Since the introduction of hexapod-type motion systems for flight simulation in the 1970s, Motion Drive Algorithm tuning has been primarily based on the subjective judgement of experienced pilots. This subjective method is often not transparent and often leads to ambiguous process of adjustment of the tuning parameters. Consequently, there are large variations in the motion cueing characteristics of flight training devices, a variability that subsequently raises questions regarding the value of motion cueing for pilot training itself. The third revision of ICAO 9625 Manual of Criteria for the Qualification of Flight Simulation Training Devices offered the opportunity to take a closer look at simulator motion cueing requirements in general. This led to the concept of the objective motion cueing test (OMCT), which was reported in 2006. After the method was evaluated on three research flight simulators, the results were published in 2007, demonstrating a larger spread in dynamic behaviour of cueing algorithms than expected. After discussions with the simulator industry regarding the form and methodology of the OMCT, an evaluation of the test in cooperation with the industry started in 2011. This led to the final form of the OMCT and cueing parameter criterion for the in-flight mode of transport aircraft. This paper describes the OMCT, the evaluation results and the criterion.

When the stagnation temperature of a perfect gas increases, the specific heats and their ratio do not remain constant and start to vary with the temperature. The gas remains perfect; its state equations remain valid, so it can be named as calorifically imperfect gas. The aim of this research is to develop the necessary thermodynamic and geometrical equations and to study the supersonic flow at high temperature, lower than the dissociation threshold. The results are found by the resolution of nonlinear algebraic equations and integration of complex analytical functions where the exact calculation is impossible. The dichotomy method is used to solve the nonlinear equations and Simpson’s algorithm for the numerical integration applied. A condensation of the nodes is used. The functions to be integrated have a high gradient at the extremity of the interval of integration. The comparison is made with the calorifically perfect gas to determine the error. The application is made for air in a supersonic nozzle.

This paper reviews findings on the type, morphology and constitutive behaviour of impact damage zones during loading after impact and their effect on the laminate strength and stability. The paper is limited to tape prepreg based monolithic laminates, although some similarities exist with impact damage in textile based laminates. Damage zones have a complex geometry with several damage types, which results in an interaction and competition between different failure mechanisms, e.g. local and global buckling, compressive failure, and delamination growth. Hence, simplified damage models may provide incorrect predictions of the failure load and failure mechanisms after impact. The constitutive behaviour of damage zones has been studied experimentally in tension and compression using an inverse method, and the results have been compared with detailed FE models of a generic impact damage. The paper is concluded with a discussion on analytical and computational models to predict the resulting strength of impacted laminates.

Inlet unstart boundary is one of the most important issues of the hypersonic inlet and is also the foundation of the protection control of a scramjet. To solve this problem, the 2D internal steady flow of a 2D mixed internal/external compression hypersonic inlet was numerically simulated at different freestream conditions and backpressures with a RANS (Reynolds-Averaged Navier-Stokes) solver using a RNG (Renormalisation Group) k-ε turbulence model, and two different inlet unstart phenomena were analysed. The dimensional analysis method was introduced to find the essence variables describing the inlet unstart boundary based on “numerical experimental” data in this paper. The dimensionless pressure ratios of the forebody and isolator were analysed respectively. The results show that the unstart boundary of the 2D mixed hypersonic inlet is determined by M0, α and Re0. Pressure ratio π increases with M0 increasing, and it increases firstly and decreases then with α increasing. Pressure ratio π increases with Re0 increasing. Re0 (Re0 < 2 × 107) has a major effect on π and Re0 (Re0 > 2×107) has little effect on π.

This paper considers the dynamic stall vortex of importance in helicopter rotor aerodynamics and discusses previous measurements of its convection speed. It emerges that an anomaly exists between the available data sets, i.e. that some workers find that the convection speed is dependent upon the aerofoil motion, while others find that this is not the case. Measurements of the convection speed from data gathered at Glasgow University for a variety of aerofoil shapes and motion types are then presented, which support the conclusion that the dynamic stall vortex convection speed is independent of aerofoil type and motion type to a first order.

