Trajectory optimization is an important part of launch vehicle design and operation. With the high costs of launching payload into orbit, every pound that can be saved increases affordability. One way to save weight in launch vehicle design and operation is by optimizing the ascent trajectory. Launch vehicle trajectory optimization is a field that has been studied since the 1950’s. Originally, analytic solutions were sought because computers were slow and inefficient. With the advent of computers, however, different algorithms were developed for the purpose of trajectory optimization. Computer resources were still limited, and as such the algorithms were limited to local optimization methods, which can get stuck in specific regions of the design space. Local methods for trajectory optimization have been well studied and developed. Computer technology continues to advance, and in recent years global optimization has become available for application to a wide variety of problems, including trajectory optimization. The aim of this thesis is to create a methodology that applies global optimization to the trajectory optimization problem. Using information from a global search, the optimization design space can be reduced and a much smaller design space can be analyzed using already existing local methods. This allows for areas of interest in the design space to be identified and further studied and helps overcome the fact that many local methods can get stuck in local optima. The design space included in trajectory optimization is also considered in this thesis. The typical optimization variables are initial conditions and flight control variables. For direct optimization methods, the trajectory phase structure is currently chosen a priori. Including trajectory phase structure variables in the optimization process can yield better solutions. The methodology and phase structure optimization is demonstrated using an earth-to-orbit trajectory of a Delta IV Medium launch vehicle. Different methods of performing the global search and reducing the design space are compared. Local optimization is performed using the industry standard trajectory optimization tool POST. Finally, methods for varying the trajectory phase structure are presented and the results are compared.

Orientadores: João Mauricio Rosario, Francisco Carlos Parquet Bizarria. / Tese (doutorado) - Universidade Estadual de Campinas, Faculdade de Engenharia Mecanica / Made available in DSpace on 2018-08-12T03:58:51Z (GMT). No. of bitstreams: 1 Barbosa_SilvanaAparecida_D.pdf: 3254271 bytes, checksum: dfd0e2d57b3560c893dad3675337cfe9 (MD5) Previous issue date: 2008 / Resumo: A tarefa contínua de detecção de mudança de cenários em determinado ambiente pode ocasionar estresse físico e mental para o supervisor. Com o uso da visão computacional associada à automação é possível oferecer recursos para que essa tarefa seja executada de forma automatizada. Neste contexto, este trabalho apresenta uma proposta de arquitetura para um sistema detectar mudanças de cenários por meio da comparação sucessiva de imagens. Esse sistema de visão deverá possibilitar por meio da implementação de um algoritmo, utilizando recursos matemáticos específicos para a aplicação, a detecção de mudança no cenário e o registro das imagens relativas a essa situação. Para validar a referida proposta, com o enfoque na aplicação aeroespacial, é apresentado nesta tese o estudo de caso para aplicação de técnicas de Visão Artificial na Detecção de Mudança de Cenários em Plataforma de Integração de Veículos Espaciais. Esse estudo visa supervisionar áreas destinadas principalmente à integração e aos testes desse veículo espacial. A implementação desse algoritmo utiliza técnicas de processamento de imagens baseada em filtros espaciais, em especial convolução por filtro da média. Os resultados satisfatórios, obtidos nos ensaios práticos realizados em protótipo da Torre Móvel de Integração, indicam que o sistema é adequado para aplicação a qual se destina. / Abstract: The continuous task of detecting changes in scenarios can cause a mental and physical stress on supervisor. The associated use of computational vision and automation provides resources to execute this task in an automated way. This thesis presents an architectural proposal for a system to detect changes in the imaged scenes by comparing images successively. Through the development of algorithm using specific mathematical tools, this vision system detects and record changes in the imaged scenes. Focusing aerospace area to validate this proposal, Artificial Vision Applied in Detection of Change of Scenes, a Study of Case in Platform for Integration of Space Vehicles is presented in this thesis. The purpose of this research is supervising the designed areas for integration and tests of the space vehicle. Image processing techniques based on spatial filters, in particular convolution, are used for algorithm development. The results obtained through practical tests executed on a prototype of an Integration Mobile Tower, show the system is appropriate to this application. / Doutorado / Mecanica dos Sólidos e Projeto Mecanico / Doutor em Engenharia Mecânica

