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 19

Tanker Argus: Re-supply for a LEO Cryogenic Propellant Depot

(Georgia Institute of Technology, 2002-10) St. Germain, Brad David ; Kokan, Timothy Salim ; Marcus, Leland R. ; Miller, Jeff ; Rohrschneider, Reuben R. ; Staton, Eric ; Olds, John R.

The Argus reusable launch vehicle (RLV) concept is a single-stage-to-orbit (SSTO) conical, wingedbodied vehicle powered by two liquid hydrogen (LH2)/liquid oxygen (LOX) supercharged ejector ramjets (SERJ). The 3rd generation Argus launch vehicle utilizes advanced vehicle technologies along with a magnetic levitation (Maglev) launch assist track. A tanker version of the Argus RLV is envisioned to provide an economical means of providing liquid fuel and oxidizer to an orbiting low Earth orbit (LEO) propellant depot. This depot could then provide propellant to various spacecraft, including reusable orbital transfer vehicles used to ferry space solar power (SSP) satellites to geo-stationary orbit. Two different tanker Argus configurations were analyzed. The first simply places additional propellant tanks inside the payload bay of an existing Argus reusable launch vehicle. The second concept is a modified pure tanker version of the Argus RLV in which the payload bay is removed and the vehicle propellant tanks are extended to hold additional propellant. An economic analysis was performed for this study that involved the calculation of the design/development and recurring costs of each vehicle. The goal of this analysis was to determine at what flight rate it would be economically beneficial to spend additional development funds to change an existing, sunk cost, payload bay tanker vehicle into a pure tanker design. The results show that for yearly flight rates greater than ~50 flts/yr it is cheaper, on a $/lb basis , to develop and operate a dedicated tanker.
Solar Electric Propulsion Module Concept for the BiFrost Architecture

(Georgia Institute of Technology, 2002-10) Rohrschneider, Reuben R. ; Sakai, Tadashi ; Steffes, Stephen R. ; Grillmayer, Georg ; St. Germain, Brad David ; Olds, John R.

This paper describes the design of a solar electric propulsion module for the Bifrost architecture. Bifrost consists of a magnetic levitation launch tube with the exit end elevated to 20 km. A 35,000 kg hybrid logistics module (HLM) is designed to attach to an array of propulsion modules that accommodate different missions. The solar electric propulsion (SEP) module is designed to circularize a payload in Geosynchronous Earth orbit (GEO) from a highly elliptic transfer orbit. A configuration consisting of a central spacecraft body propelling itself with electric thrusters and gathering solar power from two inflatable concentrating reflectors was chosen. Concentrating reflectors were chosen over thin film arrays due to the large mass savings. Details of the conceptual design process are presented. Disciplines include trajectory, power system, propulsion, and weights & sizing. A computational framework was used to wrap the disciplinary analysis to speed the design process, and optimization was performed to minimize the initial mass of the vehicle from within the design framework. The resulting vehicle has an initial mass in orbit of 40,780 kg. A demonstration model was then designed and constructed from the conceptual design. The manufacturing process for the inflatable reflector and the spacecraft body are described in detail. The demonstration model shows that an inflatable reflector is a feasible method of generating large amounts of power in space.
A Design of Experiments-Based Method for Point Selection in Approximating Output Distributions

(Georgia Institute of Technology, 2002-09) McCormick, David Jeremy ; Olds, John R.

The goal of this research is to find a computationally efficient and easy-to-use alternative to current approximation or direct Monte Carlo methods for robust design. More specifically, a technique is sought to use selected deterministic analyses to obtain probability distributions for analyses with large inherent uncertainties. Previous research by the authors has presented a promising class of methods known as Discrete Probability Matching Distributions (DPOMD). This paper introduces a new type of DPOMD better suited to problems with larger numbers of random variables. This new type utilizes a fractional factorial design of experiments array in combination with an inverse Hasofer-Lind standard normal space transform. The method defines points in the problem space that represent the moment characteristics of the input random variables. This new method is compared to two other approximation techniques, Descriptive Sampling and Response Surface/Monte Carlo Simulation, for three common aerospace analyses (Mass Properties and Sizing, Propulsion Analysis and Trajectory Simulation). A Monte Carlo analysis with corresponding error bands is used for reference. Preferences for probabilistic analysis each of these problems are determined based on the speed and accuracy of analysis. These results are presented here. The new DPOMD technique is shown to be advantageous in terms of speed and accuracy for two of the three problems tested.
A Distributed Framework for Probabilistic Analysis

(Georgia Institute of Technology, 2002-09) McCormick, David Jeremy ; Olds, John R.

