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Daniel Guggenheim School of Aerospace Engineering

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End date: 2014 × search.filters.applied.f.isOrgUnitOfPublicationTitle: Daniel Guggenheim School of Aerospace Engineering × search.filters.applied.f.isSeriesOfPublicationTitle: Doctor of Philosophy with a Major in Aerospace Engineering × Type: Text ×

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Design strategies for rotorcraft blades and HALE aircraft wings applied to damage tolerant wind turbine blade design

2014-12-15 , Richards, Phillip W.

Offshore wind power production is an attractive clean energy option, but the difficulty of access can lead to expensive and rare opportunities for maintenance. Smart loads management (controls) are investigated for their potential to increase the fatigue life of damaged offshore wind turbine rotor blades. This study will consider two commonly encountered damage types for wind turbine blades, the trailing edge disbond (bond line failure) and shear web disbond, and show how 3D finite element modeling can be used to quantify the effect of operations and control strategies designed to extend the fatigue life of damaged blades. Modern wind turbine blades are advanced composite structures, and blade optimization problems can be complex with many structural design variables and a wide variety of aeroelastic design requirements. The multi-level design method is an aeroelastic structural design technique for beam-like structures in which the general design problem is divided into a 1D beam optimization and a 2D section optimization. As a demonstration of aeroelastic design, the multi-level design method is demonstrated for the internal structural design of a modern composite rotor blade. Aeroelastic design involves optimization of system geometry features as well as internal features, and this is demonstrated in the design of a flying wing aircraft. Control methods such as feedback control also have the capability alleviate aeroelastic design requirements and this is also demonstrated in the flying wing aircraft example. In the case of damaged wind turbine blades, load mitigation control strategies have the potential to mitigate the effects of damage, and allow partial operation to avoid shutdown. The load mitigation strategies will be demonstrated for a representative state-of-the-art wind turbine (126m rotor diameter). An economic incentive will be provided for the proposed operations strategies, in terms of weighing the cost and risk of implementation against the benefits of increased revenue due to operation of damaged turbines. The industry trend in wind turbine design is moving towards very large blades, causing the basic design criterion to change as aeroelastic effects become more important. An ongoing 100 m blade (205 m rotor diameter) design effort intends to investigate these design challenges. As a part of that effort, this thesis will investigate damage tolerant design strategies to ensure next-generation blades are more reliable.

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CONTRAST: A conceptual reliability growth approach for comparison of launch vehicle architectures

2014-11-17 , Zwack, Mathew R.