Proper engine selection for Reusable Launch Vehicles (RLVs) is a key factor in the design of low cost reusable launch systems for routine access to space. RLVs typically use combinations of different types of engines used in sequence over the duration of the flight. Also, in order to properly choose which engines are best for an RLV design concept and mission, the optimal trajectory that maximizes or minimizes the mission objective must be found for that engine configuration. Typically this is done by the designer iteratively choosing engine combinations based on his/her judgment and running each individual combination through a full trajectory optimization to find out how well the engine configuration performed on board the desired RLV design. This thesis presents a new method to reliably predict the optimal engine configuration and optimal trajectory for a fixed design of a conceptual RLV in an automated manner. This method is accomplished using the original code Steele-Flight. This code uses a combination of a Genetic Algorithm (GA) and a Non-Linear Programming (NLP) based trajectory optimizer known as GPOPS II to simultaneously find the optimal engine configuration from a user provided selection pool of engine models and the matching optimal trajectory. This method allows the user to explore a broad range of possible engine configurations that they wouldn't have time to consider and do so in less time than if they attempted to manually select and analyze each possible engine combination. This method was validated in two separate ways. The codes ability to optimize trajectories was compared to the German trajectory optimizer suite known as ASTOS where only minimal differences in the output trajectory were noticed. Afterwards another test was performed to verify the method used by Steele-Flight for engine selection. In this test, Steele-Flight was provided a vehicle model based on the German Saenger TSTO RLV concept and models of turbofans, turbojets, ramjets, scramjets and rockets. Steele-Flight explored the design space through the use of a Genetic Algorithm to find the optimal engine combination to maximize payload. The results output by Steele-Flight were verified by a study in which the designer manually chose the engine combinations one at a time, running each through the trajectory optimization routine to determine the best engine combination. For the most part, these methods yielded the same optimal engine configurations with only minor variation. The code itself provides RLV researchers with a new tool to perform conceptual level engine selection from a gathering of user provided conceptual engine data models and RLV structural designs and trajectory optimization for fixed RLV designs and fixed mission requirement. / Master of Science

O sistema de controle por alocação dos pólos com filtro foi utilizado para controlar a atitude de um modelo de veículo lançador de satélites. Com este intuito, foram confeccionados um modelo e uma base de fixação que permite a movimentação nos três graus de liberdade. Utilizando a resposta à entrada degrau em conjunto com um sistema de controle PID obtido de forma empírica para estabilizar o sistema, as características da planta foram identificadas e então o sistema de controle por alocação de pólos foi projetado. Este sistema apresentou uma oscilação em torno da referência com amplitude menor do que 0,5° e tempo de pico para a entrada degrau na ordem de 2,17 segundos. Um segundo controlador PID foi projetado de forma analítica para se obter uma referência, porém apresentou resposta com características inferiores ao controlador por alocação de pólos. Os dois sistemas de controle projetados conseguem manter o modelo estável mesmo quando um dos motores é desligado. / Pole placement control system with filter was implemented to control the attitude of a satellite launch vehicle model. With this purpose, a model and a fixing base with three degrees of freedom was made. Utilizing the system response to step input with PID controller empirically designed to stabilize the system, the model characteristics were identified and the pole placement control system was designed. This system oscillated around the reference with amplitude smaller than 0.5° and peak time around 2.17 seconds. Another PID controller was designed analytically for reference, however the pole placement controller had better response characteristics than the PID controller. Both controllers can stabilize the system even when one engine is shut off.

