Values are important in understanding the managerial behaviour. Values are the unique criteria that enable people to become conscious of social relations and duties. We contribute to this understanding through determining the values which affect an organisation’s business approach by providing evidence from a comparative study of various airports through a questionnaire method. The study was carried out with 163 participants and factor analysis was used to reduce the complexity of a data set so that it becomes easier to use the data in applied settings. Ranking analysis was used to get the values hierarchy of managers. This hierarchy-addicted culture helps to understand corporate sustainability and loyalty. Managing values increases quality and retains sustainability. Further suggestions are made regarding values that should be taken into consideration for achieving corporate strategies, whether operating regionally or globally. This study contributes towards improving awareness on the effects of values in business management in both theory and practice, along with their limitations. The analysis shows that there is a conformity between organisational and individual values.

A new type of design domain for a sounding rocket is suggested in this study, which is more intuitive, simplified, and conducive to a single-stage sounding rocket development process. Among the various operation parameters, this study identifies several effective variables, which are also among the most practical in a sounding rocket design process. Of the many design variables considered for peak altitude optimisation, arguably the three most effective in shaping the whole system and conferring the most practicality in the design process of a sounding rocket were determined. A simulation-based study was conducted to establish: the effect of the selected parameters on flight performance, and the optimum design conditions of a single-stage sounding rocket in terms of its peak altitude. The simulation result was compared with randomly chosen experimentally tested flight data and validated. With the clear performance curves varying depending on the variables, the combination of the considered design inputs was effective. The new type of design domain and design procedure proposed is expected to constitute a useful reference and deliver practical benefits to the development process for a target-altitude-optimised single-stage sounding rocket.

Aerodynamic characterisation from flight testing is an integral subroutine for evaluating a new flight vehicle’s aerodynamic performance, stability and controllability. The estimation of aerodynamic parameters from flight test data has extensively been explored, in the past, using estimation methods such as the equation error method, output error method and filter error method. However, in the current era, non-gradient-based estimation techniques are gaining attention from researchers due to their inherent data-driven optimisation capability to find the global best solution. In this paper, a novel non-gradient-based estimation method is proposed for the aerodynamic characterisation of unmanned aerial vehicles from flight data, which relies on the maximum likelihood method augmented with particle swarm optimisation. Flight data sets of a wing-alone unmanned aerial vehicle are used to demonstrate the capabilities of the proposed method in estimating aerodynamic derivatives. Estimates from the proposed method are corroborated with the wind tunnel test and output error method results. It has been observed that simulated flight vehicle responses using estimated parameters are in good agreement with measured data in most of the manoeuvers considered. Confidence in the estimates of linear and nonlinear aerodynamic parameters is well established with the lower limit of Cramer-Rao bounds, which are minimal. The proposed method also demonstrates good predictability of the quasi-steady stall aerodynamic model by estimating stall characteristic parameters such as aerofoil static stall characteristics parameter, hysteresis time constant and breakpoint. The overall performance of the proposed estimation method is on par with the output error method and is validated with the proof-of-match exercise.

A combined experimental and theoretical study has been made of radial compressible flow without swirl between parallel discs when the fluid velocity is everywhere subsonic. It constitutes an extension of a previous analysis, wherein the flow was considered to be incompressible.

A similarity solution for the radial pressure distribution is shown to be possible only in special cases where certain terms in the equations of motion can be neglected. Approximate solutions are obtained for the laminar and the turbulent radial pressure distributions, using an integral momentum method. Both theories agree well with experiment. The critical Reynolds number for reverse transition is found to be approximately the same as that for incompressible flow in radial channels and circular pipes. The radial pressure level for compressible radial flow between parallel discs is found to be less than that for incompressible flow at the same mass-flow rate. The non-linear form, in pressure, of the compressible solution can be shown to account for this result.

