Structural Design Considerations for a 50 kW-Class Solar Array for NASA’s Asteroid Redirect Mission Thomas W. Kerslake1, Thomas G. Kraft2, John T. Yim3 and Dzu K. Le4 NASA Glenn Research Center, Cleveland, Ohio, 44135 NASA is planning an Asteroid Redirect Mission (ARM) to take place in the 2020s. To enable this multi-year mission, a 40 kW class solar electric propulsion (SEP) system powered by an advanced 50 kW class solar array will be required. Powered by the SEP module (SEPM), the ARM vehicle will travel to a large near-Earth asteroid, descend to its surface, capture a multi-metric ton (t) asteroid boulder, ascend from the surface and return to the Earth-moon system to ultimately place the ARM vehicle and its captured asteroid boulder into a stable distant orbit. During the years that follow, astronauts flying in the Orion multipurpose crew vehicle (MPCV) will dock with the ARM vehicle and conduct extra- vehicular activity (EVA) operations to explore and sample the asteroid boulder. This paper will review the top structural design considerations to successfully implement this 50 kW class solar array that must meet unprecedented performance levels. These considerations include beyond state-of-the-art metrics for specific mass, specific volume, deployed area, deployed solar array wing (SAW) keep in zone (KIZ), deployed strength and deployed frequency. Analytical and design results are presented that support definition of stowed KIZ and launch restraint interface definition. An offset boom is defined to meet the deployed SAW KIZ. The resulting parametric impact of the offset boom length on spacecraft moment of inertias and deployed SAW quasistatic and dynamic load cases are also presented. Load cases include ARM spacecraft thruster plume impingement, asteroid surface operations and Orion docking operations which drive the required SAW deployed strength and damping. The authors conclude that to support NASA’s ARM power needs, an advanced SAW is required with mass performance better than 125 W/kg, stowed volume better than 40 kW/m3, a deployed area of 200 m2 (100 m2 for each of two SAWs), a deployed SAW offset distance of nominally 3-4 m, a deployed SAW quasistatic strength of nominally 0.1 g in any direction, a deployed loading displacement under 2 m, a deployed fundamental frequency above 0.1 Hz and deployed damping of at least 1%. These parameters must be met on top of challenging mission environments and ground testing requirements unique to the ARM project. Nomenclature ° = angular degree ACS = attitude control system ARM = asteroid redirect mission ARV = asteroid redirect vehicle B-frame = body frame °C or C = Celsius CAD = computer aided design 1 ARM SEP Module solar array lead, Power Architecture and Analysis Branch, 21000 Brookpark Rd., Mail Stop 142- 6. 2 Structures and Mechanisms Senior Engineer, Mechanisms and Tribology Branch, 21000 Brookpark Rd., Mail Stop 86-12, and AIAA Member. 3 Propulsion Engineer, In-Space Propulsion Systems Branch, 21000 Brookpark Rd., Mail Stop 301-3, and AIAA Member. 4 Structural Dynamics & Control Senior Engineer, Intelligent Control and Autonomy Branch, 21000 Brookpark Rd., Mail Stop 77-1, and AIAA Member. 1 American Institute of Aeronautics and Astronautics CG = center of gravity CM = center of mass conops = concept of operations CSI = controls structures interactions dB = decibel EP = electric propulsion eV = electron volt EVA = extra-vehicular activity FEM = finite element model g = acceleration due to gravity GSE = ground support equipment Hz = hertz IDS = international docking system ISS = international space station IXX = moment of inertia about the X axis Ixy = moment of inertia about the X-Y plane IYY = moment of inertia about the Y axis Iyx = moment of inertia about the Y-X plane Iyz = moment of inertia about the Y-Z plane Izy = moment of inertia about the Z-Y plane IZZ = moment of inertia about the Z axis KIZ = keep in zone kg = kilogram kW = kilowatt m = meter mm = millimeter μm = micrometer MBD = multibody dynamics MCR = mission concept review MOI = moment of inertia MPCV = multipurpose crew vehicle N = newton NDS = NASA docking system OD = outer diameter PV = photovoltaic RAMP2 = reacting and multi-phase 2 computer code PLIMP = plume impingement computer code RCS = reaction control system ROSA = roll out solar array SADA = solar array drive assembly SAW = solar array wing SCS = soft capture system Sec = second SEP = solar electric