In 2004, the NASA Astronaut Office produced a memo regarding the safety of next generation launch vehicles. The memo requested that these vehicles have a probability of loss of crew of at most 1 in 1000 flights, which represents nearly an order of magnitude decrease from current vehicles. The goal of LOC of 1 in 1000 flights has since been adopted by the launch vehicle design community as a requirement for the safety of future vehicles. This research addresses the gap between current vehicles and future goals by improving the capture of vehicle architecture effects on reliability and safety. Vehicle architecture pertains to the physical description of the vehicle itself, which includes manned or unmanned, number of stages, number of engines per stage, engine cycle types, redundancy, etc. During the operations phase of the vehicle life-cycle it is clear that each of these parameters will have an inherent effect on the reliability and safety of the vehicle. However, the vehicle architecture is typically determined during the early conceptual design phase when a baseline vehicle is selected. Unless a great amount of money and effort is spent, the architecture will remain relatively constant from conceptual design through operations. Due to the fact that the vehicle architecture is essentially “locked-in” during early design, it is expected that much of the vehicle's reliability potential will also be locked-in. This observation leads to the conclusion that improvement of vehicle reliability and safety in the area of vehicle architecture must be completed during early design. Evaluation of the effects of different architecture decisions must be performed prior to baseline selection, which helps to identify a vehicle that is most likely to meet the reliability and safety requirements when it reaches operations. Although methods exist for evaluating reliability and safety during early design, weaknesses exist when trying to evaluate all architecture effects simultaneously. The goal of this research was therefore to formulate and implement a method that is capable of quantitatively evaluating vehicle architecture effects on reliability and safety during early conceptual design. The ConcepTual Reliability Growth Approach for CompariSon of Launch Vehicle ArchiTectures (CONTRAST) was developed to meet this goal. Using the strengths of existing techniques a hybrid approach was developed, which utilizes a reliability growth projection to evaluate the vehicles. The growth models are first applied at the subsystem level and then a vehicle level projection is generated using a simple system level fault tree. This approach allows for the capture of all trades of interest at the subsystem level as well as many possible trades at the assembly level. The CONTRAST method is first tested on an example problem, which compares the method output to actual data from the Space Transportation System (STS). This example problem illustrates the ability of the CONTRAST method to capture reliability growth trends seen during vehicle operations. It also serves as a validation for the development of the reliability growth model assumptions for future applications of the method. The final chapter of the thesis applies the CONTRAST method to a relevant launch vehicle, the Space Launch System (SLS), which is currently under development. Within the application problem, the output of the method is first used to check that the primary research objective has been met. Next, the output is compared to a state-of-the-art tool in order to demonstrate the ability of the CONTRAST method to alleviate one of the primary consequences of using existing techniques. The final section within this chapter presents an analysis of the booster and upper stage block upgrade options for the SLS vehicle. A study of the upgrade options was carried out because the CONTRAST method is uniquely suited to look at the effects of such strategies. The results from the study of SLS block upgrades give interesting observations regarding the desired development order and upgrade strategy. Ultimately this application problem demonstrates the merits of applying the CONTRAST method during early design. This approach provides the designer with more information in regard to the expected reliability of the vehicle, which will ultimately enable the selection of a vehicle baseline that is most likely to meet the future requirements.

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An adaptive atmospheric prediction algorithm to improve density forecasting for aerocapture guidance processes

2014-11-12 , Wagner, John Joseph

Many modern entry guidance systems depend on predictions of atmospheric parameters, notably atmospheric density, in order to guide the entry vehicle to some desired final state. However, in highly dynamic atmospheric environments such as the Martian atmosphere, the density may vary by as much as 200% from predicted pre-entry trends. This high level of atmospheric density uncertainty can cause significant complications for entry guidance processes and may in extreme scenarios cause complete failure of the entry. In the face of this uncertainty, mission designers are compelled to apply large trajectory and design safety margins which typically drive the system design towards less efficient solutions with smaller delivered payloads. The margins necessary to combat the high levels of atmospheric uncertainty may even preclude scientifically interesting destinations or architecturally useful mission modes such as aerocapture. Aerocapture is a method for inserting a spacecraft into an orbit about a planetary body with an atmosphere without the need for significant propulsive maneuvers. This can reduce the required propellant and propulsion hardware for a given mission which lowers mission costs and increases the available payload fraction. However, large density dispersions have a particularly acute effect on aerocapture trajectories due to the interaction of the high required speeds and relatively low densities encountered at aerocapture altitudes. Therefore, while the potential system level benefits of aerocapture are great, so too are the risks associated with this mission mode in highly uncertain atmospheric environments such as Mars. Contemporary entry guidance systems utilize static atmospheric density models for trajectory prediction and control. These static models are unable to alter the fundamental nature of the underlying state equations which are used to predict atmospheric density. This limits both the fidelity and adaptive freedom of these models and forces the guidance system to retroactively correct for the density prediction errors after those errors have already impacted the trajectory. A new class of dynamic density estimator called a Plastic Ensemble Neural System (PENS) is introduced which is able to generate high fidelity, adaptable density forecast models by altering the underlying atmospheric state equations to better agree with observed atmospheric trends. A new construct called an ensemble echo is also introduced which creates an associative learning architecture, permitting PENS to evolve with increasing atmospheric exposure. The PENS estimator is applied to a numerical guidance system and the performance of the composite system is investigated with over 144,000 guided trajectory simulations. The results demonstrate that the PENS algorithm achieves significant reductions in both the required post-aerocapture performance, and the aerocapture failure rates relative to historical density estimators.

