Spacecraft bus stresses during launch. 

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... solar arrays designed for 10 years. Since the telescope systems will not be robotically serviceable, the components are generally 2-fault tolerant for a 20 year life. The power system is designed to provide 3.5kW for science instruments and 2kW of supplemental power for heating the primary mirror. When instruments are powered down, the EPS system will provide 4kW of power for heating the PM. The total observatory EPS provides 11.2kW of power with 30% margin. The SC battery is sized to provide 3 hours of power to the observatory prior to solar array deployment and during servicing operations as necessary. The batteries are 100Wh/kg Li-ion. Solar arrays are GaAs 100 W/kg triple junction. Total solar array area is 72 m 2 . The solar arrays are deployed on adjustable booms to counter solar pressure counter on the telescope tube; this balancing system can theoretically nullify all solar torque, providing unlimited observation times. The solar arrays are also doubled gimbaled to accommodate all slews and roll maneuvers, while maintaining the solar pressure balance. EPS elements are arranged in ORUs to allow for replacement during servicing missions, and incorporate blind mate connectors. While some solar array mechanisms will need to be space qualified, all EPS components are TRL 8. The passive spacecraft thermal control system consists of radiators, cold plate/heat pipe assemblies, heaters, and multilayer insulation. Figure 5 illustrates a notional thermal control scheme for the spacecraft. The thermal control system is required to maintain allowable operating temperatures for all subsystem and experiment components. The solar flux at an altitude of 1.5 x 10 6 km (consistent with an orbit about L2) is assumed to be 1296 W/m 2 . Radiators (30 m ) are body mounted around the circumference of the optical telescope assembly and spacecraft and will accommodate 4660 watts of heat dissipation. Spacecraft subsystems components and science instruments are mounted to cold plate/heat pipe assemblies in order to dissipate heat. The spacecraft will be wrapped in a 50 layer insulation blanket to help minimize the impact of solar heating on the interior (effective emissivity = 0.004). The outer layer of MLI is assumed to be 10 mil silverized Teflon for low solar flux absorbance. The objective is to maintain the on-orbit operational steady state temperature of the spacecraft interior at 300° K (±10° K). This requires 2500 watts of heat, supplied radiatively by science instruments and spacecraft components. The spacecraft propulsion system consists of a conventional pressure-fed bipropellant MMH-NTO system with two fuel tanks, two oxidizer tanks, and two pressurant tanks, all located inside the insulated spacecraft compartment. Table 7 provides details of the power consumption/heat dissipation for the spacecraft subsystems and science instruments. The spacecraft astrionics heat dissipation is assumed equal to the power requirement for the attitude control system (ACS), command and data system (CDS), instrumentation and monitoring (I&M), & communications (COMM) systems. The attitude control system will require 1630 watts, the command and data system will require 850 watts, the instrumentation and monitoring will require 806 watts, and the communications system will require 539 watts of power, on average to operate. The power system (batteries and power system enclosure) will dissipate an estimated 3835 watts and the science instruments will dissipate 3500 watts. A total of 11,160 watts must be managed by the spacecraft thermal control system. The primary mirror is to be maintained at 280 K, and is thermally managed by wrapping the support structure with a 50 layer MLI blanket. Active heating of the mirror requires up to 4000 watts, to be supplied as 2000 watts of heater power and scavenging up to 2000 watts of excess heat from the instruments via a heat pipe/cold plate assembly. The telescope MLI, heat pipes located behind the mirror, and the heaters and heater structure located behind the mirror are not budgeted with the spacecraft. The preliminary design thermal control mass summary for the ATLAST-8m is shown in Table 8. All thermal control components are commonly used in spacecraft design and considered to be a high TRL. Heat pipes are always custom designed to specifications and the integration of the heat pipes into the cold plate and primary mirror thermal control design will be a unique application and will require rigorous design and testing. During launch, the spacecraft is attached to the rear of the telescope structure, and does not support the observatory nor transfer launch loads to the Ares V launch vehicle (LV). Therefore, the structure is a lightweight and simple design, with aluminum-lithium plate elements and aluminum tubing being the main components. The SC Bus is hard-mounted to the observatory in sixteen locations, and supports all of the internal components