This paper focuses on the three dimensional flight path planning for an unmanned aerial vehicle (UAV) on a low altitude terrain following\terrain avoidance mission. The UAV trajectory planning problem is to compute an optimal or near-optimal trajectory for a UAV to do its mission objectives in a surviving penetration through the hostile enemy environment, considering the shape of the earth and the kinematics constraints of the UAV. Using the three dimensional terrain information, three dimensional flight path from a starting point to an end point, minimising a cost function and regarding the kinematics constraints of the UAV is calculated. The geographic information of the earth shape and enemy locations is generated using digital terrain model (DTM) and geographic information system (GIS), and is displayed in a 3D environment. Using 3D-maps containing the geographic data accompanied by DTM, and GIS, the problem is modelled by deriving the motion equations of the UAV. Two heuristic algorithms are proposed for this problem: genetic and particle swarm algorithms. Genetic and particle swarm algorithms are general purposes algorithms, because they can solve a wide range of problems, so they have to be adjusted to solve the trajectory planning problem. To test and compare the paths obtained from these algorithms, a software program is built using GIS tools and the programming languages C# and MATLAB.

Numerical analysis of the flow in weapon bays modelled as open rectangular cavities of length-to-depth (L/D) ratio of 5 and width-to-depth (W/D) ratio of 1 with doors-on and doors-off is presented. Flow conditions correspond to Mach and Reynolds numbers (based on cavity length) of 0·85 and 6·783m respectively. Results from unsteady Reynolds-averaged Navier-Stokes (URANS), large-eddy simulation (LES) and detached-eddy simulation (DES) are compared with the simulation methods demonstrating the best prediction of this complex flow. It was found that URANS was not able to predict the change of flow characteristics between the doors-on and doors-off configurations. In addition, the energy content of the cavity flow modes was much better resolved with DES and LES. Further, the DES was found to be quite capable for this problem giving accurate results (within 3dB of) experiments and appears to be a promising alternative to LES for modelling massively separated flows.

In moving from the realm of stationary wing flight towards the complexity of flapping wing flight as evolved in birds, we wish to introduce the simplest means of flapping as a first step in an attempt to simplify as much as possible our wind tunnel model and also to confine our attention to the dominant thrust producing mechanism. Our model half wings flap only in the plane normal to flight. The flapping motion is sinusoidal with time and symmetrical about the mid position (Fig. 6). No wing bending was allowed. The wing section shape had no camber and also was rigid in shape but free to twist in torsion about its shear centre. However, the model root support was fixed and therefore could not simulate vertical plunging of a bird associated with ideal elliptic span loading (Ref. 1).

Recent advances in coaxial rotor design have shown benefits of this configuration. Nevertheless, issues related to rotor-head drag, aerodynamic performance, wake interference, and vibration should also be considered. Simulating the unsteady aerodynamic loads for a coaxial rotor, including the aerodynamic interactions between rotors and rotor blades, is an essential part of analysing their vibration characteristics. In this article, an unsteady aerodynamic analysis based on a vortex particle method is presented. In this method, a reversed-flow model for the retreating side of the coaxial rotor is proposed based on an unsteady panel technique. To account for reversed flow, shedding a vortex from the leading edge is used rather than from the trailing edge. Moreover, vortex-blade aerodynamic interactions are accounted for. The model considers the unsteady pressure term induced on a blade by tip vortices of other blades, and thus accounts for the aerodynamic interaction between the rotors and its contribution to the unsteady airloads. Coupling the reversed-flow model and the vortex-blade aerodynamic interaction model with the viscous vortex-particle method is used to simulate the complex wake of the coaxial rotor. The unsteady aerodynamic loads on the X2 coaxial rotor are simulated in forward flight, and compared with the results of PRASADUM (Parallelized Rotorcraft Analysis for Simulation And Design, developed at the University of Maryland) and CFD/CSD computations with the OVERFLOW and the CREATE-AV Helios tools. The results of the present method agree with the results of the CFD/CSD method, and compare to it better than the PRASADUM solutions. Furthermore, the influence of the aerodynamic interaction between the coaxial rotors on the unsteady airloads, frequency, wake structure, induced flow, and force distributions are analysed. Additionally, the results are also compared against computations for a single-rotor case, simulated at similar conditions as the coaxial rotor. It is shown that the effect of tip vortex interaction plays a significant role in unsteady airloads of coaxial rotors at low speeds, while the rotor blade passing effect is obviously strengthened at high-speed.

In commercially available gas turbine performance simulation tools, individual engine components are typically represented with non-dimensional maps of experimental or default data. In those cases where actual component characteristics are not available and default characteristics are used instead, conventional tools can deviate substantially at off-design and transient conditions. Similarly, when real component characteristics are available, conventional engine cycle simulation tools can not predict the performance of the engine at other than nominal conditions satisfactorily, or account for the impact of changes in component geometry.

