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This Defense Intelligence Reference Document from the Defense Intelligence Agency, dated 1 November 2010, was produced under the Advanced Aerospace Weapon System Applications (AAWSA) Program. It describes nanosatellite technologies and proposes launching nanosats into orbit with laser Lightcraft propulsion. It also covers a weapon mission selection study and multi-megawatt laser options. The author recommends that the Department of Defense and NASA bring Lightcraft R&D back to the United States and restart the X-50LR flight demonstration program.
“Masse”33 pages
UNCLASSIFIED/ il"9" 91"l"!e!lltt tl!IL 9HL I • Weight: 0.25 kg. • Input data rate: 2 kbits/sec. • Output data rate: 100 kbits/sec. • Data storage: 2 Gbits. • Encoding: advanced convolutional. • Processing speed: 12 MIPS. • Radiation tolerance: > 100 krads total dose. In order to develop a low mass C&DH, a lightweight and low power electronics packaging method must be used. The packaging method that will be chosen must have a small volume and small footprint (6 cm x 6 cm x variable height). The packaging technique must provide data on programmable substrates and data on a compliant interconnects for space use. A multi-chip module (MCM) has been successfully produced by Pico Systems Inc. A combined effort to reduce mass, power, size and cost led to the development of the CMOS Ultra Low Power Radiation Tolerant (CULPRiT) system on a chip, and "C&DH in your Palm" are technologies that enable the power reduction required for nanosats. The goals of these technologies are a 20: 1 power reduction over current 5-volt technology, foundry independence of die production, and radiation tolerance. Another technology enabling a decrease in volume is the RHrFPGA, which reduces volume by replacing many logic functions/circuits with one die. The RHrFPGA also allows concurrent design by decoupling the logic design from the module, shortens the design schedule, lowers the part count, and eases rework. The above technologies allow for higher levels of electronic integration, effectively combining spacecraft subsystem electronics and instrument electronics into the smallest possible mass, power, and volume. POWER SYSTEMS Total spacecraft power is limited by the small satellite size. The Sun's power density is 1.35 kW/m 2. Assuming 15% conversion efficiency for a 0.3 m x 0.1 m disk shaped nanosat (cross section of 0.03 m 2 ), with a 67% area coverage, this results in a total electric power of only 4.0 watts. Lightweight, efficient solar array panels that minimize the effective array mounting area are needed. Dual or triple junction GaAs solar cells that give 18% conversion efficiency at end of life (EOL), and assuming a more optimistic area factor of 85%, will result in only 6.2 Wat EOL. Small satellites that do not have extended solar panels simply do not intercept a large solar power density and must use the available power very efficiently. For a small spinning satellite, it is expected that three solar cells will be connected in series along the spin axis, and groups of three will be connected in parallel around the circumference. Each section will generate 3.3 volts and rotate into and out of sunlight as a unit. Voltage drops at 3.3 volts, bus regulation, circuit protection (e.g., fuse or circuit breaker) and Lithium ion battery discharge characteristics are being studied. Highly elliptical orbits in the ecliptic plane where the apogee velocity is very low will cause a several hour eclipse during part of the year. Spacecraft batteries to cover this eclipse period presents a significant mass impact. However, only a 10° orbit plane UNCLASSIFIED//F811. 8FFl81*L tl!IL 8HLY 5
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Report, from the dia collection. The PDF is mirrored here; the original link is under it. 77 pages are in the text index: search them above, or from the library's search.