Organizational Unit:

Space Systems Design Laboratory (SSDL)

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https://hdl.handle.net/1853/70747

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Organizational Unit

Daniel Guggenheim School of Aerospace Engineering

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https://finding-aids.library.gatech.edu/agents/corporate_entities/1138

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Publication Search Results

Now showing 1 - 10 of 106

Optimal trajectories for soft landing on asteroids

(Georgia Institute of Technology, 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.
Lazarus: A SSTO Hypersonic Vehicle Concept Utilizing RBCC and HEDM Propulsion Technologies

(Georgia Institute of Technology, 2006-11) Young, David Anthony ; Kokan, Timothy Salim ; Clark, Ian G. ; Tanner, Christopher ; Wilhite, Alan W.

Lazarus is an unmanned single stage reusable launch vehicle concept utilizing advanced propulsion concepts such as rocket based combined cycle engine (RBCC) and high energy density material (HEDM) propellants. These advanced propulsion elements make the Lazarus launch vehicle both feasible and viable in today's highly competitive market. The Lazarus concept is powered by six rocket based combined cycle engines. These engines are designed to operate with HEDM fuel and liquid oxygen (LOX). During atmospheric flight the LOX is augmented by air traveling through the engines and the resulting propellant mass fractions make single stage to orbit (SSTO) possible. A typical hindrance to SSTO vehicles are the large wings and landing gear necessary for takeoff of a fully fueled vehicle. The Lazarus concept addresses this problem by using a sled to take off horizontally. This sled accelerates the vehicle to over 500 mph using the launch vehicle engines and a propellant cross feed system. This propellant feed system allows the vehicle to accelerate using its own propulsion system without carrying the necessary fuel required while it is attached to the sled. Lazarus is designed to deliver 5,000 lbs of payload to a 100 nmi x 100 nmi x 28.5° orbit due East out of Kennedy Space Center (KSC). This mission design allows for rapid redeployment of small orbital assets with little launch preparation. Lazarus is also designed for a secondary strike mission. The high speed and long range inherent in a SSTO launch vehicle make it an ideal global strike platform. Details of the conceptual design process used for Lazarus are included in this paper. The disciplines used in the design include aerodynamics, configuration, propulsion design, trajectory, mass properties, cost, operations, reliability and safety. Each of these disciplines was computed using a conceptual design tool similar to that used in industry. These disciplines were then combined into an integrated design process and used to minimize the gross weight of the Lazarus design.
Characterizing High-Energy-Density Propellants for Space Propulsion Apllications

(Georgia Institute of Technology, 2006-10) Kokan, Timothy Salim ; Olds, John R.

A technique for determining the thermophysical properties of high-energy-density matter (HEDM) propellants is presented. HEDM compounds are of interest in the liquid rocket engine industry due to their high density and high energy content relative to existing industry standard propellants (liquid hydrogen, kerosene, and hydrazine). In order to model rocket engine performance, cost, and weight in a conceptual design environment, several thermodynamic and physical properties are needed. These properties include enthalpy, entropy, density, viscosity, and thermal conductivity. These properties need to be known over a wide range of temperature and pressure. A technique using a combination of quantum mechanics and molecular dynamics is used to determine these properties for quadricyclane, a HEDM compound of interest. Good agreement is shown with experimentally measured thermophysical properties. A vehicle case study is provided to quantify the system level benefits of using quadricyclane instead of hydrazine for the lunar lander ascent stage of the Exploration Systems Architecture Study. The results show that the use of HEDM propellants can significantly reduce the lunar lander mass and indicate that HEDM propellants are an attractive technology to pursue for future lunar missions.
Cost of Safety for Space Transportation

(Georgia Institute of Technology, 2006-10) Krevor, Zachary C. ; Wilhite, Alan W.

This paper proposes a methodology that explores the tradeoff between increasing component reliability and utilizing component redundancy as the strategy to meet space transportation reliability requirements. This technique would be employed by design engineers to make decisions about a reliability approach. The tradeoff between component redundancy and making parts more reliable warrants more investigation. System level reliability decisions are being made without a thorough exploration of cost saving opportunities. The impact of using redundancy on a system, including how it affects metrics such as development and operations cost, is presented. Additionally, there is little understood about the resources required to improve component reliability to acceptable levels. The process of making parts more reliable is studied and quantified. To incorporate the uncertainty that exists from reliability applications, a stochastic approach is used. Case studies of historical space systems are presented to demonstrate how this methodology is applicable. The findings show how a different reliability approach may have resulted in significant cost reductions. Conclusions are drawn about how to best meet reliability requirements while remaining within strict budgetary guidelines.
A Method for Allocating Reliability and Cost in a Lunar Exploration Architecture

(Georgia Institute of Technology, 2006-10) Young, David Anthony ; Wilhite, Alan W.

