The stress distribution in rotating circular discs containing a central hole and a symmetrical array of non-central holes is determined by numerical solution of the equations of generalised plane stress. Particular attention is given to an annulus containing the holes and of width approximately eight hole diameters, in which the full two-dimensional equations are solved. The region outside this annulus is treated as radially symmetric and the stresses there are determined from a simpler one-dimensional model. Stress distributions are reported for uniform discs of fixed geometry containing 10, 20 and 45 holes. Results are also obtained for 20-hole discs of non-uniform thickness comprising a uniformly tapered disc, a disc with a thickened annulus containing the holes, and a uniform disc with each hole surrounded by thickened bosses. As a check on the numerical method, calculations have been carried out on a disc of identical geometry to one examined photoelastically bv Leist and Weber with good agreement. The effect of changing Poisson's ratio for this particular disc is also examined.

The validity and applicability of the static margin (stick fixed) Kn,

where

as defined by Gates and Lyon is shown to be restricted to the conventional flexible aircraft. Alternative suggestions for the definition of static margin are put forward which can be equally applied to the conventional flexible aircraft of the past and the integrated flexible aircraft of the future. Calculations have been carried out on simple slender plate models with both linear and non-linear aerodynamic forces to assess their static stability characteristics.

The concept of a fully-developed flow based on the hypothesis of selective memory is here applied to general wall-jet type flows. In the presence of a (constant) external stream, the free-stream velocity and the jet momentum flux are taken to be the chief quantities governing the development of the wall jet: two additional non-dimensional parameters, representing a momentum flux Reynolds number and the relative momentum defect in the initial boundary layer, are shown to have only a secondary effect on the fully-developed flow. The standard correlations so determined are also found to predict quite well the flow development in Gartshore and Newman’s experiments on wall jets in adverse pressure gradients; possible reasons for this somewhat surprising result are discussed. Finally it is shown, by application to the still-air case, that the parameters discovered in incompressible flow are, with appropriate but straightforward modification, successful in describing compressible wall jets also.

An analysis, based upon exact relations and previously published approximate relations, is presented. It describes the connection between changes in aircraft weight and changes in the energy to revenue work ratio (ETRW), which, for a given aircraft on a given route, correspond to changes in trip fuel burn. This is used to establish the link between weight saving and fuel burn improvement at both the total aircraft and the component levels. The analysis is then extended to address the impact of trading weight savings anywhere on the aircraft for increased wing aspect ratio, whilst the aircraft total weight remains the same. It is shown that, for flights in excess of about 350 km, if saving fuel is the objective and provided that all the aerodynamic design and airworthiness requirements can be met, it is better to trade weight saving for increased aspect ratio. In general, the ratio of fuel burn reductions for traded to non-traded weight varies with aircraft size, design range, distance flown and payload carried, with the maximum values, for typical operational payloads, ranging from 2•8, for the smaller aircraft, to 2•2 for the largest aircraft, with medium haul operations deriving the largest benefit. It is estimated that, over the past 50 years, about 10% of the operational empty weight has been traded for increased aspect ratio, giving close to a 20% improvement in ETRW. Finally, estimates are produced for the impact of weight reduction and traded weight reduction on the fuel burn for the current global feet.

With the advent of recreational sports like kite surfing and buggying, the performance of kites has become a market driven item. Producers increasingly require methods to measure and improve the performance of the kites they manufacture. The Mechanical Engineering Department at the University of Canterbury has been working with a local kite producer to develop testing procedures suitable for kite manufacturers. The primary performance measurement is the lift to drag ratio. An early test rig was mounted in the top of a car, but limitations inherent in the design meant that it lost precision as the lift to drag ratio approached that of more advanced kites. This led the investigators to look for alternatives, and resulted in the development of the circular flight method. This method allows the test apparatus to be tuned to the performance of each kite, significantly improving the precision of the results while reducing the time taken for each test. In their raw form, the L/D results are not quite the same as those of the more traditional methods. But they reflect the underlying aerodynamic characteristics, and when used comparatively they can be used in the kite development process. Alternatively, with suitable processing the circular flight results can be converted to the traditional forms.

Any vehicle that is required to travel at high speed over the surface of the earth (either land or water) should have

• (a) Low drag and

• (b) Some form of spring suspension.

Anyone who doubts the need for the second of these, for travel over water, should consider the accelerations involved in following the surface of the sea at high speed and observe the behaviour of a high-speed motor boat on rough water.

