Three-dimensional RANS calculations and comparisons with experimental data are presented for subsonic and transonic flow past a non-axisymmetric (rectangular) nozzle/afterbody typical of those found in fast-jet aircraft. The full details of the geometry have been modelled, and the flow domain includes the internal nozzle flow and the jet exhaust plume. The calculations relate to two free-stream Mach numbers of 0-6 and 0-94 and have been performed during the course of a collaborative research programme involving a number of UK universities and industrial organisations. The close interaction between partners contributed greatly to the elimination of computational inconsistencies and to rational decisions on common grids and boundary conditions, based on a range of preliminary computations. The turbulence models used in the study include linear and non-linear eddy-viscosity models. For the lower Mach number case, the flow remains attached and all of the turbulence models yield satisfactory pressure predictions. However, for the higher Mach number, the flow over the afterbody is massively separated, and the effect of turbulence model performance is pronounced. It is observed that non-linear eddy-viscosity modelling provides improved shock capturing and demonstrates significant turbulence anisotropy. Among the linear eddy-viscosity models, the SST model predicts the best surface pressure distributions. The standard k -ε model gives reasonable results, but returns a shock location which is too far downstream and displays a delayed recovery. The flow field inside the jet nozzle is not influenced by turbulence modelling, highlighting the essentially inviscid nature of the flow in this region. However, the resolution of internal shock cells for identical grids is found to be dependent on the solution algorithm -specifically, whether it solves for pressure or density as a main dependent variable. Density-based time-marching schemes are found to return a better resolution of shock reflection. The paper also highlights the urgent need for more detailed experimental data in this type of flow.

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.

Lobed mixers are widely used in gas turbine engines to increase mixing between hot and cold streams and consequently reduce jet noise. CFD predictions are presented for a simplified experimental configuration of a planar, three lobe geometry. Results are shown for a standard linear k–ε turbulence model, the same model with a time scale limitation invoked and a non-linear quadratic model also employing a time scale limitation. Comparisons are presented between the three models for axial velocity, velocity vectors, shear stress and turbulence kinetic energy at a selected plane in the mixing region. The non-linear model was found to have little influence on the mean flow but some effect on the turbulence structure was observed. Comparison with measurements showed that all major features were reproduced but detail differences were evident. The use of a time scale limit reduced peak values of predicted turbulence quantities by 20-30%. As compared to the standard linear model, the time scale limited non-linear model moved the position of the zero streamwise circulation location by about one lobe wavelength upstream so giving better agreement with experiment.

This paper presents the first results from research into active control of structural load alleviation (SLA) for tiltrotor aircraft carried out in the European ‘critical technology’ RHILP project. The importance of and the need for SLA in tiltrotors are discussed, drawing on previous US experience reported in the open literature. The paper addresses the modelling aspects in some detail; hence forming the foundation for both the FLIGHTLAB simulated XV-15 and EUROTILT configurations. The primary focus of attention is the suppression of in-plane rotor yoke loads for pitch manoeuvres in airplane mode; without suppression these loads would result in a very high level of fatigue damage. Multi-variable control law design methods are used to develop controller schemes and load suppression of 80-90% is demonstrated using rotor cyclic control, albeit at a 20-30% performance penalty. However, rotor flapping transients tend to increase by the action of the SLA system. A dual-objective control design approach demonstrates the effectiveness of suppressing both loads and flapping simultaneously.

Time-critical targets will generally be observed or ‘sensed’ by personnel inadequately equipped to shoot at them. Since time may be a critical parameter, the passage of information from the ‘sensor’ to the ‘shooter’ must be rapid and accurate. QinetiQ has developed and integrated airborne and ground equipment to facilitate the passage of targeting data digitally over standard air and ground radios. The UK and Holland have selected an integrated rugged PC, internal modem card and purpose-designed software to equip Forward Air Controllers (FACs) to enable digital forward air control. This gives direct interoperability with UK and NATO assets equipped with Improved Data Modem (IDM) technology.

A large, and growing, number of air, land and sea assets employ Link-16. This entails entirely different frequency, data rate and, generally, protocol, from the IDM. The UK MoD’s recently-ended applied research EXTENDOR programme was a three-year operational concept demonstrator for the purpose of translating in near real-time between selected key Link-16 and IDM-based messages. The intention was to demonstrate the viability and utility of translation between disparate data links. This staged programme centred around a translation engine designed to be flown in a Predator UAV. The programme was entirely successful, demonstrating the capability to translate both own position and nine-line target data. Thus the position of a non-Link-16-equipped asset could be seen in real time in the Link-16 recognised air picture (RAP). Also, a Link-16 asset was shown to be able to task an IDM-equipped asset with a nine-line brief via the translation engine. The high equipment readiness level offers considerable potential for near-to-medium term operational exploitation.