propulsion SEPM = solar electric propulsion module SLS = space launch system SR = structural reference t = metric ton TCS = thermal control system W = watt I. Introduction O N April 15, 2010, President Obama instructed NASA to develop the spacecraft and technologies needed to enable human exploration of a near Earth asteroid1. In response to this call, NASA plans to make use of a heavy lift 2 American Institute of Aeronautics and Astronautics launch vehicle, such as the Space Launch System (SLS) rocket under development to launch large payloads into space, and the Orion multipurpose crew vehicle (MPCV), being developed to carry human explorers on missions beyond low Figure 1. ARM spacecraft concept with mission module (right) and the SEPM (left). Earth orbit. NASA and its commercial partners are also developing solar electric propulsion (SEP) and power technologies that enable efficient in-space transportation and ultimately, the Asteroid Redirect Mission (ARM). ARM pulls together all of these elements, the SLS, Orion and SEP technologies to meet the President’s goal of human asteroid exploration by the 2020s. To enable this multi-year mission, a 40 kW class solar electric propulsion (SEP) system powered by an advanced 50 kW class solar array will be required. Using the solar electric propulsion module (SEPM) in-space propulsion system and mission module, the ARM vehicle, such as the concept shown in Figure 1, Figure 2. Two candidate SAW technology concepts shown side by side for comparison: the Roll Out Solar Array (ROSA), left, and MegaFlexTM, right. Ultimately, one SAW type (from competing advanced technologies) will be selected for both SAWs that will power the SEPM. will travel to a large near-Earth asteroid. Once there, the vehicle will descend to the asteroid surface, capture a multi- metric ton (t) class asteroid boulder, ascend from the surface and return to the Earth-moon system. At this phase of the mission, the ARM vehicle and its captured asteroid boulder will be placed into a stable distant Earth-moon orbit. During the years that follow, astronauts flying in the Orion multipurpose crew vehicle (MPCV) will dock with the ARM vehicle and conduct extra-vehicular activity (EVA) operations to explore and sample the asteroid boulder. Two solar array wings (SAWs) will comprise the advanced solar array, with possible technology options including, but not limited to, those recently funded by NASA2: the Roll Out Solar Array (ROSA) and the MegaFlexTM, as shown in Figure 2. Both of these options use flexible blankets on which the photovoltaic cells are mounted. Ultimately, only one advanced SAW technology option will be selected for both SAWs that power the SEPM. Each SAW offset boom is mounted to a single-axis solar array drive assembly (SADA). 3 American Institute of Aeronautics and Astronautics This paper will review the top structural design considerations to successfully implement this 50 kW class solar array (25 kW class SAW) that must meet unprecedented performance levels. These considerations include beyond state-of-the-art metrics for specific mass, specific volume, deployed area, deployed keep in zone (KIZ), deployed strength and deployed frequency. Analytical and design results are presented that support definition of stowed KIZ and launch restraint interface definition. Also presented, is an offset boom with parametric length required to meet the deployed SAW KIZ. The resulting impact of boom length on spacecraft moment of inertias and deployed SAW quasistatic and dynamic load cases is discussed. The SAW deployed strength and damping requirements arise from loads from ARM spacecraft thruster plume impingement, asteroid surface operations and Orion docking operations. II. General Considerations A. Mass At the Mission Concept Review (MCR) timeframe in early 2015, the mass allocation for each SAW, including offset boom, was 200 kg. This translates to a specific power value for the SAW of about 130 W/kg at beginning of life with a SAW power level of 26 kW at the SAW to SADA interface. Compared to a state-of-the-art rigid panel SAW meeting the same requirements, this specific power value is 2-3X higher. The requirement for low SAW mass will most likely drive the design solution to one of an advanced, flexible blanket SAW. B. Deployed SAW flexible blanket area At the ARM MCR timeframe, the spacecraft load power, dominated by the electric propulsion subsystem power draw, led