A comparison is made of two recently published integral equation methods for calculating the high subsonic flow around lifting and non-lifting aerofoils. It is shown that one method leads to a non-unique solution.

This paper describes some work, mainly of a theoretical nature, done at the Royal Aircraft Establishment, Farnborough, during the years 1946-1947. The immediate object of this work was to establish some general principles for the guidance of the Ministry of Civil Aviation Airfield Lighting Committee, which is the body charged with the task of deciding what visual aids shall be installed on the civil airports of this country. The work is far from completion and from the scientific point of view, this account of it must necessarily leave some loose ends, but it is felt that the method of attack and the underlying ideas may prove useful to other workers in this field. If this should prove to be the case, then international standardisation will have been brought a little nearer, and as will be made clear in this paper, this is almost as necessary for the visual aids as for the radio aids.

The general problem of determining cruise fuel burn is addressed by considering the variation of the product of engine overall efficiency and airframe lift-to-drag ratio, (ηoL/D), with Mach number and lift coefficient. This quantity is the aerothermodynamic determinant of fuel burn rate. Using a small amount of real aircraft data and exploiting normalisation, it is found that near universal relationships exist between the key variables. With this major simplification, an analytic, near exact solution is derived in which the aircraft-related input data are reduced to just three parameters, namely the optimum value of (ηoL/D) and the lift coefficient and Mach number combination at which it occurs. These are quantities that are either available from open information sources or can be estimated using established analytic methods. By introducing models of the take-off and climb and the descent and landing phases, the method is extended to provide accurate trip fuel estimates.

It is shown that there is an ideal flight level (IFL) at which the fuel consumption rate is a minimum for all speeds in the normal cruise operating range. Furthermore, the IFL at the end of cruise is approximately the same for all aircraft, whilst the IFL at the beginning of cruise depends, primarily, on the distance to be flown. There is little dependence on the size of the aircraft, or its take-off mass.

In the context of the ‘fuel-based’ assessment of operational inefficiency, the method is further developed to determine the sensitivity of the trip fuel requirement to changes in the operational parameters governed by air traffic management (ATM). The result is two simple equations. These are used to estimate the current ATM fuel burden. At the global level, this is about 20% with more than half being attributable to flights of less than 1,200 km.

Finally, the method is used to estimate the fuel burn penalty associated with reducing contrail formation by simply avoiding those regions of the atmosphere that are supersaturated with respect to ice. If the aircraft is at the IFL when avoiding action is required, flying below the supersaturated region minimises the additional fuel use. Even when multiple evasive manoeuvres are needed, the additional trip fuel requirement is expected to be less than 0.5%.

Two unconventional micro air vehicles developed by the Naval Research Laboratory are described. One of the vehicles employs flapping wings which is inspired by the flight of birds or insects but does not copy it directly. The second vehicle is a stop-rotor hybrid vehicle employing a pair of single blade, rotary/fixed wing panels, attached at their roots to separate coaxial shafts. An unstructured grid based incompressible flow solver, called feflo, is used to simulate the flow past these novel configurations in order to determine the flight characteristics of these vehicles.

A transport equation for the intermittency factor is employed to predict transitional flows under the effects of pressure gradients, freestream turbulence intensities, Reynolds number variations, flow separation and reattachment, and unsteady wake-blade interactions representing diverse operating conditions encountered in low-pressure turbines. The intermittent behaviour of the transitional flows is taken into account and incorporated into computations by modifying the eddy viscosity, μτ with the intermittency factor, γ. Turbulent quantities are predicted by using Menter’s two-equation turbulence model (SST). The onset location of transition is obtained from correlations based on boundary-layer momentum thickness, accelaration parameter, and turbulence intensity. The intermittency factor is obtained from a transport model which can produce both the experimentally observed streamwise variation of intermittency and a realistic profile in the cross stream direction.

The intermittency transport model is tested and validated against several well documented low pressure turbine experiments ranging from flat plate cases to unsteady wake-blade interaction experiments. Overall, good agreement between the experimental data and computational results is obtained illustrating the predicting capabilities of the model and the current intermittency transport modelling approach for transitional flow simulations.

