Climate change is increasing global-mean tropospheric temperatures, but the localised trends are uneven, including cooling the lower stratosphere and lifting the tropopause. The wind speeds are also being modified, both at the surface and aloft. A further effect, additional to wind and temperature alone, is of increasing fluctuations and severity of extreme weather. These are impacting air transport, and this will continue. The effects are known to include increased take-off distances where excess runway lengths exist and reduced payloads where they do not, increased en-route flight times, increased frequency and severity of encounters with clear air turbulence in some regions, changed patterns of wildlife — particularly bird — activity in some regions (potentially also for other anthropogenic reasons) are shifting locations of flight safety hazards, and increased burdens upon airport and associated infrastructure. There is increasing understanding and acknowledgment by companies and authorities of these effects and the importance of mitigating them, although this is not universal and there are as yet no universally understood best practices for air transport climate change mitigation.

From an analysis of the work of Crocco and others, semi-empirical formulae are derived for the skin friction on a flat plate at zero incidence with a laminar boundary layer. These formulae are

for the general case of heat transfer, and

when there is no heat transfer.

The problem of heat transfer and the effect of radiation are discussed in the light of these formula;. The second formula is then utilised in the development of an approximate method for solving the momentum equation of the boundary layer on a cylinder without heat transfer. The method indicates that with increase of Mach number there is a marked forward movement of separation from a flat plate in the presence of a constant adverse velocity gradient.

An analysis of an extensive series of unsteady and steady pressure measurements at transonic speeds on a rigid 12% thick NACA 16 series aerofoil reveals a useful classification of the type of excitation characteristics of particular flow regions.

The pressure fluctuation measurements suggest that when the shock from the aerofoil reaches the tunnel roof, serious interference (both dynamic and static) occurs.

The RAE 2 ft × 1½ ft Transonic Tunnel used for the tests has low levels of flow unsteadiness at low frequencies, making it a useful facility for basic research in time-dependent aerodynamics.

The flow field of a turbulent free jet issuing from a contoured rectangular nozzle of aspect ratio 2 has been studied experimentally, using hot-wire anemometry. The study was undertaken to gain some understanding of the mixing process within the jet. Results of the measured and derived mean flow and turbulence quantities are presented. The three components of the mean velocity vector, the three Reynolds normal stresses, the two Reynolds primary shear stresses, and the flatness factor of the streamwise fluctuating velocity were measured. Mass entrainment, turbulence kinetic energy, and the intermittency factor were derived from the mean streamwise velocity data, the Reynolds normal stress data, and the flatness factor data, respectively. The derived results and the Reynolds primary shear stress data indicate enhanced near-field mixing of the present rectangular jet compared to a round turbulent free jet. The mean streamwise velocity field changes from a rectangular to an oval and then to an approximately circular shape at about twenty equivalent nozzle diameters downstream of the nozzle exit plane.

Major results and findings of the numerical work package of the European high lift programme EUROLIFT are outlined. The main objective of these studies is to validate and test numerical methods for the prediction of high lift flows for transport aircraft configurations. The activities comprise the assessment of current CFD methods for 3D flows, evaluation of means for code improvement, and transition prediction. All aspects are especially devoted to high lift flow problems. A general capability to predict maximum lift on a simplified wing/fuselage high lift configuration is demonstrated by a variety of different numerical approaches. In general, major shortcomings are the reliability and the accurate simulation of large separation areas and the turn-around time to compute 3D lift polars. Advanced turbulence modelling and numerical solver features, such as the preconditioning technique, show a potential to overcome these deficiencies. Promising results with respect to transition prediction were obtained on a swept high lift wing using a database method. The results obtained in the numerical activities represent major ingredients on the way to a consistent numerical approach for the simulation of transport aircraft high lift configurations including all maximum lift determining effects.