through all aspects of the observatory’s mission. Figure 6 shows the SC Bus stress plot during launch. The SMS also supports SC Bus servicing by incorporating ORUs that contain most of the active SC Bus components. After separating from the LV, the SMS deploys the Solar Array (SA) and the steerable High Gain Antenna (HGA). Each of the two solar arrays is connected to a telescoping boom by an articulating mechanism, thus enabling the solar arrays to always be perpendicular to the sun but allowing their distance to the observatory’s center of mass to vary. This allows for optimal power from the arrays while utilizing the solar radiation pressure to help balance the solar radiation torque from the telescope’s optical tube. Except for the solar array deployment mechanisms, all SMS components are considered to be at TRL 6 or higher. For the solar array deployment, the telescoping boom and the articulation mechanism have both flown in space, but both have not been used simultaneously for solar array control. The ATLAST team estimates this component has a TRL of 5. An animation of the solar array deployment is included in Section 6. The S&MS masses are listed in Table 9. An earlier large monolithic telescope study very similar to the ATLAST 8-meter study showed that designing spacecraft components to survive 20 years in space is very expensive, and designing the spacecraft such that components can be replaced during servicing missions can greatly reduce the cost of the spacecraft. Servicing also allows replacement of science instruments as well as allowing technological upgrades to spacecraft components. Given the obvious benefits that servicing has had on the Hubble Space Telescope (HST), the study team decided to design the spacecraft to make servicing a viable option. While human servicing of spacecraft has been demonstrated with HST, robotic servicing is in the early stages of development. Nevertheless, the study team decided to assume that by the time that ATLAST is operational, robotic servicing will be a viable option. The general approach for servicing is to launch a servicing spacecraft aboard an EELV approximately every five years, bringing replacement spacecraft components and new science instruments to the observatory. The science instruments are designed to be inserted/removed from the rear of the instrument volume, and key spacecraft subsystem components or groups of components are configured in on-orbit replaceable units (ORUs) to allow removal from the rear of the spacecraft. The servicing spacecraft would use AR&D technology and dock with one of several passive docking stations on the observatory. A robotic arm, either part of the servicing spacecraft or fixed to the observatory, could facilitate servicing. Figure 7 shows a notional servicing concept for the observatory, with the robotic arm removing an ORU from the spacecraft. An additional benefit that the servicing spacecraft can provide is storage of extra science instruments. Since the initial propellant load is sufficient for a 20-year mission, on-orbit propellant transfer is not required. The design of the solar array in both deployment and functionality are important features on the ATLAST spacecraft. These concepts required pre-visualization to provide clarity and understanding. The visualization strategy was to animate the solar array deployment and show its balancing capabilities during maneuvering. Additionally, the array boom retraction characteristics were depicted during both the main propulsion burns and science observing scenarios. Careful consideration was made during the animation process. The goal of this work was to provide realistic solar array deployment and positioning. Considerations such as the spacecraft center of gravity and solar pressure were also captured in the animation. The modeling, texturing, and animation work was created in Autodesk 3ds Max 2009. The rendered scenes were imported into Autodesk Combustion 2008 for background compositing and labeling of the on-screen events. Once these scenes were rendered they were edited together into one final animated video using Sony Vegas Pro 8. The final video provided clarity through visualization that aided the solar array design process. Video 1 provides a link to an animation showing the solar array deployment. Video 2 illustrates how the solar arrays are used to help offset the torque due to solar radiation pressure on the optical tube enclosure. This study resulted in a conceptual spacecraft design capable of placing and keeping the ATLAST 8-meter observatory in a halo orbit about the second Sun-Earth Lagrange point. The Ares V heavy lift vehicle is capable of placing 65000 kg onto the L2 transfer trajectory with a C3 of -0.7 km 2 /s 2 . Even though the proposed NASA budget calls for canceling the Constellation Program and thus the Ares V, the new heavy lift vehicle design study at Marshall Space Flight Center is considering launch vehicles with volume and mass capabilities comparable to the Ares V. Therefore, ...