to a required 52 kW solar array approximate power level at beginning of life. Using state of the art, triple junction solar cells with 29% conversion efficiency, the ARM SEPM application requires a SAW flexible blanket area of about 100 m2. This translates to a deployed MegaFlexTM SAW diameter of about 12.5 m and a ROSA SAW about 5 m wide by 24 m long. The flexible blanket area sizing accounts for the appropriate power performance loss mechanisms including harnessing voltage drop, SAW integration factors, SAW operational factors and all natural and induced environmental degradation factors for the ARM mission. C. Others Based on the concept of operations (conops) for ARM, the SAW must be deployed autonomously without the assistance of ground operators or on-orbit crew. During deployed operation, the SAWs must achieve a high level of strength and stiffness (further discussed below). Given the SAW deployed loading orientation cannot be controlled following credible failure modes of the SADA, SAW deployed acceleration requirements must be both in-plane and out-of-plane directions. The SEPM coordinate system has the Y axis along Figure 3a. Stowed SAW KIZ – shown as orange regions. the longitudinal axis of the vehicle with the X- Z plane coincident with the SEPM to upper stage separation plane. SAWs are located in the direction of the +Z and –Z axes. See Figures 3a, 3b and 4 depicting this coordinate system. 4 American Institute of Aeronautics and Astronautics III. Stowed SAW Considerations D. Stowed SAW Keep in Zone Given a structural optimized SEPM is short and squat, much of the available launch fairing volume is filled up leaving only a narrow radial gap into which the SAWs must fit. At the SEPM module base, longitudinal length is limited by the upper stage to spacecraft structural adaptor support. At the SEPM top, longitudinal length is constrained Figure 3b. Stowed SAW KIZ dimensions in mm. by the ARM mission module structure. A preliminary drawing has been prepared showing the required KIZ configuration (orange colored volume, see Figure 3a) and dimensional constraints (see Figure 3b). To meet tight stowage volume requirements of the ARM SEPM application, SAW technologies tend to need specific volumes of about 40 kW/m3, or about 3X better than state of the art SAWs. Even if a SAW technology exceeds 40 kW/m3 metric, it may not be stowable on the SEPM given the combination of KIZ dimensional constraints, tie down location constraints (discussed in the next section) and kinematic deployment trajectory constraints. E. Stowed SAW Structural Interface To meet launch acceleration loading and stowed fundamental frequency requirements, massive and large dimension SAWs require many (perhaps even 8 or more) structural tie downs per SAW. The tie downs provide a load path to safely react SAW inertial loads into the spacecraft primary structure without stress and displacement 5 American Institute of Aeronautics and Astronautics Figure 4. Stowed SAW tie down structural interface keep in area for tie down attachments, dimensions in mm, +Z axis is out of the paper. exceedances. Given the SAW technology has not been selected, ARM SEPM designers specified a generalized SAW tie down structural configuration using secondary structure. Several options for the secondary structure have been assessed, but all must have a gossamer configuration. This feature is required to minimize view factor blockage for the SEPM heat rejection radiator surface located below. The allowable area for SAW tie downs is shown in Figure 4. The SADA interface with the SAW offset boom, envisioned as a circular bolt plate with approximately 0.2 m outer diameter, is located on the stowed SAW KIZ inner plane surfaces allocated for SAW tie downs (Figure 3b) and along the dashed line in Figure 4 (exact location under study). Given the stowed SAW tie down locations will not be generally located in optimum locations, SAW structures may incur a mass penalty associated with stiffer cores, thicker facesheets and/or the need for greater localized panel reinforcements. The stowed SAW KIZ, in combination with the SADA interface attachment plus deployed SAW KIZ, discussed below, will drive stowed offset boom dimensions and overall SAW deployment kinematics to accomplish the needed articulations and trajectories to transition from stowed, to phased, to fully deployed SAW configuration. F. Others A detailed coupled loads analysis has not been