Effects of boundary conditions on the performance of compact oscillatory momentum and vorticity generators, commonly known as ‘synthetic jet’, ejecting a train of vortex-pairs into still air, were studied experimentally. The different boundary conditions altered the near-device entrainment process of the zero-net-mass-flux actuator. The measurements included hot-wire and Particle Image Velocimetry, cavity pressures and temperatures.

When the actuator operates in still air, a quasi-2D vortex pair is generated due to the extreme shear at the edges of the fluid slug ejected during the blowing stage of each excitation cycle. The vorticity flux exiting the slot determines the resulting vortex-pair circulation. The threshold slot exit velocity, for the current configuration and operating conditions, determines if the vortices will be sucked into the actuator’s cavity or be released. Once released, the vortex convection speed approximately scales with the peak velocity at the slot exit. However, the normalised convection velocity increases with the slot Reynolds number.

When even a very short extension is attached to one ‘lip’ of the actuator exit, the jet is deflected in the direction opposite the extended lip, due to the restriction on the entrainment process. When long, one ‘lip’ extension is attached, such that the vortex pair is ejected parallel to a plate, the coherence of the vortices improve, their phase speed and magnitude decrease.

The effects of high-frequency excitation, ejected perpendicular to a wall into still air, were also investigated. It was found that the presence of the plate does not have a measurable effect on the wall normal excitation, indicating that the majority of the entrainment is taking place from the forward 180° of the actuator exit plane. When the slot is inclined to the surface at a shallow angle of 30 degrees, an unsteady wall jet is formed, transferring momentum along the wall. This is a direct result of the symmetry break, altering the relative magnitudes of the vortex pair.

The degree of complexity and randomness of aircraft tracking which can be safely managed is a fundamental question for a given airspace design. To answer this question and to provide optimum conflict resolution strategies, the nature of conflict needs to be thoroughly understood. This paper reviews traditional conflict models showing how these models may be integrated into a more unified model based on computational geometry. This mathematical approach is being investigated for the specification and management of future airspace, achieving a more solid architectural basis for airspace design. This approach has the potential to significantly reduce the subjectivity when attempting to model controller and pilot workload, an important requirement in future designs.

Reduction of aircraft manufacturing cost benefits aircraft direct operating cost (DOC). The degree of stringency in specifying aircraft smoothness influences cost, i.e. the tighter the tolerance, the higher is the manufacturing cost. Discrete surface roughness arising from manufacturing tolerance at the wetted surface may be seen as an ‘aerodynamic’ defect. Features such as steps, gaps, waviness and fastener flushness (termed excrescence), seen as defects, contribute to aircraft parasitic drag.

The study is conducted on an isolated nacelle which is considered to be representative of an entire aircraft. Eleven key manufacturing features at the wetted surface of a generic long duct nacelle are identified, each associated with surface roughness. The influence of tolerance allocation at each of the key features is investigated to establish a relationship between aircraft aerodynamics and associated costs. The initial results offer considerable insight to a relatively complex problem in a multi-disciplinary environment.

Excrescence drag arising out of these ‘aerodynamic’ defects is assessed by using CFD and semi-empirical methods. Cost versus tolerance relationships are established through in-house methods using industrial data.

The aircraft unit price typically contributes from two to four times more than the fuel burn to aircraft direct operating costs. A trade-off study between manufacturing cost and aircraft drag indicates that, in general, there is scope for some relaxation of present-day tolerance allocation, to reduce aircraft acquisition cost, which would in turn reduce direct operating costs.

A recent paper gave an elementary analysis of the effect of control procedures on the flow of air traffic. It was shown that the average delay experienced by aircraft landing at an airport increases rapidly as the arrival rate approaches the maximum handling capacity of the airport. It is the object of the present paper to give a more rigorous treatment of the problem and to derive the distribution of delays under different traffic conditions.

The maximum handling capacity of an airport is the greatest rate at which aircraft can enter the airport with absolute safety. It is determined by the safe minimum separation between successive landings. An aircraft arriving in the central zone with less than the safe minimum separation is withheld from landing until the minimum interval has elapsed.

In the following it is generally assumed that aircraft cross the boundaries of the control zone from random directions with random separations, the mean arrival rate being constant. Aircraft which follow their predecessors too closely are assumed to be delayed for a minimum period so that none lands with an interval less than the safe minimum. As a simplification it is assumed that no aircraft takes off from the airport.

