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1
Torres, Calderon Rosa Maria. „Multi-objective environmentally-friendly departure procedures for civil aircraf : modeling, optimization and robustness assessment“. Toulouse 3, 2011. http://www.theses.fr/2011TOU30320.
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Cette thèse présente le concept MCDP (Multi-Criteria Departure Procedure), qui a comme objectif l’optimisation des procédures de départ afin de minimiser l’impact environnemental des avions. Ces impacts portent essentiellement sur les nuisances acoustiques, la qualité de l’air et l’émission des gaz d’effet de serre. Il n’existe pas une seule procédure capable de minimiser les trois critères simultanément. Le concept MCDP est donc modélisé par un problème d’optimisation multi-objectif sous contraintes. La technique de recherche directe MADS est utilisée pour obtenir les trajectoires Pareto optimales associées au problème d’optimisation. Des indicateurs de robustesse sont ensuite construits pour analyser l’impact des incertitudes sur le front de Pareto. L’application des méthodologies développées aux avions Airbus montre le potentiel intérêt d’optimiser la phase de départ en termes environnementauxThis Ph. D. Work introduces the MCDP (Multi-Criteria Departure Procedure) concept, the aim of which is to optimize civil aircraft departure procedures in order to minimize their environmental impact, mainly represented by noise, local air quality and greenhouse gas emissions. It is not possible to determine a single departure procedure minimizing the three objectives simultaneously. Therefore, the concept is modeled as a multi-objective, constrained optimization problem. The MADS direct search optimization technique is used to obtain the Pareto optimal departure procedures associated with this problem. Robustness indicators have then been built to determine the impact of uncertainty on the Pareto front. Application to Airbus aircraft of the methodologies developed shows the potential interest of optimizing the departure phase in terms of environmental benefits
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Urbano, Simone. „Detection and diagnostic of freeplay induced limit cycle oscillation in the flight control system of a civil aircraf“. Thesis, Toulouse, INPT, 2019. http://www.theses.fr/2019INPT0023/document.
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Cette étude est le résultat d’une thèse CIFRE de trois ans entre le bureau d’étude d’Airbus (domaine du contrôle de l’avion) et le laboratoire TéSA à Toulouse. L’objectif principal est de proposer, développer et valider une solution logicielle pour la détection et le diagnostic d’un type spécifique de vibrations des gouvernes de profondeur et direction, appelée oscillation en cycle limite (limit cycle oscillation ou LCO en anglais), basée sur les signaux existants dans les avions civils. LCO est un terme mathématique générique définissant un mode périodique indépendant de conditions initiales et se produisant dans des systèmes non linéaires non conservatifs. Dans cette étude, nous nous intéressons au phénomène de LCO induit par les jeux mécaniques dans les gouvernes d’un avion civil. Les conséquences du LCO sont l’augmentation locale de la charge structurelle, la dégradation des qualités de vol, la réduction de la durée de vie de l’actionneur, la dégradation du confort du poste de pilotage et de la cabine, ainsi que l’augmentation des coûts de maintenance. L’état de l’art en matière de détection et de diagnostic du LCO induit par le jeu mécanique est basé sur la sensibilité du pilote aux vibrations et sur le contrôle périodique du jeu sur les gouvernes. Cette étude propose une solution basée sur les données (issues de la boucle d’asservissement des actionneurs qui agissent sur les gouvernes) pour aider au diagnostic du LCO et à l’isolement du jeu mécanique. L’objectif est d’améliorer encore plus la disponibilité des avions et de réduire les coûts de maintenance en fournissant aux compagnies aériennes un signal de contrôle pour le LCO et les jeux mécaniques. Pour cette raison, deux solutions algorithmiques pour le diagnostic des vibrations et des jeux ont été proposées. Un détecteur en temps réel pour la détection du LCO est tout d’abord proposé basé sur la théorie du rapport de vraisemblance généralisé (generalized likelihood ratio test ou GLRT en anglais). Certaines variantes et simplifications sont également proposées pour satisfaire les contraintes industrielles. Un détecteur de jeu mécanique est introduit basé sur l’identification d’un modèle de Wiener. Des approches paramétrique (estimateur de maximum de vraisemblance) et non paramétrique (régression par noyau) sont explorées, ainsi que certaines variantes des méthodes non paramétriques. En particulier, le problème de l’estimation d’un cycle d’hystérésis (choisi comme la non-linéarité de sortie d’un modèle de Wiener) est abordé. Ainsi, les problèmes avec et sans contraintes sont étudiés. Une analyse théorique, numérique (sur simulateur) et expérimentale (données de vol et laboratoire) est réalisée pour étudier les performances des détecteurs proposés et pour identifier les limitations et la faisabilité industrielle. Les résultats numériques et expérimentaux obtenus confirment que le GLRT proposé (et ses variantes / simplifications) est une méthode très efficace pour le diagnostic du LCO en termes de performance, robustesse et coût calculatoire. D’autre part, l’algorithme de diagnostic des jeux mécaniques est capable de détecter des niveaux de jeu relativement importants, mais il ne fournit pas de résultats cohérents pour des niveaux de jeu relativement faibles. En outre, des types d’entrée spécifiques sont nécessaires pour garantir des résultats répétitifs et cohérents. Des études complémentaires pourraient être menées afin de comparer les résultats de GLRT avec une approche Bayésienne et pour approfondir les possibilités et les limites de la méthode paramétrique proposée pour l’identification du modèle de WienerThis research study is the result of a 3 years CIFRE PhD thesis between the Airbus design office(Aircraft Control domain) and TéSA laboratory in Toulouse. The main goal is to propose, developand validate a software solution for the detection and diagnosis of a specific type of elevator andrudder vibration, called limit cycle oscillation (LCO), based on existing signals available in flightcontrol computers on board in-series aircraft. LCO is a generic mathematical term defining aninitial condition-independent periodic mode occurring in nonconservative nonlinear systems. Thisstudy focuses on the LCO phenomenon induced by mechanical freeplays in the control surface ofa civil aircraft. The LCO consequences are local structural load augmentation, flight handlingqualities deterioration, actuator operational life reduction, cockpit and cabin comfort deteriorationand maintenance cost augmentation. The state-of-the-art for freeplay induced LCO detection anddiagnosis is based on the pilot sensitivity to vibration and to periodic freeplay check on the controlsurfaces. This study is thought to propose a data-driven solution to help LCO and freeplaydiagnosis. The goal is to improve even more aircraft availability and reduce the maintenance costsby providing to the airlines a condition monitoring signal for LCO and freeplays. For this reason,two algorithmic solutions for vibration and freeplay diagnosis are investigated in this PhD thesis. Areal time detector for LCO diagnosis is first proposed based on the theory of the generalized likeli hood ratio test (GLRT). Some variants and simplifications are also proposed to be compliantwith the industrial constraints. In a second part of this work, a mechanical freeplay detector isintroduced based on the theory of Wiener model identification. Parametric (maximum likelihoodestimator) and non parametric (kernel regression) approaches are investigated, as well as somevariants to well-known nonparametric methods. In particular, the problem of hysteresis cycleestimation (as the output nonlinearity of a Wiener model) is tackled. Moreover, the constrainedand unconstrained problems are studied. A theoretical, numerical (simulator) and experimental(flight data and laboratory) analysis is carried out to investigate the performance of the proposeddetectors and to identify limitations and industrial feasibility. The obtained numerical andexperimental results confirm that the proposed GLR test (and its variants/simplifications) is a very appealing method for LCO diagnostic in terms of performance, robustness and computationalcost. On the other hand, the proposed freeplay diagnostic algorithm is able to detect relativelylarge freeplay levels, but it does not provide consistent results for relatively small freeplay levels. Moreover, specific input types are needed to guarantee repetitive and consistent results. Further studies should be carried out in order to compare the GLRT results with a Bayesian approach and to investigate more deeply the possibilities and limitations of the proposed parametric method for Wiener model identification
3
Dvořáček, Zdeněk. „Návrh hlavního podvozku pro letouny řady Zlín 40“. Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2021. http://www.nusl.cz/ntk/nusl-442850.
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This master thesis deals with the modernization of the main landing gear of the Z 143 LSi aircraft. The main emphasis is placed on optimizing of the landing gear shape and its deformation characteristics. Along with this change, variants made of different materials are examined. The landing load is calculated analytically and the load-bearing capacity of the structure is verified using the numerical FEM method. This work brings a possible replacement of the current chassis, which is production-intensive not only in time and money, but also in terms of cooperation with external companies. The objective is to design a new landing gear and verify the feasibility for the case of future implementation.
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Bauer, David O. „A historical perspective of aircrew systems effects on aircraft design“. Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 1996. http://handle.dtic.mil/100.2/ADA320281.
