The paper presents the estimation of longitudinal aerodynamic parameters from simulated as well as real flight data using an unsteady downwash model. The task of parameter estimation was accomplished using the Maximum Likelihood (ML) method in the frequency domain for the Hansa-3 aircraft. The presented work pertains to the aircraft having a semi-sweptback wing configuration, however, the work reported earlier considered a rectangular wing configuration. A simple vortex system was used to model downwash effects in the longitudinal equations of motion. First, it was established that the indicial lift function had a minor or negligible influence on aircraft motion and on the estimated aerodynamic parameters using simulated data for the configuration of Hansa-3 aircraft. Next, two parameter extraction algorithms were used for parameter estimation using simulated data (with unsteadiness only in downwash) and real flight data (Hansa-3 aircraft). The first algorithm included unsteadiness only in downwash while the second neglected all the unsteady effects. It was observed from the results obtained that if unsteady effects were omitted from the parameter estimation algorithm, the extracted parameters included the effects of unsteady aerodynamics. Therefore, application of the extraction algorithm in the two modes provides a convenient means of separating steady-state and acceleration derivatives.

Results of a series of oscillatory dynamic stall tests of a rotor aerofoil fitted with a pulsed, trailing-edge flap are presented. Flap deflection amplitude, motion profile, duration and starting phase were investigated to assess the potential of the flap for mitigating the adverse effects of dynamic stall, which is one of the limiting factors for rotor blades on the retreating side of a helicopter rotor. The tests were a continuation of the investigations by Ref. 1 who used a computational fluid dynamics method on a symmetric NACA section, and our results broadly confirm their conclusions by experimental test, using a modern rotor section. The results presented in this paper also confirm the observations from experimental work by Refs 2 and 3, which were undertaken at lower Reynolds number and with a larger flap. In the present study, the flap mitigates the high negative pitching moment and negative pitch damping seen in dynamic stall by strong suction being generated over the aerofoil lower surface, and it is the modification to the lower surface shape by the flap that creates this effect. The dynamic stall vortex acts to enhance the lower surface suction, and careful flap phasing and flap motion profile shaping can make the control more effective.

The dynamics of the wing tip vortex in the near-field of a NACA 0012 aerofoil has been analysed by means of flow visualisations in a water tunnel. Different axial distances near the wing up to four chords, Reynolds numbers up to 42,000 and three angles-of-attack are studied to characterise the behaviour of the vortex meandering. The spatio-temporal vortex centre positions show distorted elliptical shapes in a (x,y)-plane. The Reynolds number has no significant influence on the axial evolution of the meandering amplitude. In addition, the flow visualisations obtained with a low speed camera are analysed by the singular value or proper orthogonal decomposition. Thus, the most energetic displacement modes are obtained. The frequency associated to these modes is computed by FFT. In all the cases studied, our results show that the most unstable mode corresponds to the azimuthal wavenumber |n| = 1 in the so-called Kelvin helical modes and the frequency is lower or close to 1Hz.

An integrated artificial immune system-based scheme that can operate over extended areas of the flight envelope is proposed in this paper for the detection and identification of a variety of aircraft sensor, actuator, propulsion, and structural failures/damages. A hierarchical multi-self strategy has been developed in which different self configurations are selected for detection and identification of specific abnormal conditions. Data collected using a motion-based flight simulator were used to define the self for a wide area of the flight envelope and to test and validate the scheme. The aircraft model represents a supersonic fighter, including model-following direct adaptive control laws based on non-linear dynamic inversion and artificial neural network augmentation. The proposed detection scheme achieves low false alarm rates and high detection and identification rates for all the categories of failures considered.

The GOAHEAD (Generation of an Advanced Helicopter Experimental Aerodynamic Database for CFD code validation) consortium was created in the frame of an EU-project in order to create an experimental database for the validation of 3D-CFD and comprehensive aeromechanics methods for the prediction of unsteady viscous flows. This included the rotor dynamics for complete helicopter configurations, i.e. main rotor – fuselage – tail rotor configurations with emphasis on viscous phenomena like flow separation and transition from laminar to turbulent flow. The wind tunnel experiments have been performed during two weeks in the DNW-LLF on a Mach-scaled model of a modern transport helicopter consisting of the main rotor, the fuselage, control surfaces and the tail rotor. For the sake of controlled boundary conditions for later CFD validation, a closed test section has been used. The measurement comprised global forces of the main rotor and the fuselage, steady and unsteady pressures, transition positions, stream lines, position of flow separation, velocity profiles at the test section inlet, velocity fields in the model wake, vortex trajectories and elastic deformations of the main and tail rotor blades.

