The design of a thrust chamber for a rocket propulsion system is a challenging task. The thrust chamber has to be designed for minimum structural weight with an adequate factor of safety. This calls for a thorough knowledge of various structural loads, both internal and external, and the behaviour of the structure in flight. This paper presents the design and structural analysis of a pressure vessel used as thrust chamber for a rocket propulsion unit. The effects of kinetic heating, thermal stress, spinning and various aerodynamic loads and their mutual interactions are accounted for in this analysis. Based on a detailed stress analysis of the components and a modal/structural dynamic analysis of the flight vehicle as a whole, in which the thrust chamber is the main load bearing member, the required mechanical properties of the thrust chamber are obtained.

This paper develops a three-dimensional guidance and control algorithm to ensure that a manoeuverable target is preserved by a quadrotor in a long-term tracking scenario. The proposed guidance approach determines the desired altitude of the quadrotor to adjust the field of view (FOV) to the union of two desired trusted and critical regions. The dimensions of the desired trusted region depend on the controller performance that is evaluated by the distance of the target from the center of the FOV. The critical region is a predefined margin around the trusted region that is defined by the operator based on the upper bounds of the quadrotor and target localisation errors. It also depends on the duration and magnitude of the temporal increase in the target velocity compared to the quadrotor velocity. A sufficient condition is provided for the minimum desired altitude of the quadrotor to ensure that the target is maintained in the FOV. Furthermore, a model predictive control (MPC) is employed to preserve the target at the center of the aerial image and the desired altitude determined by the guidance law. Also, the integrals of the position errors are used to achieve null steady-state errors in the presence of wind disturbances. The simulation results show the effectiveness of the proposed approach in preserving the manoeuverable target in the FOV in the presence of the wind, the uncertainty of the target and quadrotor localisation, accelerations estimation errors, and terrain altitude variation.

Elastomeric and Fluidlastic® absorbers are proposed to be embedded in blade cavity to reduce troublesome second harmonic flap-wise blade loads of rigid rotors. A simple model of a blade with a flap-wise dynamic absorber is derived to investigate centrifugal force effects on the absorber's rotating frequency. The system's dynamic equations of motion indicate that centrifugal force has no effect on absorber's frequency under rotation. Aeroelastic simulation of the coupled rotor and absorber system based on an elastic beam model in generalised force formulation is used to explore load reduction extent and dynamic absorber behaviour. At a forward speed of 200 km/h, the elastomeric absorber reduces the second harmonic flap-wise root-bending moment by 88.8% for unacceptably large stroke; the Fluidlastic® absorber reduces the moment by 78.5% with a stroke 4.26% of blade chord length. Results indicate that placing an embedded Fluidlastic® absorber in the blade's flap-wise direction has potential for reducing load in different flight states with relatively small strokes. Increasing tuning port area ratio or Fluidlastic® absorber mass can distinctly decrease strokes without increasing loads. However, deviation from normal rotor speed significantly lowers Fluidlastic® absorber performance.

An analysis is presented which allows the engineer to quantitatively estimate the validity bounds of aerodynamic methods based in linear potential flows a-priori. The development is limited to quasi-steady planar flows with attached shocks and small body curvature. Perturbation velocities are parameterised in terms of Mach number and flow turning angle by means of a series-expansion for flow velocity based in the method of characteristics. The parameterisation is used to assess the magnitude of non-linear term-groupings relative to linear groups in the full potential equation. This quantification is used to identify dominant nonlinear terms and to estimate the validity of linearising the potential flow equation at a given Mach number and flow turning angle. Example applications include the a-priori estimation of the validity bounds for linear aerodynamic models for supersonic aeroelastic analysis of lifting surfaces and panels.

Fatigue management in military aviation remains problematic and can be easily neglected by planners and aircrew in a busy operational environment. The basic assessment tool (BAT) is an objective and educational computer tool for evaluating and combining various fatigue factors. It provides an impartial performance comparison while at the same time educating and reminding its users of the importance of fatigue management. Individual factors of the BAT were chosen because they affect fatigue, sleepiness, alertness and performance. A short explanation of components not chosen is also included. An example of how the BAT can be used is described and figures of the resulting outputs (screenshots) are included. Most commercial fatigue estimation products are subjective and do not consider specific issues associated with tactical operations in a field environment. This novel approach can be used to help plan missions by minimizing the BAT score while still maximising the tactical advantage and keeping fatigue management fresh in the minds of military rotary wing pilots while in training and on operations. Although the BAT is military aviation focused and objective, it can be easily modified to suit particular fatigue concerns.