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Thesis (M.S. in Aeronautical Engineering) Naval Postgraduate School, September 1996.Thesis advisor(s): Conrad F. Newberry. "September 1996." Includes bibliographical references (p. 99-100). Also available online.
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Megas, Vasileios. „Aircraft to Aircraft Connectivity Analysis“. Thesis, KTH, Skolan för elektroteknik och datavetenskap (EECS), 2018. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-254887.
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It is very common to feel the need to be connected, especially in cases where we are idle, such as in airplanes. Aircraft connectivity has already been implemented by using direct air-to-ground communication and satellite communication. However, since the current solutions are not satisfying, more ways to provide connectivity to aircraft are being researched. One solution is to use aircraft as relay nodes and forward signals to aircraft which are out of range of the ground base stations, creating an ad-hoc network.This thesis aims in evaluating the performance of such networks over the North Atlantic ocean in terms of connectivity and achievable data rate by means of simulations. First, an aircraft mobility model is created by analyzing past flight data. Then, the topology of the network is created with the purpose of maximizing the number of connected aircraft and minimizing interference. Finally, the available bandwidth is allocated to all connected aircraft in a maxmin way. The simulations were repeated to evaluate the impact of: threshold in data rates to assume connectivity, maximum number of connections per aircraft, maximum antenna steering angle, beamwidth and bandwidth.Our results show that aircraft to aircraft connectivity is feasible if all airlines cooperate, and more than 80% of the aircraft can find a path to a base station, when the aircraft density is higher than 1 aircraft per approximately 150000 km2, which is 88% of a 24 hours period. The median data rate of the connected aircraft, achieved in the middle of the ocean is limited to 25 Mbps with a 20MHz bandwidth, which is not high enough to provide access to demanding applications, such as video streaming. Increasing the bandwidth to 200 MHz can increase the data rate to more than 184 Mbps, for 50% of the connected aircraft, which surpasses the performance of current satellite communication systems.Det är ganska vanligt att känna behovet av att sammankopplas, särskilt i fall där vi är inaktiva, exempelvis i flygplan. Flygplansanslutning har redan implementerats genom att utnyttja direkt luft till markkommunikation och satellitkommunikation. Men eftersom de befintliga lösningarna inte är tillfredsställande, undersöks fler sätt att tillhandahålla anslutning till flygplan. En lösning är att använda flygplan som relä knutpunkter och framåtsignaler till flygplan som ligger utanför markbasen, skapar ett adhocnätverk.Den här avhandlingen syftar till att utvärdera prestanda för sådana nätverk över Nordatlanten när det gäller anslutning och uppnåelig datahastighet genom simuleringar. För det första frambringas en flygplansmobilitet genom att analysera tidigare flygdata. Därefter åstadkommas nätets topologi med syftet att maximera antalet anslutna flygplan och minimera störningar. Slutligen tilldelas den tillgängliga bandbredden till alla anslutna flygplan på maximal väg. Simuleringarna upprepas för att utvärdera effekten av: tröskeln i datahastigheter för att mottaga anslutning, maximalt antal anslutningar per flygplan, max antennstyrvinkel, strålbredd och bandbredd.Våra resultat visar att flygplan till flygplansförbindelser är genomförbart om alla flygbolag samarbetar och mer än 80% av flygplanet kan sammankoppla, när flygplansdensiteten är högre än 1 flygplan per ungefär 150000 km2. Median datahastigheten uppnådd i mitten av havet är begränsad till 25 Mbps med en 20 MHz bandbredd, vilket inte är tillräckligt hög för att bringa tillgång till krävande applikationer, såsom videostreaming. Ö kning av bandbredd till 200 MHz kan öka datahastigheten mer än 184 Mbps, för 50% av de anslutna flygplanen, vilketöverträffar prestanda för de nuvarande satellitkommunikationssystemen.
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Federico, Claudio Antonio. „Dosimetria da radiação cósmica no interior de aeronaves no espaço aéreo Brasileiro“. Universidade de São Paulo, 2011. http://www.teses.usp.br/teses/disponiveis/85/85131/tde-26012012-104504/.
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Neste trabalho, cujo objetivo foi fornecer subsídios que permitam estabelecer um sistema de dosimetria para as tripulações de aeronaves que operam no território nacional, foi desenvolvida uma técnica para realizar medições do equivalente de dose ambiente em aeronaves. Foi testado um detector ativo, para uso embarcado em aeronaves, avaliando-se sua adequação a este tipo específico de medição, assim como sua suscetibilidade às interferências magnética e eletromagnética. O equipamento foi calibrado em campos metrologicamente rastreados e foi verificado em um campo especial do laboratório europeu CERN, que reproduz com grande proximidade o espectro de partículas existentes em altitudes de voo de aeronaves, sendo a seguir testado em diversos ensaios em voo, em uma aeronave da Força Aérea Brasileira. Os resultados foram avaliados e comparados com aqueles obtidos por diversos programas computacionais de estimativa de dose oriunda de radiação cósmica, com respeito à sua adequação para uso na região da América do Sul e foi selecionado o programa CARI-6 para avaliar as doses efetivas médias estimadas a que estão sujeitas as tripulações que operam nesta região. Foi efetuada uma avaliação estatística da distribuição de doses efetivas das tripulações dos voos que operam na região da América do Sul e Caribe e concluiu-se que grande parte destas tripulações estão sujeitas a doses efetivas anuais que ultrapassam o limite estabelecido para membros do público. Adicionalmente, foi proposto um dosímetro passivo preliminar, baseado em detectores termoluminescentes, e foram ainda estabelecidas colaborações internacionais com grupos de pesquisa do Reino Unido e da Itália, para a realização de medições conjuntas do equivalente de dose ambiente em aeronaves.The objective of this work is the establishment of a dosimetric system for the aircrew in the domestic territory. A technique to perform measurements of ambient dose equivalent in aircrafts was developed. An active detector was evaluated for onboard aircraft use, testing its adequacy to this specific type of measurement as well as its susceptibility to the magnetic and electromagnetic interferences. The equipment was calibrated in standard radiation beams and in a special field of the European Laboratory CERN, that reproduces with great proximity the real spectrum in aircraft flight altitudes; it was also tested in several flights, in an Brazilian Air Force\'s aircraft. The results were evaluated and compared with those obtained from several computational programs for cosmic radiation estimatives, with respect to its adequacy for use in the South American region. The program CARI-6 was selected to evaluate the estimated averaged effective doses for the aircrew who operate in this region. A statistical distribution of aircrew effective doses in South America and Caribe was made, and the results show that a great part of this aircrew members are subjected to annual effective doses that exceed the dose limits for the members of the public. Additionally, a preliminary passive dosemeter, based in thermoluminescent detectors, was proposed; international collaborations with United Kingdom and Italy were established for joint measurements of the ambient equivalent doses in aircrafts.
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Serrano, Ignacio. „Unmanned Aircraft System (UAS) vs. Manned Aircraft System (MAS): A Military Aircraft Study“. Digital Commons at Loyola Marymount University and Loyola Law School, 2015. https://digitalcommons.lmu.edu/etd/430.
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Unmanned Aerial Vehicles (UAVs) are common place in the 21st century, whether they are small to medium sized remotely piloted vehicles (aka drones) or large advanced Unmanned Aerial Systems with a preprogrammed flight path. There is anticipation that these Unmanned Systems will, in the future assume the roles of their traditional manned aircraft counterparts. There is also the perception that these Unmanned Systems should be developed partly because they would be less expensive when compared to their manned aircraft. This integrative paper asserts that this perception is not reality with regards to developing a newly designed UAV to replace its manned counterpart, for the same mission. Through the examination of systems engineering principles between the unmanned RQ-4 Global Hawk and the manned U-2 Dragon Lady one will understand why this perception is not correct. Both aircraft perform the same mission of providing High Altitude Intelligence, Surveillance, and Reconnaissance (ISR). Through evaluation of requirements analysis both aircraft flowed down the requirements to all the various subsystems in a similar manner, creating similar subsystems for Imagery Intelligence (!MINT) and Signals Intelligence (SIGINT). However, the additional requirement for long endurance required that the Global Hawk systems engineers had additional requirements to flow down to the software, communications, data processing, and ground support subsystems in order to control an unmanned aircraft for greater than 24 hours. This one additional requirement had various derived requirements that needed to be verified, and validated during analysis, manufacturing, subsystem build and test, and final system integration. By using both System Integration Laboratories (SIL) and Flight Tests both systems requirements were verified and validated by the systems engineers. The Global Hawk since it was unmanned was required to perform more verification of subsystems and software as it was the first UAV to achieve flight airworthiness. The future of ISR missions requires that the aircraft become more adaptable to future technologies and situations. The U-2 has a modular configuration to change out to and from different subsystems depending on the mission. However, these subsystems were designed 20 to 30 years ago, and were not designed for lower level modularity or interoperability. The Global Hawk systems engineering team understood the future needs and the high level demand and data to be gathered and processed. The SE's developed modularity and interoperability requirements and flowed them down to the various subsystems. The Global Hawk system is more useful in highly contested areas of interest as there is no pilot; however resilient communications of the data and data link must be robust with anti-jamming capabilities to ensure the data is secure from cyber-attack. However,the U-2 is more survivable since it has a defense system, and can provide greater situational awareness. Taking all the general ISR requirements into consideration a trade study using a matrix was performed indicating that the Global Hawk is the most optimal solution to meeting both the current and future requirements for ISR missions. Even though the overall acquisition cost of the Global Hawk is equivalent to the U-2, systems engineering for Global Hawk had the responsibility to flow down requirements to all subsystems with consideration of the entire systems lifecycle. This is exemplified in that the Global Hawk is cost effective to fly in terms of cost per flight hour. Therefore, the Global Hawk can fulfill all the requirements of the given stakeholders with the lowest operational cost.