This paper explores the issues related to the application of fuzzy logic techniques to aid the process of weight estimation for aircraft structures at preliminary design stages. The focus lays on the design of a neuro-fuzzy system for the weight analysis, through the use of the Neuro-Fuzzy Function Approximator (NEFPROX) algorithm. The paper introduces a three-level process designed around Mamdani fuzzy systems derived through NEFPROX and analyses its application to the sizing and weight estimation of spoiler attachment ribs. Problems such as structure parameterisation, variable selection and model optimisation are explored as part of the process validation phase on the selected structural case study. The model performance is evaluated with respect to modelling accuracy, generalisation capabilities and the interdependencies between the variables the model is able to derive from the analysis of the given structural examples. The results are then compared to the performance obtained from the application of Takagi-Sugeno-Kang (TSK) fuzzy models derived using Adaptive Network-based Fuzzy Inference System (ANFIS) on the same sample of spoiler attachment ribs. Results highlight the benefits of adopting NEFPROX for the derivation of fuzzy systems to be applied in structural weight estimation problems, from the point of view of both accuracy of the approximation provided and the quality and interpretability of the final rulebase derived by the system.

This paper deals with the design and implementation of the nonlinear control algorithm for the attitude tracking of a four-rotor helicopter known as quadrotor. The choice of this algorithm, which is based on the Backstepping sliding mode technique with adaptive gain, was justified by the fact that it ensures robustness with respect to modelling errors and external disturbances while reducing the chattering phenomenon caused by the sign function in the first order sliding mode based controllers with fixed gains. In order to show the effectiveness of the controller, experimental tests were carried out on a quadrotor. The obtained results show good performance of the proposed controller in terms of stabilisation, tracking and robustness.

Micro air vehicles (MAVs) are typically of low mass and moment of inertia and have flight speeds comparable to birds and the larger insects. Such craft traverse the lower levels of the atmospheric boundary layer (ABL) which is a significantly different environment than that experienced by larger manned aircraft, which spend the majority of their time in relatively clean air and fly at speeds significantly higher than typical wind speeds in the ABL. Here a new series of measurements dedicated to understanding spatial and temporal velocity fields that MAVs experience are presented. Atmospheric wind measurements were taken by sampling four multi-hole dynamic pressure probes spanned perpendicular to the oncoming wind at spans of between 0·014m and up to 0·45m. It was noted that the variation of both longitudinal velocity and flow pitch angle against spacing followed a fractional power law and as such large variations were present even for the smallest inter-probe separations. This effect is thought to explain the increasing piloting difficulties experienced in maintaining good roll control for decreasing scales of craft.

Three-component laser Doppler anemometry (LDA) has been used to measure the complex flow distributions over the suction surface of a symmetrical 40°swept wing model at an angle of incidence of 9° and for the relatively low Reynolds number (based on the root chord) of 2·1 × 105. Emphasis was placed on the separation and reattachment of the boundary-layer flow, and the formation of vortical structures.

The experimental programme now described utilised a large wind tunnel and several advanced measurement techniques to produce an unusually detailed collection of results. Thus, data are presented here for the behaviour of the three-dimensional boundary layer developing on the suction surface at positions from 30% to 90% semi-span. These results show the spatial variations of the time-averaged mean and fluctuating velocity components in three orthogonal directions, including the distributions of the normal and shear stress levels. Further analysis has enabled the time-averaged vortical structures to be identified and compared with the results of surface flow visualisation.

Flow over a swept wing poses many challenges for the computational modeller. The underlying aim, here, therefore, was to produce a data archive for the validation of computer predictions. As will be shown, some of the experimental results have already been used in this way and the data base is available for use by other workers.

