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 141

A Method for Modeling System-Driven Uncertainty during Probabilistic Part Life Analyses

(Georgia Institute of Technology, 2004-11) Wallace, Jon Michael ; Volovoi, Vitali V. ; Mavris, Dimitri N.

Probabilistic part life analyses of turbine components have typically been conducted in an ad-hoc fashion with respect to the influence of the system. While this approach greatly simplifes the analysis, signifcant errors and misleading results are possible. However, directly modeling the system analyses in a fully probabilistic and integrated fashion can be prohibitive in terms of the infrastructure required. An effcient approach to characterizing and quantifying the system-driven input for probabilistic part life assessments is proposed. The approach is demonstrated for a turbine blade operating in a medium size commercial transport jet. The results of this demonstra- tion illustrate how the component parameters and failure mechanisms can be qualitatively identifed and the complex probabilistic input modeled as driven by the system behavior.
A Methodology for Assessing Business Models of Future Air Transportation in the Atlanta Regional Transportation System

(Georgia Institute of Technology, 2004-09) Lim, Choon Giap ; Lewe, Jung-Ho ; DeLaurentis, Daniel A. ; Mavris, Dimitri N.

A methodology employing physics-based and economics-based tools in conjunction with probabilistic treatment is developed to study Personal Air Vehicle business model. In the context of the paper, a business model is a mathematical representation of a service provider business operation. Vehicle concepts and hypothesized metrics such as mobility freedom and 'value of time'are embedded in the methodology. Market behavior of the complex transportation environment is captured as part of the equation through Agent-based Modeling and Monte Carlo Simulation techniques. This simulation platform for the transportation environment facilitates the case study of the Atlanta Regional Transportation System. The establishment of this model lays the foundation for creating a robust and adaptive design methodology that allows experts in fields other than aerospace engineering to contribute their expertise towards the realization of this very diverse and dynamic future air transportation system.
Use of Probability of Success as an Independent Variable for Decision-Making

(Georgia Institute of Technology, 2004-09) Frits, Andrew P. ; Mavris, Dimitri N.

Early phases of design are characterized by risk and uncertainty. Appropriate accounting for this uncertainty is an important requirement for any designer. This work suggests collapsing risk and uncertainty into a single metric called the probability of success, which accounts for the probability of a given design simultaneously meeting all of the design requirements. Optimal or lowest cost designs can then be found for various levels of probability of success. These designs can be compared to each other, creating a trade-off between the cost of a design and its risk. These risk versus cost figures can be generated before a decision-maker commits to the design. Thus, the decision-maker will have all the information regarding the cost and risk of potential designs before making any design decisions. The decision-maker can thus treat the probability of success, or risk, as an independent variable, choosing the level of risk that he or she finds acceptable based upon the cost of the system, with the corresponding
Flight System Options for a Long Duration Mars Airplane

(Georgia Institute of Technology, 2004-09) Rohrschneider, Reuben R. ; Olds, John R. ; Kuhl, Christopher A. ; Braun, Robert D. ; Steffes, Stephen R. ; Hutchinson, Virgil L., Jr.

The goal of this study was to explore the flight system options for the design of a long endurance Mars airplane mission. The mission model was built in the design framework ModelCenter and a combination of a hybrid and user-driven fixed point iteration optimization method was used to determine the maximum endurance solution of each configuration. Five different propulsion systems were examined: a bipropellant rocket, a battery powered propeller, a direct methanol fuel cell powered propeller, and beamed solar and microwave powered propeller systems. Five airplane configurations were also studied. The best configuration has a straight wing with two vertical tails. The direct methanol fuel cell proved to be the best onboard power system for a long endurance airplane and the solar beamed power system showed potential for indefinite flight. The combination of the best configuration and the methanol fuel cell resulted in an airplane capable of cruising for 17.8 hours on Mars.
Inclusion of Tactical Considerations for System-of-Systems Optimization of Torpedoes

(Georgia Institute of Technology, 2004-08) Frits, Andrew P. ; Weston, Neil R. ; Mavris, Dimitri N.

In the current torpedo design process, torpedoes are often designed independently from the tactics with which they are employed. This serial design process, of first developing tactics, then designing the torpedo, then re-developing tactics leads to torpedo designs that are sub-optimal when viewed from the greater system-of-systems perspective. This paper looks at the effects that tactics have on the design of torpedoes. It proposes a new paradigm, of simultaneous tactics development and torpedo design, and looks at the implications of various tactics on the optimal design of torpedo systems.
Benefits of Non-Dimensionalization in Creation of Designs of Experiments for Sizing Torpedo Systems

(Georgia Institute of Technology, 2004-08) Frits, Andrew P. ; Reynolds, Kristen ; Weston, Neil R. ; Mavris, Dimitri N.