The pressure oscillation and terminal shock motion in a two dimensional inlet, which was designed for tandem configuration turbine-based combined cycle propulsion systems was investigated experimentally and numerically, respectively. The inlet was characterised by a bleed cavity upstream the inlet throat, an S-shape rectangular-to-circular diffuser and flowpaths for a turbine and a ramjet engine. The terminal shock motion was calculated through a second-order unsteady Reynolds-averaged Navier-Stokes scheme. The pressure and the terminal shock were unsteady when the combined cycle inlet operated at different conditions. With the terminal shock located in the throat and at the shoulder of the third ramp of the TBCC inlet, the pressure oscillation was significant and the shock exhibited unsteady streamwise motion with an oscillatory pattern. The amplitude of shock oscillation at these two conditions was 6mm and 12mm, respectively. When the shock was located downstream of the throat and upstream of the cowl lip, it oscillated in a small range. We defined this motion as the “shake” of the shock. This unsteady behaviour of the shock was caused by flow separation in the combined cycle inlet diffuser.

The importance of fuel injection methods and their atomisation characteristics has been well described in literature. Also, there are a large number of studies in literature detailing the impact of bulk properties of conventional fuels on atomisation and spray characteristics. However, there is a lack of knowledge on how different alternative fuels, and specifically aromatic species, impact spray and atomisation characteristics. In this investigation, the impact of alternative fuels, selected aromatic species and their properties on spray atomisation was studied. Details of how different aromatic structures, bonds and other micro properties of aromatics species impact spray and atomisation characteristics were investigated. To achieve this objective, testing was conducted using a Rolls-Royce Tay engine fuel injector. It was found that the droplet sizes in the form of the Sauter Mean Diameter (SMD) correlate well with fuel properties including density, viscosity and surface tension, which is in line with other published literature. Moreover, it was found that there are several aromatic species (ethylbenzene, cumene and tert-butylbenzene) which display improved atomisation. This indicates that the size of the aliphatic groups attached to the benzene ring in the aromatic molecule impact on the drop size and thereby the combustion characteristics. The worst performers were polyaromatic naphthalene compounds. With the increasing push for fuels with selected aromatic species in the future, this study provides knowledge on the impact of their on the atomisation characteristics.

The maximum attainable performance of small gas turbines represents a strong limitation to the operating altitude and endurance of high-altitude unmanned aerial vehicles (UAVs). Significant improvement of the cycle thermal efficiency can be achieved through the introduction of heat exchangers, with the consequent increase of the overall engine weight. Since semi-closed cycle engines can achieve a superior degree of compactness compared to their open cycle counterparts, their use can offset the additional weight of the heat exchangers. This paper applies semi-closed cycles to a high-altitude UAV propulsion system, with the objective of assessing the benefits introduced on the engine performance and weight. A detailed model has been created to account for component performance and size variation as function of thermodynamic parameters. The sizing has been coupled with a multi-objective optimisation algorithm for minimum specific fuel consumption and weight. Results of two different semi-closed cycle configurations are compared with equivalent state-of-the-art open cycles, represented by a recuperated and an intercooled-recuperated engine. The results show that, for a fixed design power output, engine weight is approximately halved compared to state-of-the-art open cycles, with slightly improved specific fuel consumption performance. Optimum semi-closed cycles furthermore operate at higher overall pressure ratios than open cycles and make use of recuperators with higher effectiveness as the mass penalty of the recuperator is smaller due to the lower engine mass flow rates.

The stresses around the passenger door, emergency door and window of an aircraft, Airbus A300B, are considered in this paper. By idealising these opening as rectangular holes with rounded corners in an infinite plate subjected to inplane loads and considering the presence of adjacent holes the stresses are determined using the complex variable method along with Schwarz Alternating Method. The results of the analytical solution are found to be in good agreement with those of the finite element analysis using ANSYS 5.0 software within 5%.

For the engineer, the “Concorde challenge” could be defined as follows:

A passenger carrying aircraft twice as fast as the existing aircraft and at least as safe.

