on a Ball-built BCP-100 spacecraft.

The Earth's natural environment can impact the Space Vehicle and the operation of the Space Vehicle can impact the natural environment. The Solar Electric Propulsion (SEP) Technology Demonstration Mission (TDM) is an ideal opportunity to assess the impact of the natural environment on the Space Vehicle and assess the impact of SEP TDM Space Vehicle operations on the natural environment using a proper-ly scoped suite of instruments. Furthermore, the slow transit from LEO out to GEO and back provides an unique oppor-tunity to establish detailed map of Earth's natural environ-ment and detailed structure within the radiation belts during non-EP thrusting periods (all eclipse periods and using pur-poseful shut-downs during sunlight). An example of a low mass and low power space environments monitoring suite (SEMS), using largely existing instruments is described.

A 30 kW solar electric propulsion (SEP) technology demonstration mission (TDM) concept is described. The mission concept is formulated in response to NASA's BAA and demonstrates a modular and extensible solar electric propulsion system. It launches in early 2018 and flies multiple LEO-GEO transits over the 1–2 year-long operations period. The 30 kW SEP TDM Space Vehicle (SV) is based on integrated SEP and Bus Modules allowing parallel development and efficient integration. The SEP Module includes three 12 kW Hall thruster strings (3 + 0) which can be operated singly, in pairs or simultaneously for full power operations of all 3 together. Advanced, light-weight, blanket solar array technology is employed for the SEP TDM instead of typically used, rigid panel technology. MegaFlex technology, using two 10 m-diameter wings, is baselined. The power and propulsion systems are at sufficient specific power to demonstrate the movement of large payloads from LEO to higher energy orbits at performance values consistent with future higher power electric propulsion capabilities (Isp, thrust-to-power, power-to-mass). The SEP TDM, and its SEP Module concept, represents a key infusion point to a reusable electric propulsion stage by demonstrating transfers from LEO to GEO and back to LEO. This set of high ?V trajectories demonstrates long-term SEP operations and flies the SEP TDM Space Vehicle through the radiation belts, sustained plasma environments, diverse distributed inertia spacecraft control environments and repeated spacecraft occultations. The space vehicle hosts several secondary payloads to enhance science and technology return from the mission.

This paper describes an experimental platform that will demonstrate the major technologies required for the handling and storage of cryogenic propellants in a low-to-zero-g environment. In order to develop a cost-effective, high value-added demonstration mission, a review of the complete mission concept of operations (CONOPS) was performed. The overall cost of such a mission is driven not only by the spacecraft platform and on-orbit experiments themselves, but also by the complexities of handling cryogenic propellants during ground-processing operations.

The Multiband Imaging Photometer for Spitzer (MIPS) provides long wavelength capability for the mission, in imaging bands at 24, 70, and 160 microns and measurements of spectral energy distributions between 52 and 100 microns at a spectral resolution of about 7%. By using true detector arrays in each band, it provides both critical sampling of the Spitzer point spread function and relatively large imaging fields of view, allowing for substantial advances in sensitivity, angular resolution, and efficiency of areal coverage compared with previous space far-infrared capabilities. The Si:As BIB 24 micron array has excellent photometric properties, and measurements with rms relative errors of 1% or better can be obtained. The two longer wavelength arrays use Ge:Ga detectors with poor photometric stability. However, the use of 1.) a scan mirror to modulate the signals rapidly on these arrays, 2.) a system of on-board stimulators used for a relative calibration approximately every two minutes, and 3.) specialized reduction software result in good photometry with these arrays also, with rms relative errors of less than 10%.