A simplified method to analyse propellers based on vortex theory is presented. Small-angle approximations are implemented to eliminate the need for iteration in the determination of the induced angle-of-attack. A stall model is developed and combined with analytic relations describing the blade aerofoil characteristics, eliminating the need for look-up tables of aerofoil behaviour. The method is also extended to serve as an optimal propeller design tool. Comparisons of the approach with experiment are presented for validation as an analysis tool. Use of the theory as a design tool is also demonstrated through contrast with an existing blade design methodology.

The unsteady air flow over and around the helicopter landing deck of a naval vessel is known to cause high pilot workload and to limit the helicopter's operational envelope for launch and recovery. Previous research has suggested that modifications to the ship's hangar edges can beneficially modify the flow over the deck. This paper examines the effectiveness of five hangar-edge modifications using computational fluid dynamics–generated airwakes and flight mechanics modelling, as well as piloted flight trials in a motion-base simulator. Results are presented, in terms of unsteady helicopter loads and pilot workload ratings, for modifications to the windward vertical rear edge of the hangar and with an oblique wind. The results demonstrate that while the airwake can be altered by superstructure modifications, the integrated effect of the altered airwake on the entire helicopter does not necessarily give the desired result; indeed of the five modifications tested, two were seen to be beneficial while three caused an increase in pilot workload compared with the unmodified hangar. Overall, the paper shows that the airwake can be modified by superstructure design changes, and that the effect on the helicopter can be measured through modelling and simulation. It is also demonstrated that making judgements on the severity of the airwake based on the aerodynamic flow field alone can be misleading. The benefit of these simulation tools is that they can be used during the ship design process to evaluate the effect of the superstructure aerodynamics, rather than wait until after the ship is built.

A method based on the Eulerian two-phase flow theory to numerically simulate three-dimensional rime ice accretions on an aircraft wing is presented in this paper. The governing equations for supercooled droplet motion under Eulerian framework are established using the droplet pseudo-fluid model. A permeable wall boundary condition is proposed to simulate the phenomenon of droplets impinging on the wing in solving the governing equations for droplets. The local droplet collection efficiency is readily obtained from the droplet flowfield solution in the control volume adjacent to the wing surface. The rime ice accretion can be simulated under the assumption that the droplets freeze immediately as they impinge on the wing surface since the environment temperature is low enough (typically below –15°C). A method to build the ice shape is proposed based on the assumption that ice grows in the direction normal to the wing surface. The rime ice accretion on a GLC-305 swept wing model under some specific conditions has been simulated to validate the present method. Furthermore, different flight conditions, namely, different angles of attack and different angles of sideslip, have been dealt with to investigate their effects on rime ice accretion as well as the corresponding aerodynamic effects.

This paper introduces an observer strategy, namely a Sliding Mode Observer (SMO), to realise the fault detection and estimation of general uncertain non-linear systems. The use of a non-linear observer is considered for monitoring the states of a high incidence research model (HIRM) aircraft system. For a special class of Lipschitz non-linear system, a fault reconstruction scheme is presented where the reconstructed signal can approximate the fault signal to any accuracy. The proposed method is based only on the available plant input-output information and can be calculated online. Moreover, the globally asymptotic stability of the closed-loop system is mathematically proved. Finally, an HIRM aircraft system example is given to illustrate the efficiency of the proposed approach.

This paper is intended as a general introduction to the requirements for future passenger aircraft design. The needs of the 21st century are addressed to meet the important requirements of the customer airlines as well as those of the general public. In particular, the impact on two traditional major requirements are reviewed, the Design Mission and the operating costs. The effect of aircraft on the environment and the increases in the cost of fuel will have a substantial effect on the way future aircraft are optimised. These demands are summarised before moving on to the basic equations affecting how the aircraft design must respond. Very similar targets driving research work have been set in both Europe and the United States, and some of the new technologies that we can expect to be incorporated are outlined. Finally, a glimpse is given of the possible future aircraft configurations we may see in the skies in response to the new demands.

