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Space Systems Design Laboratory (SSDL)

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Start date: 1980 × End date: 2009 × Author: Lafleur, Jarret M. ×

Publication Search Results

Now showing 1 - 10 of 19
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Exploring the F6 Fractionated Spacecraft Trade Space with GT-FAST

2009-11-12 , Lafleur, Jarret M.

Released in July 2007, the Broad Agency Announcement for DARPA’s System F6 outlined goals for 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. This paper applies the recently developed Georgia Tech F6 Architecture Synthesis Tool (GT-FAST) to the exploration of the System F6 trade space. GT-FAST is described in detail, after which a combinatorial analysis of the architectural trade space is presented to provide a theoretical contribution applicable to future analyses clearly showing the explosion of the trade space as the number of fractionatable components increases. Several output metrics of interest are defined, and Pareto fronts are used to visualize the trade space. The first set of these Pareto fronts allows direct visualization of one output against another, and the second set presents cost plotted against a Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) score aggregating performance objectives. These techniques allow for the identification of a handful of Pareto-optimal designs from an original pool of over 3,000 potential designs. Conclusions are drawn on salient features of the resulting Pareto fronts, important competing objectives which have been captured, and the potential suitability of a particularly interesting design designated PF0248. A variety of potential avenues for future work are also identified.

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GT-FAST: A Point Design Tool for Rapid Fractionated Spacecraft Sizing and Synthesis

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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Survey of Flexibility in Space Exploration Systems

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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Feasibility Assessment of Microwave Power Beaming for Small Satellites

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.

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Comparative Reliability of GEO, LEO, and MEO Satellites

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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Angle of Attack Modulation for Mars Entry Terminal State Optimization

2009-08 , Lafleur, Jarret M. , Cerimele, Chris J.

From the perspective of atmospheric entry, descent, and landing (EDL), one of the most foreboding destinations in the solar system is Mars due in part to its exceedingly thin atmosphere. To benchmark best possible scenarios for evaluation of potential Mars EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities and vehicle masses, focusing on the identification of potential benefits of enabling angle of attack modulation. The terminal state is envisioned as one appropriate for the initiation of terminal descent via parachute or other means. A particle swarm optimizer varies entry flight path angle, ten bank profile points, and ten angle of attack profile points to find maximum-final-altitude trajectories for a 10 x 30 m ellipsled at 180 different combinations of values for entry mass, entry velocity, terminal Mach number, and minimum allowable altitude. Parametric plots of maximum achievable altitude are shown, as are examples of optimized trajectories. It is shown that appreciable terminal state altitude gains (2.5-4.0 km) over pure bank angle control may be possible if angle of attack modulation is enabled for Mars entry vehicles. Gains of this magnitude could prove to be enabling for missions requiring high-altitude landing sites. Conclusions are also drawn regarding trends in the bank and angle of attack profiles that produce the optimal trajectories in this study, and directions for future work are identified.

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A Series of Unforeseen Events: The Space Shuttle Mission Evolution Flexibility

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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Response Surface Equations for Expendable Launch Vehicle Payload Capability

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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Trading Robustness Requirements in Mars Entry Trajectory Design

2009-08 , Lafleur, Jarret M.

One of the most important metrics characterizing an atmospheric entry trajectory in preliminary design is the size of its predicted landing ellipse. Often, requirements for this ellipse are set early in design and significantly influence both the expected scientific return from a particular mission and the cost of development. Requirements typically specify a certain probability level (sigma-level) for the prescribed ellipse, and frequently this latter requirement is taken at 3sigma. However, searches for the justification of 3sigma as a robustness requirement suggest it is an empirical rule of thumb borrowed from non-aerospace fields. This paper presents an investigation into the sensitivity of trajectory performance to varying robustness (sigma-level) requirements. The treatment of robustness as a distinct objective is discussed, and an analysis framework is presented involving the manipulation of design variables to effect trades between performance and robustness objectives. The scenario for which this method is illustrated is the ballistic entry of an MSL-class Mars entry vehicle. Here, the design variable is entry flight path angle, and objectives are parachute deploy altitude performance and error ellipse robustness. Resulting plots show the sensitivities between these objectives and trends in the entry flight path angles required to design to these objectives. Relevance to the trajectory designer is discussed, as are potential steps for further development and use of this type of analysis.

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Mars Entry Bank Profile Design for Terminal State Optimization

2008-08 , Lafleur, Jarret M. , Cerimele, Chris J.

One challenge examined in NASA's DRM 5.0 study is that of entry, descent, and landing (EDL) on Mars for high-ballistic-coefficient, human-class payloads. To define best-case entry scenarios for the evaluation of potential EDL system designs, a study is conducted to optimize the entry-to-terminal-state portion of EDL for a variety of entry velocities, vehicle ballistic coefficients (), and lift-to-drag ratios (L/D). The terminal state is envisioned as one appropriate for the initiation of terminal descent via parachute or other means. A particle swarm optimizer varies entry flight path angle and ten bank profile points to find maximum-final-altitude trajectories. A baseline set of optimizations is performed, as are full lift- up and relaxed-deceleration-constraint sets for comparison. In total, an estimated 9 million trajectories are analyzed to yield 1800 optimal trajectories. Parametric plots of maximum achievable altitude are shown, as are examples of optimized trajectories. Characteristic vehicle contours are overlaid on the parametric plots, and conclusions are drawn on the feasibility of vehicles in the L/D vs. design space. It is shown that entry bank angle control is highly deserving of consideration early in design, particularly for vehicles with mid- or high-L/D values, high entry velocities, and deceleration-limited trajectories. Key conclusions are also drawn regarding trends in optimal bank profiles and in the constraints which impose particularly severe limits on the design of these trajectories.

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