O sistema de controle por alocação dos pólos com filtro foi utilizado para controlar a atitude de um modelo de veículo lançador de satélites. Com este intuito, foram confeccionados um modelo e uma base de fixação que permite a movimentação nos três graus de liberdade. Utilizando a resposta à entrada degrau em conjunto com um sistema de controle PID obtido de forma empírica para estabilizar o sistema, as características da planta foram identificadas e então o sistema de controle por alocação de pólos foi projetado. Este sistema apresentou uma oscilação em torno da referência com amplitude menor do que 0,5° e tempo de pico para a entrada degrau na ordem de 2,17 segundos. Um segundo controlador PID foi projetado de forma analítica para se obter uma referência, porém apresentou resposta com características inferiores ao controlador por alocação de pólos. Os dois sistemas de controle projetados conseguem manter o modelo estável mesmo quando um dos motores é desligado. / Pole placement control system with filter was implemented to control the attitude of a satellite launch vehicle model. With this purpose, a model and a fixing base with three degrees of freedom was made. Utilizing the system response to step input with PID controller empirically designed to stabilize the system, the model characteristics were identified and the pole placement control system was designed. This system oscillated around the reference with amplitude smaller than 0.5° and peak time around 2.17 seconds. Another PID controller was designed analytically for reference, however the pole placement controller had better response characteristics than the PID controller. Both controllers can stabilize the system even when one engine is shut off.

No description available.

A novel trajectory and attitude control and navigation analysis tool for powered ascent is developed. The tool is capable of rapid trade-space analysis and is designed to ultimately reduce turnaround time for launch vehicle design, mission planning, and redesign work. It is streamlined to quickly determine trajectory and attitude control dispersions, propellant dispersions, orbit insertion dispersions, and navigation errors and their sensitivities to sensor errors, actuator execution uncertainties, and random disturbances. The tool is developed by applying both Monte Carlo and linear covariance analysis techniques to a closed-loop, launch vehicle guidance, navigation, and control (GN&C) system. The nonlinear dynamics and flight GN&C software models of a closed-loop, six-degree-of-freedom (6-DOF), Monte Carlo simulation are formulated and developed. The nominal reference trajectory (NRT) for the proposed lunar ascent trajectory is defined and generated. The Monte Carlo truth models and GN&C algorithms are linearized about the NRT, the linear covariance equations are formulated, and the linear covariance simulation is developed. The performance of the launch vehicle GN&C system is evaluated using both Monte Carlo and linear covariance techniques and their trajectory and attitude control dispersion, propellant dispersion, orbit insertion dispersion, and navigation error results are validated and compared. Statistical results from linear covariance analysis are generally within 10% of Monte Carlo results, and in most cases the differences are less than 5%. This is an excellent result given the many complex nonlinearities that are embedded in the ascent GN&C problem. Moreover, the real value of this tool lies in its speed, where the linear covariance simulation is 1036.62 times faster than the Monte Carlo simulation. Although the application and results presented are for a lunar, single-stage-to-orbit (SSTO), ascent vehicle, the tools, techniques, and mathematical formulations that are discussed are applicable to ascent on Earth or other planets as well as other rocket-powered systems such as sounding rockets and ballistic missiles.