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Iskender, Hani. „Aircraft Simulator“. Thesis, Linköping University, Department of Electrical Engineering, 2005. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-2936.
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At Saab Bofors Dynamics there are projects running which purpose are to develop simulators for various weapon systems like RBS 70. In order to manage creating real working simulators Saab Bofors Dynamics has to do more research and this final thesis is a part of this process.
This final thesis has been performed at Saab Bofors Dynamics in the department of modelling and simulation, RTRKM, in Karlskoga. The purpose was to develop a control algorithm which makes it possible for an aircraft to behave real when controlling through a joystick.
The conclusions show that further improvements are needed before the aircraft behaves entirely by the laws of physics. Among other things it is necessary to decrease the number of delimitations that have been done.
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Thompson, Brian G. „Aircraft agility“. Thesis, This resource online, 1992. http://scholar.lib.vt.edu/theses/available/etd-09192009-040436/.
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Hacker, Earl W. „One-year UDP: a cost/benefit analysis of a proposed alternative to the Marine Corps' Unit Deployment Program for fighter aviation“. Thesis, Monterey, California. Naval Postgraduate School, 1988. http://hdl.handle.net/10945/23337.
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Approved for public release; distribution is unlimitedThe author examines the incremental costs and benefits associated with a change from six-month unit deployments to one-year unit deployments. The analysis is based primarily on five fighter squadrons participating in the Marine Corps' Unit Deployment Program and takes in the period July 1976 to October 1988. Regression analysis is used to project transportation cost savings of $4 million in real terms from FY 1989 through FY 1993. With a change to a one-year Unit Deployment Program, fighter squadrons should experience net increases in aircraft readiness, aircrew training readiness, and personnel retention.
http://archive.org/details/oneyearudpcostbe00hack
Lieutenant Colonel, United States Marine Corps
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Wan, Husain Wan Mohd Sufian Bin. „Maintainability prediction for aircraft mechanical components utilising aircraft feedback information“. Thesis, Cranfield University, 2011. http://dspace.lib.cranfield.ac.uk/handle/1826/7272.
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The aim of this research is to propose an alternative approach to determine the
maintainability prediction for aircraft components. In this research, the author looks at
certain areas of the maintainability prediction process where missteps or
misapplications most commonly occur. The first of these is during the early stage of
the Design for Maintainability (DfMt) process. The author discovered the importance
of utilising historical information or feedback information. The second area is during
the maintainability prediction where the maintenance of components is quantified;
here, the author proposes having the maximum target for each individual
maintainability component.
This research attempts to utilise aircraft maintenance historical data and
information (i.e. feedback information systems). Aircraft feedback information
contains various types of information that could be used for future improvement
rather than just the failure elements. Literature shows that feedback information such
as Service Difficulty Reporting System (SDRS) and Air Accidents Investigation Branch,
(AAIB) reports have helped to identify the critical and sensitive components that need
more attention for further improvement.
This research consists of two elements. The first is to identity and analyse
historical data. The second is to identify existing maintainability prediction
methodologies and propose an improved methodology. The 10 years’ data from
Federal Aviation Administration (FAA) SDRS data of all aircraft were collected and
analysed in accordance with the proposed methodology before the processes of
maintainability allocation and prediction were carried out.
The maintainability was predicted to identify the potential task time for each
individual aircraft component. The predicted tasks time in this research has to be in
accordance with industrial real tasks time were possible. One of the identified
solutions is by using maintainability allocation methodology. The existing
maintainability allocation methodology was improved, tested, and validated by using
several case studies. The outcomes were found to be very successful.
Overall, this research has proposed a new methodology for maintainability
prediction by integrating two important elements: historical data information, and
maintainability allocation. The study shows that the aircraft maintenance related
feedback information systems analyses were very useful for deciding maintainabilityeffectiveness; these include planning, organising maintenance and design
improvement. There is no doubt that historical data information has the ability to
contribute an important role in design activities. The results also show that
maintainability is an importance measure that can be used as a guideline for managing
efforts made for the improvement of aircraft components.
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Black, Gary Douglas. „Aircraft configuration study for experimental 2-place aircraft and RPVs“. Thesis, Monterey, California : Naval Postgraduate School, 1990. http://handle.dtic.mil/100.2/ADA227604.
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Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, March 1990.Thesis Advisor(s): Howard, Richard M. Second Reader: Hauser, James P. "March 1990." Description based on signature page as viewed on August 25, 2009. DTIC Descriptor(s): Aerodynamic configurations, remotely piloted vehicles, military aircraft, lightweight, research aircraft, canard configurations, wings, aerodynamic drag, computer programs, scaling factor, aerodynamic stability, performance(engineering), comparison, microcomputers, trade off analysis, aeronautical engineering, theses. DTIC Identifier(s): Pusher propellers, Tandem wing aircraft, Joined wing aircraft, Three surface aircraft. Author(s) subject terms: Aircraft design, remotely piloted vehicle, RPV, configuration, experimental Includes bibliographical references (p. 121-123). Also available online.
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Recktenwald, Bryan David Ahmed Anwar. „Aerodynamic testing of a circular planform concept aircraft“. Auburn, Ala, 2008. http://repo.lib.auburn.edu/EtdRoot/2008/SUMMER/Aerospace_Engineering/Thesis/Recktenwald_Bryan_48.pdf.
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Ionescu, Irina Gabriela. „Aircraft noise regulation“. Thesis, McGill University, 2004. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=82660.
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Aircraft noise is one of the most controversial environmental concerns in the aviation industry, partly due to the difficulty in harmonizing countries' regulation regarding this issue. The purpose of this thesis is to analyze the ways in which aircraft noise is regulated at the national and international levels, and to compare the legislative responses to aircraft noise issues in Europe and North America. Each of the four main chapters of the thesis takes into consideration a different aspect of the problem. The first chapter describes the objective and subjective ways of measuring aircraft noise. This process is necessary in order to allow the legislation to meet its purpose, namely, to protect the environment, the sources of the aircraft noise, and the effects of the aircraft noise on people. The second chapter describes the evolution of aircraft noise issues at the national levels in the US and throughout the EU, respectively, as well as at the international level, such as at the ICAO. The third chapter analyses the EU Regulation 925/1999, which created tension between the EU and the US due to its alleged discriminatory nature. This thesis examines the arguments of both sides. Finally, the fourth chapter analyses the noise certification standards developed by ICAO, namely the "balanced approach".
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彭遠輝 und Yuen-fai Alson Pang. „Managing aircraft noise“. Thesis, The University of Hong Kong (Pokfulam, Hong Kong), 2002. http://hub.hku.hk/bib/B31255280.
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Wagner, Michael J. „AEW aircraft design“. Thesis, Monterey, California. Naval Postgraduate School, 1992. http://hdl.handle.net/10945/23815.
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Brentnall, Adam Robert. „Aircraft arrival management“. Thesis, University of Southampton, 2006. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.432622.
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Pang, Yuen-fai Alson. „Managing aircraft noise /“. Hong Kong : University of Hong Kong, 2002. http://sunzi.lib.hku.hk/hkuto/record.jsp?B2543598x.
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Elsayed, M. A. N. „Aircraft trajectory optimization“. Thesis, Loughborough University, 1985. https://dspace.lboro.ac.uk/2134/32873.
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A typical aircraft flight consists of three phases, namely climb, cruise and descent. The purpose of this research was to study the control schedules for a transport aircraft which would result in least fuel expenditure in each flight-segment.
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GongZhang, Hanlin, und Eric Axtelius. „Aircraft Winglet Design“. Thesis, KTH, Skolan för teknikvetenskap (SCI), 2020. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-276586.
