System level thermal analysis of space vehicles carrying ablative shields is usually performed by coupling the Thermal Mathematical Model (TMM) of the re-entry vehicle with the ablative bond line temperature resulting from thermo-ablative simulation. Such simulation is typically conducted by superimposing an ablative back wall condition. During these years, thanks also to the direct experience gained on ESA programs involving ablative shields like IXV, the need of a more strongly coupled thermal analysis has become higher in order to optimize both shield and thermal control designs. Being this numerical procedure involving different software, hence requiring the management of a complex iterative approach, the claim for a "one shot run" increased, this to remove time consuming and error prone iterations and improve model management/configuration control. In this frame TAS-I and Politecnico di Torino developed a new tool able to run directly intoESATAN the ablative analysis, allowing a full integration of ablative shield model with the host vehicle TMM. This paper describes the analytical model and 1D numerical scheme adopted and the verification and validation runs performed. Nomenclature β = degree of reaction m & = mass flow e = emissivity Ψ = convective flux reduction factor ρ = material/reactant density a = flow type coefficient B = ratio between the activation energy of the degradation reaction, E a , and the universal gas constant ℜ BB = blocking parameter C = thermal capacitance c = reactant weight coefficient c p = specific heat at constant pressure e = specific internal energy f = harmonic interpolation weighting parameter h = specific enthalpy k = pre-exponential factor n = reaction order 2 q = surface heat flux Q = constant recession enthalpy R = thermal resistance s = heat shield thickness S ij = surface between the cell i and the cell j T = temperature t = time V = cell volume x = in depth coordinate y = mass fraction Subscripts bl = boundary layer c = fully charred state cw = cold wall hw = hot wall i = index/counter mat = material rec = recovery condition ref = reference value rr = re-radiated v = virgin state w = external surface value

In the complex domain of the space technologies and among the different available applications, the possibility to launch spacecrafts, establish permanent outcomes and explore the universe with automated satellites to meet various scientific aspects has been extensively investigated and the related technologies have been consolidated. At the same time, re-entry technologies, due to the overall increased level of complexity, still have wide development margins, for new technologies and new mission concepts. In this perspective, the full exploitation of space applications focused to planetary exploration, sample return, crew and cargo transportation and launchers are a challenge to future. Among the different achievements obtained in this field it is well worth recalling the experience of the Atmospheric Re-entry Vehicle flight in 1998 and a certain number of important investments performed at Agency and national levels like Hermes, MSTP, Festip, X-38, FLPP, TRP, GSTP, HSTS, AREV, Pre-X. IXV (Intermediate eXperimental Vehicle) builds on these past experiences and studies and is conceived as a technological step forward toward the consolidation of an autonomous re-entry vehicle. The IXV (see Figure 4) is an atmospheric re-entry demonstrator that will be launched by the ESA/Vega from Kourou, will perform a suborbital flight, and will re-enter in the atmosphere. It will experience typical Low Earth Orbit re-entry thermal loads while performing experiments related to thermal protection system validation (based upon new materials), aerodynamics, aerothermodynamics and GNC. It will also validate the engineering approach, the margin policy, and tools used to design and develop the demonstrator; in addition it will be used to plan the mission. The IXV project Phase-A/B1 study has been started in early 2005 under ESA FLPP Period-1 Phase-1&2 contracts, which will defined the IXV mission objectives and matured the IXV design. Phase B2 of the program was completed with System PDR in early 2009 with NGL (Next Generation Launcher) Prime Spa as the prime contractor. After PDR, the IXV design has been revisited after the project consortium was redefined. As a consequence, a new design approach passed through a successful system design review (CDR, April 2011). Subsequently, this led to the official kick-off to manufacturing, Phase D. The refinement of the thermal architecture for the protection of the reaction control system (RCS), the propellant (hydrazine) thermal control, and the safeguard of the thrusters during the mission is the subject of this presentation.