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 55

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.
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.
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.
An Evaluation of Ballute Entry Systems for Lunar Return Missions

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

This study investigates the advantages and feasibility of using ballutes for Earth entry at lunar return velocities. Using analysis methods suitable for conceptual design and assuming a CEV type entry vehicle, multiple entry strategies were investigated. Entries that jettison the ballute after achieving low Earth orbit conditions were shown to reduce heating rates to within reusable thermal protection system limits. Deceleration was mitigated to approximately four g's when a moderate amount of lift was applied subsequent to ballute jettison. Primary ballute size drivers are 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. Lastly, preliminary investigation of a lower mass cargo variant of the CEV allowed for additional reduction of ballute system mass. However, ballute system mass as a percentage of the total entry mass was shown to be relatively independent of the entry mass.
Ultra Lightweight Ballutes for Return to Earth from the Moon

(Georgia Institute of Technology, 2006-05) Masciarelli, James P. ; Lin, John K. H. ; Ware, Joanne S. ; Rohrschneider, Reuben R. ; Braun, Robert D. ; Bartels, Robert E. ; Moses, Robert W. ; Hall, Jeffery L.

Ultra lightweight ballutes offer revolutionary mass and cost benefits along with flexibility in flight system design compared to traditional entry system technologies. Under funding provided by NASA's Exploration Systems Research & Technology program, our team was able to make progress in developing this technology through systems analysis and design, evaluation of materials and construction methods, and development of critical analysis tools. Results show that once this technology is mature, significant launch mass savings, operational simplicity, and mission robustness will be available to help carry out NASA's Vision for Space Exploration.
Static Aeroelastic Analysis of a Thin-Film Clamped Ballute for Titan Aerocapture

(Georgia Institute of Technology, 2006-05) Rohrschneider, Reuben R. ; Braun, Robert D.

Many authors have shown the potential mass savings that a ballute can offer for both aerocapture and entry. This mass savings could enhance or even enable many scientific and human exploration missions. Prior to flight of a ballute several technical issues need to be addressed, including aeroelastic behavior. This paper begins to address the issue of aeroelastic behavior by developing and validating the Ballute Aeroelastic Analysis Tool (BAAT). The validation effort uses wind tunnel tests of clamped ballute models constructed of Kapton supported by a rigid nose and floating aft ring. Good correlation is obtained using modified Newtonian aerodynamics and non-linear structural analysis with temperature dependent material properties and thermal expansion. BAAT is then used to compute the deformed shape of a clamped ballute for Titan aerocapture in both the continuum and transitional regimes using impact method aerodynamics and direct simulation Monte Carlo.