Varying missile configurations may create uncertainty for a missile control algorithm developed with linear control theory, for instance the control system performance requirements may not be satisfied anymore. Missile configuration may change during the missile design period due to variations in subsystem locations, subsystem weights and missile geometry. Likewise, burning propellant, deployment of aerodynamic surfaces and wings with varying sweep angle can be considered as in-flight missile configuration changes. This thesis study addresses development and analysis of an adaptive missile control algorithm to account for the uncertain effects caused by varying missile configuration. Control algorithms, designed using pole placement, are augmented with adaptive neural networks. The resulting controller is a type of model reference adaptive controller. Adaptation characteristics of the augmented control algorithms are investigated to changing center of pressure location and missile geometry. Analyses are performed for three different missile configurations using simulation.

A parametric study on Secondary Injection Thrust Vector Control (SITVC) has been accomplished numerically with the help of a commercial Computational Fluid Dynamics (CFD) code called FLUENT&reg / . This study consists of two parts / the first part includes the simulation of three dimensional flowfield inside a test case nozzle for the selection of parameters associated with both computational grid and the CFD solver such as mesh size, turbulence model accompanied with two different wall treatment approaches, and solver type. This part revealed that simulation of internal flowfield by a segregated solver with Realizable k-&amp / #949 / (Rke) turbulence model accompanied by enhanced wall treatment approach is accurate enough to resolve this kind of complex three dimensional fluid flow problems. In the second part a typical rocket nozzle with conical diverging section is picked for the parametric study on injection mass flow rate, injection location and injection angle. A test matrix is constructed / several numerical simulations are run to yield the assessment of performance of SITVC system. The results stated that for a nozzle with a small divergence angle, downstream injections with distances of 2.5-3.5 throat diameters from the nozzle throat lead to higher efficiencies over a certain range of total pressure ratios, i.e., mass flow rate ratios, upstream injections should be aligned more to the nozzle axis, i.e., higher injection angles, to prevent reflection of shock waves from the opposite wall and thus low efficiencies. Injection locations that are too much downstream may result reversed flows on nozzle exit.

Multidisciplinary design optimization of aerospace vehicles has attracted interest of many researchers. Well known aerospace companies are developing tools for the mutlidisciplinary design optimization. However, the multiobjective optimization of the design is a new and important area investigated very little by the researchers. This thesis will examine the approaches to the multiobjective and mutlidisciplinary design optimization of rockets and missiles. In the study, multiobjective optimization method called MC-MOSA will be used.

In this study, a detailed service life assessment procedure for solid propellant rockets under random environmental temperature and transportation loads is introduced. During storage and deployment of rocket motors, uncontrolled thermal environments and random vibratory loads due to transportation induce random stresses and strains in the propellant which provoke mechanical damage. In addition, structural capability degrades due to environmental conditions and induced stresses and strains as well as material capability parameters have inherent uncertainties. In this proposed probabilistic service life prediction, uncertainties along with degradation mechanisms are taken into consideration. Vibration loads are accounted by utilizing acceleration spectral density values which are induced during various deployment scenarios of ground, air and sea transportation. Furthermore, thermal loads are represented with a mathematical model being a harmonic function of time. Throughout the finite element analyses, a linear viscoelastic material model is to be used for the propellant. Change in the structural capability of the propellant with time is calculated using Laheru&#039 / s cumulative damage model. Moreover, to include aging effect of the propellant, Layton model is used. To determine the effects of induced stress and strains under variations and uncertainties in the random loads and material constants, mathematical surrogate models are constructed using response surface method. Limit state functions are utilized to predict failure modes of the solid rocket motor. First order reliability method is used to calculate reliability and probability of failure of the propellant grain. With the proposed methodology, instantaneous reliability of the propellant grain is determined within a confidence interval.

Solid propellant rocket motors are the most widely used propulsion systems for military applications that require high thrust to weight ratio for relatively short time intervals. Very wide range of magnitude and duration of the thrust can be obtained from solid propellant rocket motors by making some small changes at the design of the rocket motor. The most effective of these design criteria is the geometry of the solid propellant grain. So the most important step in designing the solid propellant rocket motor is determination of the geometry of the solid propellant grain. The performance prediction of the solid rocket motor can be achieved easily if the burnback steps of the rocket motor are known. In this study, grain burnback analysis for some 3-D grain geometries is investigated. The method used is solid modeling of the propellant grain for some predefined intervals of burnback. In this method, the initial grain geometry is modeled parametrically using commercial software. For every burn step, the parameters are adapted. So the new grain geometry for every burnback step is modeled. By analyzing these geometries, burn area change of the grain geometry is obtained. Using this data and internal ballistics parameters, the performance of the solid propellant rocket motor is achieved. To verify the outputs obtained from this study, rocket motor tests are performed. The results obtained from this study shows that, the procedure that was developed, can be successfully used for the preliminary design of a solid propellant rocket motor where a lot of different geometries are examined.