performed to define stowed SAW loads during launch. Until that time, preliminary requirements will include an ascent quasistatic acceleration of 20 g, acoustic excitation as shown in Figure 5 and random vibration power spectral density levels as shown in Figure 6. The 20 g quasistatic acceleration includes the launch vehicle mass acceleration curve design load as multiplied by a 1.25 load uncertainty factor and 2.0 distributed load factor to account for the multi-meter span of a stowed SAW. These environments encompass those anticipated for state of the art heavy lift vehicles and for the SLS under development by NASA. On top of these environments and design factors, a SAW to SEPM interface load factor of 1.5 must be applied for the design of the SAW interface components such as tie down mechanisms and their associated panel fittings. Stowed SAW shock spectrums, maximum levels that the SAW must accommodate from spacecraft events and the maximum levels the SAW can produce during tie down release, have been assessed and defined. However, these shock levels are not expected to be SAW design drivers and hence are not discussed further. 6 American Institute of Aeronautics and Astronautics Internal Acoustic Requirements for Spacecraft 145.0 140.0 135.0 )a P µ 0 2 e 130.0 r( B d ,le 125.0 v eL Protoflight (148.0 dB OASPL) e rus 120.0 Acceptance (145.0 dB OASPL) s e rP d n 115.0 u o S 110.0 105.0 100.0 10 100 1000 10000 One-Third Octave Band Frequency, Hz Figure 5. Ascent acoustic loading for stowed SAW. Random Vibration Requirement for Solar Array 1 Solar Array (Protoflight, 9.6 Grms) )z H 2g/ Solar Array (Acceptance, 6.8 Grms) ( y tis n e D la rtc 0.1 e p S n o ita re le c c A 0.01 10 100 1000 Frequency (Hz) Figure 6. Ascent random vibration loading levels. To avoid excessive coupling with ascent vibroacoustic loads, the stowed SAW fundamental frequency must exceed 7 American Institute of Aeronautics and Astronautics 25 Hz. This requirement would apply to dynamic modes with >10% mass participation and for those modes that drive local SAW stress levels. The additional assumption can be made that the SAW is mounted to an infinitely stiff spacecraft structure. It is expected that subsequent, system level structural dynamic assessments will be performed that include SEPM structure and SADA interface stiffnesses. SAW tie down area (see Figure 4) enforced displacements, with respect to the SADA mechanical interface, have been calculated. Results show displacements of +0.8/-1.1 mm in the X direction, -0.7 to -2.6 mm in the Y direction and Z direction displacements are shown in Figure 7. The SEPM cylindrical bus structure is very stiff leading to Figure 7. Stowed +Z side SAW tie down enforced displacements relative to the SADA along the Z axis (+Z out of the page). relatively small displacements compared to that expected with more traditional spacecraft structural topology. The stowed SAW must be designed to either take up these displacements within the structure or implement tie down flexures or attachments configured to release the required displacement degrees of freedom. For large SAWs that cannot use kinematic attachments to the bus, the small displacements of this stiff SEPM structure become all the more important. Some mission applications require strength for acceleration loading during SAW deployment prior to mechanism latching when the SAW may have reduced strength. As of now, the ARM vehicle will be in free drift mode during SAW deployment and does not have credible single fault scenarios resulting in vehicle acceleration (such as an inadvertent reaction control system (RCS) thruster firing). Thus, we foresee ARM SEPM SAWs will not be required to handle acceleration loading events during the brief period of SAW tie down release, phasing and deployment. This period is anticipated to be <20 minutes in total time. In addition, the deploying SAW must not extend outside the stowed SAW KIZ XY plane in the Z direction towards the SEPM centerline as show in Figures 3a and 3b. High power SAWs can have >100 strings of solar cells that leads to a large power harness that must be managed to avoid entanglement during deployment. The large power harness also introduces proportionally larger deployment parasitic torques compared to conventional smaller SAWs. 8 American Institute of Aeronautics and Astronautics IV. Deployed SAW Considerations In this section, deployed SAW structural considerations