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CNES Launchers Directorate (DLA) has to maintain a high level of expertise to achieve its missions regarding the exploitation of the Ariane 5 family launchers and to prepare the future of launchers. For this purpose, developing and using numerical simulation tools is of key importance. CARMEN software is CNES reference tool for simulating propulsion systems from their design to their exploitation. CARMEN regroups several CNES software packages. Among them are: CARDIM, for engine thermodynamics and geometric design, and CARSTAT/CARINS, for analyzing functional behavior of propulsion systems in stationary (CAR-STAT) or transient (CARINS) modes. This paper extensively presents CARMEN software development status and capabilities.

Recent advances in electric propulsion technologies such as magnetoplasma rockets gave a new momentum to the study of nuclear electric propulsion concepts for Mars missions. Some recent works have been focused on very short Earth-to-Mars transfers of about 40 days with high-power, variable speci¦c impulse propulsion systems [1]. While the interest of nuclear electric propulsion appears clearly with regard to the payload mass ratio (due to a high level of speci¦c impulse), its interest with regard to the transfer time is more complex to de¦ne, as it depends on many design parameters. In this paper, a general analysis of the capability of nuclear electric propulsion systems considering both criteria (the payload mass ratio and the transfer time) is performed, and the technological requirements for fast EarthMars transfers are studied. This analysis has been performed in two steps. First, complete trajectory optimizations have been performed by CNES-DCT in order to obtain the propulsion requirements of the mission for di¨erent technological hypotheses regarding the engine technology (speci¦c impulse levels and the throttling capability) and di¨erent mission requirements. The methodology used for designing fuel-optimal heliocentric trajectories, based on the Pontryagin£s Maximum Principle will be presented. Trajectories have been computed for various power levels combined with either variable or ¦xed I sp . The second step consisted in evaluating a simpler method that could easily link the main mission requirements , 2013 This is an Open Access article distributed under the terms of the Creative Commons Attribution License 2.0, which permits unrestricted use, distribution, and reproduction in any medium, provided the original work is properly cited. Progress in Propulsion Physics4 (2013) 313-336 DOI: 10.1051/eucass/201304313 © Owned by the authors, published by EDP SciencesArticle available at http://www.eucass-proceedings.eu or http://dx.doi.org/10.1051/eucass/201304313 PROGRESS IN PROPULSION PHYSICS(the transfer time and the payload fraction) to the main technological requirements (the speci¦c mass of the power generation system and the structure mass ratio of the whole vehicle, excluding the power generation system). Indeed, for power-limited systems, propulsion requirements can be characterized through the ¤trajectory characteristic¥ parameter, de¦ned as the integral over time of the squared thrust acceleration. Technological requirements for the vehicle can then be derived from the propulsion requirements using a simpli¦ed performance model designed by Onera [2]. This model yields the optimum vehicle design in terms of the payload mass ratio as well as the theoretical upper limit of the power source£s speci¦c mass as a function of the transfer time. Both studies show that the key to very fast EarthMars transfers (40 days, or less) is the reduction of the power source speci¦c mass below 1 kg/kW. On-going French studies [3] tend to show that speci¦c masses of nuclear reactors for exploration mission are expect...

Recent advances in electric propulsion technologies such as magnetoplasma rockets gave new momentum to the study of nuclear electric propulsion concepts for Mars missions.While the interest of nuclear electric propulsion appears clearly with regard to the payload mass ratio (due to a high level of specific impulse), its interest with regard to transfer time is more complex to define, as it depends on many design parameters. In this paper, we perform a general analysis of the mid-term capability of nuclear electric propulsion systems considering both payload mass ratio and transfer time. This study has been performed through a multidisciplinary analysis which combines general performance calculations for power-limited systems, an analysis of nuclear power-source that could be available in the future, and a series of mission analysis including trajectory optimization. The results obtained in this study emphasize and -most importantly -quantify the importance of the specific mass of the power and propulsion system with regards to the objective of a fast transfer. The objective of a very fast Earth-to-Mars transfer in less than six weeks appears unrealistic in a mid-term context, as it depends on a hypothetical breakthrough on the nuclear electric power source, with a specific mass typically lower than 1 kg/kW. The results obtained in this paper draw the contour of performance improvements that could be obtained in a mid-term horizon, using power generation technologies that are challenging but more commonly considered as reasonably optimistic, for example using a high temperature Brayton or Rankine conversion cycle. It appears that the shorter Earth-toMars transfer time that could be expected for missions with sufficient payload is about 120 days, compared to 180 days with chemical propulsion. Nomenclatureg 0 = Earth gravity acceleration at sea-level (9.80665 m/s 2 ) Isp = specific impulse k = tank structural mass ratio = M S / M P q = mass flow-rate (kg/s) M = mass (kg) M f = total final mass (kg) M i = total initial mass (kg) M P = propellant mass (kg) M PL = payload mass (kg) M W = power source mass (kg) M S = structure mass (kg) 2 Pe = input power delivered by the source (W) r = position vector t = time T = thrust U = thrust direction vector v = velocity vector α = specific mass of the power source and propulsion system (kg/kW) γ T = thrust acceleration (m/s 2 ) λ = trajectory characteristic parameter (m 2 /s 3 ) μ = Earth gravity constant (3.986.10 14 km 3 /s 2 ) η = global efficiency of the propulsion system