The anti-torque mechanism of a NOTAR helicopter is a complex system including vertical tail and pressurized tail boom which provides air ejection used for both circulation control around the boom and creating directed jet air at the end of the boom. This thesis targets the modeling of this mechanism and integrating it to a helicopter simulation model. Flight tests are performed on the MD 600N helicopter to verify the results. Finally, the simulation is compared with flight test data.

The deployment of unmanned aerial vehicles (UAV) in military applications increased the research about them and the importance of them. The unmanned helicopters are the most agile and maneuverable vehicles among the unmanned aerial vehicles (UAV). The ability of hovering and low speed cruise makes them even more attractive. Such abilities supply more areas to deploy the usage of the unmanned helicopters like search &amp / rescue, mapping, surveillance. Autonomy is the key property for these vehicles. In order to provide autonomy to an unmanned vehicle, the guidance and the autopilot units are designed in the first step. Waypoints are used to track the desired trajectories. The line of sight guidance is used to reach an active waypoint. In order to realize the guidance commands controllers are designed by using LQR. In addition, position and heading controllers are designed by root-locus method. The trimming and linearization are implemented in order to extract linear models used for controller design. Keywords: Helicopter, control, guidance

By applying pattern recognition analysis, the organised eddy structures in the near and intermediate fields (10-40 jet diameters) of a numerically simulated turbulent plane jet have been investigated. The Reynolds number based on the jet width is 1, 600. The results of the nearly full simulation show good agreement with the experimental measurements in the near and intermediate regions (< 40 diameters) of the jet. Patterns of organised eddy structures have been extracted from the three-dimensional simulation databases. Ensemble averaged fields of three velocity components and pressure in the spanwise and transverse planes suggest the existence of inclined vortex pairs. These highly three-dimensional vortical structures are confined to one side of the jet centreplane and approximately align with the principal axis of rate-of-strain. Some of the topological features of these averaged eddy structures have been examined.

In this thesis, the effects of split canard on the aerodynamic characteristics of missiles at high angles of attack are numerically investigated. Moreover, an enhanced semi-empirical engineering-level method is developed for prediction of normal force and pitching moment of split canard mounted missiles. In order to analyze the effects of split canard, a generic test case model is created by mounting a split canard to a generic test case model, NASA Dual Control Missile (NDCM), which was previously modeled and analyzed for the validation of CFD modeling. After obtaining a generic missile model with split canard, the effects of split canard on the aerodynamic characteristics of this missile in case of no control, pitch control, yaw control, and roll control deflections are numerically investigated. It is seen that the split canard decreases the local angle of attack of existing canard, increases the normal force and the maneuverability of the missile, and reduces the induced rolling moment at high angles of attack. Five different aerodynamic design parameters are determined for split canard and the effects of each parameter on missile aerodynamics are numerically investigated. It is seen that the roll orientation, deflection angle, size of the split canards have strong effects on missile&rsquo / s aerodynamic performance whereas longitudinal position of the split canards only affects the pitching moment of the missile. Finally, an enhanced semi-empirical engineering-level method, CFD-CBU, is developed for split canard mounted missiles in order to predict the normal force and the pitching moment coefficients. The developed method is validated with NDCM test case model. After this validation, the method is applied to the split canard mounted generic missile in case of no control deflection and pitch control deflection. The results of this method are compared with CFD results and it is seen that the results are in good agreement with each other.

The aim of this thesis is to investigate the general design and analysis principles of sabots and to develop a sabot design tool. Structures which support and align the models in gun bore, and separate without disturbing the flight path of models are called sabots. In the scope of this study, structurally critical regions and loads acting on sabots due to acceleration in the gun are determined. To calculate the loads acting and to size the sabots, approximate relations are derived by the help of strength of materials approach and finite element solutions. Conventional sabots are investigated and new sabot geometries are designed to resist high accelerations. To achieve the desired test velocity without affecting the stability of the model is the main objective. Sabots should be as light as possible, to reach the desired velocity with minimum inertial load, in other words minimum gun chamber pressure. To obtain the less weight sabot geometry with enough strength to resist the loads acting, a computer tool is developed. Structural analyses are automatically performed by the help of the sabot design tool. The advantage of the design tool is to reduce the design engineer&amp / #8217 / s work time spent for routine analyses processes. The output of the tool, which is sabot geometry, should be evaluated as a result of preliminary design process, and can be used as an input for detailed design process. Detailed geometric modifications required for production can be applied on the tool output, and final product can be manufactured reliably and in the shortest possible time.