The problem of aircraft formation dynamics and control is investigated from the viewpoint of formation architecture. Three different formation structures, leader-wingman, virtual leader and behavioural approaches are introduced. A comparative study is made using a unified approach through a suitable control law. The formation systems are analyzed on a quantitative basis and objective results are made available for the designer. The trade-off between system performance and complexity is indicated. A complete nonlinear simulation involving a flight-path change and a heading change manoeuvre is discussed. Results show the superiority of a behavioural approach to maintain close formations of vehicles.

In the field of rotorcraft dynamics, it is significant that the induced inflow field is well understood and modeled. A large number of methodologies have been developed in the past years, among which the state-space model is recognised for its advantage in real-time simulation, preliminary design, and dynamic eigenvalue analysis. Recent studies have shown success in representing the induced flow field everywhere above the rotor plane even with mass source terms on the disk as long as they have zero net flux of mass injection when integrated over the disk. Nevertheless, non-zero net mass influx is expected in numerous situations, such as ground effect, tip drive rotors, etc; and the incapability of previous models limits the utilisation of the methodology in these cases. This work presents an extended potential-flow, state-space model derived from the potential-flow momentum equation by means of a Galerkin approach. The induced velocity and pressure perturbation are expanded in terms of closed-form, time-dependent coefficients and space-dependent associated Legendre functions and harmonics. Non-zero net mass flux terms are represented by the involvement of associated Legendre functions with equal degrees and orders. Validation, as well as discrepancies, of the inclusion of such terms is investigated. Numerical simulation of frequency response in axial and skew-angle flight is presented and compared with exact solutions obtained by the convolution integral. Also the study shows that, unlike other pressure distribution responses, non-zero mass influx exhibits a high sensitivity to the choice of the number of states in the velocity expansion. Error analyses are performed to show this sensitivity.

Recent developments in delta wing aerodynamics are reviewed. For slender delta wings, recent investigations shed more light on the unsteady aspects of shear-layer structure, vortex core, breakdown and its instabilities. For nonslender delta wings, substantial differences in the structure of vortical flow and breakdown may exist. Vortex interactions are generic to both slender and nonslender wings. Various unsteady flow phenomena may cause buffeting of wings and fins, however, vortex breakdown, vortex shedding, and shear layer reattachment are the most dominant sources. Dynamic response of vortex breakdown over delta wings in unsteady flows can be characterised by large time lags and hysteresis, whose physical mechanisms need further studies. Unusual flow–structure interactions for nonslender wings in the form of self-excited roll oscillations have been observed. Recent experiments showed that substantial lift enhancement is possible on a flexible delta wing.

Aeroelastic analysis is a critical area of the aircraft design process, as a good understanding of the dynamic behaviour of the wing structure is essential to safe operation of the vehicle. The inevitable inaccuracies present in the modelling of such phenomena impose mass penalties, as large safety margins are necessitated, which in turn lead to overly stiff designs. In an effort to reduce the uncertainty in analysis methods, fully coupled CFD and structural models are under widespread development. This paper describes the results produced by such a system for a series of test cases based on the AGARD445.6 and MDO wings. Results relating to the latter are of particular interest, as significant variations were found to be produced by the different methodologies used in previous studies, the precise cause of which could not be isolated. In an effort to provide this isolation, a detailed description of the method used is given, including the interpolation scheme between the structural model and the aerodynamic surface, and particular attention is given to the issue of aerofoil shape preservation.

The New Zealand Government takes airspace safety very seriously. The level of safety in New Zealand airspace is measured by the number of recorded airspace incidents. An airspace incident can be thought of as a failure in the chain of operations in the air traffic system when it is provided with an air traffic service (ATS). Some of these incidents result in a loss of separation between aircraft, varying from slight to a very serious loss with a significant risk of collision, known as a near collision. New Zealand’s Civil Aviation Authority (CAA) identifies the causal factors for all airspace incidents using the Reason model of human error, divided into three areas; active failures committed by individuals involved in the incident, local factors relating to the task and the ATS environment and organisational factors originating in the managerial and organisation spheres of the ATS provider. Based upon the CAA’s database, this paper analyses trends in controller caused incidents in the NZ airspace during the period 1994 to 2002 in six controller caused categories. The results indicate that for controller caused incidents, execution errors dominate the active failures category, while in the case of local factors, poor concentration/lack of attention, inadequate checking and controller workload are the dominant factors. For the organisation category, poor planning and inadequate control and monitoring dominate. These results should form the basis of a robust and transparent framework for intervention mechanisms by the New Zealand Civil Aviation Authority for enhanced airspace safety.

