The Thermal Protection System (TPS) protects (insulates) a body from the severe heating encountered during hypersonic flight through a planetary or the earth’s atmosphere. In general, there are two classes of TPS: Reusable TPS, where after exposure to re-entry environment there are no changes in the mass or properties of the materials, and ablative TPS that, in contrast, accommodate high heating rates and heat loads through phase change and mass loss. Most ablative TPS materials are reinforced composites employing organic resins as binders. When heated, the resin pyrolyzes producing gaseous products (mostly hydrocarbons) that percolate toward the heated surface and are injected into the boundary layer. Resin pyrolysis also produces a carbonaceous residue that deposits on the reinforcement. The resulting surface material is termed “char.” In the framework of the ESA (European Space Agency) CSTS Programme, the Department of Chemical and Materials Engineering (ICMA) of “Sapienza” University and Thales Alenia Space are involved in the development of an innovative ablative TPS for a possible application in a new crew space transportation vehicle. Experimental activities were focused on selection and characterization of materials for an ablative thermal protection system able to withstand a heat flux range of 2 ÷ 9 MW/m2, in order to analyze material response in heat flux conditions consistent with the moon-earth re-entry. The selected raw materials (carbon based substrate, thermosetting resin) were used for the manufacture at lab scale of samples requested for the first stage material testing and analyses. To this purpose an infiltration technique, vacuum assisted, was developed. Impregnation was then used in order to obtain uniform and non uniform samples to match the specified maximum overall mass density. Moreover, after characterization and selection of materials and after laboratory tests on thermal protection samples, an ablative specimen (frustum of cone, 80 mm high) was manufactured and tested in the Plasma Wind Tunnel “Scirocco”, a facility of Italian Aerospace Research Centre (CIRA) that simulates the hypersonic re-entry environment

Low density ablative thermal shields for ballistic re-entry from lunar missions / Pulci, Giovanni; Tirillo', Jacopo; Marra, Francesco; F., Morici; E., Bonifaci; A., Simone; F., Fossati; Valente, Teodoro. - STAMPA. - (2009). (Intervento presentato al convegno VII CONVEGNO NAZIONALE INSTM SULLA SCIENZA E TECNOLOGIA DEI MATERIALI tenutosi a Tirrenia (PI) nel 9-12 Giugno 2009).

Low density ablative thermal shields for ballistic re-entry from lunar missions

PULCI, Giovanni;TIRILLO', Jacopo;MARRA, FRANCESCO;VALENTE, Teodoro
2009

Abstract

The Thermal Protection System (TPS) protects (insulates) a body from the severe heating encountered during hypersonic flight through a planetary or the earth’s atmosphere. In general, there are two classes of TPS: Reusable TPS, where after exposure to re-entry environment there are no changes in the mass or properties of the materials, and ablative TPS that, in contrast, accommodate high heating rates and heat loads through phase change and mass loss. Most ablative TPS materials are reinforced composites employing organic resins as binders. When heated, the resin pyrolyzes producing gaseous products (mostly hydrocarbons) that percolate toward the heated surface and are injected into the boundary layer. Resin pyrolysis also produces a carbonaceous residue that deposits on the reinforcement. The resulting surface material is termed “char.” In the framework of the ESA (European Space Agency) CSTS Programme, the Department of Chemical and Materials Engineering (ICMA) of “Sapienza” University and Thales Alenia Space are involved in the development of an innovative ablative TPS for a possible application in a new crew space transportation vehicle. Experimental activities were focused on selection and characterization of materials for an ablative thermal protection system able to withstand a heat flux range of 2 ÷ 9 MW/m2, in order to analyze material response in heat flux conditions consistent with the moon-earth re-entry. The selected raw materials (carbon based substrate, thermosetting resin) were used for the manufacture at lab scale of samples requested for the first stage material testing and analyses. To this purpose an infiltration technique, vacuum assisted, was developed. Impregnation was then used in order to obtain uniform and non uniform samples to match the specified maximum overall mass density. Moreover, after characterization and selection of materials and after laboratory tests on thermal protection samples, an ablative specimen (frustum of cone, 80 mm high) was manufactured and tested in the Plasma Wind Tunnel “Scirocco”, a facility of Italian Aerospace Research Centre (CIRA) that simulates the hypersonic re-entry environment
2009
VII CONVEGNO NAZIONALE INSTM SULLA SCIENZA E TECNOLOGIA DEI MATERIALI
04 Pubblicazione in atti di convegno::04b Atto di convegno in volume
Low density ablative thermal shields for ballistic re-entry from lunar missions / Pulci, Giovanni; Tirillo', Jacopo; Marra, Francesco; F., Morici; E., Bonifaci; A., Simone; F., Fossati; Valente, Teodoro. - STAMPA. - (2009). (Intervento presentato al convegno VII CONVEGNO NAZIONALE INSTM SULLA SCIENZA E TECNOLOGIA DEI MATERIALI tenutosi a Tirrenia (PI) nel 9-12 Giugno 2009).
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Utilizza questo identificativo per citare o creare un link a questo documento: https://hdl.handle.net/11573/412383
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