The introduction of the original time delay method of Beddoes, an engineering model for the unsteady response of an aerofoil section including dynamic stall, had a profound effect on the design and development of rotor systems in the UK. Over the years, the model expanded to include more and more features of the unsteady flow, with many contributors. It is now in use throughout the world as part of rotor analysis packages. Nevertheless, it retains its essential simplicity. Work to confirm the ability of the most recent version of the dynamic stall model from the University of Glasgow to replicate the complicated behaviour of an advanced rotor aerofoil section at full scale Reynolds and Mach numbers provides an opportunity to review the use of this new engineering model in the helicopter rotor design environment.

This note discusses the application of dynamic stall synthesis methods to the problem of classifying and comparing aerofoil sections when designing rotors for the retreating blade stall condition that determines the rotor blade area requirement of the helicopter. The development of the dynamic stall models employed in UK rotor designs is reviewed in this paper and their use in the design process explained, with emphasis on the assumptions that overcome the limitations of the models and exploit their simplicity, enabling accurate and conservative rotor designs. The paper shows how the model may be used to structure the analysis of complex sets of dynamic aerofoil data. It illustrates how structured comparison between the model and the data yields a concise appreciation of the behaviour of the aerofoil and an understanding of the physical processes involved. Some previously unappreciated effects are identified and the model is used to transfer experience of the aerofoil section behaviour from the non-rotating wind-tunnel environment to that of the rotor. Finally, the application of the new engineering model developed at Glasgow University in the rotor design process is outlined. Some remarks on the use of engineering models in comparison with CFD models in the design context are included.

This paper presents the Adaptive Torsion Wing (ATW) concept and performs two multidisciplinary design optimisation (MDO) studies by employing this novel concept across the wing of a representative UAV. The ATW concept varies the torsional stiffness of a two-spar wingbox by changing the enclosed area through the relative chordwise positions of the front and rear spar webs. The first study investigates the use of the ATW concept to improve the aerodynamic efficiency (lift-to-drag ratio) of the UAV. In contrast, the second study investigates the use of the concept to replace conventional ailerons and provide roll control. In both studies, the semi-span of the wing is split into five equal partitions and the concept is employed in each of them. The partitions are connected through thick ribs that allow the spar webs of each partition to translate independently of the webs of adjacent partitions and maintain a continuous load path across the wing span. An MDO suite consisting of a Genetic Algorithm (GA) optimiser coupled with a high-end low-fidelity aero-structural model was developed and employed in this paper.

Aircraft upset and Loss of Control (LOC) is a leading cause of accidents in commercial and general aviation aircraft operations. A number of measures have been taken in the commercial segment to improve training and awareness of this problem and several organisations offer in-flight training to enhance awareness. In relative terms, in both commercial and general aviation sectors, the use of Full Flight Simulators (FFS) and Flight Training Devices (FTD) for this purpose is minimal. A key reason for this is the limited capability and coverage of flight models used in these devices. This paper provides examples of the limitations in civilian simulators in contrast to military efforts that have been using full envelope modelling for decades to enhance pilot training. The paper also presents techniques used in full-envelope modelling, their validity, and a regulatory vehicle that is in-line with current international guidelines for application to civilian trainer development.

The use of the linearised potential model for the analysis of compressible flows is quite widespread, and provides good results for subsonic and supersonic flows. However, the calculation of aerofoils and wings subject to transonic flows requires a non-linear model, such as the transonic small-disturbance (TSD) potential equation. The solution of the problem by a singularity distribution requires singularities over the field, as well as panels on the boundary, characterising the procedure known as field panel method. The present work shows results of calculations of the transonic small-disturbance potential equation for flows without shock waves using the dual reciprocity method (DRM), which permits calculation of integrals only at the boundary of the problem, without the need of field distributions. This approach, compared to the field panel methods, takes considerably less computer time, and shows a significant improvement when compared to results of linear theory without much additional computer time, making this technique adequate to design phases of aircraft. Pressure distribution results show good agreement with other methods found in litterature. The low computational cost of the present method allows us to perform parametric tests and explore the effects of thickness and Mach number on the lift and pitching moment coefficients. A discussion of the physical effect of these parameters on the problem is presented, and the thickness of the aerofoil is shown to increase the lift and change the position of the aerodynamic centre. However, this non-linear effect depends on the precise shape of the thickness distribution.