This study looked into the full integration of two-dimensional streamline curvature component models with a low fidelity cycle program. Firstly, the obtained engine performance was compared against the one calculated based on default component characteristics. As a second case study, a range of flight Mach numbers and angles of attack were examined together with the effect of three different intake lip geometries on the performance of a notional, two-spool, low-bypass ratio, military engine. Two-dimensional models were used in the engine cycle analysis to provide a more accurate, physics- and geometry-based estimate of intake and fan performances.

The analysis carried out by this study demonstrated relative changes in the predicted engine performance larger than 1%. For briefness, representative results are presented and discussed in this paper for one flight Mach number and angle of attack setting. More importantly, this research effort established the necessary methodology and technology required towards a full, two-dimensional engine cycle analysis at an affordable computational resource in the very short term.

The motion of liquid in the containers of a vehicle may have severe influence upon the stability of the vehicle. For this reason many authors have investigated the behaviour of liquid with a free surface due to various modes of excitation. In the earlier development of small vehicles, the frequencies of the elastic structures were relatively high compared to the propellant frequencies and the effect of their interaction could be neglected. With the increase of modern space vehicles this interaction of the propellant with the elastic structure is becoming a pronounced influence upon the design of the vehicle and its control system. This has some influence upon the stability and response behaviour of a large space vehicle, since the close grouping of propellant sloshing and structural frequencies around the control frequency of the vehicle creates continuous interaction during flight.

In this paper a non-linear theory of the supersonic bangs from an accelerating body is worked out, using an extension of Whitham’s method. The strength of the bow shock is determined, and it is shown that the effect of acceleration on the pressure rise is appreciable at large distances from the nose for Mach numbers near unity. The results are compared with those of linear theory; the estimate for the pressure rise is found to be about a half of the linear value. The results also include an expression for the curvature of an attached shock at the nose, and a formula for the decay of a shock after it is detached from a decelerating body.

A novel, non-intrusive, high spatial and temporal resolution thermal imaging system, based on the fluorescence properties of a dysprosium doped yttrium-aluminum-garnet (Dy3+:YAG) thermographic phosphor, has been developed. In this system, the phosphor coating on the test surface is excited by a pulsed Nd:YAG laser. The resulting fluorescent emission of the temperature sensitive 456nm transition and that of the temperature independent 496nm transition, are acquired using a pair of image-intensified charged coupled device cameras. The ratio of the acquired emissions is then correlated to temperature. The Dy:YAG phosphor displays temperature sensitivity in the range of 295–l,350K (with possibility up to 1,800K). Temperature measurements obtained from the surface of a disk heated by an impinging jet, demonstrate the feasibility of this technique for general heat transfer studies involving significant unsteadiness and transient phenomena.

This paper deals with the force coefficients on the leeward of two bluff bodies, in particular with the origin and nature of the lift force. Initially a general description of the characteristics of the force coefficients, plus the effect of parameters such as free-stream turbulence and surface roughness in the form of stranding, is given.

A resume of all the known suggestions for the origin of the lift force is listed and a series of experiments, attempting to test these suggestions using simple methods, is described. These methods consist mainly of force measurement, involving force measurement on the leeward conductor using different shapes for the leeward and windward bodies. However, static pressure tapping and flow visualisation are also employed to a small degree.

It is shown that none of the suggestions gives a satisfactory explanation for the origin of the lift force. However, several interesting points were noted; for instance the dependence of the lift profiles for the leeward body on its own characteristics in the wake and not particularly on the wake characteristics. Hence it is shown that the use of wake parameters, measured without the presence of the leeward body when attempting to assess the forces thereon, is an over-simplification of the situation as far as lift is concerned while appearing to work quite well for drag.

Research is at present being carried out at the Turin Polytechnic University with the aim of designing an HAVE/UAV (high altitude very-long endurance/unmanned air vehicle). The vehicle should climb to 17-20km by mainly taking advantage of direct Sun radiation and thereafter maintain a level flight; during the night, a fuel cells energy storage system would be used. A computer program has been developed to carry out a parametric study for the platform design. The solar radiation change over one year, the altitude, masses and efficiencies of the solar and fuel cells, and the aerodynamic performances have all been taken into account. The parametric studies have shown how fuel cells and solar cells efficiency and mass have the most influence on the platform dimensions. A wide use of high modulus CFRP has been made in designing the structure in order to minimise the airframe weight. A first configuration of HELIPLAT® (HELIos PLATform) was worked out, following a preliminary parametric study. The platform is a monoplane with eight brushless electric motors, a twin-boom tail type with an oversized horizontal stabiliser and two rudders. The co-ordinates at the root and along the wing span as well as the wing planform were optimised to achieve the best efficiency. Several profiles and wing plans have been analysed using the CFD software Xfoil and Vsaero. Several wind-tunnel tests were carried out to compare the analytically predicted performances. A preliminary design of a scale-sized technological demonstrator was completed with the aim of manufacturing a proof-of-concept structure. A FEM analysis was carried by using the Msc/Patran/Nastran code to predict the static and dynamic behaviour of the UAV structure.