The results of an investigation into the effect of both the number of orifices and the geometry of these orifices on the fluid structures exiting from a synthetic jet actuator are presented. The experiments have used smoke and laser visualisation and single hotwire measurement techniques to establish the flow field exiting the various orifices and the fluid velocities in the exit plane of the orifice. The results indicate that the use of two small orifices can perhaps produce two coherent vortex rings with a total circulation greater than that found in the ring produced by a large, single orifice of equivalent area. However, they have also shown that there are three different types of interaction between two adjacent orifices, related to the spacing between them and the level of excitation applied to the diaphragm. In addition, these interactions may affect the overall level of circulation in different ways. Finally the results have shown that a rectangular orifice of a given exit plane area is more likely to produce a turbulent ring than the equivalent circular orifice due to a ‘smoothing’ process which endeavours to create an axisymmetric ring from a non-axisymmetric slug of fluid. This work has shown that when using synthetic jet actuators in flow control applications, care should be taken in selecting the spacing and geometry of the orifices used, as certain set-ups may be more suited to certain applications than others.

Contra-Rotating Open Rotor (CROR) propulsion systems have seen renewed interest as an economic and environmentally friendly powerplant for future transport aircraft. Installation effects, i.e. the mutual interactions between airframe components and the rotors, have a pronounced impact on the aerodynamic and aeroacoustic performance for this type of engine. In the past five years, DLR’s Institute of Aerodynamics and Flow Technology has performed a number of numerical studies investigating important aspects relating to engine-airframe integration of CROR engines. In this article an overview of coupled aerodynamic and aeroacoustic simulations investigating representative pusher-configuration CROR engines will be given, focusing on the impact on aerodynamic performance and aeroacoustics caused by the presence of a pylon, the potential for noise reduction by employing trailing-edge blowing at the pylon trailing edge as well as the performance and noise variations caused by different senses of rotation of the rotors. It is shown that the interaction with the pylon strongly impacts blade performance and front rotor noise emissions but that the use of active flow control in the form of pylon trailing-edge blowing can alleviate these adverse installation effects to a notable extent.

A method for performing three-dimensional, multipoint, lift-constrained drag minimisation for complex aircraft configurations is presented. Parameters representing the aircraft geometry are the design variables used in the solution of an optimisation problem. The compressible Euler equations for the flow are discretised on automatically generated unstructured meshes, and the sensitivities of the objective function and the constraints with respect to the design parameters are efficiently calculated using the discrete adjoint method. In addition, the solution algorithm has been parallelised making the approach feasible for practical applications. Several constrained minimisation strategies are discussed and some numerical tests are carried out for a lift-constrained, two-dimensional problem. The strategy found to work best is demonstrated for the three-dimensional optimisation of a wing-body configuration operating in transonic and supersonic conditions. It is concluded that Euler-based optimisation can be useful as a first step in the design process but that most applications involving transonic flows do require viscous flow modelling to avoid unrealistic pressure distributions.

Passive imaging using mm-waves offers very significant advantages in scientific and military surveillance. However, the relatively long wavelengths mean that for the resolutions that are typically sought, the input aperture of the imager needs to be quite large; typically in excess of one metre. Deployment of conventional dish antennae of these dimensions on small aircraft is highly problematic. The use of snapshot synthetic aperture interferometric radiometry (SAIR) offers an attractive route to integrating a two-dimensional antenna array into the structure of an aircraft so that the transverse effective dimensions of the antenna can be almost as large as the aircraft. We report here a study into the feasibility of deployment of a SAIR on unmanned airborne vehicles and the achievable performance parameters. Critical considerations are the achievement of acceptable sensitivity and angular resolution from a SAIR that does not require excessive complexity. It is shown that traditional approaches based on fully sampling the spatial frequencies in the scene are unable to simultaneously meet all of these criteria, but that a SAIR based on thinned sampling of the spatial frequencies shows promise.