A flight simulator study has been carried out to evaluate the performance of modern flight control systems encountering curved wake vortices. During the decay process the shape of wake vortices alters significantly which has an influence on the encounter characteristics and thus on the encounter hazard. To analyse most realistic wake encounters, flow fields of matured vortices have been generated with large-eddy-simulations. These were used for the determination of induced force and moment histories during the encounter. The force and moment histories have been implemented into the equations of motions of a 6 DoF flight simulation. For the sake of comparison encounters have also been simulated with straight vortices as they were mostly used for many other investigations. The most important goal of the study was to analyse the difference between these ideally straight vortices and vortices with a more realistic shape regarding encounter acceptance. The simulator study was conducted in an A330 motion-based full-flight simulator with pilots-in-the-loop. The analysed scenario was a wake encounter during final approach. The encounter conditions corresponded to a heavy-behind-heavy situation for Instrument Flight Rules (IFR) operations. The aircraft was flown either manually (in normal law) or with autopilot engaged. Altogether 93 encounters have been simulated, 38 with straight and 55 encounters with curved vortices. For encounters under manual control the simulator study revealed a potential risk of pilot induced oscillations (PIO) during encounters with curved vortices. With autopilot engaged not even one encounter with curved vortices was classified to be unacceptable. Although significant aircraft response was experienced the autopilot never disengaged automatically in any encounter. Altogether about 12 percent of the encounters were not accepted by the pilots. This is indeed a significant number, especially as the analysed scenarios can be regarded to be realistic situations which can occur in reality even if the applicable separation distances are applied.

This paper describes a study that investigated the efficacy of modifications to the trailing end of the externally mounted advanced targeting forward looking infrared pod (ATFLIR) on the store separation characteristics of the F/A-18C aircraft. Prior work by Godiksen suggests that the trailing end of the geometrically similar targeting forward looking infrared pod (TFLIR) is the likely source of shock waves that can adversely impact the trajectory of a recently released store. In our study five different modifications to the aft end of the ATFLIR were analysed using computational fluid dynamics (CFD). The two most promising designs, an ogive shape such as that used in artillery shells and rockets, and a simpler extended but truncated cone shape were then further investigated. The moments that these trailing shapes produced on an adjacent released store were compared. CFD analysis revealed that the simpler cone shape resulted in weaker shocks from the aft end of the pod with a resultant smaller adverse moment on the store. While there is an extensive history of using CFD to predict store separation behavior, results from our study should be compared with wind tunnel data in order to validate the CFD simulations.

Aero engine designs can have a life time of over 45 years, which is long enough for the understanding of environmental problems to change significantly. This places the aero engine designer in a position of uncertainty, as unforeseen environmental problems could affect the viability of a design. ‘Risk’ is used to describe future uncertainties that can lead to undesirable consequences. This paper presents a framework for environmental risk management that allows the designer to answer the question: what is the risk to a design from its environmental impacts over the life cycle? The framework provides a process for turning complex environmental business hazards into a form that can be used to develop mitigating actions within the design process. The paper demonstrates the framework through two examples and discusses findings, leading to conclusions on what is required to implement the framework into a business.

Recent work by Tsitas and Kingston(1) has demonstrated that an 8kg 6U CubeSat can be designed to perform Earth observation missions equivalent to those of 50-150kg microsatellites. Their design is reviewed and its commercial potential is compared to the 156kg RapidEye spacecraft. Three other commercial applications of this design are described. These are: a standard spacecraft for space scientists and astronomers; the spacecraft component of an N nation 5 spectral band disaster monitoring constellation and a night imaging satellite. Nanosatellites should now be considered for commercial missions previously thought to require microsatellites.

An experimental and numerical study was performed on a Clark Y aerofoil with a 10% chord leading edge Kruger flap to examine its aerodynamic performance at Reynolds numbers of 0·6 × 106, 1 × 106, and 1·6 × 106, to help to identify the forces and moments acting on a basic configuration. A detailed comparison of the numerical and experimental data is presented in this paper. The leading edge flap was effective at high angles of attack with an increase in CL of up to 18% over a conventional no flap configuration and delayed separation by up to 3°. The moments around the Kruger flap rotation point were calculated from the numerical analysis as an initial stage in the design of a UAV passive flap system and they are also presented in the paper.