As a consequence of its aerodynamic lift, each aircraft is generating wake vortices that can become a serious hazard for the succeeding aircraft and in the worst case lead to loss of control. Separation standards were introduced to prevent dangerous wake encounters that have proven sufficiently safe but also very conservative, thus limiting airspace and airport capacity.

The objective of contemporary wake vortex research is regaining capacity in favourable atmospheric conditions, at the same time preserving or even improving the current safety level. The wake vortex advisory systems developed for this purpose are nearing experimental implementation and all include a forecasting component and additional online detection sensors, with both elements having their benefits and limitations.

The wake vortex research objectives at the Institute of Flight Guidance (IFF) of the Technische Universitaet Braunschweig focus on a close coupling approach between model predictions and sensor measurements. Using the complementary capabilities of each, the aim of improved performance of the overall system shall be reached.

After giving a short overview on the wake vortex phenomenon and current approaches to wake vortex alerting and advisory systems, this paper will introduce the fusion concept. The benefits of this approach will be demonstrated by means of examples. An outlook on further development and envisaged advanced implementation concepts will be given.

The use of machine vision to guide robotic spacecraft is being considered for a wide range of missions, such as planetary approach and landing, asteroid and small body sampling operations and in-orbit rendezvous and docking. Numerical simulation plays an essential role in the development and testing of such systems, which in the context of vision-guidance means that realistic sequences of navigation images are required, together with knowledge of the ground-truth camera motion. Computer generated imagery (CGI) offers a variety of benefits over real images, such as availability, cost, flexibility and knowledge of the ground truth camera motion to high precision. However, standard CGI methods developed for terrestrial applications lack the realism, fidelity and performance required for engineering simulations.

In this paper, we present the results of our ongoing work to develop a suitable CGI-based test environment for spacecraft vision guidance systems. We focus on the various issues involved with image simulation, including the selection of standard CGI techniques and the adaptations required for use in space applications. We also describe our approach to integration with high-fidelity end-to-end mission simulators, and summarise a variety of European Space Agency research and development projects that used our test environment.

Distributed real-time avionics networks have been subjected to a great evolution in terms of functionality and complexity. A direct consequence of this evolution is a continual growth of data exchange. AFDX standardised as ARINC 664 is chosen as the backbone network for those distributed real-time avionics networks as it offers high throughput and does not require global clock synchronisation. For certification reasons and engineering research, a deterministic upper bound of the end-to-end transmission delay for packets of each flow should be guaranteed. The Forward Approach (FA) is proposed for the computation of the worst-case end-to-end transmission delay. This approach iteratively estimates the maximum backlog (amount of the pending packets) in each visited switch along the transmission path, and the worst-case end-to-end transmission delay can be computed. Although it is pessimistic (overestimated), the Forward Approach can provide a tighter upper bound of the end-to-end transmission delay while considering the serialisation effect. Recently, our research finds the computation of the serialisation effect might induce an optimistic (underestimated) upper bound. In this paper, we analyse the pessimism in the Forward Approach and the optimism induced by the computation of the serialisation effect, and then we provide a new computation of the serialisation effect. We compare this new computation with the original one, the experiments show that the new computation reduces the optimism and the upper bound of the end-to-end transmission delay can be computed more accurately.

This article presents a unified approach to solve steady-state conjugate heat-transfer problem including simultaneously gas, liquid and solid regions in just one 3D domain, distinguished by their particular properties. This approach reduces approximation errors and the time to solve the problem, which characterise iterative methods based on separated domains. The formulation employs RANS equations, realisable k-ε turbulence model and near-wall treatment model. A commercial CFD code solves the pressure-based segregated algorithm combined with spatial discretisation of second order upwind. The problem consists of a convergent-divergent metallic nozzle that contains cooling channels divided in two segments along the wall. The nozzle wall insulates the high-speed hot air flow, dealt as perfect gas, from the two low-speed cold water flows, dealt as compressed liquid, both influenced by transport properties dependent of the local temperature. The verification process uses three meshes with increasing resolutions to demonstrate the independence of the results. The validation process compares the simulation results with experimental data obtained in high-enthalpy wind tunnel, demonstrating good compliance between them. Results for the bulk temperature rise of the water in the second cooling segment of the nozzle showed good agreement with available experimental data. Numerical simulations also provided wall temperature and heat flux for the gas and liquid sides. Besides, distribution of temperature, pressure, density and Mach number were plotted along the nozzle centerline showing a little disturbance downstream the throat. This phenomenon has been better visualised by means of 2D maps of those variables. The analysis of results indicates that the unified approach herein presented can make easier the task of simulating the conjugate convection-conduction heat-transfer in a class of problems related to regeneratively cooled thrust chambers.