Rocket Factory Augsburg (RFA) a German New Space Startup is developing a three-stage rocket launcher aiming at LEO/SSO orbits. A fundamental responsibility of the GNC team is the development of the rocket navigation algorithm to estimate the attitude, position, and velocity allowing the guidance and control loops to autonomously steer the rocket. This thesis focuses on the analysis and design of a Hybrid Navigation system able to satisfy the various necessities of a launch vehicle, such as delay compensation and GNSS outages. The navigation architecture was chosen to be a Closed Loop, Loosely Coupled, Delayed Error State Kalman Filter thanks to the proven capability of COTS receivers to autonomously provide a consistent PVT solution throughout the flight. A preliminary analysis used a reference trajectory to evaluate the effect of the sensor grade on inertial performances and choose an appropriate integration scheme. The filter’s system model was explored using approximate analytical results on observability. The developed navigation module was then tested within a Monte Carlo simulation environment by perturbating the sensor parameter in accordance with the sensor datasheet. As a further verification, the modeled IMU output was compared to the engineering model, to assure that the simulation result would yield conservative errors. Due to concern over the visibility of GNSS satellites during flight, a simplified Almanac-based GPS model has been developed, proving that enough satellite visibility is available along the trajectory. The estimation error was compared with the filter’s estimated covariance and found well within the bounds. Through the study of the covariance evolution, it was determined that given the reference dynamics, the sensor misalignments are the least observable states. Realistic signal outages were introduced in the most critical flight intervals. The filter was indeed found to be robust and the tuning proved to be adequate to capture the dead reckoning drift. Finally, the entire navigation module was deployed onto the avionics engineering model, including the flight computer, IMU, GNSS, and antennas, in a configuration equivalent to flight. The navigation module was then tested to ensure that the execution was in performance under severe multipath errors and prolonged GNSS outages with the covariance estimates correctly covering the uncertainty. / Rocket Factory Augsburg (RFA), ett tyskt nystartat rymdföretag, utvecklar en trestegsraket som siktar på LEO/SSO-banor. Ett grundläggande ansvar för GNC-teamet är utvecklingen av raketnavigationsalgoritmen för att uppskatta attityd, position och hastighet så att styr- och kontrollslingorna kan styra raketen autonomt. Avhandlingen fokuserar på analys och design av ett hybridnavigeringssystem som kan uppfylla de olika krav som ställs på en bärraket, såsom kompensation för fördröjningar och GNSS-avbrott. Navigationsarkitekturen valdes att vara ett Closed Loop, Loosely Coupled, Delayed Error State Kalman Filter tack vare den bevisade förmågan hos COTS-mottagare att autonomt tillhandahålla en konsekvent PVT-lösning under hela flygningen. En preliminär analys använde en referensbana för att utvärdera effekten av sensorkvaliteten på tröghetsprestanda och välja ett lämpligt integrationsschema. Filtrets systemmodell undersöktes med hjälp av approximativa analytiska resultat om observerbarhet. Den utvecklade navigeringsmodulen testades sedan i en Monte Carlo-simuleringsmiljö genom att störa sensorparametern i enlighet med sensorns datablad. Som en ytterligare verifiering jämfördes den modellerade IMU-utgången med den tekniska modellen, för att säkerställa att simuleringsresultatet skulle ge konservativa fel. På grund av oro över GNSS-satelliternas synlighet under flygning har en förenklad Almanac-baserad GPS-modell utvecklats, som bevisar att tillräcklig satellitsikt finns tillgänglig längs banan. Uppskattningsfelet jämfördes med filtrets uppskattade kovarians och låg väl inom gränserna. Genom att studera kovariansutvecklingen fastställdes det att givet referensdynamiken är sensorernas feljusteringar de minst observerbara tillstånden. Realistiska signalavbrott infördes i de mest kritiska flygintervallen. Filtret visade sig verkligen vara robust och inställningen visade sig vara tillräcklig för att fånga upp dödberäkningens drift. Slutligen installerades hela navigeringsmodulen på den flygtekniska modellen, inklusive flygdator, IMU, GNSS och antenner, i en konfiguration som motsvarar en flygning. Navigationsmodulen testades sedan för att säkerställa att utförandet var i prestanda under allvarliga multipath-fel och långvariga GNSS-avbrott med kovariansuppskattningarna som korrekt täcker osäkerheten.