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Aerodynamic drag can be decreased with respect to a wing’s geometry, and wingtip devices, so called winglets, play a vital role in wing design. The focus has been laid on studying the lift and drag forces generated by merging various winglet designs with a constrained aircraft wing. By using computational fluid dynamic (CFD) simulations alongside wind tunnel testing of scaled down 3D-printed models, one can evaluate such forces and determine each respective winglet’s contribution to the total lift and drag forces of the wing. At last, the efficiency of the wing was furtherly determined by evaluating its lift-to-drag ratios with the obtained lift and drag forces. The result from this study showed that the overall efficiency of the wing varied depending on the winglet design, with some designs noticeable more efficient than others according to the CFD-simulations. The shark fin-alike winglet was overall the most efficient design, followed shortly by the famous blended design found in many mid-sized airliners. The worst performing designs were surprisingly the fenced and spiroid designs, which had efficiencies on par with the wing without winglet.Det aerodynamiska luftmotståndet kan minskas genom justeringar i vingens geometri, och vingtipsenheter så kallade virveldämpare spelar en vital roll inom vingdesign och utveckling. Projektet hade fokuset på att undersöka de lyft -och motståndskrafterna som genererades av en förutbestämd vinge vid användandet av olika varianter av virveldämpare. Genom beräkningsströmningsdynamiska simuleringar (CFD) och vindtunneltester av nerskalade 3D-utskrivna modeller kan dessa krafter beräknas. Med hjälp av dessa kan respektive virveldämparens bidrag till de sammanlagda lyft -och motståndskrafterna på vingen vidare bestämmas. Genom att beräkna förhållandet mellan de erhållna lyft -och motståndskrafterna kan vingens glidtal sist bestämmas, som är ett dimensionslöst mått på vingens effektivitet. Resultatet från denna studie visade att vingens prestanda varierade efter på valet av virveldämpare, där vissa designer gav avsevärda förbättringar över andra enligt CFD-simuleringar. Den hajfensliknande varianten (så kallad sharklet på engelska) var den mest effektiva designen, tätt följt av den ordinära uppåtriktade typen (blended på engelska) som återfinns på många av dagens medelstora flygplan. De sämst presterande designerna var förvånande nog den triangel -och spiralformade (fenced och spiroid på engelska), vilka presterade på samma nivå som den virveldämparfria vingen.
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Vallone, Michael. „Parameter Estimation of Fundamental Technical Aircraft Information Applied to Aircraft Performance“. DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/382.
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Inverse problems can be applied to aircraft in many areas. One of the disciplines within the aerospace industry with the most openly published data is in the area of aircraft performance. Many aircraft manufacturers publish performance claims, flight manuals and Standard Aircraft Characteristics (SAC) charts without any mention of the more fundamental technical information of the drag and engine data. With accurate tools, generalized aircraft models and a few curve-fitting techniques, it is possible to evaluate vehicle performance and estimate the drag, thrust and fuel consumption (TSFC) with some accuracy. This thesis is intended to research the use of aircraft performance information to deduce these aircraft--specific drag and engine models. The proposed method incorporates models for each performance metric, modeling options for drag, thrust and TSFC, and an inverse method to match the predicted performance to the actual performance. Each of the aircraft models is parametric in nature, allowing for individual parameters to be varied to determine the optimal result. The method discussed in this work shows both the benefits and pitfalls of using performance data to deduce engine and drag characteristics. The results of this method, applied to the McDonnell Douglas DC-10 and Northrop F-5, highlight many of these benefits and pitfalls, and show varied levels of success. A groundwork has been laid to show that this concept is viable, and extension of this work to additional aircraft is possible with recommendations on how to improve this technique.
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Hintze, Joshua. „Autonomous landing of a rotary unmanned aerial vehicle in a non-cooperative environment using machine vision /“. Diss., CLICK HERE for online access, 2004. http://contentdm.lib.byu.edu/ETD/image/etd359.pdf.
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Thate, Timothy J. Michels Adam S. „Requirements for digitized aircraft spotting (Ouija) board for use on U.S. Navy aircraft carriers /“. Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02sep%5FThate.pdf.
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Thesis (M.S. in Information Systems Management)--Naval Postgraduate School, September 2002.Thesis advisor(s): Alex Bordetsky, Glenn Cook. Includes bibliographical references. Also available online.
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Li, Bo Lim Alvin S. „Multiple UAV simulation with multiresolution multistage models and decision support“. Auburn, Ala., 2008. http://hdl.handle.net/10415/1547.
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Miao, Zhisong. „Aircraft engine performance and integration in a flying wing aircraft conceptual design“. Thesis, Cranfield University, 2012. http://dspace.lib.cranfield.ac.uk/handle/1826/7249.
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The increasing demand of more economical and environmentally friendly aero
engines leads to the proposal of a new concept – geared turbofan. In this thesis,
the characteristics of this kind of engine and relevant considerations of
integration on a flying wing aircraft were studied.
The studies can be divided into four levels: GTF-11 engine modelling and
performance simulation; aircraft performance calculation; nacelle design and
aerodynamic performance evaluation; preliminary engine installation.
Firstly, a geared concept engine model was constructed using TURBOMATCH
software. Based on parametric analysis and SFC target, the main cycle
parameters were selected. Then, the maximum take-off thrust was verified and
corrected from 195.56kN to 212kN to meet the requirements of take-off field
length and second segment climb. Besides, the engine performance at offdesign
points was simulated for aircraft performance calculation.
Secondly, an aircraft performance model was developed and the performance
of FW-11 was calculated on the basis of GTF-11 simulation results. Then, the
effect of GTF-11 characteristics performance on aircraft performance was
evaluated. A comparison between GTF-11 and conventional turbofan, RB211-
524B4, indicated that the aircraft can achieve a 13.1% improvement in fuel
efficiency by using the new concept engine.
Thirdly, a nacelle was designed for GTF-11 based on NACA 1-series and
empirical methods while the nacelle dimensions of conventional turbofan
RB211-525B4 were obtained by measure approach. Then, the installation thrust
losses caused by nacelle drags of the two engines were evaluated using ESDU
81024a. The results showed that the nacelle drags account for about 4.08%
and 3.09% of net thrust for GTF-11 and RB211-525B4, respectively.
Finally, the considerations of engine installation on a flying wing aircraft were
discussed and a preliminary disposition of GTF-11 on FW-11 was presented.
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Seresinhe, R. „Impact of aircraft systems within aircraft operation : a MEA trajectory optimisation study“. Thesis, Cranfield University, 2014. http://dspace.lib.cranfield.ac.uk/handle/1826/9261.
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Air transport has been a key component of the socio-economic globalisation. The ever increasing demand for air travel and air transport is a testament to the success of the aircraft. But this growing demand presents many challenges. One of which is the environmental impact due to aviation. The scope of the environmental impact of aircraft can be discussed from many viewpoints. This research focuses on the environmental impact due to aircraft operation. Aircraft operation causes many environmental penalties. The most obvious is the fossil fuel based fuel burn and the consequent greenhouse gas emissions. Aircraft operations directly contribute to the CO2 and NOX emissions among others. The dependency on a limited natural resource such as fossil fuel presents the case for fuel optimised operation. The by-products of burning fossil fuel some of which are considered pollutants and greenhouse gases, presents the case for emissions optimised operations. Moreover, when considering the local impact of aircraft operation, aircraft noise is recognised as a pollutant. Hence noise optimised aircraft operation needs to be considered with regards to local impacts. It is clear whichever the objective is, optimised operation is key to improving the efficiency of the aircraft. The operational penalties have many different contributors. The most obvious of which is the way an aircraft is flown. This covers the scope of aircraft trajectory and trajectory optimisation. However, the design of the aircraft contributes to the operational penalties as well. For example the more-electric aircraft is an improvement over the conventional aircraft in terms of overall efficiency. It has been proven by many studies that the more-electric concept is more fuel efficient than a comparable conventional aircraft. The classical approach to aircraft trajectory optimisation does not account for the fuel penalties caused due to airframe systems operation. Hence the classical approach cannot define a conventional aircraft from a more-electric aircraft. With the more-electric aircraft expected to be more fuel efficient it was clear that optimal operation for the two concepts would be different. This research presents a methodology that can be used to study optimised trajectories for more-electric aircraft. The study present preliminary evidence of the environmental impact due to airframe systems operation and establishes the basis for an enhanced approach to aircraft trajectory optimisation which include airframe system penalties within the optimisation loop. It then presents a suite of models, the individual modelling approaches and the validation to conduct the study. Finally the research presents analysis and comparisons between the classical approach where the aircraft has no penalty due to systems, the conventional aircraft and the more-electric aircraft. When the case studies were optimised for the minimum fuel burn operation, the conventional airframe systems accounted for a 16.6% increase in fuel burn for a short haul flight and 6.24% increase in fuel burn for a long haul flight. Compared to the conventional aircraft, the more electric aircraft had a 9.9% lower fuel burn in the short haul flight and 5.35% lower fuel burn in the long haul flight. However, the key result was that the optimised operation for the moreelectric aircraft was significantly different than the conventional aircraft. Hence this research contributes by presenting a methodology to bridge the gap between theoretical and real aircraft-applicable trajectory optimisation.
