A static model of the hose and drogue used to provide data in a realtime flight simulation of air-to-air refuelling is presented. The model predicts (a) the interference effect of the flow around the nose of a combat receiver aircraft on the pre-contact position of the drogue and (b) the hose shape and the loads induced on the receiver probe during contact. The use of the static model in the pre-contact case is validated for a normal closure speed by comparison with the results from a small perturbation dynamic analysis. At higher closure speed the static model is found to be less satisfactory. In a typical receiver approach a movement of the drogue similar to the drogue radius is predicted.

There is a popular misconception that synthetic training is only concerned with time spent in the simulator. If this were so, a training exercise which had focused on the simulation experience alone, might fail to provide the rich learning environment needed to maximise training opportunities and so fall short of expectations in terms of delivery of a complete training experience.

This is particularly true when considering collective training for military aircrew; where all mission phases are equally important, and all roles and tasks must be carefully considered when assessing the training need. The training designers must create a scenario that meets the requirements of the total training audience (the collective) throughout all the mission phases; planning, briefing, flying (traditionally the key simulation event) and debriefing. This scenario must allow the trainees to become fully immersed and provide a ‘total experience’ in order for them to obtain maximum benefit from the synthetic training event.

Maintenance of such an environment during the training process places an extra burden on those who are providing the training experience. As the number and mix of trainees participating in a synthetic exercise grows so too do the demands placed upon the exercise management team, the trainers and the assessors to provide a rich and rewarding training experience. Not only must these key personnel be carefully selected, their needs must also be taken into account during the training design process. This will help to ensure that transformation of data to information and information flow is optimised for all participants.

Recent experience in mission training via distributed simulation has shown that greater training opportunities can be afforded if the training designers address the needs of all the above. Using these experiences as reference points, this paper will discuss synthetic training issues along with the expectations of the trainees and the training providers and suggest some guidelines for creating a total training experience.

Although many researchers, from a number of industries, have studied the aerodynamics of ground effect conditions, published research remains inconclusive. Published results are frequently found to conflict and modeling procedures (in both computational and experimental situations) vary. In this paper, existing research is reviewed and the methods used in obtaining the results are compared. Possible causes for the discrepancies in the published work are outlined, including the inappropriate specifications of boundary conditions and the neglect of viscosity in some computational analyses. Computational and experimental analyses are then performed to determine if these identified areas do cause inconsistencies in results and to also determine accurate simulation procedures.

Routine periodic audiometry performed as part of the statutory medical examination of flight crew employed by a major international airline indicated that up to 27–5% may suffer noise induced hearing loss, with 10-8% showing unilateral hearing loss. This was considered to result from the practice of listening to radio communications through a headset covering one ear only, leaving the other ear uncovered to allow direct conversation between crew members on the flightdeck. Ambient noise levels at the crew seats and noise levels from crew headsets were measured on a sample of aircraft and routes using a Kemar acoustic manikin. Noise levels on the flightdecks were 70-79dB(A) Leqaveraged over a flight of typically seven to ten hours duration. Noise levels from headsets were 77-89dB(A) Leq being 80dB(A) or higher on 80% of flights, and 85dB(A) or higher on 40% of flights. Five prototype or production active noise reducing (ANR) headsets were evaluated in flight and in the laboratory. It was shown that the better examples, worn to cover both ears, would allow flight crew to reduce headset volumes so reducing risk to their hearing.

In this paper two numerical procedures are presented: the first algorithm allows for the determination of the geometric characteristics of the maximum efficiency propeller for a given operative condition and profile distribution along the blade; the output of this numerical procedure is the chord distribution and twist angle of the blade, together with its efficiency and its torque and thrust coefficients for the prescribed operative condition. The aerodynamic characteristics of the optimum propeller when operating in a condition different from the design one are obtained by a second algorithm that allows for the evaluation of the efficiency, the thrust and torque coefficients of a propeller of known geometry, when the blade pitch and operative condition are varied.

In the paper the formulation used for deriving the geometry of the optimum propeller and determining its performances when operating off-design is described in detail. The results obtained from the proposed propeller model have been validated by comparison with experimental data.