This dissertation seeks to merge the ideas from robust control theory such as H-Infinity control design and the Small Gain Theorem, L stability theory and Lyapunov stability from nonlinear control, and recent theoretical achievements in adaptive control. The fusion of frequency domain and linear time domain ideas allows the derivation of an H-Infinity Norm Minimization Approach (H-Infinity-NMA) for adaptive control architecture that permits a control designer to simplify the adaptive tuning process and tune the uncertainty compensation characteristics via linear control design techniques, band limit the adaptive control signal, efficiently handle redundant actuators, and handle unmatched uncertainty and matched uncertainty in a single design framework. The two stage design framework is similar to that used in robust control, but without sacrificing performance. The first stage of the design considers an ideal system with the system uncertainty completely known. For this system, a control law is designed using linear H-Infinity theory. Then in the second stage, an adaptive process is implemented that emulates the behavior of the ideal system. If the linear H-Infinity design is applied to control the emulated system, it then guarantees closed loop system stability of the actual system. All of this is accomplished while providing notions of transient performance bounds between the ideal system and the true system. Extensions to the theory include architectures for a class of output feedback systems, limiting the authority of an adaptive control system, and a method for improving the performance of an adaptive system with slow dynamics without any modification terms. Applications focus on using aerodynamic flow control for aircraft flight control and the Crew Launch Vehicle.

Satellites have become absolute necessity in the growing modern space technology. At present, launch pads are the only means for launching of satellites or any other space vehicles. Due to the large magnitude of current and the associated rate of rise, a lightning strike to launch pads can be quite disastrous. Satellite launch complex forms typically the tallest object in that region. This makes them the more vulnerable to cloud-to-ground lightning. In addition, most of the launch pads are situated near the coastal area, where the isokeraunic levels are quite high. In view of these, almost all the satellite launch pads are provided with suitable Lightning Protection Systems (LPS). The LPS is basically intended for protecting against a direct lightning hit. The present work is related with the LPS to Indian satellite launch pads, Pad-I and Pad-II. The protection system for Pad-I consists of three 120 m tall towers placed approximately at the vertices of an equilateral triangle of 180 m. The same for Pad-II consist of 120 m tall towers placed at vertices of rectangle of size 90 m x 105 m. Towers are interconnected by 6 shield wires at the top. A mast of 10m length forms the top of the tower. Significant work on the analysis of interception efficacy of these protection systems has been reported in the literature. The lightning surge response of these systems have also been analysed and reported. The interception efficacy of these LPS in field can be ascertained by pertinent measurements. Measuring the lightning current on LPS seems to be one of the most suitable choices for this purpose. It would also greatly facilitate collection of local lightning current statistics, data on which is almost absent. Several considerations suggest that the tower bases form ideal place for such measurement. However, such lightning current records would involve mainly the current resulting from stroke interception, as well as, induced current due to strokes nearby. Literature on categorisation of measured currents to the type of stroke and correlation of measured currents to the incident stroke currents is rather limited. This is especially true for interconnected protection system of the type dealt in the present work. Considering these the present work is taken up and its scope is defined as: (i) Evolve a suitable model for study of current distribution in LPS due to Lightning and using the same deduce the current due to stroke interception and that due to stroke nearby. (ii) For the purpose of categorization identify the salient characteristics of current due to the intercepted strokes and that due to bypass/nearby strokes (iii) For the intercepted strokes, develop a processor for estimating the injected stroke current from the measured tower base currents. Lightning event, apart from other associated physical phenomena, is strongly governed by electromagnetic fields. Any method employed for the analysis, either theoretical or experimental, should satisfy the governing electromagnetic equations. As experimentation on actual system, as well as, their laboratory simulation is nearly impossible, theoretical modelling approach is selected. Modelling involves modelling of the channel along with its excitation, modelling of the LPS and modelling of the ground. Channel, following the literature, is represented as a loaded conductor with a lumped current source at the junction point. Such models have quite successfully predicted the electromagnetic fields and current in other places on the down conductor. For the LPS, some simplifications on the geometry are very essential. Tower lattice elements of dimensions much smaller than the wavelength of highest dominant frequency component of lightning current spectrum are neglected. Suitable modification is made for the tower top involving a plate and interconnection of several short members. For the close range within 200 – 400 m, even for the induced currents, the influence of ground in the literature has been reported to be small. Also, there