This paper presents the optimum design of a PID controller for the Adaptive Torsion Wing (ATW) using the genetic algorithm (GA) optimiser. The ATW is a thin-wall, two-spar wingbox whose torsional stiffness can be adjusted by translating the spar webs in the chordwise direction inward and towards each. The reduction in torsional stiffness allows external aerodynamic loads to deform the wing and maintain its shape. The ATW is integrated within the wing of a representative UAV to replace conventional ailerons and provide roll control. The ATW is modelled as a two-dimensional equivalent aerofoil using bending and torsion shape functions to express the equations of motion in terms of the twist angle and plunge displacement at the wingtip. The full equations of motion for the ATW equivalent aerofoil were derived using Lagrangian mechanics. The aerodynamic lift and moment acting on the aerofoil were modelled using Theodorsen’s unsteady aerodynamic theory. The equations of motion are then linearised around an equilibrium position and the GA is employed to design a PID controller for the linearised system to minimise the actuation power require. Finally, the sizing and selection of a suitable actuator is performed.

This paper presents an integrated design and costing method for large stiffened panels for the purpose of investigating the influence and interaction of lay-up technology and production rate on manufacturing cost. A series of wing cover panels (≈586kg, 19·9m2) have been sized with realistic requirements considering manual and automated lay-up routes. The integrated method has enabled the quantification of component unit cost sensitivity to changes in annual production rate and employed equipment maximum deposition rate. Moreover the results demonstrate the interconnected relationship between lay-up process and panel design, and unit cost. The optimum unit cost solution when using automated lay-up is a combination of the minimum deposition rate and minimum number of lay-up machines to meet the required production rate. However, the location of the optimum unit cost, at the boundaries between the number of lay-up machines required, can make unit cost very sensitive to small changes in component design, production rate, and equipment maximum deposition rate.

The establishment of the UK Applied Aerodynamics Consortium in 2004 brought together many of the UK’s leading research groups to tackle challenging aerodynamic problems on the national computing facility, HPCx. This paper provides a brief history of some early pioneers of numerical simulation and highlights some key contributions to development in parallel processing that laid the foundations for today’s researchers. The transition from vector to massively parallel processing is discussed from a UK viewpoint along with technological barriers that could have a significant impact on future systems. Solutions to these barriers are already being sought and the paper discussed some of the novel technologies that may be deployed in the future. In its short history, the consortium has made substantial progress and this is briefly discussed with several highlights that illustrate the scientific output. Although a number of challenges are identified, particularly with respect to developing a comprehensive visualisation capability, the consortium is well placed to build upon its initial success.

A statistical approach is followed for prediction of tolerences of notched strength of composite laminates using the recently proposed improved inherent flaw model (IFM). In order to examine the validity of this approach, the existing fracture data on graphite/epoxy composite laminates containing central holes and cracks were used. The notched strength estimations are found to be within the range of tested values.

The aim of the present study is to develop a relatively simple flight dynamic model which should have the ability to analyse trim, stability and response characteristics of a rotorcraft under various manoeuvring conditions. This study further addresses the influence of numerical aspects of perturbation step size in linearised model identification and integration timestep on non-linear model response. In addition, the effects of inflow models on the non-linear response are analysed. A new updated Drees inflow model is proposed in this study and the applicability of this model in rotorcraft flight dynamics is studied. It is noted that the updated Drees inflow model predicts the control response characteristics fairly close to control response characteristics obtained using dynamic inflow for a wide range of flight conditions such as hover, forward flight and recovery from steady level turn. A comparison is shown between flight test data, the control response obtained from the simple flight dynamic model, and the response obtained using a more detailed aeroelastic and flight dynamic model.

High-precision response of the surrogate model is desired in the process of optimisation. An excessive number of sampling points will increase the cost of the calculation. The appropriate number of sampling points cannot only guarantee the accuracy of the surrogate model but also save the calculation cost. The purpose of this research is to demonstrate the eventuality of using an adaptive surrogate model for optimisation problems. The adaptive surrogate model is built on an adaptive sampling approach and an extended radial basis function (ERBF). The adaptive sampling is an approach that new sampling points are placed in the blank area and all the sampling points are uniformly distributed in the design region using Multi-Island GA. The precision of the ERBF surrogate model is checked using standard error measure to determine whether the surrogate model should be updated or not. This adaptive surrogate model is used to optimise a cruise missile head shape. Aerodynamic and stealthy performance of the cruise missile head shape are considered in this research. Different global objective function and different weight factor are used to research the aerodynamic and stealthy performance in this optimisation process. The results show that the drag is reduced with a slender head shape and the radar-cross section (RCS) value is reduced with a short head shape.