Active control of single-frequency interior noise in a realistic composite aircraft fuselage is experimentally studied. The control inputs are due to piezoelectric actuators bonded to the cylinder wall while error information from the interior acoustic field is sensed by microphones. A preliminary analytical development was conducted to investigate the mechanisms of the structural/acoustic coupling exhibited by a simple cylindrical model in order to gain more inSight into the coupling properties of the piezoelectric actuators with the cylinder. Therefore, the response of a homogeneous, simply-supported cylindrical shell to the excitation of piezoelectric actuators was presented. The analytical results show that a piezoelectric actuator pair excited in-phase (stretching model) created a lower order circumferential interior acoustic field, more suitable to control the interior noise at low frequency than the same piezoelectric actuator pair excited in out-of-phase (bending model). The experiments were performed in the large anechoic chamber of the Acoustics Research Laboratory at NASA Langley Research Center at Hampton, Virginia and utilized a 1.68 m diameter, 3.66 m length composite aircraft structure, equipped with stringers, ring frames and a cabin floor. The narrowband controller used in these experiments was a four channel adaptive LMS algorithm implemented on a TMS320C25 system board. Results showed that global reduc1:ion of the interior sound pressure level of the order of 12 dB could be obtained using piezoelectric actuators. The influence of the sensor/actuator location and configuration as well as the frequency of excitation was studied. In general this investigation validates active control using piezoelectric actuators bonded to the fuselage to reduce the interior noise inside realistic aircraft structures. / Master of Science

This diploma thesis deals with the design of all-metal construction of fuselage of the replica L-40 Meta Sokol aircraft on the basis of CS-VLA regulation. As far as the aircraft is concerned, the thesis establishes the load of tail surfaces, landing gear and engine bed. Futhermore, it designs the construction of fuselage, tail landing gear, including the system of its retraction, wing-fuselage connection, engine bed and the primary design of cockpit. In the end, the thesis focuses on the fuselage production technology.

This work's objective is to create a type design of a new commuter class airplane according to CS-23 certification specifications. The design will take under account the EV-55 design to reduce the number of new parts and subassemblies. Part of this thesis are mass balance of the new design, basic airplane dimensions and fuselage dimensions estimates. The critical crossection of the fuselage is analyzed (stress analysis) and the system drawings is worked out.

This master's thesis deals with design, strength calculation and technological solution of wing-fuselage joint for composite four-seater aeroplane certificated by CS-23. Design is focused on optimal space utilization, low weight and simple manufacturing and assembly. Main output is technical documentation consisting of drawings and technological lay-ups. Conceptual design was chosen after analysis and weight estimation. With strenght calculation for composite materials it was possible to design composition of primary parts and also specify manufacture processes. Assigment was solved by unusuall conceptual design, which claims higher precision manufacturing, but it saves weight.

Le SA-160 est un avion monomoteur biplace dont le fuselage est constitué d’un châssis en tubulure d’acier recouvert d’une peau en composites non structurale. Quoi que robuste et fiable, cet assemblage est coûteux à fabriquer et pèse près de 180 lbs (80 kg). Aviatech Services Techniques (AST) désire réduire le poids de la structure de l’appareil ainsi que son coût de fabrication en retirant l’arrière du châssis d’acier pesant à lui seul près de 40 lbs. La conception de la peau en composites a donc été revue dans le cadre de cette recherche, afin qu’elle puisse assurer un rôle structural pouvant compenser le retrait de l’arrière de la structure d’acier. Mis à part les modifications apportées au châssis, les nouvelles caractéristiques géométriques du fuselage sont les suivantes : - Ajout d’un plancher permettant de déposer et fixer les bagages; - Intégration du stabilisateur vertical au fuselage; - Ajout de cloisons internes augmentant la raideur globale du fuselage; - Points de fixation des mécanismes et surfaces de contrôle préservés. La fabrication de la peau de ce nouveau fuselage se fera par infusion au moyen de renforts en fibres de carbone et de verre ainsi qu’une résine d’époxyde pour l’infusion. Les propriétés structurales de ces matériaux ont été caractérisées et ont servi à faire la conception du fuselage au moyen de simulations par éléments finis. Selon ces simulations, le concept final peut résister aux cas de chargement statiques ultimes prescrits par Transport Canada, sans rupture et sans subir de flambage, et ce avec une rigidité similaire à celle du précédent fuselage. Le poids projeté du concept final est 151 lbs ce qui constitue une réduction de 26 lbs en incluant l’économie engendrée par l’intégration du stabilisateur vertical. La nouvelle structure permet donc de réaliser un gain intéressant avec une réduction de poids de 15% par rapport à la structure originale. Cette économie de poids permettra l’augmentation de la charge utile comme de l’essence, des bagages ou d’autres équipements. Une économie substantielle devrait être réalisée lors de la fabrication du châssis puisque le nombre de pièces soudées sera grandement réduit. La fabrication de la peau pourra être faite dans les anciens moules réutilisés mais nécessitera l’utilisation de matériaux avancés plus onéreux tels la fibre de carbone. De plus, la résine thermodurcissable choisie requiert une cuisson de 16 heures à 80°C ce qui augmentera la cadence de production par rapport à la cuisson à température ambiante. Ce procédé nécessitera toutefois l’acquisition d’un four de dimensions imposantes pour réaliser cette cuisson. Ce mémoire présente la démarche ayant mené à l’obtention du nouveau concept de fuselage pour le SA-160. Les résultats de cette recherche permettront à Aviatech d’effectuer a fabrication et la certification de cette nouvelle structure en vue de sa mise en production.