This paper describes a multidisciplinary robust optimisation framework for UAV conceptual design. An in-house configuration designer system is implemented to generate the full sets of configuration data for a well-developed advanced UAV analysis tool. A fully integrated configuration designer along with the UAV analysis tool ensures that full sets of configuration data are provided simultaneously while the UAV configuration changes during optimisation. The computational strategy for probabilistic analysis is proposed by implementing a central difference method and fitting distribution for a reduced number of Monte Carlo Simulation sampling points. The minimisation of a new robust design objective function helps to enhance the reliability while other UAV performance criteria are satisfied. In addition, the fully integrated process and a probabilistic analysis strategy method demonstrate a reduction in the probability of failure under noise factors without any noticeable increase in design turnaround time. The proposed robust optimisation framework for UAV conceptual design case study yields a more trustworthy prediction of the optimal configuration and is preferable to the traditional deterministic design approach. The high fidelity analysis ANSYS Fluent 13 is performed to demonstrate the accuracy of proposed framework on baseline, deterministic and RDO configuration.

The aerodynamic interference between typical tanker and large receiver aircraft during air-to-air refuelling was modelled using a simple horseshoe vortex representation of the tanker wake whilst the aerodynamic forces and moments acting on the receiver were calculated using the vortex lattice method. For each combination of tanker and receiver, the aerodynamic loads depend mainly on the vertical separation, whilst the receiver tailplane position on the fin strongly effects the pitching moment acting on the receiver due to the tanker downwash. Particular attention was given to a high T-tail receiver aircraft. The predicted increase in the receiver drag compares favourably with available flight test data. Changes in trim in pitch of the receiver were analysed and related to the static stability of the aircraft. Aerodynamic derivatives due to the position and attitude of the receiver aircraft within the tanker downwash field were calculated and solutions of the linearised equations of motion obtained for the high T-tail receiver refuelling from different tanker aircraft at different flight conditions. Typically the receiver exhibited a divergent oscillation at the normal refuelling position. The dynamic stability characteristics are consistent with the trends observed in flight tests on the tendency of the receiver to oscillate in pitch.

The matrix methods of structural analysis developed specifically for use on modern digital computers have now become universally accepted in structural design. These methods provide a means for rapid and accurate stress and deflection analysis of complex structures under static and dynamic loading conditions and they can also be used very effectively for the stability analysis. In the conventional stability analysis two possible approaches are normally used; either the differential equations describing the structural deflections are formulated and the lowest eigenvalue representing the buckling load condition is found for a given set of boundary conditions, or alternatively, if the differential equations are too difficult to prescribe, approximate deflection shapes are used in the strain energy expression for large deflections which is subsequently minimised, leading to the stability determinant whose lowest root represents the instability condition. When designing complex structures the conventional methods of finding buckling load conditions are extremely difficult to apply, and therefore in such cases we have to rely on the matrix methods of stability analysis.

The present energy situation has caused engineers to look more closely at methods of energy recovery which have previously received only limited attention due to the availability of cheap and plentiful fossil fuels. One of these methods is the use of an airscrew, or windmill, to convert wind energy into mechanical energy. There are a number of disadvantages in the utilisation of wind energy: the power available is relatively low for the size of plant involved and the supply is intermittent, but on the other hand the energy is free and available in most areas.

The economic feasibility of using windmills on a scale large enough to supply the national grid system is debatable, but there is certainly a demand for small units suitable for the individual domestic consumer.

An approximate theory is described which takes into account the influence of the wake on the aerodynamic derivatives of an oscillating rotor blade. It is shown that the most important parameter is the number of cycles of the oscillation which occur during one revolution of the rotor. For integral values of this parameter the flexural damping is shown to become very small at small angles of incidence. The flutter characteristics of a simple rotor blade are calculated on the basis of this theory, and it is found that the wake causes radical changes in the critical speeds. The theory is in good qualitative agreement with experiment.

The climate impacts of three fleets of supersonic small-scale transport aircraft (S4TA) are simulated. Based on characteristic aircraft parameters, which were developed within the EU-project HISAC, emissions along geographically representative trajectories are calculated and in addition the resulting changes in the atmospheric composition (carbon dioxide, ozone layer, water vapour) and climate (near surface global mean temperature) are deduced. We assume a fleet development with an entry in service in 2015, a full fleet in 2050. The results show a temperature increase of 0·08mK (0·07-0·10mK) with only small but statistically significant variations between the configurations, leading to a minimum climate impact for a weight optimised and hence lower flying aircraft. A climate impact ratio of 3·0 ± 0·4 between a S4TA and its subsonic counterpart is calculated, which is considerable less than for previous supersonic fleets because of a lower flight altitude, leading to smaller water vapour impacts.

