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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.

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.

MicroSat Systems, Inc. (MSI) has developed a low cost, lightweight, solar array system using thin-film photovoltaic (TFPV) material to meet power generation needs for future responsive space missions. The Fold Integrated Thin Film Stiffener (FITS) is the deployment portion of the system. FITS is an integrated, passively deployed solar array structure designed specifically for TFPV, however a variety of photovoltaic (PV) options can be utilized by using the FITS deployment technology. FITS extends the boundaries of space PV systems by eliminating conventional rigid structures and mechanisms to maximize the lightweight and low stowage volume advantages of TFPV. FITS uses multifunctional, foldable components that store energy to provide deployment force and deployed stiffness, and have integrated power cabling to meet the demanding mass, cost and power requirements of programs like the TacSat series and anticipated future responsive space missions. MSI has completed the build and qualification test program for a two wing experimental solar array for the Air Force Research Laboratory (AFRL) TacSat-2 mission scheduled for launch in November of 2006. The array utilizes amorphous silicon (a-Si) thinfilm photovoltaics on a 1-mil stainless steel substrate from United Solar Ovonic (USOC), integrated with MSI’s patented FITS solar array deployment system. The experimental solar array will provide 120 W of additional power to the spacecraft on top of the primary solar arrays while providing valuable on-orbit performance data of the TFPV to the aerospace community for future mission planning. This characterization will be done by monitoring current, voltage, and temperature of the array wings over time using an I-V electronics box built by Lockheed Martin in Littleton, CO. Currently, MSI is under contract with AFRL to design, fabricate, and test a 380 W EOL FITS wing while focusing on the scalability and modularity of the FITS design. For the current program, MSI is designing a FITS wing consisting of four modular strings using USOC a-Si TFPV material on polymer substrate of approximately 95 W EOL each. This wing design will result in a 2 wing FITS solar array of 760 W EOL, however, because of the module array design, another string could easily be added to each wing resulting in a 950 W EOL array. This paper will discuss the current status of the design, build, and test of the TacSat-2 experimental FITS solar arrays and how the lessons learned from that program are being applied to the 380 W EOL FITS solar array design. It will outline the current status of the 380 W EOL FITS solar array as well as the benefits of the FITS solar array technology compared to state of the art conventional rigid arrays.

This paper presents an analytical, computational and experimental study of the deployment dynamics of an elastically deployable solar array. The thin film array is folded in multiple stages with elastic hinge and deployed depth stiffening elements and then allowed to deploy under its own elastic strain energy. A computational model of the geometrically nonlinear deployment is assembled using reduced order models of the elastic hinge elements. Restoring torque models developed for each hinge line are validated through isolated testing of each of three deployment stages. Both linear and nonlinear models of the corresponding elastic mechanisms are updated from the results of these experiments. The ground tests use a simple low-stiffness suspension system and videometry to measure the angular displacement of the deploying panel to be measured. The angular velocity and acceleration are computed from this data for use in the model updating process. The testing is performed on an early version of the array, called the engineering model (EM), and on the flight unit. For the first stage, the root hinge, the classical and computational models adequately predicted experimental results. The second stage, or z-fold deployment, experimental results indicated a stiffness that was 2.3 times smaller than the predicted stiffness for the EM. Flight unit results more closely matched predictions. For the final stage of deployment, the tri-fold, the nonlinear elastic response of the shallow shell, or "tape" hinges, is predicted by analytical models of their post-buckled mechanics. For this stage, the observed long range stiffness is greater than the expected results based on thin plate theory and computational simulation. Further testing of flight unit components and updating model parameters will improve the understanding of the deployment dynamics of the array.