A numerical procedure, which utilises polynomial power series expansions for the optimisation of multipanel wing structures in idealised critical flutter conditions, is introduced and developed. It arises from the Rayleigh-Ritz method and employes trial polynomial describing functions both for the flexural displacement and for the thickness variation over the multipanel surface. An idealised structural plate model, according to the Kirchhoff’s theory, together with a linearised supersonic aerodynamic approach, are supposed. The classical Euler-Lagrange optimality criterion, based on variational principles, has been utilised for the optimisation operations, where by imposing the stationary conditions of the Lagrangian functional expression, a nonlinear algebraic equations system is obtained, whose solution is found by an appropriate algorithm. By utilising an iterative process it is possible to reach the reference structure critical conditions, with an optimised thickness distribution throughout the multipanel surface. The final part of the work consists in searching the minimum weight of the multipanel planform wing structure with optimised thickness profile vs the flutter frequency, considered as a variable imput parameter, for fixed flutter speed and equal to the critical one of the reference uniform structure.

In a gas turbine, ingestion of hot gas into the high-pressure turbine disc cavities could cause metal overheat. To prevent this, cool air is taken from the compressor and ejected through the cavities. However, this sealing flow also reduces the overall efficiency, and a compromise has to be found between the level of ingestion tolerated and the losses. Recent advances made in applying Computational Fluid Dynamics to such configurations are presented, with the aim of better understanding the physical phenomena and providing reliable design tools. First, results showing the pumping effect of the rotating disc are presented, including the influence of flow instabilities observed in both computational and experimental results. Second, the influence of the main annulus pressure asymmetries are analysed on a simplified representation of an available experiment, showing the combined influence of asymmetries generated by vanes and struts. Finally, a rim seal geometry representative of aero-engine design is studied in comparison to experiment, exhibiting the coupled influence of the cavity instabilities and annulus asymmetries.

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.

Mean velocity and first order turbulence measurements were obtained from a three-dimensional upwash fountain flow generated by the impingement of two compressible axisymmetric turbulent jets onto a normal plane. The jet impingement area and fountain formation regions were examined with data obtained through the use of particle image velocimetry. Seven configurations with different nozzle pressure ratios were considered to ascertain the influence of jet compressibility on the fountain development. Results indicate that the mixing of the fountain is dependent on the nozzle pressure ratio, leading to an increase in the fountain spreading rate with increase in nozzle pressure ratio.

In this paper the development and flight testing of flapping-wing propelled, radio-controlled micro air vehicles are described. The unconventional vehicles consist of a low aspect ratio fixed-wing with a trailing pair of higher aspect ratio flapping wings which flap in counterphase. The symmetric flapping-wing pair provides a mechanically and aerodynamically balanced platform, increases efficiency by emulating flight in ground effect, and suppresses stall over the main wing by entraining flow. The models weigh as little as 11g, with a 23cm span and 18cm length and will fly for about 20 minutes on a rechargeable battery. Stable flight at speeds between 2 and 5ms–1 has been demonstrated, and the models are essentially stall-proof while under power. The static-thrust figure of merit for the device is 60% higher than propellers with a similar scale and disk loading.

Thermopile sensors detect electromagnetic radiation as a function of the object’s temperature. Because there is a temperature difference between the cold ground and the warm sky, these sensors could be used to detect the horizon and thus be used as a reference to stabilise a small aircraft, such as an unmanned aerial vehicle (UAV) in visual meteorological conditions (VMC). To verify this hypothesis, a system has been developed providing horizon detection using thermopile sensors to stabilise an R/C model aircraft. The aircraft has gone through a number of flight trials using remote control to enable and disable the stabilisation system. During the flight trials the aircraft was set at various attitudes when the system was enabled. The stabilisation system was able to assume wing level under various bank angles and weather conditions with minimum overshoot and oscillation.