When the jets came out 20 years ago, there was a sudden increase of the speeds used up to then, but their envelope stopped at the limit where the difficulties associated with very high speeds start

In much of the recent work on the flow of an incompressible fluid past a swept-back symmetrical wing of small thickness at zero incidence it has been assumed that the disturbance introduced by the wing can be represented by means of a distribution of sources and sinks over the centre plane of the wing, the source strength being determined from the equation of the wing profile. It was found by Neumark that on the wing itself this procedure was liable to lead to serious errors. Neumark and Kuchemann studied certain special types of wings for which they were able to find the correct value of the pressure at the surface by evaluation of the pressure at points off the wing and continuity considerations. This process is rather cumbersome; it may therefore be convenient to regard the problem from a rather different point of view.

This work presents a study of acoustic oscillations generated and the wave structure associated with supersonic flow past wall mounted 3D open cavities of varying length-to-width (L/W) ratio. Experiments were conducted to investigate the acoustic signature generated by the cavities at freestream Mach number of 1·7. The effect of L/W ratio of the cavity on the dominant modes of the acoustic signature registered on different walls of the cavities is investigated for an L/W range of 0·83-4. Shift in the dominant acoustic mode is observed as L/W ratio changes from 3 to 4. Statistical analysis of pressure data showed existence of acoustic waves and spreading of acoustic energy over different modes with change in cavity width. Time averaged schlieren visualisation indicated variation of shock and shear layer structure in the mainstream for the different cavities. Acoustic waves generated by the presence of the cavity and the dynamic behaviour of the shear layer were observed during instantaneous shadowgraph visualisation. Numerical simulation was done to make a prior assessment of the flow structure and the results are in good agreement with those from experiments. Ratio of mass exchange between cavity and mainflow and the cavity volume was observed to have profound effect on the magnitude of pressure oscillations generated by the cavities.

Special forms of the receptance functions of mechanical systems are shown to indicate the existence of nodes. these are associated with horizontal straight lines in receptance diagrams. the effects of this nodal condition are explained in terms of simple torsional and translational systems for the sake of simplicity; but the theory may be extended to any type of system and any set of generalised co-ordinates. a particular instance arises if principal co-ordinates are used. the paper is a sequel to an earlier paper.

Flying a light aeroplane involves a combination of pilot and aeroplane performing a set task, within a specific environment. The pilot is continuously sampling and selecting available sensory cues, interpreting those cues, making decisions and manipulating the primary controls (stick and rudder) to safely achieve flying objectives. The ‘feel’ of an aeroplane (a flying quality) is directly associated with the stick and rudder forces and how the aeroplane responds to control inputs. Classical theory has been applied to estimate the apparent (as felt by the pilot) longitudinal stick-free static stability (change of stick force with airspeed) of a typical, two-seat, high-wing light aeroplane. The theory has been extended to consider the effects of tail downwash and flap deflection. The results are compared with actual flight tests and show that the method may be used for the initial assessment of longitudinal stick-free static stability and more importantly, tendencies towards neutral or negative stability affecting flight safety.

This paper describes and presents results from static wind tunnel tests conducted on a 60° delta wing at a root chord Reynolds number of 2·7 x 106. In these tests, the wing was instrumented with 192 miniature pressure transducers which, in conjunction with a powerful multi-channel data-logging system, allowed the distribution of time-varying surface pressures to be measured at high temporal resolution. Analysis indicates that the distribution of root mean square pressure on the leeward surface of the wing can provide considerable insight into the behaviour of both the primary and secondary vortex structures. In addition, it has been established that the frequency content of pressure signals measured in the vicinity of these vortex structures is sensitive to the vortex state. It is suggested that these data features can be directly attributed to previously observed behavioral characteristics of the vortex breakdown process.

This note assesses the benefits of active aeroelastic structures (AAS) in enhancing flight performance and control authority. A representative AAS concept, whose torsional stiffness and shear centre position can be altered depending on the instantaneous flight condition, is employed in the wing of a medium altitude long endurance (MALE) UAV. A multidisciplinary design optimisation (MDO) suite is used in this study. It turns out that AAS can be very effective when used for enhancing control authority of the vehicle but have limited benefits in terms of flight performance (lift to drag).

