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End date: 2006 × Start date: 2006 × search.filters.applied.f.isOrgUnitOfPublicationTitle: Daniel Guggenheim School of Aerospace Engineering × search.filters.applied.f.isSeriesOfPublicationTitle: Master's Projects ×

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Now showing 1 - 6 of 6
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Optimal trajectories for soft landing on asteroids

2006-12-15 , Lantoine, Gregory

Robotic exploration of asteroids has been identified by NASA as a major long-term goal. Central to many asteroid missions is a precise soft landing to enable surface exploration or exploitation. This paper describes a technique for computing optimal autonomous controlled trajectories for soft landing in an irregular gravity field of a rotating asteroid. We will first discuss the complexity of the forces that act on the spacecraft during a landing and how we can model them. Then, we will present the numerical method used to solve the optimal control problem, and typical results are shown on case studies at asteroids Vesta and Golevka. In each example, we will identify the best mission design scenarios, as well as some operational difficulties. Finally, we will investigate sensitivity to parameter uncertainties and the implementation of a real-time feedback controller to increase landing accuracy.

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Performance Evaluation of a Side Mounted Shuttle Derived Heavy Lift Launch Vehicle for Lunar Exploration

2006-05-01 , Thompson, Robert W.

The NASA Exploration Systems Architecture Study (ESAS) produced a transportation architecture for returning humans to the moon affordably and safely. ESAS determined that the best lunar exploration strategy was to separate the launch of crew from the launch of cargo, thereby requiring two launches per lunar mission. An alternate concept for the cargo launch vehicle is a side mounted Shuttle-derived heavy lift launch. This configuration is similar to previously studied concepts, except engines and structure have been added to the External Tank (ET), making it a complete first stage. The upper stage is mounted on the side of the first stage, much like the Shuttle orbiter is mounted on the side of the ET. Like the Shuttle, solid rocket boosters (SRBs) are also used. This configuration has several performance and operational benefits over an in-line heavy lift launch vehicle. According to the ESAS report, side mount configurations were not considered to be among the most promising configurations, and were not carried forward for further consideration within architectural options. The performance of this launch vehicle is independently analyzed, using multidisciplinary analysis techniques. Methods and tools used include launch trajectory optimization with POST, vehicle aerodynamic analysis using APAS, and weights and sizing using historically based estimating relationships. Principal trade studies performed include first and second stage propulsion (number of engines and engine type), solid rocket booster size (four versus five segment), and staging ∆V. The vehicle design that best meets the requirements for space exploration (lunar and future missions) is presented.

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An Evaluation of Ballute Entry Systems for Lunar Return Missions

2006-05-07 , Clark, Ian G.

A study was undertaken to assess the advantages and feasibility of using ballutes for Earth entry at lunar return velocities. Using analysis methods suitable for conceptual design, multiple entry strategies were investigated. Entries that jettison the ballute after achieving orbit were shown to reduce heating rates to within reusable thermal protection system limits and deceleration was mitigated to approximately four g’s when a moderate amount of lift was applied post-jettison. Ballute size drivers were demonstrated to be the thermal limitations and areal densities of the ballute material. Performance requirements for both of those metrics were generated over a range of total ballute system masses.

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A Scalable Orbital Propellant Depot Design

2006-05-01 , Street, David

This paper describes the design and features of a Scalable Orbital Propellant Depot Design tool. The purpose of the tool is to enable others to easily test the effectiveness of adding a propellant depot to an exploration architecture. Several options are available including zero boil-off technology, usable propellant and depot geometry. It is assumed that the depot is refillable with a total service life of 10 years and resides in low earth orbit. Examples of depots created with the tool are shown. Application to existing exploration architectures is also discussed.

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Mars Entry, Descent and Landing Parametric Trades

2006-05-01 , Wells, Grant

The purpose of this investigation is to begin forming a dataset to be the basis of a Mars entry, descent and landing mission design handbook for planetary probes. The premise of the project is that Mars entry, descent and landing can be parameterized with five variables: (1) entry mass, (2) entry velocity, (3) entry flight path angle, (4) vehicle aeroshell diameter, and (5) vertical lift-to-drag ratio. For combinations of these input parameters, the following trajectory information will be determined: peak deceleration, peak heat rate, heat load, and the altitude at which Mach 2 is reached (for parachute deployment).

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Entry System Design of the Mars Gravity Biosatellite

2006-04-19 , Francis, Scott

The Mars Gravity Biosatellite will be launched to low Earth orbit and will study the effects of partial gravity on mammalian physiology. The entry vehicle will return 15 live mice to the Earth’s surface from low Earth orbit, landing in the Woomera Prohibited Area in central South Australia. This study establishes a baseline for the entry, descent, and landing system through the comparison of various concepts. The Discoverer capsule from the military’s Corona program of the 1950’s and 60’s is chosen over other concepts as the baseline aeroshell after an analysis of static stability and payload requirements for this mission. A nominal trajectory is developed based on science requirements, the safety of the mice, and payload recovery requirements. A sensitivity study is performed on the entry trajectory to determine the effects various parameters have on the nominal entry and a Monte Carlo dispersion analysis is used to establish a 3-σ landing ellipse, which fits within the boundaries of the Woomera Prohibited Area. A discussion of potential de-orbit propulsive devices is given in relation to the required de-orbit ∆V. A 16 m parachute is chosen as the baseline due to the resulting 4.8 m/s ground impact velocity and a crushable aluminum foam is chosen as a means to attenuate the shock of ground impact.

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