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Flodin, Linus, und Oscar Hag. „Energy Efficient Concept Aircraft“. Thesis, KTH, Farkost och flyg, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-102886.
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A concept study for an environmentally friendly aircraft is made with the help of mathematical models. The aircraft has a capacity of 4 passengers and is powered by electrical engines and lithium-air batteries so that no greenhouse gases are emitted. The aim of the study is to find a concept that can compete with today’s petrol powered aircraft, with applications in entertainment flying and transportation. Flight conditions such as steady state flight and steady climbing are modeled to estimate the performance of the aircraft and to find the most efficient way to fly. A static stability analysis is made to determine a geometry for which the aircraft is statically stable in pitch. The study shows that a canard configuration together with pusher contra rotating propellers is a good way to design a very efficient aircraft. It is also found out that the most energy efficient way to climb is to use a high rate of climb and a low velocity and that it can be profitable to increase the weight of the aircraft with more batteries.
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Zhu, Zhongxi. „The aircraft rotation problem“. Diss., Georgia Institute of Technology, 1994. http://hdl.handle.net/1853/21788.
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Swift, Adam. „Simulation of aircraft aeroelasticity“. Thesis, University of Liverpool, 2011. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.569519.
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Aeroelastic phenomena such as flutter can have a detrimental effect on aircraft performance and can lead to severe damage or destruction. Buffet leads to a re- duced fatigue life and therefore higher operating costs and a limited performance envelope. As such the simulation of these aeroelastic phenomena is of utmost importance. Computational aeroelasticity couples computational fluid dynamics and computational structural dynamics solvers through the use of a transforma- tion method. There have been interesting developments over the years towards more efficient methods for predicting the flutter boundaries based upon the sta- bility of the system of equations. This thesis investigates the influence of transformation methods on the flutter boundary predition and considers the simulation of shock-induced buffet of a transport wing. This involves testing a number of transformation methods for their effect on flutter boundaries for two test cases and verifying the flow solver for shock-induced buffet over an aerofoil. This will be followed by static aeroelastic calculations of an aeroelastic wing. It is shown that the transformation methods have a significant effect on the predicted flutter boundary. Multiple transformation methods should be used to build confidence in the results obtained, and extrapolation should be avoided. CFD predictions are verified for buffet calculations and the mechanism behind shock-oscillation of the BGK No. 1 aerofoil is investigated. The use of steady calculations to assess if a case may be unsteady is considered. Finally the static aeroelastic response of the ARW-2 wing is calculated and compared against ex- perimental results.
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Sonei, Arash. „Designing A Conventional Aircraft“. Thesis, KTH, Skolan för teknikvetenskap (SCI), 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-153225.
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This paper is explaining the important design phases of dimensioning an unmanned conventional aircraft from scratch and will also design one according to a few chosen requirements. The design phases discussed will be all from wing dimensioning to stability and spin recovery, aircraft performance requirements and how to select a motor which overcomes these. As well as the optimal rate of climb for improved efficiency is discussed. In the end an aircraft which manages the set requirements and is stable in pitch managing spin recovery with no problem will have been dimensioned.
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Hartland, A. J. „Dispersed base combat aircraft“. Thesis, Cranfield University, 1989. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.237543.
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Heffernon, Timothy James. „Aircraft noise installation effects“. Thesis, University of Southampton, 2017. https://eprints.soton.ac.uk/415885/.
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Airframe noise is currently of a comparable level to engine noise for an aircraft on approach with high-lift devices and landing gears deployed. The landing gears are a large contributor to the overall airframe noise in this situation. Main landing gears are typically installed beneath a lifting wing. The wing surfaces act as scattering surfaces for the noise generated by these landing gears, and the non-uniform flow around the wing affects both the propagation and strength of the noise. This thesis focuses on investigating the propagation and scattering of installed landing gear noise sources. Boundary element methods are capable of computing acoustic scattering by large and complex geometries, such as a complete aircraft geometry. However, due to their use of Green’s functions, flow effects can only be approximated. As a result, the refraction of acoustic waves due to a non-uniform flow is not accounted for. A uniform flow formulation based on a Lorentz-type transform is typically employed with boundary element methods. The effect of neglecting refraction on the propagation and scattering of landing gear noise sources is determined in this thesis. Investigations are conducted using computational aeroacoustic methods that solve the linearised Euler equations, which account for the refraction of acoustic waves due to non-uniform flow. Using computational aeroacoustic methods, the effect of non-uniform flow due to circulation on the acoustic scattering is quantified as the difference in acoustic scattering over uniform and non-uniform base flows. These investigations are conducted using both single frequency and broadband monopole sources, and both single-element and multi-element airfoils. Increasing the angle of attack, increasing the Mach number, and deploying flaps all increase the circulation around the airfoil. The effect of varying these parameters is investigated systematically. It is shown that for a source in the approximate position of a landing gear with flow conditions similar to that of an airliner on approach, the largest difference observed is at single frequencies for an airfoil configuration with a deployed flap. Otherwise, the differences are small, and in some cases so small that they can be considered negligible. It is shown that moving the source to a position above the airfoil and using a higher Mach number gives a larger difference, although this is not representative of a landing gear source. A new method is proposed to generate a broadband input signal for use with a computational aeroacoustic solver that gives a specified power spectral density at a given radial distance from a monopole source. A signal that is equal in power across a specified range of frequencies is generated using this method. The effect on the frequency content of the scattered noise from a broadband source installed beneath a lifting wing is investigated using this generated signal. It is shown for a single-element airfoil that the major contributor to the obtained power spectral density is the distance of the source from the airfoil. Varying the angle of attack and Mach number has only a small additional effect on the power spectral density. It is then shown that flap and slat deployment has a larger effect on the computed power spectral density due to the additional reflective surfaces. Existing boundary element method formulations that estimate uniform and nonuniform flow effects are evaluated for their suitability for landing gear noise scattering predictions. It is shown that the uniform flow formulation is more suitable due to a simplifying assumption made in the derivation of the non-uniform flow formulation. An existing realistic landing gear noise model is coupled with a three-dimensional acoustic boundary element method solver. The landing gear noise model applies scaling laws to directional databases for isolated landing gear components in order to estimate the total far-field noise. The implemented coupling methodology is used to compute the sound pressure level on a ground plane beneath a realistic scattering aircraft geometry. The geometrical effect of flap deployment is investigated using sources of constant strength for each configuration. It is shown that the effect of flap deployment is to increase the sound pressure level directly below and in the region immediately surrounding the aircraft. The effect of source strength reduction due to circulation around a lifting wing is then included in the predictions. This results in a large decrease in the predicted sound pressure level on the ground plane with flap deployment.
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Tsentseris, Michail. „ARES: aircraft propulsion integration“. Thesis, Cranfield University, 2013. http://dspace.lib.cranfield.ac.uk/handle/1826/12121.
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The combat aircraft and the gas turbine engine interact with one another. These
interactions have become of even greater significance than previously due to the fact
that the modern aircraft have increased requirements in terms of thrust, efficiency,
reliability and ease of maintenance. Issues that affect the performance and operational
capability of combat aircraft are the aircraft aerodynamic characteristics, propulsion
system performance and the airframe/engine matching.
This study focuses on the evaluation of the propulsion system of modern jet
powered combat aircraft and combines computations with the existing integration
methodology. The main objective of this study is the update of ARES method (aircraftengine
performance tool for combat aircraft) so as to be feasible the incorporation of
new data in the calculations of the program. The update contains:
• the establishment of a procedure that can be used for the calculation of
the additional drag that a external load (tanks, bombs, missiles, pods) contribute in the
total aircraft drag,
• the modification of ARES method in order to integrate the additional
drag of the external loads in the program calculations,
• the creation of a procedure that will be used for the calculation of the
thrust deductions due to installation effects.
The information that can be received from this study are valuable because can be
used in the evaluation of the engine performance, can give the opportunity for further
engine studies such as the structural and thermal analysis, component sizing and
geometry, life consumption and of course can make apparent if a specific aircraftengine
system is capable of accomplishing a demanded operating requirement.
The final results of the study led to the general conclusion that the total aircraft
drag and the demanded uninstalled net thrust increased with the addition of external
tanks in the aircraft. The fuel consumption was higher for the configuration with the two
tanks and was noticed an increase of the TET which is proportional to the number of the
used external tanks (max TET increases at least 0.8% per tank). Additionally was
extracted the conclusion that the addition of an external wing tank increases the amount
of the consumed fuel at least 3.5%.