The influence of trailing-edge devices such as Gurney flaps and divergent trailing edges of different height on the aerodynamic performance of an aerofoil at transonic speeds has been investigated experimentally. The investigation has been carried out in the Transonic Wind Tunnel Göttingen (TWG) using the two-dimensional aerofoil model VC-Opt at freestream Mach numbers of M ε [0.755, 0.775, 0.790] and a Reynolds number of Re = 5.0 x 106.

The results have shown that the trailing-edge devices increase the circulation of the aerofoil leading to a lift enhancement and pitching-moment decrease as well as an increase in minimum drag compared to the baseline configuration. The maximum lift-to-drag ratio is considerably improved and the onset of trailing-edge flow separation is shifted to higher lift. Besides the increased rear-loading, a downstream displacement of the shock provides the main lift enhancement in transonic flow.

The simple Gurney flap provides the largest additional circulation of all geometries tested. The smoother turning of the flow due to the additional ramp of the divergent trailing edge leads to a smaller increase of circulation. Slightly less lift but considerably less viscous (pressure) drag is generated enhancing the maximum lift-to-drag ratio compared to the Gurney flap. The negative affect of the Gurney flap on the pitching moment is also reduced.

For the high divergent trailing edges, different ramp slopes have a significant influence on the aerodynamic performance whereas at low device heights the influence is considerably diminished.

The results show that the divergent trailing edge proves to be the better trailing-edge device at transonic speeds. The application as an element for an adaptive wing is generally possible.

This paper describes the application of the theory of pulse responses to a linearised Euler scheme on a moving grid. The impulse responses of such a linear system are memory functions or temporal representations of the manner in which, and the time over which a perturbation remains active in the response of the system. The exact response of the linear system to an arbitrary input can then be predicted via convolution. The linearised Euler scheme is derived from a full nonlinear Euler method based on Jameson’s cell-centred scheme which has been modified to make it time-accurate and has the necessary terms to account for grid motion. The discrete form of the full Euler equations are linearised about the corresponding nonlinear steady mean flow assuming the unsteadiness in the flow and grid are small. The resulting set of differential equations for the flow perturbations are solved at each time step using a dual-time scheme. These equations can be easily solved to obtain pulse responses, which capture the entire frequency content of the linear system. These pulse responses then provide a simple and efficient method for calculating the system response to many different inputs via convolution. A further use of the pulse responses is their potential to allow a reduced-order linear model to be constructed, which preserves some of the accuracyof the original system. Preliminary work carried out in this area is presented. Results of calculations for heave, pitch and ramp test cases are presented.

The US Naval Research Laboratory (NRL) is developing technologies that will enable Navy-relevant missions with the smallest practical Micro Air Vehicles (MAVs). The NRL Micro Tactical Expendable (MITE) air vehicle is a result of this research. MITE is a hand-launched, dual-propeller, fixed-wing air vehicle, with a 25cm chord and a wingspan of 25–47cm, depending on payload weight. Vehicle gross weight is 130–350g. Miniature autopilot systems, based on visual imaging techniques, are being developed for MITE. These will be used in conjunction with conventional autopilot sensors to allow the MITE to fly autonomously. This paper provides an overview of the MITE development, including aerodynamic design considerations, electric propulsion, and vision-based autopilot research. Also presented is a rationale for the development of control laws that can direct the behavior of large groups of MAVs or other vehicle agents. Dubbed ‘physicomimetics,’ this process can bring about the self-assembly of complex MAV formations, though individual MAVs have minimal onboard processing power and limited local sensing capabilities.

The topological structure of an underexpanded Mach 2·5 jet injected into a Mach 0·75 crossflow are studied experimentally and numerically. The structure of the skin friction pattern is examined in detail by localising precisely the topological singularities. Oil-flow visualisations reveal the complex nature of the wall flow and its strong three-dimensionality. Four vortical structures appear in the vicinity of the wall, where two are essentially parallel to the wall (horseshoe vortex and boundary-layer vortices) and the other two are directed perpendicular to the wall (wake vortices). Comparison of computation with experiment shows good agreement for the complex separated-flow surface shear pattern and the pressure distribution. The crossflow Mach number plays an important role in the separation mechanisms. Significant differences in the topology occur with a Mach 0·3 crossflow whereas above Mach 0·6 the mean structure of the interaction is almost conserved, except that the secondary wake vortices disappear.

