The flight of barnacle geese at airspeeds representing high-speed migrating flight is investigated using experiments and simulations. The experimental part of the work involved the filming of three barnacle geese (Branta Leucopsis) flying at different airspeeds in a wind tunnel. The video footage was analysed in order to extract the wing kinematics. Additional information, such as wing geometry and camber was obtained from a 3D scan of a dried wing. An unsteady vortex lattice method was used to simulate the aerodynamics of the measured flapping motion. The simulations were used in order to successfully reproduce the measured body motion and thus obtain estimates of the aerodynamic forces acting on the wings. It was found that the mean of the wing pitch angle variation with time has the most significant effect on lift while the difference in the durations of the upstroke and downstroke has the major effect on thrust. The power consumed by the aerodynamic forces was also estimated; it was found that increases in aerodynamic power correspond very closely to climbing motion and vice versa. Root-mean-square values of the power range from 100W to 240W. Finally, it was observed that tandem flying can be very expensive for the trailing bird.

Reduced order models representing the dynamic behaviour of symmetric aircraft are well known and can be easily derived from the standard equations of motion. In flight testing, accurate measurements of the dependent variables which describe the linearised reduced order models for a particular flight condition are vital for successful system identification. However, not all the desired measurements such as the rate of change in vertical velocity (Ẇ) can be accurately measured in practice. In order to determine such variables two possible solutions exist: reconstruction or differentiation. This paper addresses the effect of both methods on the reliability of the parameter estimates. The methods are used in the estimation of the aerodynamic derivatives for the Aerosonde UAV from a recreated flight test scenario in Simulink. Subsequently, the methods are then applied and compared using real data obtained from flight tests of the Cranfield University Jetstream 31 (G-NFLA) research aircraft.

This paper presents a mathematical approach for the simulation of rotor-fuselage aerodynamic interaction in helicopter aeroelasticity and flight dynamics applications. A Lagrangian method is utilised for the numerical analysis of rotating blades with nonuniform structural properties. A matrix/vector-based formulation is developed for the treatment of elastic blade kinematics in the time-domain. The combined method is coupled with a finite-state induced flow model, an unsteady blade element aerodynamics model, and a dynamic wake distortion model. A three-dimensional, steady-state, potential flow, source-panel method is employed for the prediction of induced flow perturbations in the vicinity of the fuselage due to its presence in the free-stream and within the rotor wake. The combined rotor-fuselage model is implemented in a nonlinear flight dynamics simulation code. The integrated approach is deployed to investigate the effects of rotor-fuselage aerodynamic interaction on trim performance, stability and control derivatives, oscillatory structural blade loads, and nonlinear control response for a hingeless rotor helicopter modelled after the Eurocopter Bo105. Good agreement is shown between flow-field predictions and experimental measurements for a scaled-down isolated fuselage model. The proposed numerical approach is shown to be suitable for real-time flight dynamics applications with simultaneous prediction of structural blade loads, including the effects of rotor-fuselage aerodynamic interaction.

Advances in space activity are linked to reductions in launch cost. Air-breathing propulsion-assisted flight systems offer the potential for revolutionary change of the space operations paradigm. Horizontal launch of a space-access system provides mission flexibility, responsiveness, and affordability. One way to reduce launch cost is to increase the Mach number at which a launch vehicle is staged from a carrier aircraft. Without exceeding the engine and airframe design limits, the pre-compressor cooling technology allows an operational aircraft to operate at Mach numbers and altitudes beyond its basic operational limits. This is an essential, near-term technology for reducing launch cost to place small-weight payloads in low Earth orbit. The advantage of this technology is assessed with a modified McDonnell Douglas QF-4C aircraft. Payloads are unachievable or marginal with an unmodified QF-4C. However, payloads weighing around 150 pounds are plausible with this aircraft when incorporating the water injection pre-compressor cooling (WIPCC) technology.