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Data transfer strategies for overset and hybrid computational methods

2014-08-25 , Quon, Eliot

Modern computational science permits the accurate solution of nonlinear partial differential equations (PDEs) on overlapping computational domains, known as an overset approach. The complex grid interconnectivity inherent in the overset method can introduce errors in the solution through “orphan” points, i.e., grid points for which reliable solution donor points cannot be located. For this reason, a variety of data transfer strategies based on scattered data interpolation techniques have been assessed with application to both overset and hybrid methodologies. Scattered data approaches are attractive because they are decoupled from solver type and topology, and may be readily applied within existing methodologies. In addition to standard radial basis function (RBF) interpolation, a novel steered radial basis function (SRBF) interpolation technique has been developed to introduce data adaptivity into the data transfer algorithm. All techniques were assessed by interpolating both continuous and discontinuous analytical test functions. For discontinuous functions, SRBF interpolation was able to maintain solution gradients with the steering technique being the scattered-data analog of a slope limiter. In comparison with linear mappings, the higher-order approaches were able to more accurately preserve flow physics for arbitrary grid configurations. Overset validation test cases included an inviscid convecting vortex, a shock tube, and a turbulent ship airwake. These were studied within unsteady Reynolds-Averaged Navier-Stokes (URANS) simulations to determine quantitative and qualitative improvements when applying RBF interpolation over current methods. The convecting vortex was also analyzed on a grid configuration which contained orphan points under the state-of-the-art overset paradigm. This was successfully solved by the RBF-based algorithm, which effectively eliminated orphans by enabling high-order extrapolation. Order-of-magnitude reductions in error compared to the exact vortex solution were observed. In addition, transient conservation errors that persisted in the original overset methodology were eliminated by the RBF approach. To assess the effect of advanced mapping techniques on the fidelity of a moving grid simulation, RBF interpolation was applied to a hybrid simulation of an isolated wind turbine rotor. The resulting blade pressure distributions were comparable to a rotor simulation with refined near-body grids.

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Architecting aircraft power distribution systems via redundancy allocation

2014-11-17 , Campbell, Angela Mari

Recently, the environmental impact of aircraft and rising fuel prices have become an increasing concern in the aviation industry. To address these problems, organizations such as NASA have set demanding goals for reducing aircraft emissions, fuel burn, and noise. In an effort to reach the goals, a movement toward more-electric aircraft and electric propulsion has emerged. With this movement, the number of critical electrical loads on an aircraft is increasing causing power system reliability to be a point of concern. Currently, power system reliability is maintained through the use of back-up power supplies such as batteries and ram-air-turbines (RATs). However, the increasing power requirements for critical loads will quickly outgrow the capacity of the emergency devices. Therefore, reliability needs to be addressed when designing the primary power distribution system. Power system reliability is a function of component reliability and redundancy. Component reliability is often not determined until detailed component design has occurred; however, the amount of redundancy in the system is often set during the system architecting phase. In order to meet the capacity and reliability requirements of future power distribution systems, a method for redundancy allocation during the system architecting phase is needed. This thesis presents an aircraft power system design methodology that is based upon the engineering decision process. The methodology provides a redundancy allocation strategy and quantitative trade-off environment to compare architecture and technology combinations based upon system capacity, weight, and reliability criteria. The methodology is demonstrated by architecting the power distribution system of an aircraft using turboelectric propulsion. The first step in the process is determining the design criteria which includes a 40 MW capacity requirement, a 20 MW capacity requirement for the an engine-out scenario, and a maximum catastrophic failure rate of one failure per billion flight hours. The next step is determining gaps between the performance of current power distribution systems and the requirements of the turboelectric system. A baseline architecture is analyzed by sizing the system using the turboelectric system power requirements and by calculating reliability using a stochastic flow network. To overcome the deficiencies discovered, new technologies and architectures are considered. Global optimization methods are used to find technology and architecture combinations that meet the system objectives and requirements. Lastly, a dynamic modeling environment is constructed to study the performance and stability of the candidate architectures. The combination of the optimization process and dynamic modeling facilitates the selection of a power system architecture that meets the system requirements and objectives.