Although the system shows good performance during flight trials, most of the original design was done using trial and error. A design tool was needed to implement further improvements to the system and to efficiently implement it on other aircraft. This required a good understanding of the physical behaviour of the system and the interaction between the sensors, aircraft and the environment. A mathematical model of the overall system was developed using the MATLAB/Simulink environment to simulate the behaviour of the system under various conditions. The simulation results were then compared with actual flight experiments. This paper describes the modelling techniques used for the different system components and the results of the simulation compared to actual flight trials.

The results are presented for a cycle of experimental investigations of flow field characteristics (static pressure distribution, static pressure fluctuations, upwash, boundary-layer parameters) in the perforated test section of the transonic TsAGI T-128 Wind Tunnel. The investigations concern the effect of nozzle shape, wall open-area ratio, Mach and Reynolds numbers on the above-outlined flow characteristics. During the tests, the main Wind-tunnel drive power is measured. Optimal parameters of the nozzle shape and test section perforation are obtained to minimise acoustic perturbations in the test section and their non-uniformity in frequency, static pressure field non-uniformity, nozzle and test section drag and, accordingly, required main Wind-tunnel drive power.

An investigation into the behaviour of clustered synthetic jet Actuators for flow-control applications is described. Experiments have been undertaken with two small-scale synthetic jet actuators in a zero-pressure gradient boundary-layer, in order to investigate the effect of configuration yaw angle and relative input signal phase. Oil-flow visualisation and hotwire anemometry techniques were used, demonstrating that changes in the downstream flow structure could be observed. Compared to a streamwise configuration, in which a symmetrical counter-rotating vortex pair was produced by the synthetic jet-boundary-layer interaction, a broader asymmetric interaction was produced in a 15° yaw configuration. Streamwise velocity contour plots, illustrating the development of the interaction downstream, over four phase angles, were presented. Significant differences in the PSD analyses of downstream streamwise velocity time histories were found, suggesting that input signal phase could influence the stability and hence effectiveness of flow structures used in flow-control applications.

The aviation community is currently faced with various approaches for the determination of Target Levels of Safety (TLS). The targets are usually derived for a specific airspace region, for a specific type of operation or a specific phase of flight. Therefore, current practices support the determination of TLS for specific aviation components in isolation. This paper argues that the setting of a new safety target for aviation has to be driven by an integrated system approach. Relevant past research on TLS is reviewed and augmented with the results from operational reports from two countries. A possible safety target for the year 2020 is suggested and scoped down to estimate the safety budget for air traffic control (ATC) equipment. The paper concludes with a discussion of the results and recommends useful practices to achieve the proposed integrated safety approach.

This paper presents a new approach to vectoring jet thrust using a miniature fluidic actuator that provided spatially distributed mass addition. The fluidic actuators used had no moving parts and produced oscillatory flow with a square wave form at frequencies up to 1·6kHz. A subsonic jet with an exit diameter of 3·81cm was controlled using single and dual fluidic actuators, each with an equivalent circular diameter of 1·06mm. The fluidic nozzle was operated at pressures between 20·68 and 165·47kPa. The objectives of the present work included documentation of the actuation characteristics of fluidic devices, assessment of the effectiveness of fluidic devices for jet thrust vectoring, and evaluation of mass flow requirements for vectoring under various conditions. Measurements were made in the flow field using a pitot probe for the vectored and unvectored cases. Some acoustic measurements were made using microphones in the near-field and for selected cases particle image velocimetry (PIV) measurements were made. Thrust vectoring was obtained in low speed jets by momentum effects with fluidic device mass flow rates of only 2 × 10–4kg/sec (0·6% of main jet mass flow per fluidic oscillator). Although a single fluidic device produced vectoring of the primary jet, the dual fluidic device configuration (with two fluidic devices on either side of the jet exit) produced mass flux enhancement of 28% with no vectoring. Our results indicate that fluidic actuators have the potential for use in thrust vectoring, flow mixing and industrial flow deflection applications.

