513 research outputs found

    Ascent trajectory optimisation for a single-stage-to-orbit vehicle with hybrid propulsion

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    This paper addresses the design of ascent trajectories for a hybrid-engine, high performance, unmanned, single-stage-to-orbit vehicle for payload deployment into low Earth orbit. A hybrid optimisation technique that couples a population-based, stochastic algorithm with a deterministic, gradient-based technique is used to maximize the nal vehicle mass in low Earth orbit after accounting for operational constraints on the dynamic pressure, Mach number and maximum axial and normal accelerations. The control search space is first explored by the population-based algorithm, which uses a single shooting method to evaluate the performance of candidate solutions. The resultant optimal control law and corresponding trajectory are then further refined by a direct collocation method based on finite elements in time. Two distinct operational phases, one using an air-breathing propulsion mode and the second using rocket propulsion, are considered. The presence of uncertainties in the atmospheric and vehicle aerodynamic models are considered in order to quantify their effect on the performance of the vehicle. Firstly, the deterministic optimal control law is re-integrated after introducing uncertainties into the models. The proximity of the final solutions to the target states are analysed statistically. A second analysis is then performed, aimed at determining the best performance of the vehicle when these uncertainties are included directly in the optimisation. The statistical analysis of the results obtained are summarized by an expectancy curve which represents the probable vehicle performance as a function of the uncertain system parameters. This analysis can be used during the preliminary phase of design to yield valuable insights into the robustness of the performance of the vehicle to uncertainties in the specification of its parameters

    Trajectory optimization and guidance law development for national aerospace plane applications

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    The work completed to date is comprised of the following: a simple vehicle model representative of the aerospace plane concept in the hypersonic flight regime, fuel-optimal climb profiles for the unconstrained and dynamic pressure constrained cases generated using a reduced order dynamic model, an analytic switching condition for transition to rocket powered flight as orbital velocity is approached, simple feedback guidance laws for both the unconstrained and dynamic pressure constrained cases derived via singular perturbation theory and a nonlinear transformation technique, and numerical simulation results for ascent to orbit in the dynamic pressure constrained case

    Multi-Objective Trajectory Optimization of a Hypersonic Reconnaissance Vehicle with Temperature Constraints

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    Temperature-constrained optimal trajectories for a scramjet-based hypersonic reconnaissance vehicle were generated by developing an optimal control formulation and solving it using a variable order Gauss-Radau quadrature collocation method. The vehicle was assumed to be an air-breathing reconnaissance aircraft that has specified takeoff/landing locations, airborne refueling constraints, specified no-fly zones, and specified targets for sensor data collections. The aircraft model included fight dynamics, aerodynamics, and thermal constraints. This model was incorporated into an optimal control formulation that includes constraints on both the vehicle as well as mission parameters, such as avoidance of no-fly zones and coverage of high-value targets. Optimal trajectories were be developed using several different performance costs in the optimal control formulation--minimum time, minimum time with control penalties, and maximum range. The resulting analysis demonstrated that optimal trajectories that meet specified mission parameters and constraints can be determined and used for larger-scale operational and campaign planning

    Algorithmic Advances to Increase the Fidelity Of Conceptual Hypersonic Mission Design

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    The contributions of this dissertation increase the fidelity of conceptual hypersonic mission design through the following innovations: 1) the introduction of coupling between the effects of ablation of the thermal protection system (TPS) and flight dynamics, 2) the introduction of rigid body dynamics into trajectory design, and 3) simplifying the design of hypersonic missions that involve multiple phases of flight. These contributions are combined into a unified conceptual mission design framework, which is in turn applicable to slender hypersonic vehicles with ablative TPS. Such vehicles are employed in military applications, wherein speed and terminal energy are of critical importance. The fundamental observation that results from these contributions is the substantial reduction in the maximum terminal energy that is achievable when compared to the state-of-the art conceptual design process. Additionally, the control history that is required to follow the maximum terminal energy trajectory is also significantly altered, which will in turn bear consequence on the design of the control actuators. The other important accomplishment of this dissertation is the demonstration of the ability to solve these class of problems using indirect methods. Despite being built on a strong foundation of the calculus of variations, the state-of-the-art entirely neglects indirect methods because of the challenge associated with solving the resulting boundary value problem (BVP) in a system of differential-algebraic equations (DAEs). Instead, it employs direct methods, wherein the optimality of the calculated trajectory is not guaranteed. The ability to employ indirect methods to solve for optimal trajectories that are comprised of multiple phases of flight while also accounting for the effects of ablation of the TPS and rigid body dynamics is a substantial advancement in the state-of-the-art

