Theory and applications of ballute aerocapture and dual-use ballute systems for exploration of the solar system
Abstract
The use of a large, lightweight inflatable device for aerocapture could provide a significant mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. Ballute aerocapture combines the benefit of non-propulsive capture with the advantages of high altitude capture, such as lower heating rates than achieved with traditional aeroshells. The dual-use ballute concept employs the ballute to capture a spacecraft and, subsequently, descend a payload to the surface. The primary objective of this thesis is to determine the feasibility of using ballute aerocapture systems and dual-use ballute systems at the atmosphere-bearing bodies in the Solar System. In anticipation of future human missions, particular attention is paid to the problem of delivering high-mass payloads to Mars. To date, no dual-use ballute studies, neither numerical nor analytical, have been published. In this investigation, new and existing analytical theories are derived and explicit expressions are found for the maximum heating rates, maximum deceleration, and maximum dynamic pressure during a dual-use ballute trajectory. Each of the resulting expressions is a function of the ballistic coefficient and of the capture trajectory, which in turn provides the required size (area/mass ratio) of the ballute for the dual-use ballute system. An aerocapture study is conducted to assess propellant mass costs for orbit insertion and compare those costs to the system masses of aerocapture ballutes and aerocapture tethers at Venus, Earth, Mars, Jupiter, Saturn, Titan, Uranus, and Neptune. Based on the mass comparison, ballutes are more favorable than tethers and propellant for Venus, Earth, Mars, Titan, Uranus, and Neptune, and least favorable for Jupiter. Aerothermodynamic and trajectory analysis is critical in determining if ballutes are feasible for high-mass Mars aerocapture and entry applications. Trajectories, thermal protection system (TPS) trades, and aerothermodynamics of ballute-assisted, low- to high-mass, Mars-entry systems are investigated. Because a towed ballute has three different characteristic lengths, associated with it (corresponding to the ballute, the spacecraft, and the tethers), it can encounter different flow regimes simultaneously. To account for rarefaction, the Direct Simulation Monte Carlo (DSMC) method for aerothermodynamic analysis, which is based on kinetic flow modeling, is employed. Lastly, a one-dimensional heat-conduction code is employed to calculate the thermal protection mass for a range of vehicles with high- to low-ballistic coefficients and for a range of payloads (from 1 ton to 100 tons). In this study the thermal protection mass includes the ballute, heat shielding, or a combination of the two. Results show that ballutes (with sufficiently low-ballistic coefficients) require substantially lower thermal protection mass than traditional aeroshells.
Degree
Ph.D.
Advisors
Longuski, Purdue University.
Subject Area
Aerospace engineering
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