Michael J. Grant

Guidance, Navigation, and Control System Performance Trades for Mars Pinpoint Landing

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 guidancealgorithms were examined. Of these four, a near propellant-optimal analytic guidance law showed promisefortheconceptualdesignofpinpointlandingvehicles.Theexistenceofapropellantoptimumwithregardto theinitiationtimeofthepropulsiveterminaldescentwasshowntoexistforvarious flightconditions.Subsonicguided 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.

High Mass Mars Entry, Descent, and Landing Architecture Assessment

As the nation sets its sight on returning humans to the Moon and going onward to Mars, landing high mass payloads ( 2 t) on the Mars surface becomes a critical technological need. Viking heritage technologies (e.g., 70 sphere-cone aeroshell, SLA-561V thermal protection system, and supersonic disk-gap-band parachutes) that have been the mainstay of the United States’ robotic Mars exploration program do not provide sucient capability to land such large payload masses. In this investigation, a parametric study of the Mars entry, descent, and landing design space has been conducted. Entry velocity, entry vehicle conguration, entry vehicle mass, and the approach to supersonic deceleration were varied. Particular focus is given to the entry vehicle shape and the supersonic deceleration technology trades. Slender bodied vehicles with a lift-to-drag ratio (L=D) of 0.68 are examined alongside blunt bodies with L=D = 0.30. Results demonstrated that while the increased L=D of a slender entry conguration allows for more favorable terminal descent staging conditions, the greater structural eciencies of blunt body systems along with the reduced acreage required for the thermal protection system aords an inherently lighter vehicle. The supersonic deceleration technology trade focuses on inatable aerodynamic decelerators (IAD) and supersonic retropropulsion, as supersonic parachute systems are shown to be excessively large for further consideration. While entry masses (the total mass at the top of the Mars atmosphere) between 20 and 100 t are considered, a maximum payload capability of 37.3 t results. Of particular note, as entry mass increases, the gain in payload mass diminishes. It is shown that blunt body vehicles provide sucient vertical L=D to decelerate all entry masses considered through the Mars atmosphere with adequate staging conditions for the propulsive terminal descent. A payload mass fraction penalty of approximately 0.3 exists for the use of slender bodied vehicles. Another observation of this investigation is that the increased aerothermal and aerodynamic loads induced from a direct entry trajectory (velocity 6.75 km/s) reduce the payload mass fraction by approximately 15% compared to entry from orbital velocity ( 4 km/s). It should be noted that while both IADs and supersonic retropropulsion were evaluated for each of the entry masses, congurations, and velocities, the IAD proved to be more mass-ecient in all instances. The sensitivity of these results to modeling assumptions was also examined. The payload mass of slender body vehicles was observed to be approximately four times more sensitive to modeling assumptions and uncertainty than blunt bodies.

Analytic Hypersonic Aerodynamics for Conceptual Design of Entry Vehicles

and stability derivatives have been developed for basic shapes, including sharp cones, spherical segments, cylindrical segments, and at plates at varying angles of attack and sideslip. Each basic shape has been generically parametrized, requiring the development of only a single set of analytic relations for each basic shape. These basic shapes can be superimposed to form common entry vehicles, such as spherecones and blunted biconics. Using Bezier curves of revolution, more general bodies of revolution are studied in which the location of the control nodes that dene the shape of the curve is also generically parametrized. Analytic relations at unshadowed total angles of attack have been developed for these congurations. Analytic aerodynamic equations are orders of magnitude faster than commonly used panel methods and were validated using the NASA-developed Conguration Based Aerodynamics tool. Consequently, rapid aerodynamic trades and shape optimization can be performed. Additionally, the relations may impact guidance design using onboard trajectory propagation, real-time ablation modeling in simulations, and simultaneous vehicle-trajectory optimization.

Mars Science Laboratory Entry Optimization Using Particle Swarm Methodology

An attempt to improve upon the manual, time consuming traditional point design method for current reference trajectory design is discussed. A single objective particle swarm optimization (SOPSO) algorithm and a multiobjective particle swarm optimization (MOPSO) algorithm were developed. The SOPSO algorithm was used to validate the capability of applied swarming theory to Mars entry optimization. The MOPSO algorithm is an extension of SOPSO, providing the capability to generate Pareto fronts in the environment of competing objectives. The Pareto fronts generated by MOPSO provided quick insight into entry trajectory design characteristics that took years to understand through the traditional point design process. The optimal trade associated with the conflicting objectives of supersonic parachute deployment altitude, range error ellipse length, and g-loading were quantified in a visual environment. It is shown that the analysis of the Pareto front allows the user to make intelligent reference trajectory design decisions that could potentially span various disciplines in vehicle design. This work serves as a beginning of a fundamental departure from the traditional point design process that is manual, time consuming, and only allows the designer to evaluate a select few designs. This fundamental change in design allows the designer to analyze the set of best solutions via an automated, efficient, and global exploration of the design space using particle swarm methodology.

Guidance, Navigation, and Control Technology System Trades for Mars Pinpoint Landing

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.

