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Space Systems Design Laboratory (SSDL)

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https://hdl.handle.net/1853/70747

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Daniel Guggenheim School of Aerospace Engineering

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https://finding-aids.library.gatech.edu/agents/corporate_entities/1138

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Publication Search Results

Now showing 1 - 10 of 218

Exploring the F6 Fractionated Spacecraft Trade Space with GT-FAST

(Georgia Institute of Technology, 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.
Comparative 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.
Statistical Reliability Analysis of Satellites by Mass Category: Does Spacecraft Size Matter?

(Georgia Institute of Technology, 2009-10) Dubos, Gregory F. ; Castet, Jean-Francois ; Saleh, Joseph H.

Reliability has long been recognized as a critical attribute for space systems, and potential causes of on-orbit failures are carefully sought for identification and elimination through various types of testing prior to launch. From a statistical or actuarial perspective, several parameters of the spacecraft, such as mission type, orbit, or spacecraft complexity, can potentially affect the probability of failure of satellites. In this paper, we explore the correlation between satellite mass, considered here as a proxy for size, and satellite reliability, and we investigate whether different classes of satellite, defined in terms of mass, exhibit different reliability profiles. To this end, we first conduct nonparametric analysis of satellite reliability based on a sample of 1,444 satellites. The satellites are organized in three main categories defined by satellite mass (Small – Medium – Large). Three nonparametric reliability curves are thus derived. We then provide parametric fits of the reliability curves to facilitate the identification of failure trends. We proceed to the comparative analysis of failure profiles over time and clearly identify different reliability behaviors for the various satellite mass categories. Finally, we discuss possible structural and causal reasons for these trends and failure differences, in particular with respect to design, testing and procurement.
Response 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.
High Mass Mars Entry, Descent, and Landing Architecture Assessment

(Georgia Institute of Technology, 2009-09) Steinfeldt, Bradley A. ; Theisinger, John E. ; Korzun, Ashley M. ; Clark, Ian G. ; Grant, Michael J. ; Braun, Robert D.

As the nation sets its sight on returning humans to the Moon and going onward to Mars, landing high mass payloads (>/= 2 t) on the Mars surface becomes a critical technological need. Viking heritage technologies (e.g., 70degrees sphere-cone aeroshell, SLA-561V thermal protection system, and supersonic disk-gap-band parachutes) that have been the mainstay of the United States' robotic Mars exploration program do not provide sufficient capability to land such large payload masses. In this investigation, a parametric study of the Mars entry, descent, and landing design space has been conducted. Entry velocity, entry vehicle configuration, entry vehicle mass, and the approach to supersonic deceleration were varied. Particular focus is given to the entry vehicle shape and the supersonic deceleration technology trades. Slender bodied vehicles with a lift-to-drag ratio (L=D) of 0.68 are examined alongside blunt bodies with L=D = 0.30. Results demonstrated that while the increased L=D of a slender entry configuration allows for more favorable terminal descent staging conditions, the greater structural efficiencies of blunt body systems along with the reduced acreage required for the thermal protection system affords an inherently lighter vehicle. The supersonic deceleration technology trade focuses on inflatable aerodynamic decelerators (IAD) and supersonic retropropulsion, as supersonic parachute systems are shown to be excessively large for further consideration. While entry masses (the total mass at the top of the Mars atmosphere) between 20 and 100 t are considered, a maximum payload capability of 37.3 t results. Of particular note, as entry mass increases, the gain in payload mass diminishes. It is shown that blunt body vehicles provide sufficient vertical L=D to decelerate all entry masses considered through the Mars atmosphere with adequate staging conditions for the propulsive terminal descent. A payload mass fraction penalty of approximately 0.3 exists for the use of slender bodied vehicles. Another observation of this investigation is that the increased aerothermal and aerodynamic loads induced from a direct entry trajectory (velocity ~6.75 km/s) reduce the payload mass fraction by approximately 15% compared to entry from orbital velocity (~4 km/s). It should be noted that while both IADs and supersonic retropropulsion were evaluated for each of the entry masses, configurations, and velocities, the IAD proved to be more mass-efficient in all instances. The sensitivity of these results to modeling assumptions was also examined. The payload mass of slender body vehicles was observed to be approximately four times more sensitive to modeling assumptions and uncertainty than blunt bodies.
GT-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.
Fast Sensitivity Computations for Trajectory Optimization

(Georgia Institute of Technology, 2009-08) Arora, Nitin ; Russell, Ryan P. ; Vuduc, Richard W.

Gradient based trajectory optimization relies on accurate sensitivity information to robustly move a solution towards an optimum. Computational complexity of sensitivity calculations increases exponentially for higher problem dimensions and orders. Hence, the computation of these sensitivities is traditionally a major speed bottleneck in trajectory optimization and targeting algorithms. We propose to use Nvidia's GPU (Graphics Processing Unit) to rapidly calculate the derivatives in a multilayer, parallel, and heterogeneous way while the CPU (Central Processing Unit) sequentially computes the less expensive state equations. The proposed tool computes both the first and second order analytic sensitivities on the GPU with double precision accuracy. For an example trajectory propagation, we demonstrate overlapped computations such that sensitivities are calculated almost for free compared to the conventional CPU implementation.
Trading Robustness Requirements in Mars Entry Trajectory Design

(Georgia Institute of Technology, 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.
Angle of Attack Modulation for Mars Entry Terminal State Optimization

(Georgia Institute of Technology, 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.
Performance Characterization of Supersonic Retropropulsion for Application to High-Mass Mars Entry, Descent, and Landing

(Georgia Institute of Technology, 2009-08) Korzun, Ashley M. ; Braun, Robert D.

Prior high-mass Mars EDL systems studies have neglected aerodynamic-propulsive interactions and performance impacts during the supersonic phase of descent. The goal of this investigation is to accurately evaluate the performance of supersonic retropropulsion with increasing vehicle ballistic coefficient across a range of initiation conditions relevant for future high-mass Mars landed systems. Past experimental work has established supersonic retropropulsion trends in static aerodynamics as a function of retropropulsion configuration, freestream conditions, and thrust. From this experimental database, an aerodynamic-propulsive interactions model is created. EDL system performance results are developed with the potential aerodynamic drag preservation included and excluded during this phase of flight for comparison against prior studies. The results of this investigation demonstrate the significance of aerodynamic drag preservation as a function of retropropulsion initiation conditions, characterize mass optimal trajectories utilizing supersonic retropropulsion, and compare propulsion system requirements with existing propulsion systems and systems under development for future exploration missions.