The aerodynamics and acoustics of a generic coaxial helicopter with a stiff main rotor system and a tail-mounted propulsor are investigated using Brown’s Vorticity Transport Model. In particular, the model is used to capture the aerodynamic interactions that arise between the various components of the configuration. By comparing the aerodynamics of the full configuration of the helicopter to the aerodynamics of various combinations of its sub-components, the influence of these aerodynamic interactions on the behaviour of the system can be isolated. Many of the interactions follow a simple relationship between cause and effect. For instance, ingestion of the main rotor wake produces a direct effect on the unsteadiness in the thrust produced by the propulsor. The causal relationship for other interdependencies within the system is found to be more obscure. For instance, a dependence of the acoustic signature of the aircraft on the tailplane design originates in the changes in loading on the main rotor that arise from the requirement to trim the load on the tailplane that is induced by its interaction with the main rotor wake. The traditional approach to the analysis of interactional effects on the performance of the helicopter relies on characterising the system in terms of a network of possible interactions between the separate components of its configuration. This approach, although conceptually appealing, may obscure the closed-loop nature of some of the aerodynamic interactions within the helicopter system. It is suggested that modern numerical simulation techniques may be ready to supplant any overt reliance on this reductionist type approach and hence may help to forestall future repetition of the long history of unforeseen, interaction-induced dynamic problems that have arisen in various new helicopter designs.

This paper presents a systematic formulation for the simulation of rigid body dynamics, including the short period dynamics, inherent to stage separation and jettisoning parts of a satellite launcher. This also gives a review of various types of separations involved in a launch vehicle. The problem is sufficiently large and complex; the methodology involves iterations at successively lower levels of abstraction. The best choice to tackle such problems is to use state-of-the-art programming technique known as object oriented programming. The necessary classes have been identified to represent various entities in the launch vehicle separation process (e.g., gravity, aerodynamics, propulsion and separation mechanisms etc.). Simple linkages are modelled with suitable objects. This approach helps the designer to simulate a launch vehicle separation dynamics and also to analyse separation system performance. To examine the influence of the design variables on the separating bodies, statistical analyses have been performed on the upper stage separation process and pull out of ongoing stage nozzle from the spent stage of a multistage rocket carrier using retro rockets.

A procedure for developing efficient aeroelastic reduced order models (ROMs) for aerospace structures containing geometric nonlinearities is described. The structural modelling is based upon a combined modal/FE approach that describes the non-linear stiffening effects from results of non-linear static analyses for a range of prescribed inputs. Once the structural ROM has been defined, it is coupled to the rational fraction approximation of the doublet lattice aerodynamic model corresponding to the wing planform. The aeroelastic model can then be used to predict the dynamic aeroelastic behaviour of the defined structure. The methodology is demonstrated on the aeroelastic model of a flexible high aspect ratio wing with the static deflections, LCO behaviour and gust response being predicted.

There is a growing interest in the use of alternative fuels in gas turbine engines to reduce emissions. Testing of alternative fuels is expensive when done on a large-scale gas turbine engine. In this study, a re-commissioned small gas turbine auxiliary power unit (APU) has been used to test various blends of Jet A-1, synthetic paraffinic kerosene (SPK) and diesel with as well as eight other novel fuels. A detailed analysis of performance, gaseous emissions and particulate emissions has been presented in this study. It is observed that aromatic content in general as well as the particular chemical composition of the aromatic compound plays a vital role in particulate emissions generation. SPK fuel shows substantially lower particulate emissions with respect to Jet A. However, not all the species of aromatics negatively impact particulate emissions. Gaseous emissions measured are comparable for all the fuels tested in this study.