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Karlsson, Albin, und Anton Lomaeus. „Transport Aircraft Conceptual Design“. Thesis, KTH, Skolan för teknikvetenskap (SCI), 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-210778.
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A conceptual design for a transport aircraft has been created, tailored for human-itarian missions along the equator with its home base in the European Union while optimizing for fuel eciency and speed. An initial estimate of the empty weight was made using historical data and Breguet equations, based on a required payload of 60 tonnes and range of 5 500 nautical miles. A constraint diagram consisting of require-ments for stall speed, takeo distance, climb rate and landing distance was used to determine wing loading and thrust to weight ratio, resulting in a main wing area of 387m 2 and thrust to weight ratio of 0:224, for which two Rolls Royce Trent 1000-H engines were selected. A high aspect ratio wing was designed with blended winglets to optimize against lift induced drag. Wing placement and tail volume were decided by iterative calculations, resulting in a centre of lift located aft of the centre of gravity during all stages of the mission. The resulting aircraft model has a high wing with a span of 62 m, length of 49m with a takeo gross weight of 221 tonnes, of which 83 tonnes are fuel.
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Filho, Nelson Whitaker. „Aircraft Distance Measurement System“. International Foundation for Telemetering, 1994. http://hdl.handle.net/10150/611674.
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International Telemetering Conference Proceedings / October 17-20, 1994 / Town & Country Hotel and Conference Center, San Diego, CaliforniaThe Aircraft Distance Measurement System (ADMS) could be used in Flight Test application to determine the aircraft position and speed during takeoff, landing and acceleration-stop performance test within runway limits using a microwave link.
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Andersson, Henric. „Aircraft Systems Modeling : Model Based Systems Engineering in Avionics Design and Aircraft Simulation“. Licentiate thesis, Linköping University, Linköping University, Machine Design, 2009. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-17573.
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Aircraft developers like other development and manufacturing companies, are experiencing increasing complexity in their products and growing competition in the global market. One way to confront the challenges is to make the development process more efficient and to shorten time to market for new products/variants by using design and development methods based on models. Model Based Systems Engineering (MBSE) is introduced to, in a structured way, support engineers with aids and rules in order to engineer systems in a new way.
In this thesis, model based strategies for aircraft and avionics development are studied. A background to avionics architectures and in particular Integrated Modular Avionics is described. The integrating discipline Systems Engineering, MBSE and applicable standards are also described. A survey on available and emerging modeling techniques and tools, such as Hosted Simulation, is presented and Modeling Domains are defined in order to analyze the engineering environment with all its vital parts to support an MBSE approach.
Time and money may be saved by using modeling techniques that enable understanding of the engineering problem, state-of-the-art analysis and team communication, with preserved or increased quality and sense of control. Dynamic simulation is an activity increasingly used in aerospace, for several reasons; to prove the product concept, to validate stated requirements, and to verify the final implementation. Simulation is also used for end-user training, with specialized training simulators, but with the same underlying models. As models grow in complexity, and the set of simulation platforms is expanded, new needs for specification, model building and configuration support arise, which requires a modeling framework to be efficient.
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Bergkvist, Erik, und Tommy Sabbagh. „Smart Future Solutions for Maintenance of Aircraft : Enhancing Aircraft Maintenance at Saab AB“. Thesis, Linköpings universitet, Industriell Produktion, 2021. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-176561.
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This master thesis has the purpose to analyse and identify smart efficient future solutions within the Maintenance, Repair and Overhaul (MRO) process for aircraft. The efficiency solutions, in form of new technologies and tools, should present a foundation that MRO suppliers can continue to develop to enhance and streamline their maintenance processes. The project was performed as a case study at the aerospace and defense company Saab AB in Linköping, where the company's MRO process was investigated. Through identifying possibilities and alternative technologies available today and in the near future, one continues to have a competitive and future-proof position in the market. Through an own constructed course of action influenced from established methods, the thesis' purpose and aim was attained. The method was based on authentic approaches for case studies but also inspired by the so-called "Requirement Engineering". The combination of the methods resulted in a precise focus on the relevant subjects, together with a clear structure of the requirements on the technologies to reach a successful implementation. Through a detailed data collection comprised of study visits, interviews, literature studies, market analyses, and document reviews, multiple relevant technologies and requirement-lists for utilization were identified. To concretize the use and potential improvements with the technologies, the project had the objective to develop a demonstrator with one of the technologies presented. The demonstrator should focus on minimizing the use of paper, which is a common problem among many market actors today. The most promising technology was considered to be a tablet application with an accommodated application. The selection of the tablet solution was based on the motivation that it is a well-established technology and a favorable first step from paperwork. To conclude this master thesis, a tablet application was developed in Novacura Flow Studio, where the majority of the identified requirements were fulfilled. Beyond the demonstrator, an introduction and analysis of technologies, such as AR-glasses, voice guidance, additive manufacturing, and a digital twin, was presented.
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Cummins, Lane, und Tony Wilborn. „Cost-benefit analysis of the Department of the Navy's transition from C-9 Aircraft to C-40 Aircraft for logistic support aircraft“. Monterey, California: Naval Postgraduate School, 2009. http://hdl.handle.net/10945/10395.
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Approved for public release, distribution unlimitedMBA Professional Report
The Navy began transitioning from the aging C-9s, which peaked at a total fleet size of 27 C-9B/DC-9 aircraft, to the C-40A. However, in response to increasing defense budget scrutiny and competing priorities, the Navy has decided to put this program on hold. Although the C-9B is an aging airframe and will require mandatory Federal Aviation Administration mandated modifications and upgrades, the DC-9/C-9B airframe has recently been determined to have significant operational service life remaining. This project provides a Cost-benefit Analysis (CBA) of the changes associated with replacing the C-9B aircraft with the C-40A. We analyze three alternatives. The first assumes that the C-40A acquisition program will remain on-hold indefinitely. The second alternative foresees the C-40A acquisition resuming as currently projected in FY2015. The third alternative involves the original C-40A acquisition program as per Naval Air Plan 2030 (NAP 2030). The objective is to compare the three alternatives, choosing the alternative, which provides the greatest net benefit and most efficient use of resources. The analysis will involve data collection of operational costs per flight hour, and total costs over the life of the program. Our cost-benefit analysis is intended to indicate the best course of action to provide continued execution of the Navy's Unique Fleet Essential Aircraft (NUFEA) mission. We intend to document all costs incurred and potential savings from a transition to the C-40 aircraft. We find that Alternative Three has lower discounted costs, as well as lower risk and better capabilities, and therefore recommend the Navy transition back to the original C-40 program as soon as practical.
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Cummins, Lane Wilborn Tony. „Cost-benefit analysis of the Department of the Navy's transition from C-9 Aircraft to C-40 Aircraft for logistic support aircraft“. Monterey, California : Naval Postgraduate School, 2009. http://edocs.nps.edu/npspubs/scholarly/MBAPR/2009/Dec/09Dec%5FCummins%5FMBA.pdf.
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"Submitted in partial fulfillment of the requirements for the degree of Master of Business Administration from the Naval Postgraduate School, December 2009."Advisor(s): Franck, Raymond E. ; Hildebrandt, Gregory. "December 2009." "MBA Professional report"--Cover. Description based on title screen as viewed on January 27, 2010. Author(s) subject terms: Cost-benefit Analysis, C-9B Skytrain, C-40A Clipper, NUFEA, Fleet Logistic Support Wing, Acquisition. Includes bibliographical references (p. 85-86). Also available in print.
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Kehayas, N. „ASTOVL Combat Aircraft Design Synthesis and Optimization“. Thesis, Cranfield University, 1992. http://hdl.handle.net/1826/3387.
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This thesis presents the development of a Baseline
Configuration for an Advanced Short Take-Off and Vertical
Landing (ASTOVL) Combat Aircraft, the Design Synthesis
and coding of this Baseline Congiguration (Code VERTI),
the interfacing of the Design Synthesis Code VERTI with
the Optimizer code RQPMIN and the optimization of the
Baseline Configuration.
The background and the objectives of this Research
Programme are initially examined. The evaluation of the
ASTOVL Combat Aircraft Baseline Configuration is then
described, including all the problems, assumptions,
choices and compromises that led to the specific
configuration.
The development of the Design Synthesis and the Code
VERTI then follow, where the methodology used, the
techniques adopted and the code operation are explained.
A full description of the Design Synthesis is included as
an appendix.