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A methodology for evaluating fleet implications of mission specification changes

2014-11-17 , Brett, Paul S.

Civil aviation has matured to become a vital piece of the global economy, providing the rapid movement of goods and people to all regions. This has already led to significant growth and expectations of further growth are on the rate of 5% per year. Given the high projected rate of growth, environmental consequences of commercial aviation are expected to rise. To mitigate the increase of noise and emissions, governing bodies such as ICAO and the FAA have established and are considering additional regulation of noise, NOₓ, and CO₂ while the European Union has integrated aviation into their Environmental Trading Scheme. The traditional response to new regulation is to integrate technologies into the aircraft to reduce environmental footprint. While these benefits are positive on the aircraft level, fleet growth is projected to outpace benefits provided by technology alone. To further reduce environmental footprint, a number of mitigation strategies are being explored to determine the impact. One of those strategies involves changing the mission specifications of today's aircraft by reducing range, speed, or payload in an effort to reduce fuel consumption and has been predominantly focused at the vehicle level. This research proposes an approach that evaluates mission specification changes from the aircraft design level up to the fleet level, forecasted into the future, to assess the impact over a number of metrics to fully understand the implications of mission specification changes. The methodology Mission Specifications and Fleet Implications Technique (MS-FIT) identifies stakeholder requirements that will be tracked at either the vehicle or fleet level and leverages them to build an environment that will allow joint evaluation to facilitate increased knowledge about the full implications of mission specification adoption. Additionally laid out is an approach on how to select prospective routes for intermediate stops based on fuel burn and operating cost considerations. Guidance is provided on how to filter down a list of candidate airports to those most viable as well as regions of the world most likely to benefit from intermediate stops. Three sample problems were used to demonstrate the viability of MS-FIT: cruise speed reduction, design mission range reduction, and the combination of speed and range reduction. Each problem was able to demonstrate different implications from the implementation of the different specification changes. Speed reduction can negatively impacts cost while range reduction has consequences to noise at the intermediate airports. The combination of the two draws in negative implications from both even though the environmental benefits are better. Finally, an analysis of some of the assumptions was conducted to examine the sensitivity to the results of speed and range reduction. These include variation in costs, reductions in annual utilization of aircraft, and variation in intermediate stop adoption. Speed reduction is strongly sensitive to increases in crew and maintenance rates while landing fees significantly eat into the benefits of range reduction and intermediate stops. Minor utilization reductions can significantly reduce the viability of speed reduction as the increase in capital costs offset all the savings from fuel reduction while range reduction is a little less sensitive. Intermediate stop variation does not eliminate the benefits of range reduction and even can provide cost savings depending on the design range of the reduced variant but it can have consequences to airport noise to higher traffic airports. With the proposed framework, additional information is available to fully understand the implications with respect to fuel burn, NOₓ emissions, operating cost, capital cost, noise, and safety. This can then inform decision makers on whether pursuing a particular mission specification strategy is advantageous or not.

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Nonlinear pose control and estimation for space proximity operations: an approach based on dual quaternions