The Swiss Airforce is operating F/A-18C/D Aircraft since 1997. Since the aircraft’s structural design is different from the version operated by the US Navy it was necessary to carry out a structural integrity study (ASIP) which was done by The Boeing Company in St. Louis. To validate this study a full scale fatigue test facility was build at RUAG and operated from 2003 to 2005. When operating this facility difficulties were encountered with the aerodynamic loads data provided by Boeing (insufficient, not well documented, questionable data). As a result RUAG looked for alternative methods to provide the aerodynamic loads, and a large investment was made in the development of a Computational Fluid Dynamics (CFD) tool. The Navier Stokes Multi Block (NSMB) solver, which was developed in an international collaboration, was adopted. In a first phase the code was validated by comparing results of CFD calculations with wind-tunnel results, results from literature and flight test data results. In the second phase, discussed in this paper, a Fluid Structure Interaction (FSI) tool was developed to permit unsteady aero-elastic simulations. Particular attention is focused on the vertical tail since this component of the F/A-18 fighter is very sensitive to fatigue due to unsteady loads generated by buffeting phenomena.

In the current absence of comprehensive and generic UAS airworthiness regulations, the development of UAS and their introduction into non-segregated airspace pose significant challenges to the UAS industry and regulators. This paper reports on a research study that considered the problem, from an engineering perspective, beyond the limits of the airworthiness of the aircraft and remote control station. The study introduces the concept of UAS operability, which includes the safe and reliable functioning of the UAS as a system, the airworthiness of its airborne sub-systems, and the safe and reliable functioning of its non-airborne sub-systems and functional payloads. The regulatory domain for UAS operability is described and the study establishes a generic and comprehensive UAS operability framework. The framework was validated by populating its elements with engineering criteria that can be used by the UAS engineering domain for the development of engineering specifications and as guidance towards achieving the safe and reliable functioning of UAS.

Operating micro-UAVs autonomously in complex urban areas requires that the guidance algorithms on-board are robust to changes in the operating environment. Limited endurance capability demands an optimal guidance algorithm, which will change as the environment does. All optimal path-planning routines are computationally intensive, with processor load a function of the environmental complexity. This paper presents a new algorithm, the reactive route selection algorithm, for storing a bank of optimal trajectories computed off-line and blending between these optimal trajectories as the operating environment changes. An example is presented using a mixed-integer linear program to generate the optimal trajectories.

Accurate and reliable airdata systems are critical for aircraft flight control system. In this paper, both extended Kalman filter (EKF) and unscented Kalman filter (UKF) based various multi sensor data fusion methods are applied to dynamic manoeuvres with rapid variations in the aircraft motion to calibrate the angle-of-attack (AOA) and angle-of-sideslip (AOSS) and are compared. The main goal of the investigations reported is to obtain online accurate flow angles from the measured vane deflection and differential pressures from probes sensitive to flow angles even in the adverse effect of wind or turbulence. The proposed algorithms are applied to both simulated as well as flight test data. Investigations are initially made using simulated flight data that include external winds and turbulence effects. When performance of the sensor fusion methods based on both EKF and UKF are compared, UKF is found to be better. The same procedures are then applied to flight test data of a high performance fighter aircraft. The results are verified with results obtained using proven an offline method, namely, output error method (OEM) for flight-path reconstruction (FPR) using ESTIMA software package. The consistently good results obtained using sensor data fusion approaches proposed in this paper establish that these approaches are of great value for online implementations.

The influence of angle-of-attack and vortex-aerofoil vertical miss-distance on the aerofoil-vortex mechanism of interaction is investigated numerically using the large-eddy simulation approach. The study concerns the aerofoil-vortex interaction (AVI) phenomena encountered in the rotorcraft aerodynamics. The studies are performed for a Reynolds number of 1·3 × 106 based on the chord (c = 0·2m) of the aerofoil and free-stream velocity U = 100ms−1. The studies are conducted for three different angles-of-attack (α = 0°, and α = 5°, α = 10°) and vertical miss-distances (h = 0·0m, h = −0·01, h = −0·02m) respectively. The present study shows that the magnitude of lift coefficient and flow separation, at the instant of interaction, decay with the increase of angle-of-attack and vertical miss-distance.

Flying under degraded visual environments is a challenging task in helicopter operations. The brownout phenomenon, caused by the rotor downwash on a dusty or sandy underground has caused several critical incidences and accidents in the past. In contrast to an extensive free wake analysis of the rotor vortices, this paper presents a computer graphical solution to simulate brownout conditions. The animation approach builds upon a particle system with a basic aerodynamic assumption of a permanent ground vortex and additional turbulence. The created brownout effect module is a part of the RealTime Image Generator software, used in DLR’s ground based helicopter flight simulator. The object oriented structure of the underlying application programming interface OpenSceneGraph facilitates the development of a generic effect interface for future improvements. An overview of the developed module, the particle system engine and the physical dependencies is provided. To overcome the challenge of capturing the visual realism, a pilot test course with brownout typical tasks is presented. Criteria focussing on the visual and physical parameters are defined to build up a questionnaire. It is used to improve the visual standard and to meet the future pilot’s needs.