Rapid Indirect Trajectory Optimization for Conceptual Design of Hypersonic Missions

During conceptual design, trajectory optimization is often performed using direct methods. This investigation illustrates that trajectory optimization and vehicle performance characterization can be rapidly performed by coupling indirect optimization, continuation, and symbolic manipulation. Examples illustrate that the historical challenges associated with indirect optimization methods can be overcome, enabling the rapid construction of high-quality solutions. In this design methodology, highly constrained optimal trajectories are rapidly constructed by traversing optimal manifolds of varying design problems throughout the continuation process.

Smart Divert: A New Mars Robotic Entry, Descent, and Landing Architecture

This study investigates the performance and feasibility of a new entry, descent, and landing architecture onMars, termed Smart Divert, for landing in one of a number of small safe zones surrounded by hazardous terrain. Smart Divert consists of a ballistic entry followed by supersonic parachute deployment. After parachute release, the vehicle diverts to one ofmany predefined, fuel-optimal safe zone sites. The Smart Divert concept does not require hypersonic guidance or real-time terrain recognition. Instead, it relies on a priori orbital observations to identify small, multiple safe zones within a larger hazardous region and additional terminal descent propellant to land at the fuel-optimal safe zone.Before launch,mission designers could trade thenumber and size of the safe zones as part of the landing site selection process.Reasonable propellantmass fractions of 0.3 canbe achievedby initiating the divert at 5 kmaltitude, providing a 10 km horizontal divert capability. The number of safe zones is shown to be a function of landing ellipse size. Assuming Mars Science Laboratory state-of-the-art interplanetary navigation, four safe zone sites, randomly placed throughout the landing ellipse to simulate unknowndestinations of futuremissions, require a propellantmass fraction less than 0.3 for 97% of the cases analyzed. The unconstrained optimal arrangement of four safe zone sites within the same landing ellipse reduced the required propellant mass fraction from 0.3 to 0.22. The propellant mass fraction may be further reduced as the number of safe zone sites is increased. An example scenario using rock count data for the Phoenix landing site region demonstrates that Smart Divert can be implemented to land in previously unreachable terrain for a propellant mass fraction of 0.2.

Rapid Design Space Exploration for Conceptual Design of Hypersonic Missions

During conceptual design, multidisciplinary optimization is often performed using computationally intensive direct methods. Prior work has shown that rapid design studies can be performed using fast indirect methods, but several optimization techniques including discrete dynamic programming, nonlinear inversion, and pseudospectral methods are required to construct a suitable initial guess within the design space. In this investigation, a simplied methodology is developed to eliminate these optimization techniques, enabling rapid design space exploration using continuation of indirect methods alone. This is made possible by initially converging to a simple solution that is outside of the design space of interest, and solutions within the design space of interest are quickly accessed through continuation from this initial solution. As an initial step to automate this continuation process, state transition tensors are used to predict optimal solutions throughout the design space. A methodology is developed to provide accurate predictions of trajectories with varying ight times, and the error of these predictions is controlled to regulate the continuation process. This approach provides exibility to adapt to future computational capabilities and serves as an initial step to bridge the gap between conceptual trajectory design and onboard trajectory planning.

Rapid Simultaneous Hypersonic Aerodynamic and Trajectory Optimization Using Variational Methods

Traditional multidisciplinary design optimization methodologies of hypersonic missions often employ population-based global searching methods that rely on shooting methods to perform trajectory optimization. In this investigation, a rapid simultaneous hypersonic aerodynamic and trajectory optimization methodology is constructed based on variational methods. This methodology is constructed from two enabling advancements in analytic hypersonic aerodynamics and rapid trajectory optimization. Comparisons made with a single and multi-objective particle swarm optimizer highlight the computational advantages and improved solutions obtained through continuation of variational methods. The incorporation of trajectory constraints into the particle swarm optimization process through penalty functions or as additional objectives is shown to greatly increase the complexity of the design process. Alternatively, variational methods are able to precisely satisfy trajectory constraints without this added complexity. Examples demonstrate that Pareto frontiers in both vehicle and trajectory objectives can be constructed using variational methods. For convex frontiers, this is performed using a weighted sum of the objectives. For nonconvex frontiers, the optimization is performed through continuation of a set of constrained objectives.

Feasibility of Guided Entry for a Crewed Lifting Body Without Angle-of-Attack Control

The feasibility of flying a crewed lifting body, such as the HL-20, during entry from low Earth orbit without the steady-state body flap deflections required for angle-of-attack control was evaluated. This entry strategy mitigates the severity of the aerothermal environment on the vehicle’s body flaps and reserves control power for transient steering maneuvers. A real-time numeric predictor–corrector entry guidance algorithm was developed to accommodate the range of vehicle trim angle-of-attack profiles possible in the absence of angle-of-attack control. Results show that it is feasible to steer the vehicle from low Earth orbit to a desired target with real-time guidance while satisfying a reasonable suite of trajectory constraints, including limits on peak heat rate, peak sensed deceleration, and integrated heat load. Uncertainty analyses confirm this result and show that the vehicle maintains significant performance robustness to expected day-of-flight uncertainties. Additionally, parametric scans over ...