The problem of impact damage in laminated composite structures, and the consequent reduction in residual strength, has been a topic of continual research for over two decades. The number of journal papers on the subject now runs into four figures and most have been conscientiously reviewed by Abrate(1991, 1994, 1998). This review is not intended to be in the academic tradition, with emphasis on acknowledging the authorship of all the various research initiatives. Instead we present our opinions so that the reader can appreciate our current understanding of the problem, our capability of predicting by analysis, and the scope of the design tools for avoiding structural damage, or at least designing damage tolerant aerospace structures.

Separated and vortical flows are dominant over various unmanned air vehicles (UAVs). In this article, issues and challenges of vortical flows for future UAVs are reviewed. These include shear layer instabilities, vortex breakdown and wing stall, vortex interactions, nonslender vortices, multiple vortices, and manoeuvring wing vortices. There are also issues relating to vortical flows in certain flow/structure interactions, as well as in aerodynamics/propulsion interactions. Separated and vortical flows are even more dominant at low Reynolds number flows. The main features of vortical flows, unsteady aerodynamics, and propulsion related vortical flow isssues relevant to mini- and micro air vehicles, are discussed.

Flow visualisation experiments related to turbine film cooling have been conducted. These investigated the fluid mechanics of coolant ejection using a large-scale, flat-plate model at engine-representative Reynolds numbers in a low-speed tunnel with ambient-temperature mainstream flow. The coolant trajectories were captured using a fine nylon mesh covered with thermochromic liquid crystals, allowing measurement of gas temperature contours in planes perpendicular to the flow. Three injection geometries were assessed: cylindrical holes with stream-wise injection, cylindrical holes with cross-stream injection, and fan-shaped holes. The data demonstrated that the cylindrical holes produced discrete jets, which lifted off the surface at high coolant-to-mainstream momentum flux ratios; these jets were characterised by the kidney-shaped stream-tubes expected for injection into cross-flow. The jets injected with cross-stream momentum exhibited a more obvious kidney-shaped cross-section, which rotated with distance downstream of injection. The jets from the fan-shaped holes were attached to the surface even at high momentum flux ratios, were more diffuse, and exhibited two cores of high temperature. The trajectory visualisation data were used to interpret the adiabatic cooling effectiveness measured at the surface.

The interaction of a helicopter tail rotor blade with the tip vortex system from the main rotor is a significant source of noise and, in some flight states, can produce marked reductions in control effectiveness. This paper describes a series of wind-tunnel tests to simulate tail rotor blade vortex interaction with a view to providing data for the development and validation of numerical simulations of the phenomenon. In the experiments, which were carried out in the Argyll wind-tunnel of Glasgow University, a single-bladed rotor located in the tunnel’s contraction was used to generate the tip vortex which travelled downstream into the working section where it interacted with a model tail rotor. The tail rotor was instrumented with miniature pressure transducers that measured the aerodynamic response during the interaction. The results suggest that the rotor blade vortex interaction is similar in form to that measured at much higher spatial resolution on a fixed, non-rotating blade. The combination of the two datasets, therefore, provides a valuable resource for the development and validation of predictive schemes.

This paper describes an evaluation of the longitudinal flying qualities of a generic blended-wing-body (BWB) transport aircraft at low speed flight conditions. Aerodynamic data was obtained from several sources and integrated into the equations of motion of a typical BWB configuration in order to provide a reasonable basis for flying qualities assessment. The control requirements to trim are enumerated for a representative range of cg position and static margin over the typical range of approach speeds for both stable and unstable configurations. The linear dynamic characteristics of the unaugmented airframe are also described for the same range of stability margin. Subsequent work describes the development of a rate command-attitude hold command and stability augmentation system configured to comply with representative modern handling criteria. Finally, the flight dynamics of the augmented aircraft are described after refinement of the control law by means of piloted simulation in a fixed base flight simulator.