An experimental study has been carried out to understand jet flow development from plain and grooved rectangular nozzles of aspect ratio 2:1 using two-component hotwire anemometry. Grooves of square configuration (side 4mm) and length 5mm were introduced in the (i) minor-axis, (ii) major-axis and, (iii) in both minor- and major-axes directions. The equivalent diameter of the plain rectangular nozzle is 37·5mm. Studies were carried out for a nominal jet exit velocity of 20ms−1 and for Reynolds number based on equivalent diameter of 54,000. The introduction of grooves in either plane does not show any influence on the potential-core length but results in faster jet-decay thereafter. It is observed that the grooves when introduced in minor-axis direction inhibit the jet growth in that plane while promoting the jet growth along major-axis plane and hence, delays the phenomena of axis-switching. However when introduced in major-axis direction, the grooves promote jet growth along major-axis plane while inhibiting jet-growth in minor-axis plane. Cross-sectional contours of mean-velocity suggest that the grooves modify the process of overall jet development significantly in the plane in which they are introduced relative to the plain jet and hence, significantly affect the axis-switching location in each case.

Research efforts at Bombardier Aerospace have investigated design constraints that need to be addressed when considering the design of a future unconventional fuselage concept. Using a composite design platform integrated with structural health monitoring (SHM) technologies, the design space is broadened and there are many more unknowns to investigate from our current knowledge of the subject to date. The principal idea is to identify structural zones and quantify stress levels in areas of concern, referred to as ‘hot-spots’, by integrating an SHM system at a conceptual design level. This new concept would require an assessment of benefits and detriments in order to evaluate the certification processes, impact on maintenance, operation, and ownership costs.

The integration of an SHM system using onboard sensors introduces many challenges, such as the requirement for multiple sensors in the structure and additional systems weight. The investigation shows how the aircraft structural design would be impacted and how design, stress, supply-chain, manufacturing, and the systems departments need to be harmonised in order to design a feasible and integrated SHM-structural fuselage concept.

Space debris represents a significant risk to satellite operations, due to the possibility of damaging or catastrophic collisions. Consequently, many satellite operators screen the orbiting population for close approaches with their on-orbit assets and a public conjunction assessment service, Satellite Orbital Conjunction Reports Assessing Threatening Encounters in Space (SOCRATES), generates close approach predictions on a daily basis for all satellite payloads in the catalogue. These screening capabilities are used to inform operational decisions relating to risk mitigation but it is anticipated that the demands placed on these services will increase as debris becomes more prolific. This hypothesis is explored in a preliminary analysis of conjunction data for the years 2004 to 2009 and a new ‘Business as usual’ study using the Debris Analysis and Monitoring Architecture for the Geosynchronous Environment (DAMAGE) model. The results suggest a 50% increase in the number of close approaches reported by SOCRATES (or its equivalent) within the next ten years. By 2059, daily conjunction reports could contain over 50,000 close approaches below 5km, affecting the demands placed on tracking facilities and satellite resources.

The light gyroplane is a class of aircraft popular with amateur constructors and pilots. As a result, there is limited design guidance available since formal technical resources are not available to the community. Rule-of-thumb, intuition and historical experience tend to influence design evolution. Empennage configuration is a prime example of this paradigm, and the objective of this Paper is to explore those factors that influence horizontal stabiliser effectiveness with particular reference to dynamic stability. An individual-blade rotorcraft mathematical model is coupled with a vorticity-based flowfield code, necessary to capture the highly interactional aerodynamics associated with empennage location at the rear of the airframe. A parametric study of horizontal stabiliser location shows that maximum benefit from the energising influence of the propeller slipstream is obtained if the surface is placed near the edge of the propeller wake. Further, traditional design parameters such as tail volume ratio offer an incomplete indicator of empennage effectiveness without consideration of airframe blockage and propeller slipstream. It is concluded that empennage sizing calculations can be straightforward, but require due consideration of the impact of the close-coupled nature of the vehicle on stabilising surface aerodynamic effectiveness.