    A prototype computerized synthesis methodology for generic space access vehicle (SAV) conceptual design.

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    This dissertation presents the development steps required towards a generic (configuration independent) hands-on flight vehicle conceptual design synthesis methodology. This process is developed such that it can be applied to any flight vehicle class if desired. In the present context, the methodology has been put into operation for the conceptual design of a tourist Space Access Vehicle. The case study illustrates elements of the design methodology & algorithm for the class of Horizontal Takeoff and Horizontal Landing (HTHL) SAVs. The HTHL SAV design application clearly outlines how the conceptual design process can be centrally organized, executed and documented with focus on design transparency, physical understanding and the capability to reproduce results. This approach offers the project lead and creative design team a management process and tool which iteratively refines the individual design logic chosen, leading to mature design methods and algorithms. As illustrated, the HTHL SAV hands-on design methodology offers growth potential in that the same methodology can be continually updated and extended to other SAV configuration concepts, such as the Vertical Takeoff and Vertical Landing (VTVL) SAV class. Having developed, validated and calibrated the methodology for HTHL designs in the 'hands-on' mode, the report provides an outlook how the methodology will be integrated into a prototype computerized design synthesis software AVDS-PrADOSAV in a follow-on step.Today's and especially tomorrow's competitive launch vehicle design environment requires the development of a dedicated generic Space Access Vehicle (SAV) design methodology. A total of 115 industrial, research, and academic aircraft, helicopter, missile, and launch vehicle design synthesis methodologies have been evaluated. As the survey indicates, each synthesis methodology tends to focus on a specific flight vehicle configuration, thus precluding the key capability to systematically compare flight vehicle design alternatives. The aim of the research investigation is to provide decision-making bodies and the practicing engineer a design process and tool box for robust modeling and simulation of flight vehicles where the ultimate performance characteristics may hinge on numerical subtleties. This will enable the designer of a SAV for the first time to consistently compare different classes of SAV configurations on an impartial basis

    Controls and guidance research

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    The objectives of the control group are concentrated on research and education. The control problem of the hypersonic space vehicle represents an important and challenging issue in aerospace engineering. The work described in this report is part of our effort in developing advanced control strategies for such a system. In order to achieve the objectives stated in the NASA-CORE proposal, the tasks were divided among the group based upon their educational expertise. Within the educational component we are offering a Linear Systems and Control course for students in electrical and mechanical engineering. Also, we are proposing a new course in Digital Control Systems with a corresponding laboratory

    Evolution of the SpaceLiner towards a Reusable TSTO-Launcher

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    Since a couple of years the DLR launcher systems analysis division is investigating a visionary and extremely fast passenger transportation concept based on rocket propulsion. The fully reusable concept consists of two vertically launched winged stages in parallel arrangement. The space transportation role of the SpaceLiner concept as a TSTO-launcher is now, for the first time, addressed in technical detail. Different mission options to LEO and beyond are traded and necessary modifications of the passenger stage to an unmanned cargo-carrier are investigated and described in this paper. Meanwhile, technical progress of the SpaceLiner ultra-high-speed passenger transport is ongoing at Phase A conceptual design level. Iterative sizings of all major subcomponents in nominal and off-nominal flight conditions have been performed. Potential intercontinental flight routes, taking into account range-safety and sonic boom constraints as well as good reachability from major business centers, are evaluated and flight guidance schemes are established. Alternative passenger cabin and rescue capsule options with innovative morphing shapes were also investigated. The operational and business concept of the SpaceLiner is under definition. The project is on a structured development path and as one key initial step the Mission Requirements Review has been successfully concluded