Probabilistic multidisciplinary design optimization promises to incorporate critical design uncertainty in order to create optimal products with a high probability of meeting design constraints under a wide variety of circumstances. Several methods of accelerated probability analysis are available to designers. What is not available is a formal method for tying contributing analysis-level probability analysis into an integrated design framework capable of optimization. This would allow probability methods to be tailored to the characteristics of a particular contributing analysis as well as potentially reduce the dimensionality of the problems considered. This research presents such a method, and then tests it on a conceptual launch vehicle design problem. This probabilistic optimization problem consisted of 84 noise variables and four design variables. This problem setup consistently found system optimums in 6-8 hrs. It utilized several probability approximation methods run in an iterative manner to generate probabilistic vehicle sizing information. Once the probabilistic optimum was identified and confirmed using this process, a system-level Monte Carlo random simulation of the vehicle design was conducted around the optimum point to confirm the accuracy of the distributed approximation method. Because this simulation was prohibitively expensive, it was only conducted at the single optimum point. Following this accuracy confirmation, a comparison to a deterministic optimization of the same problem illustrated the difference between probabilistic and deterministic optimums.
Mission Capture Rate versus Turnaround Time and Fleet Size for the Military Spaceplane

(Georgia Institute of Technology, 2002-07) Kokan, Timothy Salim ; Olds, John R.

The United States Air Force Research Laboratory (AFRL) is conducting research into a military spaceplane (MSP) through the Military Spaceplane System Technology Program Office. The goal of this program is to provide the Air Force with safe, reliable, affordable, and routine access to space. An important mission performance metric of the MSP program is the mission capture rate. The mission capture rate is a measure of the MSP’s ability to meet mission sortie requirements. Extending this to a fleet of MSPs, the mission capture rate is defined as the total number of sorties the fleet is capable of divided by the total required number of sorties. This research analyzes the relationship between mission capture rate and both turnaround time and fleet size. The turnaround time is the time between when the vehicle lands and when it can take off again. During this time the vehicle is refueled, maintenance and repair work is done, and the payload is loaded. As turnaround time decreases and fleet size increases, the mission capture rate will increase. A precise definition of this relationship is made in order to determine the necessary fleet size for a given turnaround time subject to a desired mission capture rate. A Monte Carlo simulation is performed to probabilistically analyze the mission capture rates. This analysis takes into account uncertainties in the utilization requirements of the MSP fleet. These uncertainties include the number of wars within the simulation period, the starting date & duration of each war, and each war’s required sortie rate. This analysis utilizes Crystal Ball Pro® along with Microsoft Excel®. This gives the analysis technique compatibility with commonly used computer platforms.
Technology Assessment for Manned Mars Exploration Using a ROSETTA Model of a Bimodal Nuclear Thermal Rocket (BNTR)

(Georgia Institute of Technology, 2001-08) Way, David Wesley ; Medlin, Matt ; Sakai, Tadashi ; McIntire, James ; Olds, John R.

This paper investigates a new method of measuring the affordability of aerospace technologies. First, a new bimodal NTR Mars mission architecture was defined. Starting with brainstorming on the different ways to get to Mars, several different trade studies were investigated, the results of which defined the architecture. A Reduced-Order Simulation for Evaluating Technologies and Transportation Architectures (ROSETTA) models has been created from this architecture. This model is an Excel workbook of interconnected worksheets that represented the different disciplines used in creating the architecture. Each worksheet is based on the results of higher fidelity codes such as the Program to Optimize Simulated Trajectories (POST) and the Aerodynamic Preliminary Analysis System (APAS). These results were then reduced to simpler, parametric relations, giving rise to the 'Reduced Order' in ROSETTA. The BNTR ROSETTA model is capable of rapidly resizing the Mars transfer vehicle and landers and estimating the key cost and mass metrics as the input technology assumptions change. Future technology assessment will be done probabilistically, by assigning a distribution to each input parameter that the technology affects, then running a Monte Carlo analysis in order to generate an output distribution for each metric. Benefit-to-cost ratios and top-level uncertainties can be determined from this data.
A Comparison of Modern and Historic Mass Estimating Relationships on a Two-Stage to Orbit Launch Vehicle

(Georgia Institute of Technology, 2001-08) Rohrschneider, Reuben R. ; Olds, John R.

Traditionally mass estimation for conceptual design of advanced launch vehicles has depended on historically based mass estimating relationships (MERs). This paper compares the modern MERs used in the Space Systems Design Lab (SSDL) at Georgia Tech to the 1960s era relationships used in the NAS7-377 report on advanced propulsion design for launch vehicles. Comparisons of the weight breakdowns of a two-stage-to-orbit vehicle are made for between the Marquardt equations and the SSDL equations using two different technology assumptions. The first assumes 1970 technology for a direct comparison of the equations, while the second assumes 2015 technology. Additionally, technology and material advances are estimated in an attempt to justify the lower weight of the 2015 technology. The SSDL model using 1970 technology weighs in 7 percent heavier than the Marquardt equations for a comparable two-stage-to-orbit vehicle. When 2015 technology is applied to the same vehicle SSDL, equations show that a 33 percent savings, on the entire vehicle, could be made due to technology.
ASPEN Revisited: The Challenge of Nuclear Propulsion for ETO