The value of computational fluid dynamics, CFD, delivered to date has mainly been related to its application to high-speed cruise design. To increase its applicability CFD must apply to the full flight envelope frequently characterised by large regions of separated flows. These flows are encountered by transport aircraft at low speed with deployed high lift devices, at their structural design loads conditions, or subjected to in-flight upsets that expose them to speed and/or angle-of-attack conditions outside the envelope of normal flight conditions to name a few. Such flows can only be characterised by the Navier-Stokes equations. This paper will report the progress toward CFD for full flight envelope. The CFD methods in use at Boeing will be described. Examples presented will address high-lift, loads and stability and control concerns including Reynolds scaling from wind tunnel to flight, vortex generator simulation, spoiler and horizontal tail effectiveness. In general, results shown are in ‘good enough’ agreement with experimental data. Deficiencies and the need for further algorithm and process improvement are noted. The need for automation to enable the large scale use of CFD will also be discussed.

The development of the non-linear equations of motion for the hang glider from first principles is described, including the complex geometry of control by pilot ‘weight shift’. By making appropriate assumptions the linearised small perturbation equations are derived for the purposes of stability and control analysis. The mathematical development shows that control is effected not by pilot weight shift, but by centre of gravity shift and that lateral-directional control by this means is weak, and is accompanied by significant instantaneous adverse response.

The development of a comprehensive semi-empirical mathematical model of the flexible wing aerodynamics is described. In particular, the modelling attempts to quantify camber and twist dependencies. The performance of the model is shown to compare satisfactorily with measured hang glider wing data obtained in earlier full scale experiments. The mathematical aerodynamic model is then used to estimate the hang glider stability and control derivatives over the speed envelope for substitution into the linearised equations of motion.

Solution of the equations of motion is illustrated and the flight dynamics of the typical hang glider are described. In particular, the dynamic stability properties are very similar to those of a conventional aeroplane, but the predicted lateral directional stability margins are significantly larger. The depth of mathematical modelling employed enables the differences to be explained satisfactorily. The unique control properties of the hang glider are described in some detail. Pitch and roll control of the hang glider is an aerodynamic phenomenon and results from the pilot adjusting his position relative to the wing in order to generate out of trim aerodynamic control moments about the centre of gravity. Maximum control moments are limited by hang glider geometry which is dependent on the length of the pilot‘s arm. The pilot does not generate control moments directly by shifting his weight relative to the wing. The modelling thus described would seem to give a plausible description of the flight dynamics of the hang glider.

The ever expanding range of applications for unmanned air vehicles covers a wide variety of powerplant requirements. While many vehicles are high value assets and demand similar powerplant attributes to a conventional manned aircraft, there is an emerging requirement for powerplants which are designed to a limited life philosophy. This covers a range of applications: at the simplest level there are vehicles where the mission definition dictates that attrition rates will be high, for example low level reconnaissance, targeting or battle damage assessment, and the entire vehicle may only be expected to make a short number of missions. On the other hand there are applications where it may be more cost effective to remove the need for engine maintenance and to treat the powerplant as a disposable limited life unit.

In general, the industry trend is towards the reduction of life cycle costs, and a significant emphasis is placed upon continuing to increase aero-engine component lives though novel design and improved manufacturing techniques. A limited life application calls for a different approach, as it may be that the powerplant has a life requirement of anywhere between 50 and 250 hours. This allows scope to explore areas of design and manufacture which would not be suitable for a ‘conventional’ engine and can lead to large reductions in engine cost.

This paper seeks to identify the primary cost drivers for a small UAV powerplant in the 10kN thrust class, and to establish the relationships between cost, life and performance at both a whole engine and an individual component level. It also explores the feasibility of using new or unconventional designs and manufacturing methods, and presents a selection of examples.

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.

This paper looks at active control of the normal shock wave/turbulent boundary layer interaction (SBLI) using smart flap actuators. The actuators are manufactured by bonding piezoelectric material to an inert substrate to control the bleed/suction rate through a plenum chamber. The cavity allows rapid thickening of the boundary-layer approaching the shock, which splits into a series of weaker shocks forming a lambda shock foot, thus reducing wave drag. Active control allows optimisation of the interaction, as it would be capable of either positioning the control region around the original shock position using a series of unimorph flaps or fixing the shock position by controlling the rate of mass transfer.