are discussed. These items are under assessment and technical study in an effort to properly formulate technical specifications for the ARM SEPM advanced SAW. A. Keep In Zone (KIZ) To accomplish the ARM, a deployed SAW KIZ (see Figure 8) is required to ensure the spacecraft and SAW meet functional operation requirements. This KIZ includes 360° of SADA rotation and SAW displacements from Figure 8. Deployed SAW KIZ, dimensions in mm. thermal/structural loading requirements discussed below. The KIZ starts at the SADA structural interface plane and extends outward from the SEPM along the Z axis. The KIZ includes a 1-m separation margin against interference with SEPM, mission module and docking/docked Orion MPCV spacecraft surfaces and appendages, asteroid/boulder surfaces and avoiding the electric propulsion (EP) thruster plume keep out zone. The EP plume keep out zone is defined by a cone on the thruster centerline with a 55° cone angle (45° plume angle plus 10° thruster gimbal angle) and the cone apex at the thruster exit plane. KIZ defining surfaces also include: a 0.4 m diameter cylinder with length between 1.0 m and 2.2 m for offset boom accommodation with the intention of minimizing view factor to the radiator panels adjacent to and behind the SADA and a high gain antenna minimum separation of 1.0 m. The KIZ was derived from numerous assessments including: (1) limiting xenon ion sputtering rates from the electric propulsion (EP) plume, (2) limiting thermal and structural loading from SEPM RCS chemical thruster plume impingement, (3) limiting the thermal control system (TCS) radiator to SAW view factor, (4) limiting thermal and structural loading from Orion RCS chemical thruster plume impingement during docking, (5) maintaining a safe 9 American Institute of Aeronautics and Astronautics separation of the SAW with ARM and Orion spacecraft surfaces during dynamic docking load disturbances, (6) maintaining a safe separation of the SAW with the asteroid during near surface operation dynamic loading disturbances, (7) minimizing SADA bending loads during dynamic SAW load events, (8) minimizing deployed SAW moment of inertias, (9) minimizing line-of-sight blockage of communication antennas, star trackers and cameras, (10) alignment error stack up in actual flight hardware and (11) separation distance margins. Many of these assessments are discussed below in greater detail. Even if meeting the 55° cone angle to limit EP plume impingement, SAW composites may still suffer resin material thickness losses of about 25 μm requiring the use of an extra outer ply or a thicker outer ply to maintain structural allowables. Similar material thickness losses from EP plume ion sputtering could occur on flexible blanket mesh materials. As such, thicker flexible blanket mesh products will be required with their attendant penalties in packaging volume efficiency and mass. Please note that the EP plume thrust level is so small (about 0.5 Newton (N)) that EP plume impingement results in negligible structural loading of a deployed SAW. However, the main EP plume should be avoided by deployed SAWs to limit unwanted disturbance momentum transfer to the spacecraft and to limit excessive xenon ion sputtering of SAW surfaces and the resulting molecular contamination produced. B. Moment of Inertial (MOI) The deployed SAW MOI dominates the full spacecraft MOI and as such, will greatly impact the attitude control system (ACS) performance in terms of agility (attitude angular acceleration) and dead band. Therefore, a study was completed to evaluate and quantify MOIs for ARM spacecraft (ARV) for various SAW configurations, offset boom lengths, mission phases and spacecraft configurations to better understand the effect these different variables had on the overall vehicle MOI. Specifically, the relative MOI contributions for the SAWs as compared to overall vehicle MOI was evaluated. Two different SAW architectures, ATK-Orbital’s MegaFlexTM and Deployable Space Systems (DSS) Roll Out Solar Array (ROSA), were evaluated for the study. Two SAW orientations were evaluated, SAW photovoltaic (PV) surface edge on and normal to flight path. In addition, the impact that the spacecraft xenon propellant tankage mass and captured asteroid boulder mass/size (29 t, 3 m diameter class), had on the relative MOI Figure 9. Ratio of ARV plus ROSA MOI divided by ARV plus MegaFlexTM MOI 10 American Institute of Aeronautics and Astronautics