(Georgia Institute of Technology, 2001-07) Weglian, John E. ; Olds, John R. ; Marcus, Leland R. ; McIntire, James ; Nelson, Douglas K. ; Blevins, John

ASPEN was a study conducted by Los Alamos National Labs in the early 1960s to examine the benefits of using a Nuclear Thermal Rocket (NTR) for Earth-to-Orbit (ETO) single-stage launch vehicle applications. Using the analysis methods and assumptions of the time, this formerly classified study showed that a significant performance potential might be derived from using NTR engines for the final acceleration phase to orbit (air-breathing engines were used to Mach 11). Given the increased NASA interest in low-cost reusable space transportation, the ASPEN concept has been revisited using contemporary design assumptions and conceptual analysis techniques. The present analysis concludes with a more pessimistic view of NTR propulsion for ETO applications. Aerodynamic drag for the ASPEN configuration was found to be significantly more than that calculated in the original study. The resultant vehicle thrust-to-drag ratio is lower than necessary for high acceleration during the air-breathing acceleration phases. In addition, the NTR reactor power requirements are daunting. In most cases, reactor powers over 10 GW are required. Even with very aggressive assumptions (25 percent drag reduction and NTR thrust-to-weight ratio of 10, including shielding) a 500,000 lb gross weight ASPEN-like vehicle was found to only have a payload mass fraction of 1.6 percent. This is significantly less than the 6 to 15 percent payload mass fractions claimed in the original ASPEN study.
Moon-based Advanced Reusable Transportation Architecture

(Georgia Institute of Technology, 2001-07) Nelson, Douglas K. ; Marcus, Leland R. ; Bechtel, Ryan S. ; Cormier, Timothy A. ; Weglian, John E. ; Alexander, R.

Abstract: The Moon-based Advanced Reusable Transportation Architecture (MARTA) Project conducted an in-depth investigation of possible Low Earth Orbit (LEO) to lunar surface transportation systems capable of sending both astronauts and large masses of cargo to the Moon and back. The goal of this project was to create a profitable venture with an Internal Rate of Return (IRR) of 25%. The architecture was quickly narrowed down to a traditional chemical rocket using a liquid oxygen and liquid hydrogen. However, three additional technologies identified as potentially cost saving were: aerobraking, in-situ resource utilization (ISRU), and a mass driver on the lunar surface. The vehicle was modeled using the Simulated Probabilistic Parametric Lunar Architecture Tool (SPPLAT) that incorporated several different engineering disciplines. This tool used ISRU propellant cost, a dry weight reduction due to improved materials technology, and vehicle engine specific impulse as inputs and provides vehicle dry weight, total propellant used per trip, and price to charge the customer in order to guarantee an IRR of 25% as outputs. Estimation error, market growth, and launch cost uncertainty were also considered. The results of the project show that the desired operation is possible using current technology. Based on the stipulation that the venture be profitable, the price to charge the customer was highly dependent on ISRU propellant cost and relatively insensitive to the other inputs. With the best estimate of ISRU cost set at $1000/kg, the resulting price to charge the customer was $2600/kg of payload from LEO to the lunar surface. If ISRU cost can be reduced to $160/kg, the price to the customer is reduced to just $800/kg of payload. Additionally, the mass driver only proved to be cost effective at an ISRU propellant cost greater than $250/kg, although it reduced total propellant used by 35%.
Starsaber: A Small Payload-Class TSTO Vehicle Concept Utilizing Rocket-Based Combined Cycle Propulsion

(Georgia Institute of Technology, 2001-07) St. Germain, Brad David ; Olds, John R. ; McIntire, James ; Nelson, Douglas K. ; Weglian, John E. ; Ledsinger, Laura Anne

This paper introduces Starsaber, a new conceptual launch vehicle design. Starsaber is a two-stage-to-orbit (TSTO) vehicle capable of putting a 300 lb class payload into low Earth orbit (LEO). The vehicle is composed of a reusable winged booster, powered by two hydrocarbon fueled ejector ramjet (ERJ) engines, and a LOX/RP-1 expendable upper stage. The vehicle utilizes advanced structural and thermal protection system (TPS) materials, as well as advanced subsystems. Details of the conceptual design process used for Starsaber are given in this paper. Disciplines including mass properties, internal and external configuration, aerodynamics, propulsion, trajectory simulation, aeroheating, and cost estimation are used in this study. A baseline design was generated, and a 2-level 15-variable Taguchi L16 array was used to determine key system variables' influence on vehicle weight and cost. Based on these preliminary results, the Starsaber vehicle was optimized for both minimum weight (gross and dry weight) and recurring cost. The lowest recurring cost vehicle was estimated to have a recurring cost per flight of $2.01M, a gross liftoff weight of 168,000 lb and a booster length of 77 ft.