During early phases of wing design, analytic and low-fidelity methods are often used to identify promising design concepts. In many cases, solutions obtained using these methods provide intuition about the design space that is not easily obtained using higher-fidelity methods. This is especially true for aerostructural design. However, many analytic and low-fidelity aerostructural solutions are limited in application to wings with specific planforms and weight distributions. Here, a numerical method for minimising induced drag with structural constraints is presented that uses approximations that apply to unswept planar wings with arbitrary planforms and weight distributions. The method is applied to the National Aeronautics and Space Administration (NASA) Ikhana airframe to show how it can be used for rapid aerostructural optimisation and design-space exploration. The design space around the optimum solution is visualised, and the sensitivity of the optimum solution to changes in weight distribution, structural properties, wing loading and taper ratio is shown. The optimum lift distribution and wing-structure weight for the Ikhana airframe are shown to be in good agreement with analytic solutions. Whereas most modern high-fidelity solvers obtain solutions in a matter of hours, all of the solutions shown here can be obtained in a matter of seconds.

A multi-agent engagement scenario is considered in which a high-value aircraft launches two defenders to intercept two homing missiles aimed at the aircraft. Under the assumption that all aircrafts have first-order linear dynamic characteristics, a combined multiple-mode adaptive estimation (MMAE) and a two-way cooperative optimal guidance law are proposed for the target–defenders team. Considering the full cooperation of the target and both the two defenders, the two-way cooperative strategies provide the analytical expressions for their optimal control input, enabling the target–defenders team to intercept the missiles with minimal control effort. To successfully intercept the missiles, MMAE is used to identify the guidance laws adopted by the missiles and estimate their states. The simulation results show that the target cooperating with the defenders to perform lure manoeuvres for the missiles can improve the guidance performance of the defenders as well as reduce the control effort of the defenders for intercepting the missiles.

In this work the co-ordinates of three-dimensional nozzles of several types have been obtained as explicit functions, and some shapes have been plotted. It is hoped that some of these might be of use in cases where smooth uniform fluid flow is desired. The design of a particular nozzle shape in the presence of a plane boundary is given: previously any equations obtained have been too intractable to use.

An algorithm is described for solving the partial differential equations governing inviscid, incompressible, irrotational and axisymmetric flow in order to obtain duct wall shapes from prescribed wall velocity distributions. Equations of motion are formulated with distances from and along the axis of symmetry as dependent variables. A velocity potential and Stokes stream function are employed as independent variables. A numerical method based on the iterative solution of finite-difference equations on a non-uniform mesh is presented. The residuals obtained from calculations with differing mesh sizes, for the two examples described, indicate that the proposed method gives satisfactory solutions to the specified design problem.

Numerical simulations were performed to analyse the possibility of damping abrupt incoming free stream disturbances upon a porous aerofoil using an unsteady Reynolds-averaged Navier-Stokes (URANS) model. To mimic the turbulence disturbance levels that are typically encountered in the atmosphere, two flow configurations were considered. In the first configuration, the unsteadiness of the flow was created with vortices shed from a circular cylinder installed ahead of a WTEA-TE1 aerofoil. The continuous von Kármán shedding vortices contained within the cylinder wake were convected downstream and projected upon the aerofoil. In the second configuration, an instantaneous pair of discrete vortices was created by a rotational snapping of a flat plate, installed upstream of the aerofoil. Solid and porous aerofoil configurations, with porosity settings of 11 and 22%, were applied on 50% of the chord of the aerofoil starting from the leading edge. Both steady and unsteady flow simulations were performed to assess the performance of the porosity under steady and unsteady effects. The steady state flow simulations revealed a noticeable reduction in the aerofoil lift coefficient for the porous aerofoil. For unsteady solutions with a continuous or distinct series of vortices interacting with the aerofoil, the porosity showed insignificant damping of the lift coefficient amplitude. The porosity values investigated in the current exercise had indiscernible effect upon the unsteady lift-load alleviations caused by free stream disturbances.