The cycloidal propeller for a Micro-Aerial Vehicle (MAV)-scale cyclogyro in hover was studied using a 2D Reynolds-averaged Navier-Stokes equations solver. The effects of the blade dynamic stall, parallel Blade Vortex Interaction (BVI), inflow variation and flow curvature were discussed, based on the results of numerical simulation. The results from the 2D Computational Fluid Dynamics simulation indicated that the blade of the cycloidal rotor is actually performing a pitching oscillation, if observed in a moving reference frame. The dynamic stall vortices shed from the upstream blade cause intense parallel BVI on the downstream blade. The interaction will induce upwash and downwash on the downstream blade. This changes the effective reduced frequency and actually delays the stall of the blade, which is beneficial to the thrust generation. There is also strong downwash in the rotor cage and it changes the inflow velocity experienced by the blade. The downwash and flow curvature can either be beneficial or harmful to the thrust generation. The combined effects of dynamic stall, parallel BVI, inflow variation and flow curvature cause large aerodynamic force peaks and ensure the cycloidal rotors work at very low rotation speeds with high thrust. This guarantees that the cycloidal rotors possess at least the same level of hover efficiency as screw propellers.

Model-Order Reduction (MOR) is an important technique that allows Reduced-Order Models (ROMs) of physical systems to be generated that can capture the dominant dynamics, but at lower cost than the full order system. One approach to MOR that has been successfully implemented in fluid dynamics is the Eigensystem Realization Algorithm (ERA). This method requires only minimal changes to the inputs and outputs of a CFD code so that the linear responses of the system to unit impulses on each input channel can be extracted. One of the challenges with the method is to specify the size of the input pulse. An inappropriate size may cause a failure of the code to converge due to non-physical behaviour arising during the solution process. This paper addresses this issue by using piston theory to estimate the appropriate input pulse size.

The aviation industry is dominated by the domain of heavier-than-air, fixed-wing, subsonic flight, and central to any design in this domain is the wing itself. One of the earliest debates in aviation still centres around the usefulness of the wing volume. On the one hand it is held that the wing, as an inevitable necessity, should provide the volume also for the payload. On the other, it is argued that more efficient wings do not even have sufficient volume for the entire wing structure. This work proposes precise definitions of the Wing Density and the Inflation Factor, two parameters that can quantitatively reflect the economic and technological trends in aviation. The wing volume of a hypothetical Ideal Wing is derived from the Operational Parameters of any given Flight Objective and compared to the volume requirement of that flight objective. We conclude that the dominant aircraft configuration of the future is likely to remain within the same family of the current dominant configuration, in conflict with some older predictions.

The drag coefficient and the laminar-to-turbulent transition for the aerofoil component of a wing model are optimised using an adaptive upper surface with two actuation points. The effects of the new shaped aerofoils on the global drag coefficient of the wing model are also studied. The aerofoil was optimised with an ‘in-house’ genetic algorithm program coupled with a cubic spline aerofoil shape reconstruction and XFoil 6.96 open-source aerodynamic solver. The wing model analysis was performed with the open-source solver XFLR5 and the 3D Panel Method was used for the aerodynamic calculation. The results of the aerofoil optimisation indicate improvements of both the drag coefficient and transition delay of 2% to 4%. These improvements in the aerofoil characteristics affect the global drag of the wing model, reducing it by up to 2%. The analyses were conducted for a single Reynolds number and speed over a range of angles of attack. The same cases will also be used in the experimental testing of the manufactured morphing wing model.

Wind-tunnel tests of a heavy-class helicopter model were carried out to evaluate the effectiveness of several components optimised for drag reduction by computational fluid dynamics analysis. The optimised components included different hub-cap configurations, a fairing for blade attachments and the sponsons. Moreover, the effects of vortex generators positioned on the back ramp were investigated. The optimisation effect was evaluated by comparison of the drag measurements carried out for both the original and the optimised helicopter configurations. The comprehensive experimental campaign involved the use of different measurement techniques. Indeed, pressure measurements and stereo particle image velocimetry surveys were performed to achieve a physical insight about the results of load measurements. The test activity confirms the achievement of an overall reduction of about 6% of the original model drag at cruise attitude.

This design study applied parameterisation to rotor blade for improved performance. In the design, parametric equations were used to represent blade planform changes over the existing rotor blade model. Design variables included blade twist, sweep, dihedral and the radial control point. Updates to the blade structural properties with changes in the design variables allowed accurate evaluation of performance objectives and realistic structural constraints – blade stability, steady moments (flap bending, chord bending and torsion) and the high-g manoeuvre pitch link loads. Performance improvement was demonstrated with multiple parametric designs. Using a parametric design with advanced aerofoils, the predicted power reduction was 1.0% in hover, 10.0% at μ = 0.30 and 17.0% at μ = 0.40, relative to the baseline UH-60A rotor, but these were obtained with a 35% increase in the steady chord bending moment at μ = 0.30 and a 20% increase in the half peak-to-peak pitch link load during the UH-60A UTTAS manoeuvre. Low vibration was maintained for this design. More rigorous design efforts, such as chord tapering and/or structural redesign of the blade cross section, would enlarge the feasible design space and likely provide significant performance improvement.