Under zero-gravity conditions, the liquid surface in a container has a displaced equilibrium position due to surface tension effects; this depends upon the magnitude of the contact angle to the wall. Free and forced sloshing have been investigated for these conditions. It was found that the natural frequencies of the liquid increases with the contact angle to a maximum value at a contact angle slightly above π/2 and then decreases. This behaviour is not symmetrical within the wetting and non-wetting range. Thus, a completely non-wetting liquid has larger natural frequencies than a completely wetting liquid. It was also found that the sloshing mass of a wetting liquid is considerably larger under zero-gravity than under normal gravity conditions and that a non-wetting liquid exhibits substantially reduced sloshing mass, indicating that the influence of liquids with a small contact angle on the dynamics of a vehicle with liquid containers (with a free liquid surface) may be quite pronounced.

Nowhere in the literature of helicopter noise can one find a general understanding; only isolated, inconsistent, and often inadequately documented facts. These could, of course, be reviewed but hardly with profit; it is easier to ask critical questions of nature than of previous investigators.

For fundamental reasons the helicopter in its present state of development is not in fact matched to the bulk, low-cost transport of passengers. But for equally good reasons the phenomena of aerodynamically generated noise now bid fair to banish from inhabited localities all other forms of vertical take-off and landing aircraft so that, if a city-centre to city-centre aircraft transport system is ever to be practicable, it will most probably be based on some metamorphosis of the helicopter.

The ability to track time varying frequency and damping parameters using on-line versions of seven time domain system identification algorithms; Least Squares, Double Least Squares, Correlation Fit, Instrumental Variables, Instrumental Matrix with Delayed Observations, Extended Least Squares and Maximum Likelihood, is examined. Comparisons are made on results obtained from data generated from various simulated time varying systems, corrupted with both input and measurement noise, as a preliminary step before advancing onto real data. Only the Maximum Likelihood and Double Least Squares methods gave acceptable results on the most difficult data sets that were considered. The choice of an exponential weighting factor is crucial to the performance of the on-line methods when applied to data from time varying systems. The effect of assigning various values of the weighting factor is demonstrated. Further work is required to determine the optimum schemes for defining this parameter.

In this paper an approach to fault-tolerant flight control system design based on the integration of sensor/actuator fault detection and isolation (FDI) and reconfigurable control is presented. The effects of the sensor and actuator faults in the innovation process of the Kalman filter are investigated, and a decision approach to isolate the sensor and actuator faults is proposed.

The reconfigurable control algorithm based on the Extended Kalman Filter (EKF) is presented. The control reconfiguration procedure is executed by considering the identified control distribution matrix. In the simulations, the longitudinal dynamics of an aircraft control system are considered, and control reconfiguration is examined. In the paper, the principal block diagram of a fault tolerant aircraft control system is offered.

Two numerical methods are presented for the computation of steady and unsteady Euler flows. These are applied to steady and unsteady flows about the NACA 0012 aerofoil, using structured grids generated by the transfinite interpolation technique. An explicit central-difference scheme is produced based on the cell-vertex method of Ni modified by Hall. The method is second-order accurate in time and space, and with flow quantities stored at boundaries the boundary conditions are simple to apply. This is a definite advantage over the cell-centred approach of Jameson, where extrapolation of the flow quantities is required at the boundaries, making unsteady boundary conditions difficult to apply. An explicit upwind-biased scheme is also produced, based on the flux-vector splitting of van Leer. The method adopts a three stage Runge-Kutta time-stepping scheme and a high-order spatial discretisation which is formally third-order accurate for one-dimensional calculations. The upwind scheme is shown to be slightly more accurate than the central-difference scheme for steady aerofoil flows, but it is not clear which is the more accurate for unsteady aerofoil flows. However, the central-difference scheme requires less than half the CPU time of the upwind-difference scheme, and hence is attractive, especially when considering three-dimensional flows. The transfinite interpolation technique is ideal for generating moving structured grids due to its simplicity, and grid speeds are available algebraically by the same interpolation as grid points. The method is also ideal for use in a multi-block approach.