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Engineering

Introduction

The objective of the spacecraft design is to satisfy the Venus Express mission requirements while making extensive reuse of the Mars Express design, so deriving cost benefits through recurrence and minimising development risks. As a consequence, the Venus Express spacecraft has the following similarities to Mars Express:

  • System concept - body mounted instruments, fixed communications antennas, two solar arrays with one degree of freedom pointing mechanisms
  • Structural design - minor, local changes to accommodate the revised instrument payload
  • Propulsion subsystem - higher fuel load to meet more stringent delta V requirement
  • Avionics units - limited changes to reflect new mission profile
  • Operational concept - alternating between Venus observation during specific portions of the orbit and Earth communication and battery charging at other times

However, there are some Venus Express mission characteristics that led to design changes, primarily in the areas of thermal control, communications and electrical power:

  • Science mission - additional and/or revised instruments needed to be accommodated (MAG, VeRa, VIRTIS and VMC) and two instruments that were design drivers for Mars Express were removed (BEAGLE and MARSIS)
  • Proximity to the Sun - since Venus is closer to the Sun than Mars (0.72 AU instead of 1.5 AU), the radiant heating of the spacecraft is four times greater for Venus Express, the ionising radiation environment is harsher and the illumination of the solar panels is more intense
  • Configuration of planets - in Mars orbit, the Earth vector is always within ± 40 degrees of the Sun vector, which assists with keeping the spacecraft cold face pointed away from the Sun during Earth communication, while, in orbit around an inner planet such as Venus, this convenience is not available
  • Distance to Earth - Venus maximum distance from Earth is less than Mars (1.7 AU compared to 2.7 AU)
  • Gravity - Venus gravity is greater than that of Mars (90% of Earth gravity instead of 38%), which led to a greater change in velocity being required for orbit injection (with a consequent propellant mass increase) and led indirectly to a longer orbital period (approximately 24 hours as opposed to 7 hours) and higher velocity at pericentre (about 9 km s-1 instead of 4 km s-1)


Last Update: 24 May 2007

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