A influência da fuselagem sobre a posição do centro aerodinâmico da asa é complexa e deve ser considerada nos cálculos de equilíbrio e estabilidade estática longitudinal da aeronave. Este trabalho apresenta uma análise comparativa para indicar o mais preciso dentre sete métodos teóricos para prever esta influência, em condições de baixa velocidade, utilizando seis configurações de modelos de asa mais fuselagem em escala reduzida, com proporções dimensionais características da aviação leve. Mediram-se os coeficientes de momento e sustentação para cada configuração, através de ensaios em túnel de vento de baixa velocidade, circuito aberto e seção de testes fechada. Calcularam-se as posições experimentais do centro aerodinâmico através da distância do eixo de rotação da balança ao bordo de ataque da asa e derivadas do coeficiente de momento em relação ao coeficiente de sustentação. Aplicaram-se os métodos teóricos às configurações. Os resultados demonstram que a maioria dos métodos prevê comportamentos na variação da posição do centro aerodinâmico semelhantes aos obtidos experimentalmente e apontados na revisão da literatura. A análise dos resultados teóricos ante os experimentais aponta o método descrito em Engineering Sciences Data Unit (1996a) como o mais preciso. / The fuselage influence on the wing aerodynamic center is complex and must be considered within longitudinal static stability and equilibrium calculations of the airplane. This work presents a comparative analysis to indicate the most accurate between seven theoretical methods that predict this influence, at low speed conditions, using six configurations of wing-fuselage reduced scale models, with the dimensional proportions found in light aviation. The moment and lift coefficients have been measured by experiments in a low speed open circuit wind tunnel with a closed test section. The experimental aerodynamic center positions have been found by the distance of the balance trunnion to wing leading edge and the derivation of the moment coefficient relative to the lift coefficient. The theoretical methods have been applied to all configurations. The results show that most of the methods predict variations in aerodynamic center position in the same way as those obtained in experimental results and shown in the literature review. The analysis between theoretical and experimental results indicates the method from Engineering Sciences Data Unit (1996a) as the most accurate.

Flush repairs to military aircraft are expected to become more prevalent as more thick skin composites are used, particularly on the surface of the fuselage, wings and other external surfaces. The use of these repairs, whilst difficult to manufacture provide an aerodynamic, ???stealthy??? finish that is also more structurally efficient than overlap repairs. This research was undertaken to improve the design methodology of scarf repairs with reduced material removal and to investigate the damage tolerance of scarf repair to low velocity impact damage. Scarf repairs involve shallow bevel angles to ensure the shear stress in the adhesive does not exceed allowable strength. This is important when repairing structures that need to withstand hot and humid conditions, when the adhesive properties degrade. Therefore, considerable amounts of parent material must be machined away prior to repair. The tips of the repair patch and the parent laminate are very sharp, thus a scarf repair is susceptible to accidental damage. The original contributions include: ??? Developed analytic means of predicting the stresses within optimised scarf joints with dissimilar materials. New equations were developed and solved using numerical algorithms. ??? Verified using finite element modelling that a scarfed insert with dissimilar modulus subjected to uniaxial loading attracted the same amount of load as an insert without a scarf. As such, the simple analytic formula used to predict load attraction/diversion through a plate with an insert may be used to predict the load attraction/diversion into a scarf repair that contains a dissimilar adherend patch. ??? Developed a more efficient flush joint with a doubler insert placed near the mid line of the parent structure material. This joint configuration has a lower load eccentricity than external doubler joint. ??? Investigated the damage tolerance of scarf joints, with and without the external doubler. The results showed that scarf joints without external doublers exhibited a considerable strength reduction following low velocity impact. Based on the observations, the major damage mechanics in the scarf joint region following impact have been identified. These results demonstrated that it is important to incorporate damage tolerance in the design of scarf repairs.