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.

When the stagnation pressure of a perfect gas increases, the specific heat and their ratio do not remain constant anymore and start to vary with this pressure. The gas doesn't stay perfect. Its state equation changes and it becomes a real gas. In this case, the effects of molecular size and intermolecular attraction forces intervene to correct the state equation. The aim of this work is to determine the effect of stagnation pressure on the thermodynamic, physical and geometrical supersonic flow parameters in order to find a general form for real gas. With the assumptions that Berthelot's state equation accounts for molecular size and intermolecular force effects, expressions are developed for analysing the supersonic flow for thermally and calorically imperfect gas lower than the dissociation molecules threshold. The design parameters of the supersonic nozzle-like thrust coefficient depend directly on the stagnation parameters of the combustion chamber. The application made for air. A computation of error was made in this case to give a limit of the perfect gas model compared to the real gas model.

The past century has witnessed the rise and maturity of the flying machine, starting with the Wright brothers flyer to today’s modern passenger aircrafts and warfighters. At the start of this century, yet another achievement in flying vehicle technology was seen with the launch of the Boeing 787 aircraft, which has a significant portion by weight of polymer matrix fibre composites. This paper, therefore, addresses the effects of the manufacturing process of fibre reinforced polymer matrix composites on mechanical performance. Computations are carried out using the Finite Element (FE) method at the microscale where Representative Volume Elements (RVEs) are analysed with Periodic Boundary Conditions (PBCs). Straight fibre pre-preg-based composites and textile composites are considered. The commercial code ABAQUS is used as the solver for the FE equations, supplemented by user-written subroutines. The transition from a continuum to damage/failure is effected by using the Bažant-Oh crack band model, which preserves mesh objectivity. Results are presented for RVEs that are first subjected to curing and subsequently to mechanical loading. The effect of the fibre packing randomness on the microstructure is examined by considering multifibre RVEs where fibre volume fraction is held constant but with random packing of fibres. Plain weave textile composites are also cured first and then subjected to mechanical loads. The possibility of failure is accommodated throughout the analysis – failure can take place during the curing process even prior to the application of in-service mechanical loads. The analysis shows the differences in both the cured RVE strength and stiffness, when cure-induced damage has and has not been taken into account.

One of the important tasks that air traffic management services are faced with today is the task of maximising airport capacity. This can be achieved at the tactical level through proper organisation of air traffic around an airport. In recent years, many methods and algorithms for scheduling aircraft landings have been developed; they take into account various optimisation goals. The aim of this paper was to create a method that would allow one to evaluate landing aircraft sequences resulting from these control algorithms, especially in the presence of random disturbances. This method involves modelling the landing aircraft sequence by using Petri nets. The model and the computer tool that have been developed make it possible to take into account different kinds of disturbances and examine the effectiveness of various control strategies under these conditions. This paper presents two experiments that test disturbances with different characteristics and of different intensities. It has been shown that small but more frequent disturbances lead to the worsening of evaluation scores for a given sequence to a lesser extent than rare but larger disturbances. This is particularly important for control algorithms in which the focus is on high aircraft density. If the type of particular disturbances is properly assessed, then it will be possible to assist the decision-maker (air traffic controller) by providing him/her with quantitative evaluations of possible solutions.

