Computational Analysis of Automated Transfer Vehicle Reentry Flow and Explosion Assessment D. E. Boutamine * ESA, Noordwijk, The Netherlands Ph. Reynier † and R. Schmehl † Advanced Operations and Engineering Services, Leiden, The Netherlands and L. Marraffa * and J. Steelant * ESA, Noordwijk, The Netherlands DOI: 10.2514/1.27610 At the end of its mission to the International Space Station, during its reentry into Earth atmosphere, the automated transfer vehicle is subject to high heat fluxes leading to structural heating and fragmentation of the vehicle. It has been concluded that, depending on the mode of release, onboard residual hypergolic propellants may ignite and explode upon exposure to the hot and reactive flow environment. Because an earlier explosion of the vehicle would change drastically the impact footprint of its fragments onto the Earth surface, this study proposes a reassessment of the explosion potential. From the trajectory analysis, several points of the reentry path have been computed using a Navier–Stokes solver accounting for nonequilibrium effects. Numerical simulations have been performed with and without perforation of the structure. In parallel, a comprehensive literature survey on ignition of monomethyl hydrazine and dimethyl hydrazine vapors with pure air or air mixed with nitrogen tetroxide has been performed to assess the autoignition potential of the mixture. Finally, the results of the computational fluid dynamics computations have been used to estimate the explosion risk in the presence of a propellant leakage. Analysis confirms the risk of a destruction of the automated transfer vehicle at higher altitude, which could induce a different footprint of the fragments on the ground. Nomenclature A = area of the fissure, m 2 B = ballistic coefficent, B  m=DC d C d = drag coefficient c = sound speed, m=s D = base diameter, m dt = time step, s j = step for the Mach number integration M = Mach number M = molar mass, kg m = vehicle mass, kg _ m = mass flow rate, kg=s P = static pressure, Pa p i;min = minimum ignition pressure, Pa R = universal gas constant, J=K=mol r = gas constant per mole, J=K=kg T = temperature, K t 1 = time duration of the supersonic flow, s t 2 = time duration of the subsonic flow, s t ign = ignition delay, s v = volume of the vehicle, m 3 X = volume concentration y  = dimensionless value of the wall distance  = ratio of the gas specific heats H = enthalpy of combustion, kJ=mol  = density, kg=m 3 Subscripts i = total variable 1 = variable upstream of the bow shock 2 = variable downstream of the bow shock 3 = variable inside the vehicle * = variable at the fissure location Introduction S INCE the 1990s, the European Space Agency has conducted different studies focusing on debris engendered by space activities. These investigations were related to two main topics. The first one is coordinated by the ESOC (European Space Observation Centre) and concerns the reentry of spacecraft at the end of their mission [1–3]. The other is related to the Ariane 5 launcher and several studies have been performed by the Centre National d’Etudes Spatiales and ESA with the objective to vent propellant tanks into space at the end of the launch, thus preventing tank explosion due to solar heating [4–6]. To achieve this, tank propellants are vented into space through tubing and nozzle and submitted to a high depressurization [7]. This process can be accompanied by physical phenomena such as condensation [8,9] or vaporization [5]. One of the key points of these studies was always the prediction of the thermodynamic behavior of propellants during the venting into space. Here the topic is not to proceed to a tank passivation but to analyze the risk of vehicle explosion along its reentry trajectory path. The intensity and the altitude of vehicle fragmentation are two important parameters to assess the impact of debris on ground. The automated transfer vehicle (ATV) is a supply cargo for the International Space Station (ISS), constituted of two elements. The first one is the spacecraft subassembly (SCS) equipped with propulsion (propulsion tanks and thrusters) and avionics bays (batteries, gyroscopes, and harness). The second part is the integrated cargo carrier, containing the equipped external bay (water and gas tanks, web structure) and the equipped pressurized module (containers, cargo, and the attitude control thrusters). A view of the spacecraft is shown in Fig. 1. Received 1 September 2006; revision received 16 March 2007; accepted for publication 19 March 2007. Copyright © 2007 by European Space Agency ESA-ESTEC. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0022-4650/07 $10.00 in correspondence with the CCC. * Research Engineer, European Space Research and Technology Center. † Research Engineer, Mechanical Engineering Business Unit. JOURNAL OF SPACECRAFT AND ROCKETS Vol. 44, No. 4, July–August 2007 860 Downloaded by TECHNISCHE UNIVERSITEIT DELFT on March 3, 2014 | http://arc.aiaa.org | DOI: 10.2514/1.27610