In recent papers Messiter and Hida have proposed a first-order correction to simple Newtonian theory for the pressure distributions on the lower surfaces of lifting conical bodies with detached shocks. The method involves the solution of an integral equation which Messiter solved numerically for thin delta wings, while Hida gave an approximate solution for thick wings with diamond and bi-convex cross-sections. It is shown in the present paper that Hida’s approximate solutions give poor agreement with experiment, and a series of more precise numerical solutions of the equation are given for wings with diamond cross-sections. The pressures, and shock shapes, obtained from these solutions are in very good agreement with experiment at Mach numbers as low as 4·0.

The method has also been extended to Nonweiler wings at off-design when the shock wave is detached from the leading edges. Again the agreement with experiment is good provided the integral equation is solved numerically.

In this paper, the multi-body dynamic model of an asymmetric variable sweep wing morphing UAV is built based on Kane’s method. This model describes the UAV’s transient behaviour during morphing process and the dynamic characteristic of the variable sweep wings. An integrated design of trajectory tracking control via constrained backstepping method is presented then. The idea of aircraft roll control through asymmetric wing sweep angle changes rather than traditional aileron is explored and used in the fight control design. The control of variable sweep wings is designed as well based on the presented dynamic model. Command filters are used in the backstepping design procedure to accommodate magnitude, rate and bandwidth constraints on virtual states and actuator signals. Stability of the closed-loop system can be proved in the sense of Lyapunov. Simulation of tracking a desired trajectory which contains two manoeuvres demonstrates the feasibility of the proposed protocol and the morphing wing roll controller.

A method is described for calculating wall interference in solid-wall wind tunnels from measurements of static pressures at the walls. Since it does not require a simulation of the model flow, the technique is particularly suited to determining wall interference for complex flows such as those over VSTOL aircraft, helicopters and bluff shapes (e.g. cars and trucks). An experimental evaluation shows that the method gives wall-induced velocities which are in good agreement with those of existing methods in cases where these techniques are valid, and illustrates its effectiveness for inclined jets which are not readily modelled.

Studies begun over ten years ago have shown that the swirling air flow about the tips of a lifting wing can be redirected using small aerofoil surfaces, called sails, to reduce the overall lift-dependent drag of the wing. Tests on Paris, Jetstream and other windtunnel models, confirmed by flight tests on Paris and Centurion aircraft have shown marked reductions in drag and fuel consumption. Handling tests show improvements in roll control and stalling speeds.

Wing tip sails have the effect of reducing the peak adverse effects on other aircraft flying through the wake of a Paris aircraft but reduce the decay rate so the far field effects are worsened. Sails fitted to a crop-spraying aircraft have significantly reduced the amount of spray lifted above the height of the aircraft by the tip vortices and subsequently blown off-target by a cross-wind.

Windtunnel tests on a model with variable incidence sails show that differential sail incidence can be used effectively to provide roll power, reducing the size of the ailerons required, so allowing a greater flap span. The maximum root bending moments of the wings can be reduced by at least 10% by using negative sail movements in positive manoeuvres or gusts.

Sails can reduce the drag of an aircraft at all speeds leading to fuel savings estimated at US$225 per annum per 1000lb take-off weight per 1000 hours annual utilisation. For an airliner of 300 000lb weight and operating for 3000 hours this means a fuel saving of over a fifth of a million dollars per year!

The development of new technologies – such as rapid prototyping – and the use of materials with improved properties – such as highly resistant extruded polystyrene foam which can be easily and precisely shaped, while conserving its mechanical properties – allow researchers to improve design concepts. This article details the development of a new set of morphing wings for a 15-kg maximum take-off weight Unmanned Aerial Vehicle (UAV) from concept design, to flight tests, including modelling, design optimisation, construction and wind-tunnel tests. A set of comparator-equivalent conventional wings have been used throughout in order to be able to judge any benefits stemming from the adoption of morphing technology. This article shows that the morphing wings provide a controllable aircraft while reducing drag by a factor of 40% compared to the comparator wings with conventional ailerons in a deflected position.