<p><p>This paper is the result of a master thesis project on helicopter rotor downwash effects using computational fluid dynamics (CFD). The work was performed at the department of Aerodynamics and Flight Mechanics at Saab AB, Linköping in 2008. It completes the author’s studies for a M.Sc degree in Applied Physics and Electrical Engineering at the Department of Electrical Engineering at the Linköping institute of technology (LiTH), Linköping, Sweden.</p><p> </p><p>The aim of the project was to study the rotor downwash effects and its influence on the helicopter fuselage. To fulfil this purpose, several CFD calculations were carried out and the aerodynamic forces and moments resulting from the calculations were implemented in an existing simulation model, developed in-house at Saab. The original (existing) model was compared to the updated model by studying step responses in MATLAB, Simulink. For some step commands, the comparisions indicated that the updated model was more damped in yaw compared to the original model for the hovering helicopter. When the helicopter was trimmed for a steady turn, the states in the updated model diverged much faster than the states in the original model for any given step command.</p><p> </p><p> </p><p>In order to investigate the differences between the original helicopter model and the updated model from a controlling perspective, a linear quadratic (LQ) state feedback controller was synthesized to stabilize the vehicle in a steady turn. The LQ method was chosen as it is a modern design technique with good robustness and sensitivity properties and since it is easily implemented in MATLAB.  Before synthesising, a simplification of the helicopter model was made by reducing states and splitting them into lateral and longitudinal ones. Step responses from simulations with the original and the updated model were studied, showing an almost identical behavior.</p><p> </p><p>It can be concluded that the aerodynamic coefficients obtained from the CFD calculations can be used for determining the aerodynamic characteristics of the helicopter. Some further validation is needed though, for example by comparing the results with flight test data. In order to build an aerodynamic data base that covers the whole flight envelop, additional CFD calculations are required.</p><p> </p></p>

Design of commercial aircraft structure, composed of composite material, requires the prediction of failure loads given large scale damage. In particular, a fuselage of graphite/epoxy lamination was analyzed for damage tolerance given a standard large crack that severed both skin and internal structure. Upon loading, a zone of damage is known to develop in front of a crack-tip in composite laminates; and, its material behavior within the damage zone is characterized as strain softening. This investigation sought to develop a computational model that simulates progressive damage growth and predicts failure of complex laminated shell structures subject to combined tensile and flexural load conditions. This was accomplished by assuming a macroscopic definition of orthotropic damage that is allowed to vary linearly through the shell thickness. It was further proposed that nonlocal plate strain and curvature act to force damage growth according to a set of uniaxial criteria. Damage induced strain softening is exhibited by degradation of laminate stiffness. An expression for the damage reduced laminated plate stiffness was derived which assumed the familiar laminated plate [AM] stiffness matrix format. The model was implemented in a finite element shell program for simulation of fracture and evaluation of damage tolerance. Laminates were characterized for damage resistance according to material parameters defining nonlocal strain and the damage growth criteria. These parameters were selected using an inverse method to correlate simulation with uniaxial strength and fracture test results. A novel combined tension-plus-flexure fracture test was developed to facilitate this effort. Analysis was performed on a section of pressurized composite fuselage containing a large crack. Good agreement was found between calculations and test results. / Graduation date: 1998

This paper is the result of a master thesis project on helicopter rotor downwash effects using computational fluid dynamics (CFD). The work was performed at the department of Aerodynamics and Flight Mechanics at Saab AB, Linköping in 2008. It completes the author’s studies for a M.Sc degree in Applied Physics and Electrical Engineering at the Department of Electrical Engineering at the Linköping institute of technology (LiTH), Linköping, Sweden.   The aim of the project was to study the rotor downwash effects and its influence on the helicopter fuselage. To fulfil this purpose, several CFD calculations were carried out and the aerodynamic forces and moments resulting from the calculations were implemented in an existing simulation model, developed in-house at Saab. The original (existing) model was compared to the updated model by studying step responses in MATLAB, Simulink. For some step commands, the comparisions indicated that the updated model was more damped in yaw compared to the original model for the hovering helicopter. When the helicopter was trimmed for a steady turn, the states in the updated model diverged much faster than the states in the original model for any given step command.     In order to investigate the differences between the original helicopter model and the updated model from a controlling perspective, a linear quadratic (LQ) state feedback controller was synthesized to stabilize the vehicle in a steady turn. The LQ method was chosen as it is a modern design technique with good robustness and sensitivity properties and since it is easily implemented in MATLAB.  Before synthesising, a simplification of the helicopter model was made by reducing states and splitting them into lateral and longitudinal ones. Step responses from simulations with the original and the updated model were studied, showing an almost identical behavior.   It can be concluded that the aerodynamic coefficients obtained from the CFD calculations can be used for determining the aerodynamic characteristics of the helicopter. Some further validation is needed though, for example by comparing the results with flight test data. In order to build an aerodynamic data base that covers the whole flight envelop, additional CFD calculations are required.