In this thesis, the achievable performance and limitations of a Synthetic Aperture Radar (SAR) seeker on an Anti-Ship Ballistic Missile (ASBM) for a set of parameters has been investigated. This was achieved by modelling the dynamics of an ASBM and the constraints related to the acquisition requirements of a maritime surface target in the "Generation and Improvement Algorithm for Nonlinear Trajectories" (GIANT) optimisation tool provided by TNO. The angular velocities of the ASBM were set as the control variables for the optimal control problem where the objectives were set to the minimisation and maximisation of the exposure time and vertical end velocity of the missile and the minimisation of the required time for creating one SAR image by the SAR seeker.

Three experiments were carried out to investigate the influences for different initial conditions and resolution constraints on the optimised trajectory. In the first experiment, the trajectory was optimised for initial reentry conditions that are representative for typical ASBMs. The target was placed at the location that would match the uncorrected ballistic flight impact location of the ASBM. This allowed to compare and visualise the optimised manoeuvres of the ASBM relative to its ballistic flight trajectory. In the second experiment, the target was placed at the same location, but the initial re-entry conditions were optimised by GIANT, that provided the optimal trajectory. In the last experiment, the constraints related to the resolution of the target were lowered to analyse the maximum performance of the system.

The experiments show that a larger initial squint angle of the re-entry position of the ASBM improves the overall performance. This is because a larger squint angle allowed the missile to obtain a larger seeker look angle during the SAR phase while performing minimal manoeuvres during the re-entry phase. This permitted the dwell time of the SAR seeker and the exposure time of the missile to be as short as possible, while the missile’s vertical end velocity could be maximised. It was also shown that a ground range and crossrange resolution of 1.60 m could be achieved for an average dwell time of 0.1170 s. For better resolutions, the bandwidth of the SAR seeker appeared to be the limiting factor of the system for the chosen set of parameters.

For decades the space domain has been used for military purposes. When enemy forces establish counter-space capabilities, this could become a weakness. Responsive Space is a concept that focuses on replenishing lost capabilities and/or granting additional capabilities from space assets quickly. The continuous miniaturization of technologies has enabled small satellites to fulfill many desired functions. To launch such satellites, it might be possible to modify a missile into a small launch vehicle that can be launched from a naval ship using a shipborne missile canister launching system, the Mark 41 Vertical Launching System, thereby giving the military autonomous and quick-response space launch capability. Modifying a missile that has already been integrated in the VLS instead of developing a fully new launch vehicle could have the benefit that integration cost could be lower. Also, a fully new vehicle would require more development, which can be costly and time consuming. It is investigated what the maximum payload is of a launch vehicle developed from an already integrated missile that is launched from the Mark 41 Vertical Launching System and whether it is capable of launching microsatellites. The Block IIA version of the Standard Missile-3 anti-ballistic missile is determined to be the most capable missile to modify. A multidisciplinary design optimization tool has been created that maximizes the payload mass and the perigee altitude of the orbit of the new launch vehicle. The results show that it is possible to launch a microsatellite to various Low-Earth orbits from different launch locations. For example, for an equatorial orbit, it is possible to launch a payload mass of up to 17.8 kg to a near-circular orbit with a perigee altitude of 200 km.