A method of estimating the fatigue properties of notched test pieces when cyclic loads are applied in the presence of mean stresses is described. The method takes into account the local yielding which occurs near the base of the notch when the nominal stresses are sufficiently high. The general form of the predicted R/M diagrams closely resembles those obtained experimentally in previously published work.

Following the original work of Munk, Jones, and Ward for steady flow, a solution is given for unsteady, supersonic flow over very slender bodies of revolution and wings. The results are subject to the restriction (M2- l)δ2 loge [(M2- 1)½ δ]«1 where δ is the slenderness ratio, M is the Mach number and, in addition, to the usual restrictions imposed by linearisation. As examples, the lifts and pitching moments on flat wings and bodies of revolution executing pitching motion and the damping moment on a rolling wing are calculated.

It is shown that the order of approximation is consistent with the limitations already imposed by linearisation, at least in supersonic flow, where no Kutta condition is required.

The accurate prediction of ground effect aerodynamics is an important aspect of wing-in-ground effect vehicle (WIG) design. Computational fluid dynamics (CFD) solutions are useful alternatives to expensive (especially in the case of ground effect) wind-tunnel testing. However, the incorporation of the rigid surface effects often leads to confusion due to such a model being in a vehicle fixed reference frame (air moving, vehicle fixed) rather than the real-life situation of a ground fixed reference frame (air fixed, vehicle moving).

In a previous paper the author derived, by the Rayleigh-Ritz method, a determinantal equation for obtaining the critical magnitude of any uniform system of edge stress applied to a clamped oblique plate. In the present paper this equation is used to derive values for the buckling stresses of clamped oblique plates subjected to pure shear along and perpendicular to two edges of the plate.

As would be expected, it is found that reversal of the direction of the shear produces a change in the critical value. The lower value occurs when the shear is tending to increase the obliquity of the plate and critical values corresponding to this direction are calculated for 45° oblique plates with side ratios varying between 3/5 and 5/3. The critical values obtained for the reverse direction were obviously very inaccurate, due to the inadequacy of the series which was used to represent the buckling mode. As yet no satisfactory method has been found for overcoming this difficulty.

A formula is derived for the buckling in compression of an orthotropic sandwich panel, with all edges simply-supported. Two special types of symmetrical sandwiches, that of a sandwich having zero core flexural stiffnesses and of a corrugated core sandwich, are considered, and curves giving the critical buckling end load per inch are presented. An approximate theory is also given for the bending of sandwich panels under the action of an axial compressive load and a uniform lateral pressure. Curves are presented giving the maximum bending moments in a corrugated core sandwich panel.

Solar powered aircraft are becoming more and more interesting for future long endurance missions at high altitudes, because they could provide Earth monitoring, telecommunications, etc. without any atmospheric pollution and, hopefully in the near future, with competitive costs compared with satellites. The research activities carried out till now have been mainly focused on flying wings or conventional aircraft configurations, with a great emphasis on the technological aspects. The present paper aims to define a mathematical model for solar powered aircraft preliminary design, valid independently of the aerodynamic configuration. A preliminary analysis is carried out in order to simulate Helios and the results are compared with those available from the flights of this aircraft. The proposed mathematical model is used also to compare four different aircraft configurations, namely: a flying wing, a conventional aircraft, a twin boom aircraft and a biplane aircraft. The results obtained are discussed in the paper and an optimum aircraft is analysed.

The ionosphere is one of the main factors affecting the accuracy and integrity of satellite-based augmentation system positioning systems. This paper presents the results of a 30-day study of the accuracy and integrity of the European Geostationary Navigation Overlay Service (EGNOS) conducted at the EPOD airport belonging to the Aeroclub of Warmia and Mazury in Olsztyn, in northeastern Poland (the area until recently considered as the edge of EGNOS coverage). Analyses of the parameters characterising the accuracy and integrity of positioning were performed in three calculation variants/modes: with the original EGNOS ionospheric correction, with correction determined by means of Klobuchar algorithm, and finally with modified ionospheric coefficients developed by the CODE. Studies have shown clearly that the original EGNOS ionospheric model gives the best integrity and accuracy results allowing to use EGNOS for approach with vertical guidance procedures, while the Center for Orbit Determination in Europe and Klobuchar models could only be used for non-precision approach operations.