Gregg H. Barton

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.

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.

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.

Mars Sample Return: architecture and mission design

Mars Sample Return (MSR) architecture and mission engineering, led by Boeing for JPL, is presented. The study sought credible data to support planning a 2011 mission to return 500 g of scientifically selected samples. Phase 1 compared diverse architecture options to accomplish the mission. Seventeen theme-based architectures were conceived, quantified, measured, and scored. Two primary and three secondary architectures were recommended. Phase 2 developed engineering detail for a simple architecture specified by JPL: dual mission to two landing sites; short-range, radioisotope-powered sampling rovers; Mars orbit rendezvous; and electric return propulsion with Shuttle rendezvous. The design comprises nine system elements. Solutions for sample handling and breaking the back contamination chain are detailed. Total mission duration is five years. Technology tailoring, rather than technology creation, is required. Mission development cost, including margins and wraps, is

PredGuid entry guidance for Orion return from low Earth orbit

2.8/spl times/10/sup 9/. The study concluded that many schemes can feasibly accomplish Mars sample return, depending on program objectives adopted.

Guided Entry Performance of Low Ballistic Coefficient Vehicles at Mars

When returning from low Earth orbit, Orion will perform a lifting atmospheric entry with precision landing using the PredGuid entry guidance algorithm.12 The PredGuid algorithm is designed to guide Orion to the desired landing site while accounting for vehicle and environment uncertainties and day of flight dispersions during atmospheric entry. The low Earth orbit mode of the PredGuid entry guidance algorithm consists of three phases: the Initial Roll Phase which maintains proper entry attitude and steers the vehicle to proper transition conditions for the Final Phase; the Final Phase, a terminal point guidance algorithm that uses a stored reference trajectory to steer out range error and achieve precision landing; and the Terminal Phase, which seeks to null the remaining crossrange error through a simple proportional steering law. Simulation results indicate that the PredGuid algorithm provides precision landing capability to Orion as well as significant robustness to day-of-flight uncertainties and dispersions.

Improving Lunar Return Entry Footprints Using Enhanced Skip Trajectory Guidance

Current Mars entry, descent, and landing technology is near its performance limit and may be unable to land payloads on the surface that exceed 1 metric ton. One option for increasing landed payload mass capability is decreasing the entry vehicle’s hypersonic ballistic coefficient. A lower ballistic coefficient vehicle decelerates higher in the atmosphere, providing the additional timeline and altitude margin necessary to land more massive payloads. This study analyzed the guided entry performance of several low ballistic coefficient vehicle concepts on Mars. A terminal point controller guidance algorithm, based on the Apollo Final Phase algorithm, was used to provide precision targeting capability. Terminal accuracy, peak deceleration, peak heat rate, and integrated heat load were assessed and compared with a traditional Mars entry vehicle concept to determine the effects of lowering the vehicle ballistic coefficient on entry performance. Results indicate that, while terminal accuracy degrades slightly w...

Mars sample return: Architecture and mission design

The impending development of NASAs Crew Exploration Vehicle (CEV) 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 atmospheric skip portion of the entry trajectory. Four degree-of-freedom simulation was used to determine the footprint of the entry vehicle for the baseline Apollo entry guidance and 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 footprint of the CEV for the lunar return mission. The performance provided by the enhanced algorithm is likely to meet the entry range requirements for the CEV.

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

Abstract Mars Sample Return (MSR) architecture and mission engineering, led by Boeing for JPL, is presented. The study sought credible data to support planning a 2011 mission to return 500g of scientifically selected samples. Phase 1 compared diverse architecture options to accomplish the mission. 17 theme-based architectures were conceived, quantified, measured, and scored. Two primary and three secondary architectures were recommended. Phase 2 developed engineering detail for a simple architecture specified by JPL: dual mission to two landing sites; short-range, radioisotope-powered sampling rovers; Mars orbit rendezvous; and electric return propulsion with Shuttle rendezvous. The design comprises nine system elements. Solutions for sample handling and breaking the back contamination chain are detailed. Total mission duration is five years. Technology tailoring, rather than technology creation, is required. Mission development cost, including margins and wraps, is

Rapid Robust Design of a Deployable System for Boost-Glide Vehicles

2.8B. The study concluded that many schemes can feasibly accomplish Mars sample return, depending on program objectives adopted.