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Landing site selection is a compromise between safety concerns associated with the site's terrain and scientific interest. Therefore, technologies enabling pinpoint landing performance (sub-100-m accuracies) on the surface of Mars are of interest to increase the number of accessible sites for in situ research, as well as allow placement of vehicles nearby prepositioned assets. A survey of the performance of guidance, navigation, and control technologies that could allow pinpoint landing to occur at Mars was performed. This assessment has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at the end of the hypersonic guidance phase (parachute deployment) below approximately 3 km. Four different propulsive terminal descent guidance algorithms were examined. Of these four, a near propellant-optimal analytic guidance law showed promise for the conceptual design of pinpoint landing vehicles. The existence of a propellant optimum with regard to the initiation time of the propulsive terminal descent was shown to exist for various flight conditions. Subsonic guided parachutes were shown to provide marginal performance benefits, due to the timeline associated with descent through the thin Mars atmosphere. This investigation also demonstrates that navigation is a limiting technology for Mars pinpoint landing, with landed performance being largely driven by navigation sensor and map tie accuracy. Nomenclature a = acceleration vector, a 1 a 2 a 3 T a i = acceleration along the ith direction b = scalar weighting parameter on the terminal constraint sensitivity C j i = jth constant coefficient used in the modified Apollo lunar module guidance algorithm dt f = terminal time increment f = set of first-order differential equations of motion g = local acceleration due to gravity g = acceleration vector due to gravity i = index I JJ = partition used in the optimal control solution I J = partition used in the optimal control solution I J = partition used in the optimal control solution I = partition used in the optimal control solution J = performance index Kn = Knudsen number L = scalar objective function describing path parameters M = Mach number m prop = mass of propellant m 0 = initial mass of the vehicle p = influence function vector R = matrix of influence functions r = position vector, r 1 r 2 r 3 T r i = position along the ith direction S j = matrix defining convex state constraints t = time t go = time to go until touchdown u = control vector v = velocity vector, v 1 v 2 v 3 T v i = velocity along the ith direction W = positive definite weighting matrix x = state vector, r T v T m T = mass consumption rate j = scalar defining convex state constraints = weighting on final time to go u = control vector increment "= tolerance level = slack variable bounding thrust magnitude 1 = thrust magnitude lower bound 2 = thrust magnitude upper bound = dust tau (opacity measure of the atmosphere) c = commanded thrust vector j = vector defining convex state constraints = scalar...

The impending development of NASA's Orion crew exploration vehicle will require a new entry guidance algorithm that provides sufficient performance to meet all requirements. This study examined the effects on entry footprints of enhancing the skip trajectory entry guidance used in the Apollo program. The skip trajectory entry guidance was modified to include a numerical predictor-corrector phase during the atmospheric skip portion of the entry trajectory. A 4-degree-of-freedom simulation was used to determine the range capability of the entry vehicle for the baseline Apollo entry guidance and the predictor-corrector enhanced guidance with both high and low lofting at several lunar return entry conditions. The results show that the predictor-corrector guidance modification significantly improves the entry range capability of the crew exploration vehicle for the lunar return mission. The performance provided by the enhanced algorithm is likely to meet the entry range requirements for the crew exploration vehicle.

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Landing site selection is a compromise between safety concerns associated with the site's terrain and scientific interest. Therefore, technologies enabling pinpoint landing (sub-100 m accuracies) on the surface of Mars are of interest to increase the number of accessible sites for in-situ research as well as allow placement of vehicles nearby prepositioned assets. A survey of various guidance, navigation, and control technologies that could allow pinpoint landing to occur at Mars has shown that negligible propellant mass fraction benefits are seen for reducing the three-sigma position dispersion at parachute deployment below approximately 3 km. Four different propulsive terminal descent guidance algorithms were analyzed with varying applicability to flight. Of these four, a near propellant optimal, analytic guidance law showed promise for the conceptual design of pinpoint landing vehicles. The existence of a propellant optimum with regards to the initiation time of the propulsive terminal descent was shown to exist for various flight conditions. In addition, subsonic guided parachutes are shown to provide marginal performance benefits due to the timeline associated with Martian entries, and a low computational-cost, yet near fuel optimal propulsive terminal descent algorithm is identified. This investigation also demonstrates that navigation is a limiting technology for Mars pinpoint landing, with overall landed performance being largely driven by navigation sensor and map tie accuracy.

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