Current Ballistic Missile (BM) defence systems intercept an incoming threat in its terminal phase, and to some extend during its coast phase. During the boost phase of a BM, the initial velocity of the BM is low and the hot exhaust plume is easy to spot for infrared sensors, which provides a good opportunity for interception. To do so, the Network Centric Airborne Defense Element (NCADE) is a missile in development, which is to be launched from a 5$^{\text{th}}$ generation fighter aircraft. A tracking algorithm model has already been developed at TNO, which performs the tracking and trajectory prediction of the BM, based on the location measurements performed by the IR sensors of both the aircraft and the missile. Tracking a BM and predicting its trajectory poses a challenge, as the mass, thrust and acceleration cannot be measured directly, resulting in an uncertainty in the trajectory prediction. The NCADE therefore must be able to intercept the BM in its boost phase, considering the uncertainties in the trajectory prediction. The main research goal of this thesis is to develop a guidance algorithm for the NCADE, which plans a trajectory of an air launched missile, intercepting a BM in its boost phase. To do so, the NCADE has been modelled, and the tracking algorithm has been implemented. The actual guidance is performed using trajectory optimisation. 
The NCADE is designed to disable its target by means of kinetic penetration, meaning that there is no explosive warhead present. The missile is derived from the AIM120D AMRAAM, with a similar outer shape and suspension points. The NCADE consists of two stages, where the first stage is equipped with a solid booster for a fast acceleration. Control deflections are provided with aerodynamic surfaces. The second stage, also called Kill Vehicle (KV), is equipped with an IR sensor to determine the location of the target. Some sensor inaccuracy is present due to the amount of pixels used in the sensor. Control inputs on the second stage are performed using monopropellant pulses. Both control inputs and thrust of the KV use monopropellant from the same source, meaning that when the monopropellant tank is depleted, both control deflections and thrust cannot be delivered. To calculate different trajectories of the NCADE, the equations of motion are set up, where the NCADE is modelled as a 3 degrees of freedom point mass. Aerodynamic coefficients are obtained using software applying empirical methods, for which an extended database of projectiles is available. Verification of the equations is performed using a validated generic missile model, made by TNO using Simulink. 

The calculation of the guidance relies on a location of the target in the future. Therefore, a trajectory prediction must be performed, for which the states of the BM must be determined. There are however only position measurements of the BM available, from which a more extended set of states of the BM must be derived. This is performed using an Extended Kalman Filter (EKF), which was developed during an earlier study. The filter initiates with a guess of the states of the BM, and continuously updates those as new readings of the BM become available. Using the Kalman states, a trajectory prediction is performed. The Kalman states require a certain tracking period to converge to the correct values, to be able to calculate usable trajectory predictions. The quality of the tracking prediction is quantified by a score, which is forwarded to the guidance algorithm to be able to take the significant uncertainties in the trajectory prediction into account. The certainty score improves when the tracking duration increases and when the trajectory prediction is nearby in the future. The certainty results are forwarded to the guidance algorithm by means of coefficients of a polynomial.
Due to the complexity of the control system of the NCADE and the uncertain target trajectory, trajectory optimisation is applied in the guidance algorithm. Trajectory optimisation aims to decrease the defined performance index, which is in this case the divert cone minus the uncertainty ellipse of the trajectory prediction, to maximise the probability for interception. The divert cone of the NCADE is a volume which the missile is able to reach on a certain time, given its states and reserve fuel. The divert cone is calculated using a separate shooting optimisation algorithm, which maximises the distance in three ENU frame directions. To maximise the probability of interception, a shooting method is applied, which uses candidate solutions in the form of functions describing the control input, to calculate the performance index. Using constraints, the missile is directed towards an interception point. Constraints are also applied to bound the magnitude of the controls, and at the trajectory itself to remain physical feasible.
To investigate the behaviour of the guidance algorithm, simulations of interceptions of the NCADE have been performed on a modelled Scud BM, using a range of launch locations and tracking settings. The optimum results are presented in control deflection functions of the NCADE, which achieve the flight with the minimised performance index. When only the tracking uncertainty is minimised, flight time is minimised, as the prediction becomes less reliable when a longer trajectory prediction is performed. When only the size of the divert cone is to be maximised, the propellant of the sustainer, used for control and propulsion, is saved to increase the divert cone. Because of this, the altitude of the flight is increased, and $t_{f}$ must become larger, because less monopropellant is applied to increase the velocity at the beginning of the sustain phase. The optimum solution is a compromise between the divert cone size and time to flight. As the duration of the tracking time increases, the target trajectory prediction becomes more reliable, so the maximisation of the divert cone becomes more prominent. However, this results in the maximum range to decrease, since there is less time for interception, and the altitude of the target has increased. When the launch location is positioned further from the target, the reduction of time to flight becomes more prominent and the divert cone decreases. In conclusion, the optimisation routine performs the compromise between the amount of reserve propellant available, and the uncertainty of the trajectory prediction.