Finally, the interfacing of Code VERTI with the optimizer
RQPMIN and the optimization of the Baseline configuration
are presented. The problems and difficulties of the
RQPMIN operation are thoroughly discussed. The RQPMIN-
VERTI code is used to optimize the initial Baseline
Configuration and an optimization example is provided in
appendix form. The optimized Baseline Configuration is
partly validated against two ASTOVL combat aircraft
designs. In addition to the optimization with the
aircraft empty mass as objective function, a search for a
better objective function is attempted.
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Altman, Aaron. „A conceptual design methodology for low speed high altitude long endurance unmanned aerial vehicles“. Thesis, Cranfield University, 2000. http://hdl.handle.net/1826/3998.
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A conceptual design methodology was produced and subsequently coded into a
Visual C++ (GUI) environment to facilitate the rapid comparison of several possible
configurations to satisfy High Altitude Long Endurance (FIALE) unmanned aircraft (UAV)
missions in the Low Speed (propeller driven aircraft) regime.
Several comparative studies were performed to verify the applicability of traditional
design methods. The traditional computational design methodologies fail in several areas
such as high aspect ratio wing weight estimation and design, low Reynolds number wing
design, high altitude engine performance, low Reynolds number drag estimation, unmanned
aircraft design, and the conceptual design of unconventional configurations. The
methodology developed for this thesis was robust enough to allow not only for
consideration of these areas of inadequacy in traditional methods, but also to allow for the
inclusion of advancements in the relevant technologies as they become more widely
available.
The following configurations were evaluated for suitability to the Low Speed HALE
UAV application: conventional, canard, twin boom, multiple fuselage (conventional or
canard), tandem wing, multiple fuselage tandem wing or flying wing configuration. The
configurations were compared on the basis of aircraft endurance for takeoff weights ranging
from 2,000 to 20,000 pounds and wing loadings ranging from 5 to 25 lbs1fe.
Initial drag estimates were made using traditional parabolic drag estimation
techniques. A more refined drag buildup was performed using a vortex lattice drag
estimation for the lift induced drag (for all lifting components) and calculated skin friction
coefficients for the parasite drag. Statistically based methods were used for other
components of drag having much smaller contributions. In addition, a statistical approach
was taken to the weight estimation of the major aircraft components. However, this
approach made comparison of alternative configurations more difficult. Thus wing bending
moments trends were evaluated and utilized in the development of weight saving values for
multiple fuselage wing weight estimation.
The comparative performance of each configuration is justified with direct reference
to the terms in the Breguet Endurance equation. Validation was performed where possible
on all modules and segments associated with the methodology, as well as for the
macroscopic results. In addition, parametric studies on endurance were performed for the
conventional configuration for geometric characteristics and operating conditions directly
and indirectly effecting the calculated endurance and generalized results presented. Finally, a
case study was performed to demonstrate this capability.
A new relation was developed for aircraft empty weight prediction, a low speed
airfoil figure of merit was proposed, and new constants were offered for UAV fuselage
length prediction. In addition, horizontal and vertical tail volume coefficients were proposed
for all of the Low Speed HALE UAV configurations considered. It was determined that the
multiple fuselage configurations showed comparatively superior endurance performance
across a range of takeoff weights, with several other configurations demonstrating marginal
endurance improvements. Finally, a highly flexible and robust computer based conceptual
design methodology was developed and validated enabling the quick comparison of a greater
number of possible configurations to satisfy a given mission for Low Speed HALE UAV's
and providing detailed drag and weight breakdown data.
42
Alsina, J. „Development of an aircraft design expert system“. Thesis, Cranfield University, 1988. http://hdl.handle.net/1826/3442.
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The aircraft design process is characterised by the application
of a wide range of knowledge across many disciplines based upon a
certain degree of judgement and experience of the designer. A two
pass approach has been taken towards the development of an aircraft
design expert system based on the requirements of two conceptually
different design steps namely, wing design and aircraft configuration.
The current status of the work is one where an actual program for wing
design exists with supporting documentation, and a very effective
examination of the knowledge base performed based on the detail
investigation of overall aircraft design process with particular
emphasis on the wing design and the aircraft configuration design
steps.
The approach taken accomplishes the objectives of the current
research in defining the knowledge base, providing tools and
specifications for tools to be used within an aircraft design expert
system closely following the problem-solving techniques utilised by
the design expert.
43
Souahi, A. „Structural optimization of aircraft lifting surfaces to satisfy flutter requirements“. Thesis, Cranfield University, 1986. http://hdl.handle.net/1826/3657.
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The research reported in this thesis is concerned with
the structural weight optimization of aircraft lifting
surfaces when subjected to the satisfaction of flutter
requirements.
The main text is intended primarily as an expository
account on the work and as such it aims at introducing and
defining the subject of research and presenting the results.
Accordingly, the mathematics have been simplified to the
utmost in the main text and heavy theoretical treatments are
revealed in the appendices.
As the aim of this work is not directed at in-depth
studies of the physical nature of flutter nor for a
comprehensive treatment of structural optimization, the basic
concepts of these two subjects are touched upon in the
beginnings of chapters II and III respectively. We concluded
these two chapters by clarifying the class of flutter,
constraints and design variables for which the program we
developed is designed. We endeavored to keep the problem to
within certain practical boundaries without loosing too much
of either its generality or its applicability to structures
in realistic operational environments.
44
Broadley, Jonathan I. „The control of trailing edge separation on highly swept wings using vortex generators“. Thesis, Cranfield University, 1998. http://hdl.handle.net/1826/3966.
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The results from a series of low speed wind tunnel tests on two half model highly swept
wings (a symmetrical aerofoil section and a highly cambered aerofail section) are
presented in order to examine the trailing edge flow separation mechanism and its
development with wing sweep between 30' and 60'. The tests involved surface oil flow
visualisation, smoke flow visualisation, surface static pressure and force balance
measurements at streamwise chord Reynolds numbers from 1.5 x 105 to 5.2 x 106 and
Mach number from 0.09 to 0.17. These results are used to assess two viscous-inviscid
interaction CFD methods (BVGK and VFP) and two boundary layer methods (TAPERBL
and WAKELAG) used to predict the flow over the highly cambered wing.
A parametric study using cropped delta vane vortex generators in a co-rotating array was
conducted on the 40' swept wing to investigate the effect of vane chordwise position,
vane orientation, vane height relative to the boundary layer thickness and vane spacing on
the prevention of the trailing edge separation. The performance of these flow control
devices is assessed in terms of changes in; the wing surface flowfield, lift curve slope and
the lift-dependant drag factor. In addition comparisons are made between the clean wing
and flow control wing measured pressure distributions.
The results and analysis show that the performance of the vortex generators is improved
when the height of the vortex generator is approximately equal to that of the local
boundary layer thickness and when the vane angular deflection to the local upstream flow
direction is between 14' and 21'. The performance is also seen to depend on the vanes
position ahead of separation and on the adverse pressure gradient to be restored and may
also depend on a vane spacing made non-dimensional on the wing normal chord rather
than the vane height.
Similar performance improvements are observed with the wing swept to 50' using the
positioning guidelines from this optimisation study. The performance of concave slats,
canted cropped delta vanes, 'bent'wires and sub-boundary layer wires as vortex generating
devices are seen to be not as effective as upright cropped delta vane vortex generators.
To assist in the interpretation of the parametric vortex generator study a low speed wind
tunnel technique is developed using shear stress sensitive liquid crystals to investigate the
downstream development of vortices from cropped delta vane vortex generators. The
results show that --
i) submerged vortices have less influence on the surface flow with downstream
distance than vortices closer to the edge of the boundary layer, and
ii) the primary increase in skin ffiction arises in the flow adjacent to the upflow
side of the vortex. This area increases with vortex size.
The results from this research programme are finally shown to be applicable in two market
areas. The first is as a performance improvement on current highly swept winged military
aircraft and the second is as flight controls on future aircraft from making the vortex
generating devices active. The possible customers in these two areas are identified and
marketing strategies developed for each case.
45
Markiewicz, R. H. „A study of the aeroelastic behaviour of helicopter rotor blades featuring swept tips“. Thesis, Cranfield University, 1990. http://hdl.handle.net/1826/4188.
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The design of a model helicopter rotor blade incorporating a
swept tip is described. The swept tip is chosen to provide a coupling
between the lift at the tip and the blade twist thus achieving a
variation of blade tvist both with azimuth and with forward speed. The
design is the first stop in an investigation of aeroelastic tailoring
as a means of reducing helicopter vibration and increasing rotor
performance.
The first prototype blade encountered stability problems and
further designs were evaluated using a new modes/stability computer
program developed within RAE. Comparisons are made between the
stresses measured on a second stable swept tip blade and a dynamically
similar rectangular blade. The results show that a beneficial twisting
of the swept tip blade is achieved which enhances rotor perforawtnee
and reduces the flatvise bending and torsional moments. Comparisons
are made between the experimental results and those predicted by the
RAE/VOL coupled modes analysis. The predictions are. reasonably
accurate for the flatwise moments but a more representative model of
the hub is needed to produce acceptable predictions of the torsional
moments.