This contribution addresses the subject of engine inlet barrier filters (IBF) for rotorcraft. The purpose of an IBF is to mitigate the risk of particle ingestion by the engine, a situation that can cause irrevocable damage to key components. The risk is significantly elevated during a brownout landing, in which the rotor wake of a descending helicopter interacts with loose ground sediment, causing the generation of a dust cloud. In such a condition, an IBF successfully removes particles from the engine-bound air at the expense of a pressure drop, which grows temporally. The increase in pressure drop is caused by but not limited to an increase in incident velocity and a decrease of mean particle size; however the rate may be slowed by operating the filter at a tangential angle to the flow. The current work presents findings of a parametric study into the factors affecting particle accumulation and the consequential loss of inlet total pressure.

Sandwich panels are widely used in the aerospace industry instead of solid plates due to their high flexural stiffness-to-weight and flexural strength-to-weight ratios. However due to the mismatch of properties between the face sheets and the core, stress concentrations can occur at the face sheet/core interfaces, often leading to delamination. One possible solution to this problem is the introduction of a graded core — a core in which the properties vary gradually from the face sheets to the core centre, eliminating any abrupt changes in properties. In this paper a 3D finite element method, fully validated through comparison with results from the literature and a 3D elasticity solution, is applied to modelling of sandwich panels with graded core. The approach makes use of graded elements to study the effect of varying the boundary conditions on the elastic deformation of the panel subject to uniformly distributed loading. Comparative analysis of stress and displacement fields in sandwich panels with homogeneous and graded cores is carried out under various combinations of simply supported, clamped and free edges.

The potential positive effects of ground proximity on the aerodynamic performance of a wing or aerofoil have long been established, but at transonic speeds the formation of shock waves between the body and the ground plane would have significant consequences. A numerical study of the aerodynamics of an RAE2822 aerofoil section in ground effect flight was conducted at freestream Mach numbers from 0·5 to 0·9, at a range of ground clearances and angles of incidence. It was found that in general the aerofoil’s lifting capability was still improved with decreasing ground clearance up until the point at which a lower surface shock wave formed (most commonly at the lowest clearances). The critical Mach number for the section was reached considerably earlier in ground effect than it would be in freestream, and the buffet boundary was therefore also reached at an earlier stage. The flowfields observed were relatively sensitive to changes in any given variable, and the lower surface shock had a destabilising effect on the pitching characteristics of the wing, indicating that sudden changes in both altitude and attitude would be experienced during sustained transonic flight close to the ground plane. Since ground proximity hastens the lower surface shock formation, no gain in aerodynamic efficiency can be gained by flying in ground effect once that shock is present.

A robust and efficient computational model has been developed which is capable of modelling the dynamic non-linear behaviour of Glare panels subjected to blast loadings. High strain rate material characterisation and modelling of interfacial debonding between adjacent sublaminates have been taken into consideration. Numerical model validation have been performed considering case studies of Glare panels subjected to a blast-type pressure pulse for which experimental data on the back face- displacement and post-damage observations were available. Excellent agreement of mid-point deflections and evidence of severe yield line deformation were shown and discussed against the performed blast tests. A further parametric study identified Glare as a potential blast attenuating structure, exhibiting superior blast potential against monolithic aluminium plates. The results were normalised and showed that for a given impulse, Glare exhibited a smaller normalised displacement, outperforming monolithic Aluminium 2024-T3 plates. It was concluded that further work needed to be carried out to take into account the influence of geometry (cylindrical structures), pre-pressurisation effects and boundary conditions

This note presents the experimental method and results from a series of desktop simulation tests designed to investigate the manual control characteristics of relatively inexperienced civil pilots; with an average age and experience of 24 years and 66 hours respectively. Increased encroachment into non-linear command gearing was found to make aggressive subjects resort to high levels of crossover regression. The combined effects of rate-limiting and non-linear command gearing were observed only for demanding tasks during which over-control was a typical feature.

Accident rates for night sorties, by helicopters operating offshore, are estimated to be five times higher than for daytime. Recent attempts to redress this problem include new instrument descent procedures and changes to helideck lighting. However, there still remains a need for pilots to transition from automated flight to a manually flown night visual segment during the arrival, and three recent accidents have again highlighted the dangers of this phase. In order to explore what factors affect safety during night visual segments 33 semi-structured interviews were administered to pilots from a number of companies, and grounded theory was employed to analyse the narratives. From this, a template was developed representing pilots’ perceptions of factors affecting safety, and the relationships among factors were hypothetically mapped. Visual spatial disorientation figured as the main concern of participants, stemming from substandard pilot communication, loss of scanning skills during summer months, autopilot limitations, and the requirement to fly to too low meteorological minima. Concerted actions and further areas of research have been proposed to the different stakeholders involved in offshore nighttime operations, going beyond ongoing safety initiatives.