    Framework for the Parametric System Modeling of Space Exploration Architectures

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    This paper presents a methodology for performing architecture definition and assessment prior to, or during, program formulation that utilizes a centralized, integrated architecture modeling framework operated by a small, core team of general space architects. This framework, known as the Exploration Architecture Model for IN-space and Earth-to-orbit (EXAMINE), enables: 1) a significantly larger fraction of an architecture trade space to be assessed in a given study timeframe; and 2) the complex element-to-element and element-to-system relationships to be quantitatively explored earlier in the design process. Discussion of the methodology advantages and disadvantages with respect to the distributed study team approach typically used within NASA to perform architecture studies is presented along with an overview of EXAMINE s functional components and tools. An example Mars transportation system architecture model is used to demonstrate EXAMINE s capabilities in this paper. However, the framework is generally applicable for exploration architecture modeling with destinations to any celestial body in the solar system

    Structures and materials technology issues for reusable launch vehicles

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    Projected space missions for both civil and defense needs require significant improvements in structures and materials technology for reusable launch vehicles: reductions in structural weight compared to the Space Shuttle Orbiter of up to 25% or more, a possible factor of 5 or more increase in mission life, increases in maximum use temperature of the external surface, reusable containment of cryogenic hydrogen and oxygen, significant reductions in operational costs, and possibly less lead time between technology readiness and initial operational capability. In addition, there is increasing interest in hypersonic airbreathing propulsion for launch and transmospheric vehicles, and such systems require regeneratively cooled structure. The technology issues are addressed, giving brief assessments of the state-of-the-art and proposed activities to meet the technology requirements in a timely manner

    Optimal and Robust Control of Atmospheric Reentry Trajectories

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    Reentering the Earth’s atmosphere is one of the most difficult phases of any spaceship’s mission. During reentry, following an optimal trajectory in terms of minimal heating rate, dynamic pressure, and maximum deceleration is vital to the mission’s success. This dissertation proposes a novel design for a controller based on H8 control methods, with the goal of achieving optimal and robust control for a Reentry Lifting Vehicle. This dissertation begins by summarizing the theory and history behind the Shuttle Space Program, including the planning of its trajectories and descriptions of the Reentry Flight Dynamics Model. The analysis of the obtained reference trajectory will follow and the comparison of it to a real shuttle trajectory will be made. The design and configuration of the H8 controller comes after, beginning with the linearization of the obtained system in the previous step and ending with the recalculation of vehicle state variables after its application. The disturbance is then applied and the results of the actuation of the controller are displayed. The H8 controller for Reentry Lifting Vehicle proves itself to be a useful application presenting satisfying and significant results in this critical phase of flight.A reentrada na atmosfera terrestre é uma das fases mais difíceis da missão de qualquer nave espacial. Durante a reentrada, seguir uma trajectória óptima em termos de taxa mínima de aquecimento, pressão dinâmica e maáxima desaceleração é vital para o sucesso da missão. Esta dissertação propõe o design de um controlador baseado em métodos de H8, com o objectivo de alcançar um controlo óptimo e robusto para um Veículo de Reentrada na atmosfera capaz de produzir sustentação. Esta dissertação começa com um resumo da teoria e história por detrás do Programa Espacial do Vaivém, incluindo o planeamento de trajectórias e descrições do Modelo de Dinâmica de Voo de Reentrada. A análise da trajectória de referência obtida seguir­se­á e será feita a sua comparação com uma trajectória real do vaivém. A concepção e configuração do controlador H8 virá depois, começando com a linearização do sistema obtida no passo anterior e terminando com o recálculo das variáveis de estado do veículo após a sua aplicação. A perturbação é então aplicada e os resultados do accionamento do controlador são exibidos. O controlador H8 para Veículo de Reentrada capaz de produzir Sustentação mostrou ser uma aplicação útil apresentando resultados satisfatórios e significativos nesta fase crítica de voo
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