This paper describes part of an investigation into the prediction and characterisation of limit cycle oscillations occurring in non-linear aeroelastic systems. Through the use of a modified version of normal form theory, it is shown how it is possible to predict the limit-cycle oscillations and characterise their stability. The approach is analytical and does away with the need for an excessive amount of time marching iterative numerical simulation of the system. The methodology is demonstrated upon a simple two degrees-of-freedom aeroelastic wing model with cubic stiffness. A good agreement was obtained between the analytical prediction and numerical simulations.

For generalised Hohmann and bi-elliptic transfer orbits, we applied small tangential impulses at peri-apse and apo-apse to acquire the precise destination elliptic orbit stemming from an initial orbit. There are four cases for this correctional or improvement process. We take into consideration the first configuration for the time being. We determined the four relationships connecting the increments in major axes and eccentricities with the correctional increments in velocities ΔvA, ΔvB at points A, B for the generalised Hohmann case. We assigned the three relationships - for the bi-elliptic transfer - connecting slight changes in major axes with small increments in velocities at points A, B, C due to motor thrusts.

The effect of streamwise slots and grooves on a normal shock wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave-boundary-layer interaction to a large bifurcated lambda type shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.

The flow field produced by the orthogonal blade-vortex interaction (BVI) was examined through the application of a particle image velocimetry (PIV) technique, and acquisition of comparative unsteady surface pressure measurements. The velocity vector maps obtained revealed a significant radial flow from the vortex core after severing, in the form of an outward flow from the core when the flow is directed towards the surface (lower), and inward when directed away (upper). With progression of the vortex over the chord of the blade, the interaction led to a decrease in peak vorticity levels on the lower surface, but a steady level was maintained on the upper surface. Assessment of the effect of the interaction on the out-of-plane mass flux was performed through calculation of the divergence, revealing that there is no further variation in the axial flow on either side of the blade after the initial cut.

Civil tiltrotor technology will literally change the way man flies! With the advent of modern vertical flight aircraft, like the tiltrotor, air travel will accede to new levels of convenience, safety, and cost-effectiveness. Tiltrotors can operate in all weather conditions from airports or from smaller facilities called vertiports strategically located to provide convenient access to other transportation systems. In evaluating the economic viability of this mode of travel, the entire cost of the system must be included to realise the benefits of such a system, not just the cost of the aircraft or its cost of operation. When an improved air transportation system is developed, tiltrotors, because of their time savings and cost-competitiveness, will be able to compete in the 50–500 mile range with all types of aircraft as well as high-speed rail and other modes of transportation. While each mode has its advantages, tiltrotor will do well because it can add flexibility to the system and can complement the other modes and make the overall system more efficient. It will also be acceptable to the public at large because of its unique capability and low noise signature.

Classification of the winged aircraft according to the character of wing (blade) motion in horizontal and vertical plane is presented. Design of a rotor, the blades of which are driven into rotation via flapping motion similar to the movement of an ornithopter's wing is proposed and the principle of operation is considered. Mathematical model of the aerodynamics of the autorotating rotor with flapping blades in hover is made on the base of the helicopter rotor blade-element theory. Geometric and kinematic parameters influence on the rotor lift and efficiency is analysed. It is shown that in hover the efficiency of an autorotating rotor with flapping blades is approximately equal to the one of a helicopter rotor.

Coupling an unsteady flow-solver with a structural model offers the opportunity to simulate aeroelastic behaviour of wings and rotor blades. The moving and deforming surfaces resulting from unsteady simulations require deforming meshes during the simulation, and it is common to use simple interpolation of surface displacements and velocities onto the initial undisturbed mesh. However, aeroelastic simulations can result in large displacements and deformations of solid surfaces, and simple interpolation of perturbations results in poor grid quality and possible grid crossover. A new interpolation technique is presented which is still simple in that it is driven solely by surface motion, but represents rotational effects near the solid surface, to maintain grid quality there. Furthermore, the scheme is fully analytic, so is very cheap computationally and results in grid speeds also being available analytically. Results, in terms of unsteady grid motion and flow solution, show the scheme to be effective and efficient.