Non-dimensionalization is useful at many stages in the conceptual design process. One area of usefulness is in the creation and execution of Design of Experiments. A Design of Experiments that is run with dimensional quantities can often have a large number of failed or infeasible cases or require frustratingly small ranges on the design variables in order to execute cleanly. However, with the use of non-dimensional parameters in the Design of Experiments, the dimensional values being used in the analysis tool automatically scale themselves so that appropriate magnitudes of each parameter are always being used. This automatic scaling greatly increases the stability of Design of Experiments when non-dimensional parameters are used, limiting the number of failed cases. This paper explores potential non-dimensional parameters for use in the conceptual design of torpedo systems. The paper shows that traditional non-dimensional parameters used in propulsor design, such as advance ratio and thrust coefficient, also work well as torpedo design parameters. A short example is given where the performance of a Design of Experiments for a torpedo system is improved via the use of non-dimensional parameters.
Estimation of Launch Vehicle Propellant Tank Structural Weight Using Simplified Beam Approximation

(Georgia Institute of Technology, 2004-07-11) Olds, John R. ; Hutchinson, Virgil L., Jr.

Many conceptual launch vehicles are designed through the integration of various disciplines, such as aerodynamics, propulsion, trajectory, weights, and aeroheating. In the determination of the total vehicle weight, a large percentage of the vehicle weight is composed of the structural weight of the vehicle subsystems, such as propellant tanks. Empirical mass estimating relations (MERs) and multi-dimensional finite element analysis (FEA) are two methods commonly used by the aerospace industry to estimate the loadbearing structural weight. MERs rapidly estimate the weight by evaluating empirical equations and the high-fidelity techniques of FEA accurately calculates the structural weight. The extreme inability for either method to provide both rapid and accurate weight estimations warrants an investigation into developing an improved, intermediate method. A methodology based on fundamental beam structural analysis has been developed for the rapid estimation of the load-bearing structural weight of the launch vehicle fuselage and integral propellant tanks. By creating a simplified beam approximation model of the vehicle, the method utilizes the vehicle component weights, load conditions, and basic material properties to analytically estimate the structural shell and stability frame weight. Implementation of this methodology into a fast-acting software tool allowed for rapid estimation of the component structural weight. Using statistical techniques, an empirical relationship between the estimated and actual load-bearing structure weights was determined. The method was utilized to estimate the liquid hydrogen (LH2) and liquid oxygen (LOX) propellant tanks for an existing Evolved Expendable Launch Vehicle (EELV) and the Space Shuttle External Tank (ET) for verification and correlation.
Aztec: A TSTO Hypersonic Vehicle Concept Utilizing TBCC and HEDM Propulsion Technologies

(Georgia Institute of Technology, 2004-07) Kokan, Timothy Salim ; Hutchinson, Virgil L., Jr. ; Reeves, John Daniel ; Olds, John R.

The Aztec reusable launch vehicle (RLV) concept is a two-stage-to-orbit (TSTO) horizontal takeoff, horizontal landing (HTHL) vehicle. The first stage is powered by ten JP- 5 fueled turbine-based combined-cycle (TBCC) engines. The second stage is powered by three high energy density matter (HEDM)/liquid oxygen (LOX) staged-combustion rocket engines. The HEDM fuel is a liquid hydrogen-based propellant with a solid aluminum and methane gel additive. Aztec is designed to deliver 20,000 lbs of payload to a 100 nmi x 100 nmi x 28.5 deg orbit due East out of Kennedy Space Center (KSC). The second stage separates at Mach 8 and continues to the target orbit while the first stage flies back to KSC in ramjet mode. For the above payload and target orbit, the gross lift-off weight (GLOW) is estimated to be 690,000 lbs and the total dry weight for both stages is estimated to be 230,000 lbs. Economic analysis indicates that the Aztec recurring launch costs will be approximately 590 dollars per lb. of payload delivered to the target orbit. The total non-recurring cost including design, development, testing and evaluation (DDT&E), acquisition of the first vehicle, and the construction of launch and processing facilities is expected to be 13.6 B dollars. All cost figures are in FY2004 unless otherwise noted. Details of the Aztec design including external and internal configuration, aerodynamics, mass properties, first and second stage engine performance, ascent and flyback trajectory, aeroheating results and thermal protection system (TPS), vehicle ground operations, vehicle safety and reliability, and a cost and economics assessment are provided.
StarRunner: A Single-Stage-to-Orbit, Airbreathing, Hypersonic Propulsion System

(Georgia Institute of Technology, 2004-07) Biltgen, Patrick Thomas ; Lafleur, Jarret M. ; Loughman, Josh ; Martin, Robert ; Flaherty, Kevin W. ; Cho, Min ; Becker, Keith ; Ong, Chester ; Olds, John R.