The level of control achieved by unimorph piezoelectric actuators is not large because of small amounts of deflection possible. It is believed that to provide optimal control a piezoelectric material, which can provide greater strain and hence higher amounts of deflection is needed. However, currently such a piezoelectric material is not commercially available.

Theoretical quantification of viscous effects in fluid flows is difficult, even if turbulence is absent, except when it is legitimate to simplify the Navier-Stokes equations in some way; for example by invoking the boundary-layer approximation in appropriate cases of interacting viscous and inviscid flow. The technical importance of viscous effects was thought sufficient incentive to re-examine a very simple flow configuration — namely plane, uniform and steady flow of an incompressible, viscous fluid toward a vanishingly-thin flat plate aligned with the undisturbed stream — in search of fresh insights into the general theory for viscous-inviscid interactions.

The strategy was to exploit the analogy between vorticity transport in a viscous fluid and heat conduction in a moving solid. The key to doing so was the realization that, if the perturbation of the undisturbed flow by the plate might be represented as the sum of a series of successive approximations, then the stream function of the viscous part of the flow field — not merely the vorticity which resulted from its existence — might be expressible at every stage as the solution of an analogous heat conduction problem.

Dye flow visualisation of circular synthetic jets was carried out in laminar boundary layers developing over a flat plate at a range of actuator operating conditions and freestream velocities of 0·05 and 0·1ms–1. The purpose of this work was to study the interaction of synthetic jets with the boundary layer and the nature of vortical structures produced as a result of this interaction. The effects of Reynolds number (Re), velocity ratio (VR) and Strouhal number (St) on the behaviour of synthetic jets were studied. At low Re and VR, the vortical structures produced by synthetic jets appear as highly stretched hairpin vortices attached to the wall. At intermediate Re and VR, these structures roll up into vortex rings which experience a considerable amount of tilting and stretching as they enter the boundary layer. These vortex rings will eventually propagate outside the boundary layer hence the influence of the synthetic jets on the near wall flow will be confined in the near field of the jet exit. At high Re and VR, the vortex rings appear to experience a certain amount of tilting but no obvious stretching. They penetrate the edge of the boundary layer quickly, producing very limited impact on the near wall flow. Hence it is believed that the hairpin vortices produced at low Re and VR are likely to be the desirable structures for effective flow separation control. In this paper, a vortex model was also described to explain the mechanism of vortex tilting.

This article deals with a damage simulation of flight control systems during a flight. As a subject of flight tests the model of a classic aerodynamic aircraft scheme is considered. This aircraft has highly swept wings a moderate wing elongation and all-moving horizontal tail surfaces. The diagrams of flight parameters from flight tests of the freely flying models with loss of a control surface in flight are given. The investigation results for the failures leading to a sharp drop in the stiffness of the control units in the pitch channel are given. The damage investigation in horizontal flight and in complex manoeuvres is given.

An algorithm employing adaptive neuro-fuzzy online identification and sequential quadratic programming optimisation techniques is developed to enhance aircraft engine performance. This algorithm is implemented and tested using digital simulation for two spool, mixed exhaust, variable geometry turbofan engine. Parametric study is conducted to select the proper measurable parameter that can represent the actual thrust during online optimisation. Subtractive clustering technique is applied to generate the minimum number of fuzzy rules that can model the engine performance at various input parameters and flight conditions. The resulting neuro-fuzzy system is then optimised through training algorithm to accurately represent the engine. This system can address engine variations by relearning the network using online measurements from the actual engine. Generating the optimum schedules and comparing them with those obtained from the complete non-linear engine model verify the algorithm. Benefits from this algorithm include fuel consumption savings, reductions in turbine inlet temperature, and preventing limit exceeding.