is an extensive grounding network in these systems. In view of the same, a perfectly conducting ground along with suitable ground termination impedance is considered. Only the numerical solution of the problem is feasible and for the same, following the literature, NEC-2 is employed. All the guidelines of NEC are respected in the discretisation. Geometric mean radius is employed for modelling the complex tower elements. Fourier Transform Techniques are employed for time domain conversion of the computed frequency domain quantities. Occasionally, numerical inversion error of magnitude less than 5% is encountered. For the validation of the numerical modelling for both direct stroke and that nearby, time domain experimentation on electromagnetically reduced scale models (35:1) is employed. As the channel electrical and geometrical parameters are stochastic in nature, it is necessary to ensure that the deduction made using the model is practically relevant. For this, some parametric studies are conducted. The influence of channel length and inclination, stroke current velocity etc. has been shown to be insignificant for the case of intercepted strokes. Simulations are carried out for the stroke intercepted (i.e. direct strikes) by the LPS. The characteristics of the tower base currents are investigated. The base currents indicate a dispersive propagation along the towers and further a frequency dependent current division at the tower-shield wire junctions. Base currents contain superimposed oscillations, which basically originate from various junctions of the system. The magnitude of the oscillations is obviously dependent on the rise time of the incident currents. The tower base currents settle within about 10 -15 µs, which is shorter than that for isolated tower. Further, the full-frequency model could be limited to this time period. The corresponding current transfer functions are deduced. For the stroke interception by shield wires, based on the earlier work, only stroke to midspan is found to be relevant and hence it is considered. The nature of tower base currents for a stroke to midspan of the shield wires seem to be similar. However there are some distinct features, which are helpful in identifying the stroke location on the LPS. From the time correlated tower base currents, a suitable methodology for identifying the stroke interception location on LPS is developed. Next, simulations for induced current due to a bypass stroke, as well as, stroke to ground outside the LPS, however, within 1 km radius are taken up. In fact, it is estimated that latter is nearly 5 – 13 times higher than the strokes collected by LPS, indicating it as the most probable event. The objective here is characterization, rather than correlation. In this study, the influence of charge induced on the LPS by the descending leader is neglected and the upward leader activity is approximately considered. To the best of author’s knowledge, studies on such induced currents in down conductors are very scarce. Considering this and noting that the number of parameters is quite large, first the basic study is taken up on simple cylindrical down conductors. Many important and interesting deductions are made. The nature of the induced current is highly dependent on the rate of rise as well as the velocity of propagation of the stroke current. The magnitude and to some extent, the wave shape of the induced current is found to depend on the average as well as maximum di/dt of the stroke current. For a given wave shape, the magnitude of the induced current increases with rate of rise of the wave front; however, saturating trend will onset after some point. The height of the down conductor mainly governs the frequency of the oscillatory component of the induced current. The dependency of the induced current on the radius of the down conductor seems to be logarithmic (which is in accordance with the antenna theory). Based on these results, the parameters for the corresponding study on LPS under consideration, is chosen. The results of the investigation on the induced currents in LPS show that they have quite distinct waveform. They are basically bipolar and oscillatory in nature, with relatively short duration. These unique features facilitate clear distinction of the induced currents from that due to stroke interception. Basic characteristics are reasonably insensitive to the separation distance of the protection system and the channel, current propagation velocity along the channel, channel inclination and shape of the current front. The salient features of the induced current due to a bypass stroke are also enumerated. • The noise, if any, in the measured current can be addressed only after acquiring sufficient data. Based on the above, the following procedure is suggested for the stroke classification and estimation. • By employing the distinct features of the resulting tower base currents, analyze the measured tower base currents and classify the strokes into the intercepted stroke or stroke to ground. • For the latter case, using the salient features of the bypass strokes, further classify the strokes to bypass strokes and stroke to ground outside the protected volume. • For the intercepted strokes, using the relative strengths and wave shapes, identify the interception point to either tower top or the midspan of the shield wires. • Then by using the corresponding transfer functions and Fourier Transform techniques, compute the injected stroke current. • Using the above, other tower base currents are computed and compared with the measured currents. This gives quantification for the accuracy of the method. In summary the present work has made some original contribution to the classification and estimation of stroke currents measured on the interconnected LPS.