2014-11-11 , Salgueiro Filipe, Nuno Ricardo

The term proximity operations has been widely used in recent years to describe a wide range of space missions that require a spacecraft to remain close to another space object. Such missions include, for example, the inspection, health monitoring, surveillance, servicing, and refueling of a space asset by another spacecraft. One of the biggest challenges in autonomous space proximity operations, either cooperative or uncooperative, is the need to autonomously and accurately track time-varying relative position and attitude references, i.e., pose references, with respect to a moving target, in order to avoid on-orbit collisions and achieve the overall mission goals. In addition, if the target spacecraft is uncooperative, the Guidance, Navigation, and Control (GNC) system of the chaser spacecraft must not rely on any help from the target spacecraft. In this case, vision-based sensors, such as cameras, are typically used to measure the relative pose between the spacecraft. Although vision-based sensors have several attractive properties, they introduce new challenges, such as no direct linear and angular velocity measurements, slow update rates, and high measurement noise. This dissertation investigates the problem of autonomously controlling and estimating the pose of a chaser spacecraft with respect to a moving target spacecraft, possibly uncooperative. Since this problem is inherently hard, the standard approach in the literature is to split the attitude-tracking problem from the position-tracking problem. Whereas the attitude-tracking problem is relatively simple, since the rotational motion is independent from the translational motion, the position-tracking problem is more complicated, as the translational motion depends on the rotational motion. Hence, whereas strong theoretical results exist for the attitude problem, the position problem typically requires additional assumptions. An alternative, more general approach to the pose control and estimation problems is to consider the fully coupled 6-DOF motion. However, fewer results exist that directly address this higher dimensional problem. The main contribution of this dissertation is to show that dual quaternions can be used to extend the theoretical results that exist for the attitude motion into analogous results for the combined position and attitude motion. Moreover, this dissertation shows that this can be accomplished by (almost) just replacing quaternions by dual quaternions in the original derivations. This is because dual quaternions are built on and are an extension of classical quaternions. Dual quaternions provide a compact representation of the pose of a frame with respect to another frame. Using this approach, three new results are presented in this dissertation. First, a pose-tracking controller that does not require relative linear and angular velocity measurements is derived with vision-based sensors in mind. Compared to existing literature, the proposed velocity-free pose-tracking controller guarantees that the pose of the chaser spacecraft will converge to the desired pose independently of the initial state, even if the reference motion is not sufficiently exciting. In addition, the convergence region does not depend on the gains of the controller. Second, a Dual Quaternion Multiplicative Extended Kalman Filter (DQ-MEKF) is developed from the highly successful Quaternion MEKF (Q-MEKF) as an alternative way to achieve pose-tracking without velocity measurements. Existing dual quaternion EKFs are additive, not multiplicative, and have two additional states. The DQ-MEKF is experimentally validated and compared with two conventional EKFs on the 5-DOF platform of the Autonomous Spacecraft Testing of Robotic Operations in Space (ASTROS) facility at the School of Aerospace Engineering at Georgia Tech. Finally, the velocity-free pose-tracking controller is compared qualitatively and quantitatively to a pose-tracking controller that uses the velocity estimates produced by the DQ-MEKF through a realistic proximity operations simulation. Third, a pose-tracking controller that does not require the mass and inertia matrix of the chaser satellite is suggested. This inertia-free controller takes into account the gravitational acceleration, the gravity-gradient torque, the perturbing acceleration due to Earth's oblateness, and constant -- but otherwise unknown -- disturbance forces and torques. Sufficient conditions on the reference pose are also given that guarantee the identification of the mass and inertia matrix of the satellite. Compared to the existing literature, this controller has only as many states as unknown elements and it does not require a priori known upper bounds on any states or parameters. Finally, the inertia-free pose-tracking controller and the DQ-MEKF are tested on a high-fidelity simulation of the 5-DOF platform of the ASTROS facility and also experimentally validated on the actual platform. The equations of motion of the 5-DOF platform, on which the high-fidelity simulation is based, are derived for three distinct cases: a 3-DOF case, a 5-DOF case, and a (2+1)-DOF case. Four real-time experiments were run on the platform. In the first, a sinusoidal reference attitude with respect to the inertial frame is tracked using VSCMGs. In the second, a constant reference attitude is maintained with respect to a target object using VSCMGs and measurements from a camera. In the third, the same sinusoidal reference attitude with respect to the inertial frame tracked in the first experiment is now tracked using cold-gas thrusters. Finally, in the fourth and last experiment, a time-varying 5-DOF reference pose with respect to the inertial frame is tracked using cold-gas thrusters.

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Sustainability of multimodal intercity transportation using a hybrid system dynamics and agent-based modeling approach

2014-11-17 , Hivin, Ludovic F.