In response to the request for proposal (RFP) for the 2003 AIAA Undergraduate Team Engine Design Competition, the FAS Propulsion Design team from the Georgia Institute of Technology presents StarRunner: A Single-Stage-to-Orbit (SSTO), Airbreathing, Hypersonic Propulsion System. Low-cost, highly reliable access to low-Earth orbit (LEO) and the International Space Station (ISS) is an area of continuing research and debate. StarRunner is proposed to supplement a notional Crew Transfer Vehicle through the ability to deliver a 25,000 lb payload to the ISS. The horizontal takeoff/horizontal landing (HTHL) vehicle makes use of a turbine-based combined cycle (TBCC) propulsion system consisting of 14 low-bypass-ratio turbofan engines and a dual-mode ramjet/scramjet propulsion system for high-speed flight. The vehicle also takes advantage of ultra-high-temperature ceramic thermal protection materials and uses hydrogen fuel for regenerative cooling of engine components. StarRunner is compatible with standard runways, with a gross takeoff weight of approximately 1,000,000 lbs, and has a cost per pound to orbit of approximately $825/lb. This advanced, fully reusable space transport vehicle and integrated propulsion system design demonstrates student efforts to understand issues facing the space launch community. Future enabling and enhancing technologies for TBCC SSTO launch vehicles are explored and analyzed. The final StarRunner design addresses and proposes several innovative solutions to traditional problems.
Centurion: A Heavy-Lift Launch Vehicle Family for Cis-Lunar Exploration

(Georgia Institute of Technology, 2004-07) Young, David Anthony ; Olds, John R. ; Hutchinson, Virgil L., Jr. ; Krevor, Zachary C. ; Pimentel, Janssen ; Reeves, John Daniel ; Sakai, Tadashi ; Young, James J.

Centurion is an expendable heavy lift launch vehicle (HLLV) family for launching lunar exploration missions. Each vehicle in the family is built around a common two-stage core. The first stage of the core uses kerosene (RP-1) fuel and utilizes four staged-combustion RD-180 rocket engines. The upper stages consist of liquid oxygen (LOX)/liquid hydrogen (LH2) propellant with three 220,000 lb thrust-class expander rocket engines. The larger variants in the Centurion family will also use either one or two pairs of five-segment solid rocket motors which are now being developed by ATK Thiokol. The Centurion family consists of three vehicles denoted as C-1, C-2, and C-3. The first vehicle (C-1) is a four RD-180 core with a LOX/LH2 upper stage. The C-1 is designed to deliver a 35 metric ton (MT) CEV to a 300 km x 1000 km highly elliptical orbit (HEO). This HEO allows the CEV to more easily transfer to a lunar trajectory, while still having the ability to abort after one revolution. The C-1 also is designed to meet mission requirements with a failure of both one RD-180 and one upper stage engine. The C-2 and C-3 Centurions are both cargo carrying variants which carry 100 MT and 142 MT of cargo to a 407 km low earth orbit (LEO) respectively. The C-2 utilizes two five-segment solid rocket boosters (SRB), while the C-3 uses four SRBs. Details of the conceptual design process used for Centurion are included in this paper. The disciplines used in the design include configuration, aerodynamics, propulsion design and selection, trajectory, mass properties, structural design, aeroheating and thermal protection systems (TPS), cost, operations, and 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 team process and used to minimize the gross weight of the C-1 variant. The C-2 and C-3 variants were simulated using the C-1 optimized core with different configurations of SRBs. Each of the variants recurring and non-recurring costs were computed. The total development cost including the design, development, testing and evaluation (DDT&E) cost and a new launch pad at Kennedy Space Center (KSC), was 7.98 B FY04 dollars. The theoretical first unit (TFU) cost for the C-2 variant was 532 M FY04 dollars. A summary of design disciplines as well as the economic results are included.