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Saleh,
Joseph H.
Saleh,
Joseph H.
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ItemRegression Analysis of Launch Vehicle Payload Capability for Interplanetary Missions(Georgia Institute of Technology, 2010-09) Wise, Marcie A. ; Lafleur, Jarret M. ; Saleh, Joseph H.During the conceptual design of interplanetary space missions, it is common for engineers and mission planners to perform launch system trades. This paper provides an analytical means for facilitating these trades rapidly and efficiently using polynomial equations derived from payload planner’s guides. These equations model expendable launch vehicles’ maximum payload capability as a function of vis-viva energy (C3). This paper first presents the motivation and method for deriving these polynomial equations. Next, 34 polynomials are derived for vehicles among nine launch vehicle series: Atlas V, Delta IV, Falcon 9, and Taurus, as well as H-IIA, Long March, Proton, Soyuz, and Zenit. The quality of fit of these polynomials are assessed, and it is found that the maximum 95th percentile model fit error for all 34 vehicles analyzed is 4.43% with a mean of 1.44%, and the minimum coefficient of determination (R²) is 0.99967. As a result, the equations are suitable for launch vehicle trade studies in conceptual design and beyond. A realistic example of such a trade for the Mars Reconnaissance Orbiter mission is provided.
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ItemComparative Reliability of GEO, LEO, and MEO Satellites(Georgia Institute of Technology, 2009-10) Hiriart, Thomas ; Castet, Jean-Francois ; Lafleur, Jarret M. ; Saleh, Joseph H.Reliability has long been a major consideration in the design of space systems, and in recent years it has become an essential metric in spacecraft design trade-space exploration and optimization. The purpose of this paper is to statistically derive and compare reliability results of Earth-orbiting satellites as a function of orbit type, namely geosynchronous orbits (GEO), low Earth orbits (LEO) and medium Earth orbits (MEO). Using an extensive database of satellite launches and failures/anomalies, life data analyses are conducted over three samples of satellites within each orbit type and successfully launched between 1990 and 2008. Because the dataset is censored, the Kaplan-Meier estimator is used to estimate the reliability functions. Plots of satellite reliability as a function of orbit altitude are provided for each orbit type, as well as confidence bounds on these estimates. Using analytical techniques such as maximum likelihood estimation (MLE), parametric fits are conducted on the previous nonparametric reliability results using single Weibull and mixture distributions. Based on these parametric fits, a comparative reliability analysis is provided identifying similarities and differences in the reliability behaviors of satellites in these three types of orbits. Finally, beyond the statistical analysis, this work concludes with several hypotheses for structural/causal explanations of these trends and difference in on-orbit failure behavior.
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ItemResponse Surface Equations for Expendable Launch Vehicle Payload Capability(Georgia Institute of Technology, 2009-09) Fleming, Elizabeth S. ; Lafleur, Jarret M. ; Saleh, Joseph H.Systems analysis and conceptual design for new spacecraft commonly require the capability to perform rapid, parametric assessments of launch vehicle options. Such assessments allow engineers to incorporate launch vehicle considerations in first-order cost, mass, and orbit performance trades early during conceptual design and development phases. This paper demonstrates an efficient approach to launch vehicle analysis and selection using response surface equations (RSEs) derived directly from launch vehicle payload planner's guides. These RSEs model payload capability as a function of circular orbit altitude and inclination. Following presentation of the RSE fitting method and statistical goodness of fit tests, the RSE and model fit error statistics for the Pegasus XL are derived and presented as an example. In total, 43 RSEs are derived for the following launch vehicles and their derivatives: Pegasus, Taurus, Minotaur, and Falcon series as well as the Delta IV, Atlas V, and the foreign Ariane and Soyuz vehicles. Ranges of validity and model fit error statistics with respect to the original planner's guide data are provided for each of the 43 fits. Across all launch vehicles fit, the resulting RSEs have a maximum 90th percentile model fit error of 4.39% and a mean 90th percentile model fit error of 0.97%. In addition, of the 43 RSEs, the lowest R^2 value is 0.9715 and the mean is 0.9961. As a result, these equations are sufficiently accurate and well-suited for use in conceptual design trades. Examples of such trades are provided, including demonstrations using the RSEs to (1) select a launch vehicle given an orbit inclination and altitude, (2) visualize orbit altitude and inclination constraints given a spacecraft mass, and (3) calculate the sensitivity of orbital parameters to mass growth. Suited for a variety of applications, the set of RSEs provides a tool to the aerospace engineer allowing efficient, informed launch option trades and decisions early during design.