In January 2005, President Bush announced the Vision for Space Exploration. This vision involved a progressive expansion of human capabilities beyond Low Earth Orbit. Current design processes utilized to meet this vision employ performance based optimization schemes to determine the ideal solution. In these design processes the important aspects such as cost and reliability are currently calculated as an afterthought to the traditional performance metrics. The methodology implemented in this paper focuses on bringing these decisive variables to the forefront of the design process. To achieve this focus on cost and reliability in a lunar architecture design, a resource allocation technique from the business world will be implemented. This allocation technique optimally distributes the company's resources even though the actual performances of the products are uncertain. This method of resource allocation will be applied to a lunar architecture design to achieve the highest architecture reliability for a given budget. Once the methodology is created it will be implemented in a lunar architecture design tool. This tool will allow the decision maker to independently address the sensitivities of the lunar architecture's reliability to the overall budget of the program.
On-Orbit Propellant Re-supply Options for Mars Exploration Architectures

(Georgia Institute of Technology, 2006-10) Tanner, Christopher ; Young, James J. ; Thompson, Robert W. ; Wilhite, Alan W.

A report detailing recommendations for a transportation architecture and a roadmap for U.S. exploration of the Moon and Mars was released by the NASA Exploration Systems Architecture Study (ESAS) in November 2005. In addition to defining launch vehicles and various aspects of a lunar exploration architecture, the report also elaborated on the extent of commercial involvement in future NASA activities, such as cargo transportation to the International Space Station. Another potential area of commercial involvement under investigation is the delivery of cryogenic propellants to low-Earth orbit (LEO) to refuel NASA assets as well as commercial assets on orbit. The ability to resupply propellant to various architecture elements on-orbit opens a host of new possibilities with respect to a Mars transportation architecture – first and foremost being the ability to conduct a Martian exploration campaign without the development of expensive propulsion systems such as nuclear thermal propulsion. In-space propellant transfer in the form of an orbiting propellant depot would affect the sizing and configuration of some currently proposed vehicles such as the Earth Departure Stage (EDS) and the Mars Transit Vehicle (MTV). In addition, it would influence the overall affordability and sustainability of a long-term Mars exploration campaign. To assess these consequences, these vehicles and their various stages are modeled to approximate the ESAS performance figures using a combination of analogous systems and physics-based simulation. Well established modeling tools -- such as POST for trajectory optimization, APAS for aerodynamics, NAFCOM for cost modeling, and Monte Carlo analysis for technology advancement uncertainty -- are used to perform these analyses. To gain a more complete view of the effects of an on-orbit propellant refueling capability, a reference Mars mission is developed and compared to an equivalent mission without refueling capability. Finally, the possibility of propellant resupply in Mars orbit is also discussed along with its implications on the sustainability of a long-term Mars exploration architecture.
Improving Lunar Return Entry Footprints Using Enhanced Skip Trajectory Guidance

(Georgia Institute of Technology, 2006-09) Putnam, Zachary R. ; Braun, Robert D. ; Bairstow, S. H. ; Barton, G. H.

The impending development of NASA's 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.
Architecture Options for Propellant Re-supply of Lunar Exploration Elements

(Georgia Institute of Technology, 2006-09) Young, James J. ; Thompson, Robert W. ; Wilhite, Alan W.
Sizing of an Entry, Descent, and Landing System for Human Mars Exploration

(Georgia Institute of Technology, 2006-09) Christian, John A., III ; Wells, Grant William ; Lafleur, Jarret M. ; Manyapu, Kavya ; Verges, Amanda ; Lewis, Charity ; Braun, Robert D.

The human exploration of Mars presents many challenges, not least of which is the task of entry, descent, and landing (EDL). Because human-class missions are expected to have landed masses on the order of 40 to 80 metric tons, significant challenges arise that have not been seen to date in robotic missions. This study provides insight into the challenges encountered as well as potential solutions through parametric trade studies on vehicle size and mass. Aerocapture and entry-from-orbit analyses of 10 and 15 m diameter aeroshells with a lift-to-drag ratio of 0.3 or 0.5 were investigated. Results indicate that in the limit, a crew capsule used only for descent could have an initial mass as low as 20 t. For larger landed payloads, such as a 20 t surface power system, a vehicle with an initial mass on the order of 80 t may be required. In addition, no feasible EDL systems were obtained with the capability to deliver more than approximately 25 t of landed payload to the Mars surface for initial masses less than 100 t. This suggests that an aeroshell diameter of 15 m may not be sufficient for human Mars exploration.
Evaluation of Applications of ESAS Cargo Launch Vehicle to Manned Mars Mission

(Georgia Institute of Technology, 2006-09) Simon, Matthew ; Wilhite, Alan W. ; Young, James J.

This paper reports on an investigation of possible Launch and Trans-Mars Injection options using the Exploration Systems Architecture Study (ESAS) Cargo Launch Vehicle (CaLV) and Earth Departure Stages (EDS) for crewed Mars missions The purpose of such an investigation is to characterize some of the potential challenges and technological needs for modifying the present lunar architecture for transport to Mars. An analysis was performed to provide a relationship between the payload placed in orbit, usable propellant available in orbit, and the required propellant to perform the Trans-Mars Injection. This relationship was used to investigate several payload manifests and launch strategy options and to perform some trades to identify advantageous launch solutions. The Mars Design Reference Mission from 1998 was used to provide a representative Mars payload and mission architecture for the study. It was found that 6 and 5 launch solutions are possible without the implementation of any new technologies. It was also found that adding the ability to efficiently transfer propellants between Earth Departure Stages and developing boil-off elimination technologies improves performance to 4-launch solutions. Further reduction of launches exceeds constraints on the launch vehicle such as payload capacity and height constraints.