Demand for intercity transportation has increased significantly in the past decades and is expected to continue to follow this trend in the future. In the meantime, concern about the environmental impact and potential climate change associated with this demand has grown, resulting in an increasing importance of climate impact considerations in the overarching issue of sustainability. This results in discussions on new regulations, policies and technologies to reduce transportation's climate impact. Policies may affect the demand for the different transportation modes through increased travel costs, increased market share of more fuel efficient vehicles, or even the introduction of new modes of transportation. However, the effect of policies and technologies on mobility, demand, fleet composition and the resulting climate impact remains highly uncertain due to the many interdependencies. This motivates the creation of a parametric modeling and simulation environment to explore a wide variety of policy and technology scenarios and assess the sustainability of transportation. In order to capture total transportation demand and the potential mode shifts, a multimodal approach is necessary. The complexity of the intercity transportation System-of-Systems calls for a hybrid Agent-Based Modeling and System Dynamics paradigm to better represent both micro-level and macro-level behaviors. Various techniques for combining these paradigms are explored and classified to serve as a hybrid modeling guide. A System Dynamics approach is developed, that integrates socio-economic factors, mode performance, aggregated demand and climate impact. It is used to explore different policy and technology scenarios, and better understand the dynamic behavior of the intercity transportation System-of-Systems. In order to generate the necessary data to create and validate the System Dynamics model, an Agent-Based model is used due to its capability to better capture the behavior of a collection of sentient entities. Equivalency of both models is ensured through a rigorous cross-calibration process. Through the use of fleet models, the fuel burn and life cycle emissions from different modes of transportation are quantified. The radiative forcing from the main gaseous and aerosol species is then obtained through radiative transfer calculations and regional variations are discussed. This new simulation environment called the environmental Ground and Air Mode Explorer (eGAME) is then used to explore different policy and technology scenarios and assess their effect on transportation demand, fleet efficiencies and the resulting climate impact. The results obtained with this integrated assessment tool aim to support a scenario-based decision making approach and provide insight into the future of the U.S. transportation system in a climate constrained environment.

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Stochastic feasibility assessments of orbital propellant depot and commercial launch enabled space exploration architectures

2014-11-13 , Chai, Patrick R.