A novel approach for the multi-objective design optimisation of aerofoil profiles is presented. The proposed method aims to exploit the relative strengths of global and local optimisation algorithms, whilst using surrogate models to limit the number of computationally expensive CFD simulations required. The local search stage utilises a re-parameterisation scheme that increases the flexibility of the geometry description by iteratively increasing the number of design variables, enabling superior designs to be generated with minimal user intervention. Capability of the algorithm is demonstrated via the conceptual design of aerofoil sections for use on a lightweight laminar flow business jet. The design case is formulated to account for take-off performance while reducing sensitivity to leading edge contamination. The algorithm successfully manipulates boundary layer transition location to provide a potential set of aerofoils that represent the trade-offs between drag at cruise and climb conditions in the presence of a challenging constraint set. Variations in the underlying flow physics between Pareto-optimal aerofoils are examined to aid understanding of the mechanisms that drive the trade-offs in objective functions.

The exponentially increasing amount of information accumulated from past to current engineering projects has created an environment where repurposing existing data to support new projects is paramount to sustainable success. Strategic planning and early design decisions, specifically, occur in decision-making environments that require information support capabilities that lie outside of traditional engineering analyses. In order to advance towards a more complete planning environment, a pragmatic methodology has been developed for modern aerospace data and information collection, categorisation, and utilisation with a focus on current efforts in hypersonic vehicle research and development. The main thrust has been to provide insights into financial and technical trends that support objective programmatic and planning decision-making. The end-product is a suite of graphical decision-making interfaces, linked through a unified hypersonic database. The graphical interfaces are capable of highlighting the key project drivers along varying levels of categorisation and refinement. Aided by these newly developed data and information support interfaces covering past and present hypersonic efforts, the planner’s forecasting assessment of present and future hypersonic research and development efforts is pragmatically enriched towards a more complete managerial program-planning framework.

This paper presents the development of a mathematical model for the implementation of flexible rotor blade dynamics in real-time helicopter aeromechanics applications. A Lagrangian approach is formulated for the rapid estimation of natural vibration characteristics of nonuniform rotor blades. A matrix/vector formulation is proposed for the treatment of elastic blade kinematics in the time-domain. In order to overcome the classical hurdles of time-accurate simulation and establish applicability in real-time, a novel, second-order accurate, finite-difference scheme is employed for the numerical discretisation of elastic blade motion. The proposed rotor dynamics model is coupled with a finite-state induced flow and an unsteady blade element aerodynamics model. The combined formulation is implemented in a helicopter flight mechanics simulation code. The integrated approach is deployed in order to investigate rotor blade resonant frequencies, trim control angles, oscillatory blade loads and induced vibration for a hingeless and an articulated helicopter rotor. Extensive comparisons are carried out with wind tunnel and flight test measurements, and non-real-time comprehensive analysis methods. Good agreement with measured data is exhibited considering primarily the low-frequency harmonic components of oscillatory loading. It is shown that, the developed methodology can be utilised for real-time simulation on a typical computer with sufficient modelling fidelity for accurate estimation of oscillatory blade loads.

This paper presents the numerical simulations of flowfield over a typical Crew Module at Mach 4 for different angles-of-attack ranging from 0 to –25°. Detailed flow features such as contour of density gradient over the model, numerical oil flow and near wake vortex structures are captured very well in the present simulations. The location of the sonic line and its behaviour due to angles-of-attack is also captured in the simulations. The CP distribution on the windward and leeward side shows excellent match with the experimental results. Also, the prediction of aerodynamic coefficients shows very good agreement with the experimental results. The numerical simulation predicts CMcg, CN and CA within 8%, 4% and 3·5% respectively with respect to experimental values.

A new method for system stability analysis, the weight functions method, is applied to estimate the longitudinal and lateral stability of a Hawker 800XP aircraft. This paper assesses the application of the weight functions method to a real aircraft and a method validation with an eigenvalues stability analysis of the linear small-perturbation equations. The method consists of finding the weight functions that are equal to the number of differential equations required for system modelling. The aircraft’s stability is determined from the sign of the total weight function – the sign should be negative for a stable model. Aerodynamic coefficients and stability derivatives of the mid-size twin-engine corporate aircraft Hawker 800XP are obtained using the in-house FDerivatives code, recently developed at our laboratory of applied research in active controls, avionics and aeroservoelasticity LARCASE. The results are validated with the flight test data supplied by CAE Inc. for all considered flight cases. This aircraft model was chosen because it was part of a research project for FDerivatives code and continued with weight function method for stability analysis in order to develop a design tool, based only on the aircraft geometrical parameters for subsonic regime. The following flight cases are considered: Mach numbers = 0·4 and 0·5, altitudes = 3,000m, 5,000m, 8,000m and 10,000m, and angles-of-attack α = –5° to 20°.

