Organizational Unit:

Aerospace Systems Design Laboratory (ASDL)

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https://hdl.handle.net/1853/70744

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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/1135

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Publication Search Results

Now showing 1 - 10 of 65

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.
Mars Entry, Descent, and Landing Parametric Sizing and Design Space Visualization Trades

(Georgia Institute of Technology, 2006-08) Alemany, Kristina ; Wells, Grant William ; Theisinger, John ; Clark, Ian G. ; Braun, Robert D.

Entry, descent, and landing (EDL) is a multidimensional, complex problem, which is difficult to visualize in simple plots. The purpose of this work is to develop a systematic visualization scheme that could capture Mars EDL trades as a function of a limited number of variables, such that programmatic design decisions could be effectively made with insight of the design space. Using the Mars Science Laboratory (MSL) as a basis, contour plots have been generated for key EDL figures of merit, such as maximum landed elevation and landed mass as a function of four input parameters: entry mass, entry velocity, entry flight path angle, and vehicle L/D. Additionally, sensitivity plots have been generated in an attempt to capture the effects of varying the fixed input parameters. This set of EDL visualization data has been compiled into a Mars EDL handbook to aid in pre-phase A design space exploration and decision making.
Trajectory Options for Human Mars Missions

(Georgia Institute of Technology, 2006-08) Wooster, Paul D. ; Braun, Robert D. ; Ahn, Jaemyung ; Putnam, Zachary R.

This paper explores trajectory options for the human exploration of Mars, with an emphasis on conjunction-class missions. Conjunction-class missions are characterized by short in-space durations with long surface stays, as opposed to the long in-space durations and short surface stays characteristic of opposition-class missions. Earth-Mars and Mars-Earth trajectories are presented across a series of mission opportunities and transfer times in order to explore the space of possible crew and cargo transfer trajectories. In the specific instance of crew transfer from Earth to Mars, the potential for aborting the mission without capture into Mars orbit is also of interest. As such two additional classes of trajectories are considered: free-return trajectories, where the trajectory would return the crew to Earth after a fixed period of time; and propulsive-abort trajectories, where the propulsive capability of the transfer vehicle is used to modify the trajectory during a Mars swing-by. The propulsive requirements of a trajectory, due to their associated impact on spacecraft mass, are clearly of interest in assessing trajectories for human Mars missions. Beyond the propulsive requirements, trajectory selection can have a significant impact on the entry velocity and therefore the aeroassist system requirements. The paper suggests potential constraints for entry velocities at Earth and Mars. Based upon Mars entry velocity, the 2-year period free-return abort trajectory is shown to be less desirable than previously considered for many mission opportunities.