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ItemGT-FAST: A Point Design Tool for Rapid Fractionated Spacecraft Sizing and Synthesis(Georgia Institute of Technology, 2009-09) Lafleur, Jarret M. ; Saleh, Joseph H.In July 2007, DARPA issued a Broad Agency Announcement for the development of System F6, a flight demonstration of an architecture in which the functionality of a traditional monolithic satellite is fulfilled with a fractionated cluster of free-flying, wirelessly interconnected modules. Given the large number of possible architectural options, two challenges facing systems analysis of F6 are (1) the ability to enumerate the many potential candidate fractionated architectures and (2) the ability to analyze and quantify the cost and benefits of each architecture. One element necessary in enabling a probabilistic, valuecentric analysis of such fractionated architectures is a systematic method for sizing and costing the many candidate architectures that arise. The Georgia Tech F6 Architecture Synthesis Tool (GT-FAST) is a point design tool designed to fulfill this need by allowing rapid, automated sizing and synthesis of candidate F6 architectures. This paper presents the internal mechanics and some illustrative applications of GT-FAST. Discussed are the manner in which GT-FAST fractionated designs are specified, including discrete and continuous-variable inputs, as well as the methods, models, and assumptions used in estimating elements of mass, power, and cost. Finally, the paper concludes with sample outputs from GT-FAST for a notional fractionated architecture, an example of GT-FAST's trade study capability, and a partial validation of GT-FAST against the Jason-2 and TIMED satellites. The ease with which GT-FAST can be adapted to new fractionated spacecraft applications is highlighted, and avenues for potential future expansion of GT-FAST are discussed.
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ItemSurvey of Flexibility in Space Exploration Systems(Georgia Institute of Technology, 2008-09) Lafleur, Jarret M. ; Saleh, Joseph H.An increasingly common objective in the design of new space systems is the property of flexibility, or the capability to easily modify a system after it has been fielded in response to a changing environment or changing requirements. The body of research on this topic has been growing, but substantial work remains in developing metrics for characterizing system flexibility and trading it against other metrics of interest. This paper samples from the history of space exploration to glean heuristic insight into characteristics of flexibility in space exploration systems and their potential application to future systems. Divided into categories of intra- and inter-mission modification, examples include the Hubble Space Telescope, Mir space station, International Space Station, Apollo, Space Shuttle, and robotic Venera program. In several cases, metrics are identified which show clear performance gains due to changes after a system is fielded, and in all cases, environment or requirement changes that prompted system change are identified. Also discussed are examples where flexibility proved critical to mission success. Modular design and separation of functionality are recognized as likely flexibility-enabling characteristics. Also, briefly discussed are examples of non-configurational (e.g. software and trajectory) flexibility in space exploration applications.
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ItemA Series of Unforeseen Events: The Space Shuttle Mission Evolution Flexibility(Georgia Institute of Technology, 2008-09) Lafleur, Jarret M. ; Saleh, Joseph H.A common objective in the design of a new space system is that of flexibility, or the capability to easily modify that system in the future in response to a changing environment or changing requirements. The focus of this paper is a case study of the U.S. Space Shuttle to glean some insight into fundamental characteristics of flexibility in human space systems and how this may be applied to future systems. Data is presented on the evolution of mission requirements over time for 120 missions performed by the Space Shuttle over a period of approximately 27 years. Distinct trends in the time domain - as well causes of these trends - are identified, and early manifest plans from 1982 serve as a confirmation that these trends were not originally anticipated. Eight examples are then presented of engineering modifications that allowed the Shuttle to adapt and accommodate these requirement changes. Conclusions are drawn on the nature of flexibility as experienced by the Space Shuttle. Finally, remaining questions are posed regarding how flexibility is considered in the initial stages of design for space systems.
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ItemFeasibility Assessment of Microwave Power Beaming for Small Satellites(Georgia Institute of Technology, 2008-07) Lafleur, Jarret M. ; Saleh, Joseph H.While wireless power transmission to fulfill Earth's energy needs has been widely popularized as a potential application of microwave power beaming, one space application that has remained relatively untouched is power beaming between satellites. This paper provides a system-level analysis illustrating the feasibility and limitations of power beaming within a small-satellite cluster. To accomplish this analysis, the simple case of a two spacecraft system is examined. Parametric models of spacecraft power requirements as a function of eleven design variables allow for an extensive trade-space evaluation, and analysis is divided into four segments. First, the existence of feasible designs in the context of the small-satellite problem is verified with a Monte Carlo sweep of the design space. Next, a feasible baseline (reference) design is defined, and sensitivity of that baseline to individual variables is assessed. Finally, the design space is visualized with respect to distance between spacecraft, antenna diameter, and power independence factors. Despite optimistic assumptions in the setup of the problem, it is demonstrated that the small satellite power beaming design space is severely constrained. Only 6% of the design space falls under a suggested 250 W small satellite power constraint. Designs that are feasible involve very high transmission frequencies (>33 GHz), large antenna diameters for a small satellite (>0.93 m), and stringent proximity operations between satellites (within 740 m). Furthermore, full dependence of one spacecraft on power provided by another is shown to be effectively infeasible. These results do suggest, however, that inter-spacecraft microwave power beaming may deserve some consideration as a supplementary power mode for future small satellite clusters in short-term emergency or atypical situations.