The integration of multidisciplinary data is key to supporting decisions during the development of aerospace products. Multidimensional metamodels can be automatically constructed using limited experimental or numerical data, including data from heterogeneous sources. Recent progress in multidimensional response surface technology, for example, provides the ability to interpolate between sparse data points in a multidimensional parameter space. These analytical representations act as surrogates that are based on and complement higher fidelity models and/or experiments, and can include technical data from multiple fidelity levels and multiple disciplines. Most importantly, these representations can be constructed on-the-fly and are cumulatively enriched as more data become available. The purpose of the present paper is to highlight applications of these cumulative global metamodels (CGM), their ease of construction, and the role they can play in aerospace integration. A simple metamodel implementation based on a radial basis function network is presented. This model generalises multidimensional data while simultaneously furnishing an estimate of the uncertainty on the prediction. Four examples are discussed. The first two illustrate the efficiency of surrogate-based design/optimisation. The third applies CGM concepts to a data fusion application. The last example is used to visualise extrapolation uncertainty, based on computational fluid dynamics data.

The present paper concerns the response of a following airplane to a pair of wing tip vortices left by a leading aircraft, represented by the Hallock-Burnham model, including the effect of vorticity decay between the two aircraft. The effect of vorticity is evaluated in terms of the induced rolling moment and also the lift loss; these specify the roll acceleration and the downward acceleration, respectively. The corresponding two response equations can be put into the same dimensionless form, and integrated using exponential integrals. This specifies the roll rate and sink rate as a function of time; besides the latter, the bank angle and altitude loss, are also plotted, all also as a function of time, for all combinations of leading and following aircraft in five classes. These are the three ICAO weight categories of light, medium and heavy, plus two other cases, viz the special case of the Boeing 757, which requires larger separations distances, and the case of a future very large transport aircraft (VLTA) exceeding significantly the size of a Boeing 747.

In this article, we investigate the horizontal trajectory tracking problem for an underactuated stratospheric airship subject to nonvanishing external disturbances and model uncertainties. By transforming the tracking errors into new virtual error variables, we can specify the transient and steady-state tracking performance of the resulting nonlinear system quantitatively, which means that under the proposed control scheme, the tracking errors will converge to prescribed residual sets around the origin before a preselected finite time with decay rates no less than a preassignable value. To address unknown items, minimal learning parameter (MLP) techniques for neural networks (NNs) approximation are employed, which efficaciously relax the computational burden, enhance the robustness against dynamics uncertainties and provide an improved property for disturbances rejection. A finite-time convergent observer (FTCO) is incorporated into the control framework to realise output-feedback control, ensuring that estimation errors are bounded during operation and approach zero within a finite time. Stability analysis proves that all the closed-loop signals are uniformly bounded. The effectiveness and advantages of the proposed control strategy are verified by simulation results.

There is a need for consistency between aviation safety targets – target levels of safety (TLS). Consistent ‘risk philosophy’ ensures that resources can be allocated in areas where they will be most beneficial in reducing the number of potential accidents and fatalities. Many existing aviation targets were devised decades ago, have not been brought up to date to take account of the considerable improvement in the sector’s safety performance, and are not targeted at some specific future year. It appears feasible to construct consistent and up-to-date sets of TLSs covering the full range of aviation needs, in particular for air traffic management. However, this requires significant, more demanding, changes to some TLS values. One caveat to note is that the safety of ‘loosely coupled’ aviation systems may be better described and managed by the Health and Safety Executive version of risk assessment rather than by a TLS approach.

Occurrence of transient starting and stopping loads during tests at high Mach numbers is one of the major problems in intermittent blowdown wind tunnels. It is believed that in order to overcome this problem, the wind tunnel could be started at a low Mach number and low stagnation pressure; the desired high Mach number condition could be reached by continuously changing the nozzle contour while synchronously increasing the stagnation pressure. After completing the tests, the nozzle could be brought back to the initial low Mach number accompanied by synchronous decrease in the stagnation pressure. In such a scenario, it is important to ensure that the pressure regulating valve (PRV) of the wind tunnel delivers and maintains a specified minimum stagnation pressure at any Mach number, so that supersonic breakdown of the test section flow does not occur. In this paper, the problem is formulated based on quasi-steady one-dimensional isentropic equations and numerically solved to predict the time histories of settling chamber pressure and storage tank pressure for a given trajectory of the opening of the PRV, as the Mach number is changed from Mach 1 to 4·0 continuously in four seconds and vice versa. The effects of rate of change of PRV open area and rate of change of Mach number on the stagnation pressure characteristics in the settling chamber and storage tank are predicted. The measured trajectories of the PRV in experiments in the NAL 0·6m transonic wind tunnel are used as input to the prediction program to validate the methodology. Predictions indicate that when the nozzle throat is changed from Mach 1 to 4 in four seconds, the settling chamber stagnation pressure rapidly builds up and approaches the pressure in the storage tank. Predictions show an alarming rise in free stream dynamic pressure during transition from Mach 1 to 4 and vice versa, which needs to be verified through measurements.