This paper presents identification, control synthesis and simulation results for an YF-22 aircraft model designed, built, and instrumented at West Virginia University. The ultimate goal of the project is the experimental demonstration of formation flight for a set of 3 of the above models. In the planned flight configuration, a pilot on the ground maintains controls of the leader aircraft while a wingman aircraft is required to maintain a pre-defined position and orientation with respect to the leader. The identification of both a linear model and a nonlinear model of the aircraft from flight data is discussed first. Then, the design of the control scheme is presented and discussed with an emphasis on the amount of information, relative to the leader aircraft, needed by the wingman to maintain formation. Using the developed nonlinear model, the control laws for a maneuvered flight of the formation are then simulated with Simulink® and displayed with the Virtual Reality Toolbox®. Simulation studies have been performed to evaluate the effects of specific parameters and the system robustness to atmospheric turbulence. The conclusions from this analysis have allowed the formulation of specific guidelines for the design of the electronic payload for formation flight.

Direct optimisation techniques using different methods are presented and compared for the solution of two common flows: a two dimensional diffuser and a drag minimisation problem of a fixed area body. The methods studied are a truncated Newton algorithm (gradient method), a simplex approach (direct search method) and a genetic algorithm (stochastic method). The diffuser problem has a known solution supported by experimental data, it has one design performance measure (the pressure coefficient) and two design variables. The fixed area body also has one performance measure (the drag coefficient), but this time there are four design variables; no experimental data is available, this computation is performed to assess the speed/progression of solution.

In all cases the direct search approach (simplex method) required significantly smaller number of evaluations than the generic algorithm method. The simplest approach, the gradient method (Newton) performed equally to the simplex approach for the diffuser problem but it was unable to provide a solution to the four-variable problem of a fixed area body drag minimisation. The level of robustness obtained by the use of generic algorithm is in principle superior to the other methods, but a large price in terms of evaluations has to be paid.

In recent years, various strategies for the concurrent operation of fixed-and rotary-wing aircraft have been proposed as a means of increasing airport capacity. Some of these strategies will increase the likelihood of encounters with the wakes of aircraft operating nearby. Several studies now exist where numerical simulations have been used to assess the impact of encounters with the wakes of large transport aircraft on the safety of helicopter operations under such conditions. This paper contrasts the predictions of several commonly-used numerical simulation techniques when each is used to model the dynamics of a helicopter rotor during the same idealised wake encounter. In most previous studies the mutually-induced distortion of the wakes of the rotor and the interacting aircraft has been neglected, yielding the so-called ‘frozen vortex’ assumption. This assumption is shown to be valid only when the helicopter encounters the aircraft wake at high forward speed. At the low forward speeds most relevant to near-airfield operations, however, injudicious use of the frozen vortex assumption may lead to significant errors in predicting the severity of a helicopter’s response to a wake encounter.

Military aircraft, by definition, need to survive the onslaught of opposing forces to successfully complete their mission. From an aircraft perspective, the electromagnetic (EM) environment can be an enabler, via the use of navigation aids, radar, radio communications etc. – in fact mission success depends on its successful use. However, this environment is also potentially a disabler, as threat weapon systems and the environment itself can harm or destroy the aircraft. This paper discusses risks and hazards thus posed to aircraft survivability, partitioned into two classes – ‘direct’ and ‘indirect’ EM threats. ‘Direct’ threats are those that occur as a result of direct coupling of EM energy to the airframe and systems within, e.g. lightning strike and directed energy weapons. ‘Indirect’ threats are those that utilise EM sensors to detect, track and target the aircraft, e.g. radar-guided surface-to-air missiles. Airframe intrinsic mechanical vulnerability is also an important part of survivability, although not addressed in this EM-related paper. It is shown that risk and hazard can be minimised by gaining a thorough understanding of operational scenarios, developing holistic system-of-systems solutions to military requirements, and using best practice design and development techniques.

This paper considers the effects of both wind-tunnel walls and a downstream support structure, on the aerodynamics of a 70° delta wing. A RANS model of the flow was used with the wind-tunnel walls and supports being modelled with inviscid wall boundary conditions. A consistent discretisation of the domain was employed such that grid dependence effects were consistent in all solutions, thus any differences occurring were due to varying boundary conditions (wall and support locations). Comparing solutions from wind-tunnel simulations and simulations with farfield conditions, it has been shown that the presence of tunnel walls moves the vortex breakdown location upstream. It has also been seen that vortex strength, helix angle, and mean incidence also increase, leading to a more upstream breakdown location in wind-tunnels. The secondary separation line was also observed to move outboards. It was observed that for high Reynolds numbers, with a support downstream of the wing, vortex breakdown can be delayed due to blockage effects providing the vortices do not impinge on the support This was observed to be the case for smaller supports also.