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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...

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.

Fast, high-fidelity trajectory propagation of objects in near-Earth orbits is a key capability for space situational awareness and mitigating probability of collisions on orbit. This high-fidelity analysis requires accurate aerodynamics prediction for objects in the freemolecular regime of flight, but most tools for aerodynamic prediction for this regime either are found using assumptions or are computationally intensive. Symbolic manipulation software can be used to analytically integrate expressions for pressure and shear pressure coefficients acting on a general body in free-molecular regime to derive aerodynamic force and moment expressions. The analytical aerodynamics prediction method is described and relations have been developed for the sphere, cylinder, panel, and rectangular prism. The NASA-developed Direct Simulation Monte Carlo Analysis Code is used to validate the analytical expressions and it is shown that expressions are accurate within 0.38%. These generalized analytic expressions in terms of angle of attack, sideslip angle, freestream conditions, wall temperature, and accommodation coefficients allow near-instantaneous computation of the rarefied aerodynamics and enables space situation awareness analysis.

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a specific instance of multidisciplinary design with no feedback As opposed to the previous applications in Table 1, this investigation provides the theoretical foundations for casting the general multidisciplinary design problem as a dynamical system, including handling of equality and inequality constraints within the design. Three particular techniques from different domains of dynamical system theory are examined in depth as they directly relate to the development of a rapid robust design methodology. These techniques are:1. Stability analysis: The existence of a converged design (for a given iteration scheme) can be determined in the same way as analyzing a dynamical system's stability, where the conditions for asymptotic stability are identically equal to those required for convergence.

This paper investigates the use of state-dependent Riccati equation control for closed-loop guidance of the hypersonic phase of atmospheric entry. Included are a discussion of the development of the state-dependent Riccati equations, their outgrowth from Hamilton-Jacobi-Bellman theory, a discussion of the closed-loop nonlinear system's closed-loop stability and robustness from both a theoretical and practical viewpoint. An innovative use of sum-of-squares programming is used to solve the state-dependent Riccati equation with application of a state-dependent Riccati equation derived guidance algorithm to a high mass, robotic Mars entry example. Algorithm performance is compared to the Modified Apollo Final Phase algorithm planned for use on the Mars Science Laboratory.

The goal of this investigation is to understand the sizing and performance of supersonic inflatable aerodynamic decelerators for Earth-based sounding rocket applications. The recovery system under examination is composed of a supersonic inflatable aerodynamic decelerator and a guided parafoil system to achieve sub-100 m miss distances. Three supersonic inflatable aerodynamic decelerator configurations (tension cone, attached isotensoid, and trailing isotensoid) are examined using the metrics of decelerator mass, aerodynamic performance, and vehicle integration. In terms of aerodynamic performance, the tension cone is the preferred choice for the sizes investigated. The attached isotensoid was shown to be the most mass efficient decelerator, whereas the trailing isotensoid was found to be the more ideal decelerator for vehicle integration. A three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo uncertainty analysis to assess the landed accuracy capability of the proposed architectures. In 95% of the cases examined, the drag-modulated inflatable aerodynamic decelerator provides arrivals within the 10 km parafoil capability region, meeting the sub-100 m landed recovery goals. In 76% of the cases examined, the dragmodulated inflatable aerodynamic decelerator arrives within 5 km of this target zone. Nomenclaturenumber m = mass, kg q = dynamic pressure, N∕m 2 S = area, m 2 s = downrange, m T = reference temperature, K t = time, s u = eastward wind velocity, m∕s V = velocity, m∕s v = northward wind velocity, m∕s β = ballistic coefficient equal to m∕C D A, kg∕m 2 δDR = change in downrange, km λ = launch elevation angle, deg Subscripts D = drag deploy = deployment est = estimate f = areal max = maximum min = minimum para = parafoil ref = reference target = target