The atmospheric composition and geologic structure of Venus have been identified by the US National Research Council's Decadal Survey for Planetary Science as priority targets for scientific exploration, however the high temperature and pressure at the surface, along with the highly corrosive chemistry of the Venus atmosphere, present significant obstacles to spacecraft design that have severely limited past and proposed landed missions. Following the methodology of the NASA Innovative Advanced Concepts (NIAC) proposal regime and the Collaborative Modeling and Parametric Assessment of Space Systems (COMPASS) design protocol, this paper presents a conceptual study and initial feasibility analysis for a Discovery-class Venus lander capable of an extended-duration mission at ambient temperature and pressure, incorporating emerging technologies within the field of high temperature electronics in combination with novel configurations of proven, high Technology Readiness Level (TRL) systems. Radioisotope Thermal Power (RTG) systems and silicon carbide (SiC) communications and data handling are examined in detail, and various high-temperature instruments are proposed, including a seismometer and an advanced photodiode imager. The study combines this technological analysis with proposals for a descent instrument package and a relay orbiter to demonstrate the viability of an integrated atmospheric and in-situ geologic exploratory mission that differs from previous proposals by greatly reducing the mass, power requirements, and cost, while achieving important scientific goals.
The DLR plans to launch the microsatellite BIRD in 1999 as part of a Earth remote sensing mission. This project represents the begin of a line of small satellite missions with ambitious scientific and technological objectives by application of new technology and respecting the limitations of microsatellites. The spacecraft bus design is based on the proposed orbit and the payload requirements. The scientific payload is a novel multi-spectral sensor system, consisting of two cooled infrared sensor arrays and the Wide Angle Optoelectronic Stereo Scanner (WAOSS). A serious constrain of the satellite design is the required compatibility to a piggyback launch. The concept of the satellite bus fits to the requirements with the satellite dimensions of about 550x61Ox620 mm 3 and a total mass of approx. 85kg. The presentation describes the approach for the system design of the satellite bus with focus on the mission profile and the requirements of the payload. A special attention is paid to the bus structure, the attitude control and the thermal subsystem with its components and sensors under consideration as a low-cost mission.
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