Dynamic stall is a complex process encountered by an aerofoil in the unsteady flow environment such as a helicopter rotor in forward flight as well as a fixed wing aircraft in manoeuvres and in other unsteady situations. The onset of dynamic stall effectively determines the flight envelope of the helicopter. Significant effort are being made to develop CFD to capture the dynamic stall behaviour, however traditional engineering models based on lifting line theory still offer fast turn-round and broad understanding required for the rotor design process. This paper describes a new engineering model for dynamic stall, developed originally for wind turbine application at a typical Mach number of 0·12. The new dynamic stall model, with a better definition of stall onset, is based on improvements made to Beddoes’ original trailing edge stall model. This paper will describe and demonstrate the improvements in identifying both the stall-onset and the pitching moment break at high pitch rates, when being applied to a generic rotor aerofoil RAE 9651 at M = 0·3. Further validation against oscillatory tests and other Mach numbers are still required. However the study has provided sufficient confidence for it to be employed in a rotor analysis code.

A conventional inverse simulation does not accommodate control constraints; hence for aggressive manoeuvring flight conditions, where control inputs are close to the limits, these algorithms lose some of their applicability. A modification of the conventional inverse simulation technique that accommodates the onset of physical limits or constraints is proposed in this paper. In this way a process of constraints handling is incorporated into the inverse simulation algorithm. Therefore, the aim of this paper is to demonstrate that conventional inverse simulation can be improved in terms of the realism of the results by applying a predictive capability for applications involving manoeuvring flight. The paper gives details of the development of the predictive inverse simulation algorithm and helicopter model used and, by presenting examples of results calculated for pop-up and lateral realignment manoeuvres demonstrates that a ‘receding horizon’ predictive approach offers improvements in the realism of inverse simulation results.

Detached-eddy simulations for the 1303 UCAV geometry are undertaken aiming to investigate the flow physics of the interaction of a store with the flow inside a UCAV weapons bay. Advanced multi-block topologies had to be used to properly represent the planform of the UCAV and all the details of the weapon bay, including doors and hinges, while sliding meshes were needed to insert the store into the UCAV configuration. Results with an empty bay were encouraging for such complex configurations. Flow visualisation revealed the added turbulent content due to the door leading edges and hinges. The addition of a store in between the doors had little effect close to the front wall of the bay. However averaged flow-fields showed that the proximity of the shear layer to the apex of the store deflected it downwards into the bay and restricted its growth.

A reduced-order modelling (ROM) approach for predicting steady, turbulent aerodynamic flows based on computational fluid dynamics (CFD) and proper orthogonal decomposition (POD) is presented. Model-order reduction is achieved by parameter space sampling, solution space representation via POD and restriction of a CFD solver to the POD subspace. Solving the governing equations of fluid dynamics is replaced by solving a non-linear least-squares optimisation problem. The method will be referred to as LSQ-ROM method. Two approaches of extracting POD basis information from CFD snapshot data are discussed: POD of the full state vector (global POD) and POD of each of the partial states separately (variable-by-variable POD). The method at hand is demonstrated for a 2D aerofoil (NACA 64A010) as well as for a complete industrial aircraft configuration (NASA Common Research Model) in the transonic flow regime by computing ROMs of the compressible Reynolds-averaged Navier-Stokes equations, pursuing both the global and the variable-by-variable POD approach. The LSQ-ROM approach is tried for extrapolatory flow conditions. Results are juxtaposed with those obtained by POD-based extrapolation using Kriging and the radial basis functions spline method. As a reference, the full-order CFD solutions are considered. For the industrial aircraft configuration, the cost of computing the reduced-order solution is shown to be two orders of magnitude lower than that of computing the reference CFD solution.

A method to predict the effects of rotor icing on the flight characteristics of a UH-60A helicopter is presented. By considering both natural ice shedding and different types of ice accretion due to local temperature variations on the blade surface, an improved rotor icing model was developed. Next, the effects of icing on rotor force, torque and flapping were incorporated in a nonlinear helicopter dynamic model. Based upon icing design envelopes in cumuliform clouds, trim and stability characteristics were studied. Further development of the helicopter state-space model allowed control and handling qualities characteristics to be investigated with variation of the three icing-related cloud variables (atmospheric temperature, liquid water content, and median volumetric diameter). Results indicated that this method of evaluating rotorcraft icing is both feasible and useful.