Shock-tunnel experiments are carried out to study the strong interaction between an impinging shock wave and boundary layer on a flat plate, accompanied by large separation bubble with a length comparable to the distance of the location of shock impingement from the leading edge of the plate. For nominal freestream Mach numbers ranging from 6 to 8.5, moderate to high total enthalpies of 1.3MJ/kg to 6MJ/kg are simulated in the Indian Institute of Science's hypersonic shock tunnels HST-2 (a conventional Hypersonic Shock Tunnel) and Free Piston Shock Tunnel (FPST) with freestream Reynolds numbers ranging from 4 × 106/m to 0.3 × 106/m. The strong impinging shock is generated by a wedge (or shock generator) at an angle of 30.96° to the freestream. From the time-resolved Schlieren flow visualisations using a high-speed camera and surface pressure measurements on the flat plate using fast response sensors, a statistically steady flow field with a large separation bubble was established within the short test time of the shock tunnels (around 600µs in HST-2 and 300µs in FPST). The role of various parameters on the interaction – Mach number, location of shock impingement and flow total enthalpy – are investigated from the measured separation length and surface pressure distribution. For the nominal Mach number of 8.5, with shock impingement at 100mm from the leading edge, the separation length increased from 60mm to 70mm as the total enthalpy is increased from 1.6MJ/kg to 2.4MJ/kg; whereas it dropped drastically to 30-40mm at 6MJ/kg. This is due to the prominence of real gas effects at higher enthalpies.

Flight testing is both vital for collecting data for aeronautic research and at the same time fascinating for its contributors. Taking a glider as a versatile test bed example, this paper presents a transnational measurement campaign within the framework of a collaborative project funded by the European Commission. This project Advanced In-Flight Measurement Techniques 2 (AIM²) is a follow-up of Advanced In-Flight Measurement Techniques (AIM) and dedicated to developing and enhancing promising optical metrology for various flight test applications up to an industrial level.

The Image Pattern Correlation Technique (IPCT) and infrared thermography (IRT) are two of these modern non-intrusive measurement methods that were further developed and applied to the glider test bed within the scope of AIM². Focusing on optical deformation measurements with IPCT the experimental setup, the flight testing and results are summarily discussed. Gliders are not commonly used flight test platforms, which is why this contribution concludes with some lessons learned in general and especially related to the presented application. The experience to be shared with the flight testing community addresses equipment preparation, data collection and processing as well as how to meet official requirements and perform test flight operations in a dense controlled airspace.

A simulation study was conducted in order to investigate the influence of vortex deformation on wake encounter characteristics. Wake vortices tend to be strongly deformed during the decay process, depending on the atmospheric conditions in terms of turbulence and thermal stratification. For quantification of the influence of vortex deformation, encounters of an aircraft of the ‘Medium’ category behind a generator aircraft of the ‘Heavy’ category were simulated with straight vortices and with realistically deformed vortices derived from large-eddy simulations. All relevant parameters that influence the encounter characteristics, such as encounter angles and positions, were varied within a wide range. In order to cover all kinds of vortex deformation, encounters with different vortex ages from 16-136 seconds were simulated. Hence, all relevant phases during the vortex decay from nearly straight and wavy vortices to vortex rings were considered.

The parameter variation study revealed that on average the impact on the encountering aircraft is less with deformed vortices than with straight vortices of comparable strength. Especially with vortex rings, the encountering aircraft is exposed to a much smaller impact. However, the results also show a larger aircraft response during encounters with wavy vortices just prior to vortex linking. The maximum aircraft response with wavy vortices is stronger than with straight vortices of comparable strength. Also, the strongest encounters occur under greater encounter angles with deformed vortices than with straight ones.

The flow and noise created by sawtooth trailing-edge serrations has been studied experimentally at a low Reynolds number. Experiments have been performed on a flat-plate model with an elliptical leading edge and an asymmetrically bevelled trailing edge at Reynolds numbers of Rec = 1 × 105–1.3 × 105, based on chord. Wide serrations with a wavelength (λs) to amplitude (2h) ratio of λs/h = 0.6 were found to reduce the overall sound pressure level by up to 11dB. In contrast, narrower serrations with λs/h = 0.2 produce tonal noise and increase the overall noise level by up to 4dB. Intense vortices across the span of the trailing edge with narrow serrations are shown to be the source of tonal noise. Wide serrations reduce turbulent velocity fluctuations at low frequencies which explains the lower radiated noise. The narrow serrations that produce low Reynolds number tonal noise were shown previously to be effective at higher Reynolds numbers (Rec > 2 × 105), demonstrating that care is needed to fully understand the flow field over serrations for all intended operating conditions.

This paper compares two periodic control methods, the optimal H 2 and the periodic static output feedback (POF), to reduce the helicopter rotor vibrations. Actively twisted blades with Macro-Fibre Composite (MFC) piezoelectric actuators are used. The design model is based on a simplified aerodynamic model and on a multi-body model of the Bo 105 isolated rotor with the original blades replaced by actively twisted ones. The performance of the two controllers in alleviating hub loads is verified with improved simulations based on a free-wake model.