Designing a suitable utility helicopter cockpit for Turkish pilots is the main theme of this thesis. Helicopter cockpit is one of the ultimate human machine interface application. Consequences of pilot errors during flight in any helicopter cockpit can be catastrophic. Human errors can only be prevented by user-friendly cockpit design. In this thesis, reach compatibilities to controls in the cockpit are evaluated and the suitable positions of analogue indicators at front display panel are examined in order to obtain a user-friendly utility helicopter cockpit design. Human anthropometry is the most significant factor for evaluating cockpit reach compatibilities to controls / so all critical operational reach parameters of Turkish pilots are examined. The anthropometric study revealed vision problems and showed that the height of display panel is inappropriate for most pilots. Suitable positions of the indicators on pedestal are determined by using qualitative and quantitative approaches. As a quantitative approach Multi Criteria Decision Making (MCDM) algorithms are employed. Card sorting methodology is used for the qualitative evaluation of the aforementioned display panel design. Although there are some approaches in literature for designing of displays, a specific design methodology related with the arrangement of indicators on display panel is not offered so far. In this thesis, MCDM and Card sorting approaches are adapted and used in the design of a display panel for the first time. There are lots of similarities between the results of MCDM and Card sorting approaches. The main similarity is to provide separate locations on display panel for engine and flight system indicators. Finally the findings of these techniques are compared with the existing layout of the display panel of a utility helicopter.

In this thesis, design and analyses of honeycomb structures are investigated. Primary goal is to develop an equivalent orthotropic material model that is a good substitute for the actual honeycomb core. By replacing the actual honeycomb structure with the orthotropic model, during the finite element analyses, substantial advantages can be obtained with regard to ease of modeling and model modification, solution time and hardware resources . To figure out the best equivalent model among the approximate analytical models that can be found in the literature, a comparison is made. First sandwich beams with four different honeycomb cores are modeled in detail and these are accepted as reference models. Then a set of equivalent models with the same dimensions is generated. The material properties of the equivalent models are taken from different studies performed in the literature. Both models are analyzed under the same loading and the boundary conditions. In finite element analyses, ANSYS finite element program is used. The results are compared to find out the best performing equivalent model. After three major analyses loops, decision on the equivalent model is made. The differences between the total reaction forces calculated by the equivalent model and the actual honeycomb model are all found to be within 10%. The equivalent model gives stress results at the macro-scale, and the local stresses and the strains can not be determined. Therefore it is deemed that for stress analysis, equivalent model can be used during the preliminary design phase. However, the equivalent model can be used reliably for deflection analysis, modal analysis, stiffness determination and aero-elastic analysis.

Stacking sequence design is a combinatorial problem when the fiber orientations in each layer are restricted to certain angles. In addition, there often exist many optimal or nearoptimal designs for the stacking sequence of a composite pressure vessel under different loading conditions. Genetic algorithms are quite well suited for finding the optimal designs for such a combinatorial problem. In this thesis, a genetic algorithm code is developed in Matlab, optimizing the stacking sequence of a composite pressure vessel subjected to internal and external pressures, axial load and body force due to rotation. For testing of the code and identification of the effects of optimization parameters, a problem, whose optimum solution is obvious, is defined and the optimum design is tried to be found by using the developed code. The results have shown that the code was quite successful in finding the best design. Afterwards, the code is used for the optimization of the stacking sequence of a composite pressure vessel under different loading conditions. Again the code has proven its reliability in finding the optimal designs. The developed genetic algorithm optimization code also has the infrastructure to be easily adapted to the solution of different combinatorial problems.

This study, explores the use of an iterative solution approach for the linearized Euler equations formulated in the frequency domain for fan tone noise propagation and radiation through bypass jets. The aim is to be able to simulate high frequency propagation and radiation phenomena with this code, without excessive computational resources. All computations are performed in parallel using MPI library routines on a computer cluster. The linearized Euler equations support the Kelvin-Helmholtz type convective physical instabilities in jet shear flows. If these equations are solved directly in frequency domain, the unstable modes may be filtered out for the frequencies of interest. However, direct solutions are memory intensive and the reachable frequency is limited. Results provided shown that iterative solution of LEE is more efficient when considered memory requirement and might solve a wider scope of frequencies, if the instabilities are controlled.

In this thesis aeroelastic analysis of a typical Unmanned Aerial Vehicle (UAV) using MSC&reg / FlightLoads and Dynamics module and MSC&reg / NASTRAN Aero 1 solver was performed. The analyses were carried out at sea level, 1000m, 2000m and 4000m altitudes for Mach Numbers M=0.2, 0.4 and 0.6 for the full model of the UAV. The flutter characteristics of the UAV for different flight conditions were obtained and presented. The effect of altitude on flutter characteristics has been examined and compared with the theoretical and experimental trends in the literature. Also the divergence characteristics of the full model UAV was obtained. In the study, some verification and test cases are also included. The results of the analyses of an untapered swept-wing and AGARD 445.6 wing models were compared with wind tunnel data and a maximum error of 1.3 % in the flutter speed prediction was obtained. In two different wing models the effect of taper was investigated.