Ballute aerodynamic decelerators have been studied since early in the space age (1960's), being proposed for aerocapture in the early 1980's. Significant technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. The mass savings of this concept enables a number of high value science missions. Current studies of ballute aerocapture at Titan and Earth may lead to flight test of one or more ballute concepts within the next five years. This paper provides a survey of the literature with application to ballute aerocapture. Special attention is paid to advances in trajectory analysis, hypersonic aerothermodynamics, structural analysis, coupled analysis, and flight tests. Advances anticipated over the next 5 years are summarized. Clamped Ballutes Trailing Ballutes Single stage design Asymmetric lifting design Two stage design Circular cross-section torus design Airfoil cross-section torus design Cable Membrane Towed sphere design Clamped torus Clamped Ballutes Trailing Ballutes Single stage design Asymmetric lifting design Two stage design Circular cross-section torus design Airfoil cross-section torus design Cable Membrane Towed sphere design Clamped torus

Earth entry system options for human return missions from the Moon and Mars were analyzed and compared to identify trends among the configurations and trajectory options and to facilitate informed decision making at the exploration architecture level. Entry system options included ballistic, lifting capsule, biconic, and lifting body configurations with direct entry and aerocapture trajectories. For each configuration and trajectory option, the thermal environment, deceleration environment, crossrange and downrange performance, and entry corridor were assessed. In addition, the feasibility of a common vehicle for lunar and Mars return was investigated. The results show that a low lift-to-drag ratio (L/D = 0.3) vehicle provides sufficient performance for both lunar and Mars return missions while providing the following benefits: excellent packaging efficiency, low structural and TPS mass fraction, ease of launch vehicle integration, and system elegance and simplicity. Numerous configuration options exist that achieve this L/D.

Many authors have shown the potential mass savings that a ballute can offer for both aerocapture and entry. This mass savings could enhance or even enable many scientific and human exploration missions. Prior to flight of a ballute several technical issues need to be addressed, including aeroelastic behavior. This paper begins to address the issue of aeroelastic behavior by developing and validating the Ballute Aeroelastic Analysis Tool (BAAT). The validation effort uses wind tunnel tests of clamped ballute models constructed of Kapton supported by a rigid nose and floating aft ring. Good correlation is obtained using modified Newtonian aerodynamics and non-linear structural analysis with temperature dependent material properties and thermal expansion. BAAT is then used to compute the deformed shape of a clamped ballute for Titan aerocapture in both the continuum and transitional regimes using impact method aerodynamics and direct simulation Monte Carlo. NOMENCLATURE C p-Pressure coefficient Kn-Knudsen number nbmin-Minimum number of flow cells on the body q-Dynamic pressure (Pa) γ-Ratio of specific heats

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Traditionally mass estimation for conceptual design of advanced launch vehicles has depended on historically -based mass estimating relationships (MERs). Furthermore different organizations have different sets of MERs based on different data sets, and even formulated in different ways. This paper attempts to compare the modern MERs used in the Space Systems Design Lab (SSDL) at Georgia Tech to the 1960's era relationships used in the NAS7-377 report on advanced propulsion design for launch vehicles. Comparisons of the weight breakdowns of a two -stage -to -oibit vehicle are made for between the Marquardt equations and the SSDL equations using two different technology assumptions. The first assumes 1970 technology for a direct comparison of the equations while the second assumes 2015 technology. Additionally technology and material advances are estimated in an attempt to justify the lower weight of the 2015 technology. The SSDL model using 1970 technology weighs in 7% heavier than the Marquardt equations for a comparable two -stage -to -orbit vehicle. When 2015 technology is applied to the same vehicle SSDL equations show a 33% savings, on the entire vehicle, could be made due to technology.