The atmospheric entry of a meteor is quite complex, with the body losing kinetic energy both from atmospheric drag and from mass loss due to aerodynamic heating. Moreover, high pressures on the windward side of the body result in enormous compressive stresses which may exceed the yield strength of the material and cause rapid fragmentation of the meteor. While ablative mass loss is not important for extremely large objects, it must be accurately estimated to correctly predict the trajectories of objects that are several tens of meters in diameter. The current paper describes a computer model which performs calculations of shock layer conditions, accounting for the time varying temperature distribution, radiative cooling of the shocked gases, and blockage of surface heating by ablation products. Application of the model to the well-known Tunguska Event indicates that the responsible bolide was probably a carbonaceous chondrite, although a stony asteroid or a cometary body cannot be conclusively ruled out.
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Earth-return entry corridor analyses are performed to evaluate the atmospheric flight environment of manned return from Mars. Trajectory and performance differences between aerocapture and direct entry are assessed and quantified in terms of the required aerobrake lift-drag ratio (L/D), stagnation-point heating, and the significance of off-nominal atmospheric conditions. The Earth-return aerobraking scenarios compared are 1) aerocapture into a phasing orbit with a 24-h period, 2) aerocapture into a 500-km circular orbit, 3) and direct entry to splashdown. No significant differences between aerocapture to a 500-km circular orbit and direct entry were observed in terms of aerobrake L /D requirements, maximum deceleration, or peak stagnation-point heat rate. The importance of parking orbit selection is demonstrated for low entry velocity Earth-return missions from Mars and missions returning from the Moon. Additionally, a stagnation-point heating analysis revealed that in all cases the peak heat rate is large enough to require an ablative thermal protection system for manned return from Mars. However, for entry velocities of 12.5 km/s and less, the heating environment is of the same order of magnitude as that experienced during the Apollo program. To perform the entry analysis, two predictor-corrector guidance design strategies were developed. Use of a predictor-corrector technique was shown to provide adequate flight margin for managing off-nominal atmospheric conditions.Nomenclature reference surface area, m 2 drag coefficient gravitational constant, 9.806 m/s 2 lift-to-drag ratio vehicle mass, kg ballistic coefficient, kg/m 2 convective stagnation-point heat rate, W/cm 2 radiative stagnation-point heat rate, W/cm 2 total stagnation-point heat load, J/cm 2 total stagnation-point heat rate, W/cm 2 inertial velocity at the atmospheric interface, km/s angle of attack propulsive velocity change, m/s fly able corridor width, deg inertial flight-path angle at the atmospheric interface, deg atmospheric density, kg/m 3 1976 standard atmospheric density, kg/m 3
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