The design of a further set of blades is discussed, the aim
being to investigate the effects of introducing a strong coupling
between the flap and torsion notions of the blade by sweeping back the
shear centre. An analysis of-the results shows that there are large
gains in blade stability with no severe adverse effects on blade
loads.
A theoretical investigation has been undertaken to observe the
effects of tip sweep on the performance of a full size rotor. The
results show that aft tip sweep can reduce control loads and rotor
power for a rotor with a cambered aerofoil, section.
46
Al-Ahmed, S. M. „Integrating combat effectiveness disciplines into the aircraft conceptual/preliminary design phase“. Thesis, Cranfield University, 1996. http://hdl.handle.net/1826/3631.
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An assessment methodology has been developed for - use during the
conceptual/preliminary design phase to quantify the effectiveness of newly designed
aircraft. The effectiveness is measured by a squadron Sortie Generation Rate (SGR). Key
elements of this methodology were the establishment of link parameters between
design synthesis and the main effectiveness disciplines. These were Reliability and
Maintainability (R&M), Survivability / Vulnerability and Acquisition Cost. A
programmable solid modeller was used to create a solid CAD assembly of the aircraft
critical components. A ray tracing technique has been used to develop an interactive
vulnerability assessment tool. A Mission Simulation Model (MSM) has been developed
which typically simulates the operation of a squadron of aircraft and gives the operational
activities such as flying sorties and maintenance actions. The methodology has been
validated based on real data from recent conflicts. The application aspects of the
methodology have been demonstrated by quantifying the effectiveness of two recent
combat aircraft.
47
Macci, S. H. M. „Structural and mechanical feasibility study of a variable camber wing (VCW) for a transport aircraft“. Thesis, Cranfield University, 1992. http://hdl.handle.net/1826/4168.
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Aerodynamic investigations have shown' that variable camber wings (VCW) for
transport aircraft have considerable potential in terms of improving aircraft performance
and enhancing their operational flexibility. In order to justify these benefits it is
essential that the camber varying system is structurally and mechanically feasible.
This research examined the feasibility of providing variable camber to two supercritical
aerofoil sections of different'characteristics. The unique method of camber vaTiation
was applied by rotating the forward and aft regions of the aerofoil on a circular arc and
keeping the surface continuous and matching at their attachment to the main wing box.
The change in camber thus increased the chord due to translational motion of the
aforementioned regions.
The geometries required for varying the forward camber by this method presented
formidable design difficulties and no immediate solutions could be found. As a result,
an alternative geometry was devised which accepts camber by simply drooping the nose
region.
A novel idea was developed for aft camber variation, which is considered to be
universal for all supercritical aerofoil sections. The system utilises a tracking
mechanism which guides a trailing edge element on a continuous arc. Surface
continuity is provided by a flexible skin on the upper side and a spring loaded hinged
panel on the under side. The flexible skin remains attached to the trailing edge element
through a series of roller link arrangement which locate the skin in a separate guide
rail. The large moment arm and therefore the increased torsional loads created due to
the translational motion of the trailing edge element necessitated investigation of
alternative deployment geometries. As a result two additional geometries were
schemed. One had reduced radius of rotation and therefore reduced extension, while
the other changed camber by drooping the aft region without any chordal extension.
Since there was no aerodynamic evidence on the possible benefits offered by these
geometries it was decide to postpone them until such information was available.
Some detailed aspects of the proposed concept for aft camber variation were considered
by applying the system to a modem transport aircraft wing. This resulted in a design
which is practically feasible. Justification of this concept was made by designing and
testing a half scale structural model of one trailing edge segment. Three dimensional
(3-D) geometric investigation showed that the camber-varying elements ride on a frustum of a cone and therefore their deployment is skewed to the line of flight. The
3-D geometric implications of variable camber clearly suggested that the camber
variation by rotation on a circular arc, on a tapered wing can be possible if the rotating
element is made to flex and twist or it utilises a pin jointed arrangement. To provide
the necessary flexibility to the trailing edge element, its structural box best be made
from fibre reinforced plastic material. The deployment of the trailing edge element on
the structural m(; del was made possible by designing it in laminated wood.
Comparison of the proposed variable camber system with a conventional single slotted
flap arrangement suggests that the two systems could be equally complex but the
variable camber could be slightly heavier.,
Further systems investigations are required to quantify overall aerodynamic, mass, and
cost implications of the use of VCW on transport aircraft.
48
Gantois, K. „An MDO concept for large civil airliner wings“. Thesis, Online version, 1998. http://ethos.bl.uk/OrderDetails.do?did=1&uin=uk.bl.ethos.266980.
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49
Simpson, Mark N. „The application of semi-active control technology to aircraft landing gear“. Thesis, Loughborough University, 1988. https://dspace.lboro.ac.uk/2134/6764.
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The purpose of the research investigation was to study the application of semi-active control technology to the design of a suspension system to be used in a landing gear of a high speed military aircraft. A semi-active system was used because it will allow a system to be driven from the hydraulic systems already existing in the aircraft without extensive modification. The research work involved establishing a theoretical mathematical model for the semi-active damping system. This model involved a large number of non-linear dynamic phenomena and elements including a two-stage gas spring, lever geometry, break out friction, square law damping and the switching function needed to achieve the semi-active control. Validation of the model was carried out by means of an extensive study of the dynamic responses obtained from digital simulation. An extended programme of laboratory experiments was also carried out to confirm the theoretical and simulated results, and to demonstrate the potential benefits in performance which can be achieved with those obtained from standard and optimized passive suspension system. The experimental rig involved a physical model which used hydraulic elements of a general industry standard, but not specially approved for aircraft use. The apparatus was arranged to permit a considerable degree of freedom for implementing the control laws which facilitated the assessment of different control schemes and allowed, at the same time, the ready simulation of various passive damping arrangements. An extensive series of trials was carried out on the final design and involved frequency response tests and subjecting the experimental suspension to inputs obtained from a simulated runway profile. The profile simulation was a discrete representation of a particular runway chosen for its roughness which was characteristic of runways from which high-speed military aircraft operate. From the research investigation and these trials it was established that semi-active control of the damping function is superior to standard techniques and achieves a substantial reduction in the energy transmitted to the airframe during ground manoeuvres.
50
Guo, Shijun. „Improvement of analytical dynamic models using vibration test data“. Thesis, University of Hertfordshire, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.332846.
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Generally speaking, difficulties encountered during the improvement of an Analytical Dynamic Model (ADM) using vibration test data come from both the Spatial Coordinate Incompatibility (SCI) and especially the Modal Coordinate Incompatibility (MCI) between the ADM and the test data. Efforts were therefore made in this project to cope with these two problems by extending some of the existing methods and also by developing new methods with consideration of their feasibility, efficiency and accuracy. A general description of this part of the project and the literature survey of this study area are presented in part 1 of the thesis. In part 2, in order to solve the SCI problem, a new extended Complete Modal Expansion (CME) and a Branch Modal Expansion (BME) method were proposed especially for the case when using a branch mode method to produce the ADM. Application of these two methods and the existing physical expansion method were demonstrated in a beam example in this part and were also used in some of the examples later in this project. In part 3, efforts were made to extend the existing Direct Matrix Updating (DMU) and the Direct Parameter Identification (DPI) methods for solving the MCI problem using a direct approach. Firstly a new Direct Modal Extension (DME) method was proposed and compared with the DMU method when they were used to improve a reduced-size ADM. Secondly, in order to overcome the main limitation of the existing DPI methods in their practical use, an extended Corrected Modal Constraint (CMC) method was proposed. In part 4, in order to achieve the feasibility and accuracy of ADM improvement, efforts were then made in the study of the indirect approach. Firstly a procedure using a new Orthogonality Sensitivity Method (OSM) working together with a model reduction method was proposed. Secondly, a new Energy Error Estimation (EEE) method was also presented. The original contribution of the EEE method is that the poorly modelled stiffness and mass elements of an ADM can be identified and corrected accurately and effectively. Applications of these new proposed methods were demonstrated by taking beam examples. Further application of the EEE method was examined in a full-scale aircraft tail plane example. A general discussion, conclusion and recommendation for further study of these methods are presented in the fmal part 5. Based on the study of this project, it is concluded that the feasibility and accuracy of the direct methods described in part 3 of this thesis are at a low level for practical use. Therefore, the main efforts and contributions in this project were made in the study of the indirect methods described in part 4 of this thesis. It is concluded that both of the new proposed OSM and EEE methods provide feasible tools for ADM improvement and possess a high level of accuracy.