In this thesis the characteristics of pitch and roll damping coefficients of missile models are investigated by using Computational Fluid Dynamics (CFD) techniques. Experimental data of NACA0012 airfoil, Basic Finner (BF) and Modified Basic Finner (MBF) models are used for validation and verification studies. Numerical computations are performed from subsonic to supersonic flow regimes. Grid refinement and turbulence model selection studies are conducted before starting the dynamic motion simulations. Numerical method of dynamic motion simulation is validated with a 2D NACA0012 airfoil. After the validation of numerical method, forced-oscillation motion is given to the BF and MBF models. In order to get deeper understandings about the characteristics of dynamic pitching and rolling motions, parametric studies are performed. The amplitude and frequency of forced-oscillation motions are investigated one by one. The effects of angle of attacks are also investigated for both pitching and rolling motions. The results of CFD simulations are compared with experimental data obtained from different wind tunnel and free flight tests. It is seen from these comparisons that experimental and numerical results are in good agreement throughout the whole flow regime. In conclusion, the numerical method presented in this study is validated and can be used for the prediction of pitch and roll damping coefficient of any missile configurations.

Reliability is one of the most critical performance measures of guided missile systems. It is directly related to missile mission success. In order to have a high reliability value, reliability analysis should be carried out at all phases of the system design. Carrying out reliability analysis at all the phases of system design helps the designer to make reliability related design decisions in time and update the system design. In this study, reliability analysis process performed during the conceptual design phase of a Medium Range Anti-Tank Missile System Inertial Measurement Unit (IMU) was introduced. From the reliability requirement desired for the system, an expected IMU reliability value was derived by using reliability allocation methods. Then, reliability prediction for the IMU was calculated by using Relex Software. After that, allocated and predicted reliability values of the IMU were compared. It was seen that the predicted reliability value of the IMU did not meet the required reliability value. Therefore, reliability improvement analysis was carried out.

A three-dimensional, parallel, finite volume solver which uses Roe&#039 / s upwind flux differencing scheme for spatial and Runge-Kutta explicit multistage time stepping scheme for temporal discretization on unstructured meshes is developed for the unsteady solution of external viscous flow around rotating bodies. The main aim of this study is to evaluate the aerodynamic dynamic stability derivative coefficients for rotating missile configurations. Arbitrary Lagrangian Eulerian (ALE) formulation is adapted to the solver for the simulation of the rotation of the body. Eigenvalues of the Euler equations in ALE form has been derived. Body rotation is simply performed by rotating the entire computational domain including the body of the projectile by means of rotation matrices. Spalart-Allmaras one-euqation turbulence model is implemented to the solver. The solver developed is first verified in 3-D for inviscid flow over two missile configurations. Then inviscid flow over a rotating missile is tested. Viscous flux computation algorithms and Spalarat-Allmaras turbulence model implementation are validated in 2-D by performing calculations for viscous flow over flat plate, NACA0012 airfoil and NLR 7301 airfoil with trailing edge flap. The ALE formulation is validated in 2-D on a rapidly pitching NACA0012 airfoil. Afterwards three-dimensional validation studies for viscous, laminar and turbulent flow calculations are performed on 3-D flat plate problem. At last, as a validation test case, unsteady laminar and turbulent viscous flow calculations over a spinning M910 projectile configuration are performed. Results are qualitatively in agreement with the analytical solutions, experimental measurements and previous studies for steady and unsteady flow calculations.

Controlling unsteady combustion induced gas flow fluctuations and the resultant motor vibrations is a very significant step in rocket motor design. It occurs when the unsteady heat release due to combustion happens to feed the acoustic oscillations of the closed duct forming a feed-back system. The resultant vibrations concerned may even lead to total failure of the rocket system unless analysed and tested thoroughly. This thesis aims developing a linear numerical analysis method for the growth rate of instabilities and possible mode shape of a liquid-propelled chamber geometry. In particular, A 3-D Helmholtz code, utilizing Culicks spatial averaging linear iterative method, is developed to find the form of deformed mode shapes iteratively to obtain possible effects of heat source and impedance boundary conditions. The natural mode shape phase is solved through finite volume discretization and the open-source eigenvalue extractor, ARPACK, and its parallel implementation PARPACK. The iterative method is particularly used for analyzing the geometries with complex shapes and essentially for disturbances of small magnitudes to natural mode shapes. The developed tools are tested via two simple cases, a duct with inactive flame and a Rijke tube, used as validation cases for the code particularly with only boundary contribution and heat contribution respectively. A sample 2-D and 3-D liquid-propelled combustion chamber is also analysed with heat sources. After comparing with the expected values, it is eventually proved that the method should be only used for determining the modes instability analysis, as to whether it keeps vibrating or decays. The methodology described can be used as a preliminary design tool for the design of liquid-propellant rocket engine combustors, rapidly revealing only the onset of instabilities.

The recent interests in the necessity of high maneuverability and vertical launching triggered namely the unconventional control design techniques that are effective at high angle of attack flight regimes. For most of missile configurations, this interest required thrust vector control together with conventional aerodynamic control. In this study, nonlinear modeling and dynamical analysis of a surface to air missile with both aerodynamic and thrust vector control is investigated. Aerodynamic force and moment modeling of the presented missile includes the challenging high angle of attack aerodynamics behavior and the so called hybrid control, which utilizes both tail fins and jet vanes as control surfaces. Thrust vector and aerodynamic control effectiveness is examined during flight envelope. Different autopilot designs are accomplished with hybrid control. Midcourse and terminal guidance algorithms are implemented and performed on target sets including maneuverable targets. A different initial turnover strategy is suggested and compared with standard skid-to-turn maneuver. Comparisons of initial roll with aerodynamic and thrust vector control are examined. Afterwards, some critical maneuvers and hybrid control ratio is studied with a real coded genetic algorithm. Rapid turnover for low altitude targets, intercept maneuver analysis with hybrid control ratio and lastly, engagement initiation maneuver optimization is fulfilled.