This work presents the Topology Optimisation of the Morphing Variable Span of Tapered Wing (MVSTW) using a finite element method. This topology optimisation aims to assess the feasibility of internal wing components such as ribs, spars and other structural components. This innovative approach is proposed for the telescopic mechanism of the MVSTW, which includes the sliding of the telescopically extended wing into the fixed wing segment. The optimisation is performed using the tools within ANSYS Mechanical, which allows the solving of topology optimisation problems. This study aims to minimise overall structural compliance and maximise stiffness to enhance structural performance, and thus to meet the structural integrity requirements of the MVSTW. The study evaluates the maximum displacements, stress and strain parameters of the optimised variable span morphing wing in comparison with those of the original wing. The optimised wing analyses are conducted on four wingspan extensions, that is, 0%, 25%, 50% and 75%, of the original wingspan, and for different flight speeds to include all flight phases (17, 34, 51 and 68m/s, respectively). Topology optimisation is carried out on the solid wing built with aluminium alloy 2024-T3 to distribute the wing components within the fixed and moving segments. The results show that the fixed and moving wing segments must be designed with two spar configurations, and seven ribs with their support elements in the high-strain area. The fixed and moving wing segments’ structural weight values were reduced to 16.3 and 10.3kg from 112 to 45kg, respectively. The optimised MVSTW was tested using different mechanical parameters such as strains, displacements and von Misses stresses. The results obtained from the optimised variable span morphing wing show the optimal mechanical behaviour and the structural wing integrity needed to achieve the multi-flight missions.

At the heart of a flight simulator resides the mathematical representation of aircraft behaviour in response to control inputs, atmospheric disturbances and system inputs including failures and malfunctions. While this mathematical model can never be wholly accurate, its fidelity, in comparison with real world behaviour, underpins the usefulness of the flight simulator. The present paper examines the state of the art achieved in validating mathematical models for helicopter simulators, addressing the strengths and weaknesses of the present European standard for the qualification of helicopter flight simulators, JAR FSTD-H (previously JAR-STD-1H/2H/3H). Essential questions are examined, such as: What is the required model fidelity to guarantee a simulation is sufficiently representative to be fit for purpose? Are the tolerances set in the current standards fine enough that they lead to only minor changes in handling qualities? What is an acceptable tuning process for the simulation? What is the effect of modelling fidelity on the overall pilot control strategy? What is the relationship between the settings of the simulator cueing environment and the behaviour of the pilot? What is the industrial experience on qualification of flight simulators that might usefully inform developments? Many of these questions were addressed in Europe in a previous GARTEUR Action Group (AG) HC/AG-12 the results of which are documented in this paper. Solutions are proposed for improving the current JAR-FSTD standard with respect to validation of mathematical models.

The increasing use of unmanned aerial vehicles in areas such as rescue, mapping, and transportation have made it necessary to study more accurate techniques for calculating flight time estimates. Such calculations require knowing the battery discharge profile. Simplified flight time calculation methods provide data with uncertainties as they are based solely on manufacturer datasheet information. This study presents a setup to measure the battery discharge curve using a LabVIEW interface with a low-cost acquisition system. The acquired data passes through a nonlinear optimisation algorithm to find the battery coefficients, which enables the more precise estimation of its range and endurance. The great advantage of this model is that it makes it possible to predict how the battery will discharge at different rates using just one experimental curve. The methodology was applied to three different batteries and the model was validated with different discharge rates in a controlled environment, which resulted in endurance lower than 3.0% for most conditions and voltage estimation error lower than 3.0% in operational voltage. The work also presented a methodology for estimating cruise time based on the current used during each flight stage.

This paper compares two periodic control methods, the optimal H 2 and the periodic static output feedback (POF), to reduce the helicopter rotor vibrations. Actively twisted blades with Macro-Fibre Composite (MFC) piezoelectric actuators are used. The design model is based on a simplified aerodynamic model and on a multi-body model of the Bo 105 isolated rotor with the original blades replaced by actively twisted ones. The performance of the two controllers in alleviating hub loads is verified with improved simulations based on a free-wake model.