Examples of control problems occurring during flight tests of fighter aircraft are well documented. In many cases, the cause of the problem could be characterised as inadequate modelling or other inappropriate treatment of the aeroelastic effects on the vehicle dynamics and/or the flight-control design. However, such problems are not restricted to just aircraft. Especially long and slender bodies, such as (small) conventional launch systems, may suffer from an unwanted coupling between the rigid body and its flexible modes. In addition, due to fuel consumption during the flight, the change in atmospheric environ-ment, and aerodynamic effects, the entire flight profile should be examined, rather than a single worst-case point, to identify the stability and controllability performance of the launch vehicle. The current research treats the launch vehicle as a flexible beam with lumped masses to account for the subsystems and fuel, using a three-dimensional assumed-modes method with longitudinal and lateral effects. Sloshing is added to the system as a spring-mass-damper, where the liquid oxidiser and fuel are modelled separately. Multiple discrete flight points are considered for aeroelastic analysis at different stages of the ascent profile. These individual points are analysed as quasi-static and quasi-steady inputs for a continuous flight trajectory simulation. The stability and controllability for the entire flight under the influence of wind gust and turbulence is estimated based on this simulation. While simulating the manoeuvres with the sloshing model included it became apparent that the (very) simple PD controller was not able to stabilise the vehicle. Results concerning the influence of sloshing can therefore not be shown. The development of a more robust adaptive controller is currently on its way, and it is expected that this non-linear controller will be better suited to handle the strong non-linearities of the combination of flexibility, sloshing and a demanding wind environment.

Larger power requirements from modern spacecraft precipitate enormous solar arrays to fulfil budgetary needs. Although the solar arrays may have a relatively small influence on the total mass of the spacecraft system, large arrays with long flexible panels do affect the dynamics of the satellite bus through the Steiner terms in the moments of inertia. Unfortunately, such large arrays are broad targets for incoming orbital debris, and impacts are an increasing risk to attitude control and during guidance operations. This paper examines the effect of impacts on the response of a Rosetta-like spacecraft controlled by an adaptive controller. The control system design is based on simple adaptive control theory, and uses a stabilised, linear reference model to swiftly drive the plant output error to zero and hence achieving a commanded attitude. The adaptive controller is tuned using a linearised Euler-Bernoulli beam model as a computationally inexpensive exible system. The elastic dynamics increases the controller effort and shows the effect of the elastic bodies on the controller system by the deviations from an implementation for a rigid satellite only. A sample investigation of an impulsive force, simulating a particle debris impact, shows the control effort exerted to stabilise the spacecraft. Additional simulations covering a range of particle momenta and impact location on the spacecraft show controller efforts for a satellite in a steady-state configuration and undergoing a three-axis manoeuvre. For individual impact cases, the controller shows a robust performance, even if there is some initial saturation of the actuators. In multiple impact scenarios, the control effort is much heavier and strong oscillatory behaviour is observed, indicating the limits of the controller.

UAS will be integrated into the airspace in the near future, but the risk of UAS collision is not well understood which hampers the development of adequate regulations and standards. As risk has two constituents: frequency and consequence, collision risk analysis of UAS operations in future UTM asks for a quantitative assessment of various types of frequency and consequence. However, prior to studying such quantitative assessment, it is a prerequisite to identify the various types of collisions and consequences. Doing the latter is the objective of this paper. This paper follows a step-wise approach in identifying the various types of collision consequence under a given UTM ConOps, focusing on the very-low-level UAS operations. The first steps address the analysis of the UTM ConOps, rules, and infrastructure considered, and the identification of types of objects and UASs that will operate in the very-low-level UTM system. The follow-up steps are to characterize impact materials by applying zone of impact analysis, followed by analyzing the types of collision consequence. The result is a systematic identification and characterization of types of collision consequences as well as applicable impact materials and conditions that will form the basis for safety risk analysis in follow-on research.

Part of the work of AircraftFire, a project investigating the effects of fire and crash on aircraft survivability, is presented. This work compares the effect of changing the material model from metallic to composite on the impact damage and floor acceleration characteristics. First, the metallic two- and six-frame sections of an A320 are analysed, with drop test data to compare for reference and validation. The six-frame metallic and composite sections for a larger, A350-like aircraft are examined to compare the relative safety of newer composite fuselages. The composite model includes both a quasi-isotropic analysis with damage based on maximum allowable strain, and a ply-by-ply laminate model with Hashin damage. Energy dissipation and acceleration analyses follow, which show the potentially dangerous acceleration pulses for passengers seated in the cabin.