The 2010 National Space Policy of the United State of America introduced by President Obama directed NASA to set far reaching exploration milestones that included a crewed mission to a Near Earth Asteroid by 2025 and a crewed mission to Martian orbit by the mid-2030s. The policy was directly influenced by the recommendations of the 2009 Review of United States Human Space Flight Plans Committee, which called for an evolutionary approach to human space exploration and emphasized the criticality of budgetary, programmatic, and program sustainability. One potential method of improving the sustainability of exploration architectures is the utilization of orbital propellant depots with commercial launch services. In any exploration architecture, upwards of seventy percent of the mass required in orbit is propellant. A propellant depot based architecture allows propellant to be delivered in small increments using existing commercial launch vehicles, but will require three to five times the number of launches as compared to the using the NASA planned 70 to 130 metric ton heavy lift launch system. Past studies have shown that the utilization of propellant depots in exploration architectures have the potential of providing the sustainability that the Review of United States Human Space Flight Plans Committee emphasized. However, there is a lack of comprehensive analysis to determine the feasibility of propellant depots within the framework of human space exploration. The objective of this research is to measure the feasibility of a propellant depot and commercial launch based exploration architecture by stochastic assessment of technical, reliability, and economic risks. A propellant depot thermal model was developed to analyze the effectiveness of various thermal management systems, determine their optimal configuration, quantify the uncertainties in the system models, and stochastically compute the performance feasibility of the propellant depot system. Probabilistic cost analysis captured the uncertainty in the development cost of propellant depots and the fluctuation of commercial launch prices, and, along with the cost of launch failures, provided a metric for determining economic feasibility. Probabilistic reliability assessments using the launch schedule, launch reliability, and architecture requirements of each phase of the mission established launch success feasibility. Finally, an integrated stochastic optimization was performed to determine the feasibility of the exploration architecture. The final product of this research is an evaluation of propellant depots and commercial launch services as a practical method to achieving economic sustainability for human space exploration. A method for architecture feasibility assessment is demonstrated using stochastic system metrics and applied in the evaluation of technical, economic, and reliability feasibility of orbital propellant depots and commercial launch based exploration architectures. The results of the analysis showed the propellant depots based architectures to be technically feasible using current commercial launch vehicles, economically feasible for having a program budget less than $4 billion per year, and have launch reliability approaching the best single launch vehicle, Delta IV, with the use of redundant vehicles. These results serve to provide recommendations on the use of propellant depots in exploration architectures to the Moon, Near Earth Objects, Mars, and beyond.The 2010 National Space Policy of the United State of America introduced by President Obama directed NASA to set far reaching exploration milestones that included a crewed mission to a Near Earth Asteroid by 2025 and a crewed mission to Martian orbit by the mid-2030s. The policy was directly influenced by the recommendations of the 2009 Review of United States Human Space Flight Plans Committee, which called for an evolutionary approach to human space exploration and emphasized the criticality of budgetary, programmatic, and program sustainability. One potential method of improving the sustainability of exploration architectures is the utilization of orbital propellant depots with commercial launch services. In any exploration architecture, upwards of seventy percent of the mass required in orbit is propellant. A propellant depot based architecture allows propellant to be delivered in small increments using existing commercial launch vehicles, but will require three to five times the number of launches as compared to the using the NASA planned 70 to 130 metric ton heavy lift launch system. Past studies have shown that the utilization of propellant depots in exploration architectures have the potential of providing the sustainability that the Review of United States Human Space Flight Plans Committee emphasized. However, there is a lack of comprehensive analysis to determine the feasibility of propellant depots within the framework of human space exploration. The objective of this research is to measure the feasibility of a propellant depot and commercial launch based exploration architecture by stochastic assessment of technical, reliability, and economic risks. A propellant depot thermal model was developed to analyze the effectiveness of various thermal management systems, determine their optimal configuration, quantify the uncertainties in the system models, and stochastically compute the performance feasibility of the propellant depot system. Probabilistic cost analysis captured the uncertainty in the development cost of propellant depots and the fluctuation of commercial launch prices, and, along with the cost of launch failures, provided a metric for determining economic feasibility. Probabilistic reliability assessments using the launch schedule, launch reliability, and architecture requirements of each phase of the mission established launch success feasibility. Finally, an integrated stochastic optimization was performed to determine the feasibility of the exploration architecture. The final product of this research is an evaluation of propellant depots and commercial launch services as a practical method to achieving economic sustainability for human space exploration. A method for architecture feasibility assessment is demonstrated using stochastic system metrics and applied in the evaluation of technical, economic, and reliability feasibility of orbital propellant depots and commercial launch based exploration architectures. The results of the analysis showed the propellant depots based architectures to be technically feasible using current commercial launch vehicles, economically feasible for having a program budget less than $4 billion per year, and have launch reliability approaching the best single launch vehicle, Delta IV, with the use of redundant vehicles. These results serve to provide recommendations on the use of propellant depots in exploration architectures to the Moon, Near Earth Objects, Mars, and beyond.

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Very low earth orbit propellant collection feasibility assessment

2014-10-01 , Singh, Lake Austin

This work focuses on the concept of sustainable propellant collection. The concept consists of gathering ambient gas while on-orbit and using it as propellant. Propellant collection could potentially enable operation in very-low Earth orbits without compromising spacecraft lifetime. This work conducts a detailed analysis of propellant collection from a physics perspective in order to test the assertions of previous researchers that propellant collection can dramatically reduce the cost of propellant on-orbit. Major design factors for propellant collection are identified from the fundamental propellant collection equations, which are derived in this work from first principles. A sensitivity analysis on the parameters in these equations determines the relative importance of each parameter to the overall performance of a propellant-collecting vehicle. The propellant collection equations enable the study of where propellant collection is technically feasible as a function of orbit and vehicle performance parameters. Two case studies conducted for a very-low Earth orbit science mission and a propellant depot-type mission serve to demonstrate the application of the propellant collection equations derived in this work. The results of this work show where propellant collection is technically feasible for a wide range of orbit and vehicle performance parameters. Propellant collection can support very-low Earth operation with presently available technology, and a number of research developments can further extend propellant-collecting concepts' ability to operate at low altitudes. However, propellant collection is not presently suitable for propellant depot applications due to limitations in power.

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