Model scale tests of modern ‘open rotor’ propulsor concepts that have potential for significant fuel burn reduction for aircraft applications were completed at NASA Glenn Research Center. The recent test campaign was a collaboration between NASA, FAA, and General Electric (GE). GE was the primary industrial partner, but other organisations were involved such as Boeing and Airbus who provided additional hardware for fuselage simulations. The open rotor is a modern version of the UnDucted Fan (UDF®) that was flight tested in the late 1980s through a partnership between NASA and GE. The UDF® was memorable for its scimitar shaped propeller blades and its unique noise signature. Design methods of the time were not able to optimise for both high aerodynamic efficiency and low noise simultaneously. Contemporary CFD/CAA based design methods can produce open rotor blade designs that maintain efficiency with acceptable acoustic signatures. Tests of two generations of new open rotor designs were conducted in the 9’ × 15’ Low Speed Wind Tunnel and the 8’ × 6’ Supersonic Wind Tunnel starting in late 2009 and completed in early 2012. Aerodynamic performance and acoustic data were obtained for take-off, approach and cruise conditions in isolated and semi-installed configurations. Additional detailed flow diagnostic measurements and acoustic measurements, including canonical shielding configurations, were obtained by NASA. NASA and GE conducted joint systems analysis to evaluate the performance of the new blade designs on a Boeing 737 class aircraft. The program demonstrated a 2-3% improvement in overall net efficiency relative to the best efficiency designs of the 1980s while nominally achieving 15-17 EPNdB noise margin to Chapter 4 (at a Technology Readiness Level of 5) for a notional aircraft system defined by NASA.

The stability of a two-dimensional wall jet was studied theoretically and experimentally. As a result of the linear stability calculation, it was found that one eigenmode is separated into two modes when the Reynolds number is large, and inside a neutral stable curve in the α, R-plane there exists another neutral curve enclosing a stable region. Experimental results of small disturbances were compared with calculated results; agreement between them was satisfactory. It was found, further, that subharmonics of a predominant disturbance velocity component appear in the non-linear region.

Conditions of thermal fatigue encountered in the Aircraft Industry are usually located in engine components subjected to severe temperature and stress cycling. These components include gas turbine blading, nozzle guide vanes, flame tubes and turbine discs fabricated from a number of different materials ranging from Nimonics to sheet iron— and cobalt-base alloys. In recent years considerable experimental work has been carried out to assess material behaviour under thermal fatigue conditions and several means of testing have emerged. Rapid heating and cooling cycles using a “fluidised bed” technique have been extensively applied by Glenny and co-workers, at NGTE, and by material manufacturers to aid alloy development. Other experimenters have used axial straining devices based on expansion and contraction of heated columns followed by air cooling. One aircraft engine manufacturer has developed a thermal fatigue rig incorporating a repeated tensile axial load and temperature cycle. A further means of assessment has been developed by Forrest at NPL, in which the test specimen is strain cycled under reverse-bend conditions at constant or fluctuating temperatures.

An investigation has been carried out to examine one aspect of supersonic jet-engine inlet flow, namely shock-wave and boundary-layer interaction in the presence of an expansion corner. The experiment was undertaken at Mach numbers of 1.8 and 2.5. The local Reynolds number per metre varied from 2.79 to 3.13 × 107 for M = 1.8 and from 4.20 to 6.4-8 × 107 for M = 2.5. Adiabatic turbulent boundary layers of about 6 mm thick on an expansion-corner plate were disturbed at different streamwise positions by oblique shocks with deflection angles of 4°, 6° and 8°. The pressure distributions, wave patterns and the extent of the separation regions were strongly influenced by the presence of the expansion corner, especially when the shock impingement positions were upstream of the corner. The pressure rise induced by the incident shock began to decrease in level when the shock impingement position was about 3-4 boundary-layer thicknesses upstream of the corner. As the incident shock moved towards the corner, the pressure rise and the extent of the separation region decreased. When the shock impingement position was at the corner, the reflected shock was effectively neutralised by the Prandtl-Meyer expansion; especially when the deflection angle of incident shock matched that of the expansion corner. The three dimensionality of the incident shock and separation region is due to the interactions between the glancing oblique incident shock and the tunnel’s sidewall boundary layers.

An experimental study of supersonic flow over wall mounted cavities with different aft wall angles is carried out. Unsteady pressure measurements were made on the walls and floor of the cavity. Data analysis was performed on the experimental results using statistical methods. In the case of higher angled cavities, the presence of an upstream traveling acoustic wave could be confirmed. For lower angled cavities (60 degrees and less) where the acoustic wave could not be identified, the flow inside the cavity was more or less stable. Mode switching occurring in higher angled cavities was confirmed by spectrogram studies.