Blowing Hot & Cold: Flight Operations for Thermal Control on Mars & Venus Express

The successful launch of the Mars and Venus EXPRESS Class missions in 2003 and 2005 represents the first use of the same generic spacecraft bus on missions to orbit two different planets with very different operational environments. For Mars Express the discovery in flight of a reduction in available power of about 30% placed thermal control at the forefront in the high profile efforts to recover sufficient operating margin to fly a full science mission. This included optimizing spacecraft attitude (when possible) to achieve maximum solar contribution to heating the spacecraft, directly minimizing spacecraft heater power consumption through optimal configuration of redundant heater and extending the regime of software control (with carefully selected control ranges) to more and more heater circuits to achieve better control of temperatures and power consumption. The challenges for thermal control on Venus Express (VEX) center around its limited capacity to reject and store heat between pericenter science observations in a 24 hour highly eccentric orbit. Under some scenarios a ‘cool-down’ of 22 hours was foreseen in the design, with very little margin to recover for the next 2 hour pericenter pass. The mitigating strategies employed to maximize the science return from the mission are described, along with the operational constraints resulting from the inherited thermal design from Mars Express. This paper summarizes the history of the developments in the thermal control of the ESA interplanetary spacecraft series, the evolution of the various strategies to achieve minimum power consumption whilst maintaining a safe thermal environment (MEX), maximizing science opportunities while managing thermal loading from Sun and planetary heating (VEX) and in the development of concepts to overcome gradually evolving hardware on the spacecraft due to ageing (MEX & VEX). Major lessons learned from the flight experience of operating two very similar thermal control subsystems on two different planetary orbiters are investigated and reported. These include the traceability of operational requirements between subsystems in different mission phases (in this case CPS pressurant and thermal control during Mars Orbit Insertion), the use of on-board software existing capabilities to mitigate the effects of failed hardware (a thermostat on Mars Express), the extension of allowed flight attitudes from one purpose to another (heater power minimization to momentum control via solar sailing), and the definition and management of thermal interfaces (e.g. for Omega instrument on Mars Express). Finally the paper will explore and highlight some of the benefits and limitations of largely inheriting the design and implementation of a thermal control subsystem between spacecraft used in quite different environments.