The flow inside a constant-width wind-tunnel contraction is simulated by solving the Reynolds-Averaged Navier-Stokes equations with an eddy-viscosity turbulence model. The results show the presence of longitudinal vortices near the sidewalls centreline. This confirms a former hypothesis involving the generation of skew-induced longitudinal vorticity within the sidewalls boundary layers. Detailed analysis reveals that the flow structure is influenced by viscous effects in the boundary layers and streamline curvature in the potential flow. Three-dimensional boundary-layer profiles on the contraction sidewall are analysed in the framework of the streamline co-ordinate system and its associated hodographic diagram. The resulting profiles help understand the generation of secondary flows and the associated longitudinal vorticity.

In this study, development of a new fuel flow rate model for the climbing phase of flight was achieved using a genetic algorithm (GA) method. Two modelling approaches were performed using real flight data records (FDRs) from a medium-weight transport-category aircraft. The first model considered the dependency of fuel consumption only with respect to altitude, whereas the effects of both altitude and true airspeed (TAS) were included in the second model. The proposed models are improvements on existing models because the relationship between fuel flow rate, flight altitude, and TAS can be deduced using the derived formulations. Both modelling approaches were found to provide accurate results after performing an error analysis for fuel flow rate values. It was clear that incorporating the TAS effect into the second model enhanced the accuracy of the model, but the first model was also found to be appropriate for practical usage.

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 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 manoeuvring 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 aerodynamic shape design of a proprotor for a tiltrotor aircraft is a very complex and demanding task because it has to combine good hovering capabilities with high propeller efficiency. The aim of the present work is to describe a two-level procedure and its results for the aerodynamic shape design of a new rotor blade for a high-performance tiltwing tiltrotor aircraft taking into account the most important flight conditions in which the aircraft can operate. Span-wise distributions of twist, chord and aerofoil were chosen making use of a multi-objective genetic optimiser that worked on three objectives simultaneously. A non-linear sweep angle distribution along the blade was designed to reduce the power losses due to compressibility effects during axial flight at high speed. During the optimisation process, the aerodynamic performance of the blade was evaluated with a classical two-dimensional strip theory solver. The optimised blade was than analysed by means of a compressible Navier-Stokes solver and calculations were validated comparing numerical results with experimental data obtained from wind-tunnel tests of a scaled model of the proprotor.

Sudden release of accreted ice in fuel systems could pose a serious challenge in aircraft operation. The resultant snowshower may reach the filter and fuel-oil heat exchanger, causing a restriction in fuel flow to the engine. It is fundamental to have an appreciation of the character and the interface shear strength of the accreted ice in aircraft fuel systems. This helps to recognise factors for the sudden release of the accreted ice and the intensity of the consequential snowshower. An experimental study was carried out to quantify the character and the interface shear strength of accreted ice on subcooled surfaces submerged in jet fuel. Ice was accreted on naked aluminium, painted aluminium and carbon fibre composite surfaces at various subcooled temperatures. The accreted ice was akin to fresh snow and exhibited soft and fluffy attributes. The character may be expressed quantitatively in terms of the porosity and was found to be c. 0·95. The ice weakly adhered to the substrate surfaces, and the interface shear strength was found to be c. 0·36Pa and c. 2·19Pa at the top surface and at the vertical surface of a specimen block, respectively. It was not possible to detect any variation in the porosity and the interface shear strength for different types of surface finishes and differences in water affnity in fuels due to the crude approach in the estimation of these parameters.

The terminology ‘rotorcraft-pilot coupling’ denotes phenomena arising from interaction between pilot and rotorcraft. Among these, the present work deals with ‘pilot-assisted oscillations’ that derive from unintentional pilot actions on controls due to seat vibrations, and are strictly related to rotor-aeroelasticity/airframe-structural-dynamics coupling, with involvement of blade control actuator dynamics. Focusing the attention on helicopters, a comprehensive rotorcraft model is developed and applied, with main rotor unsteady aerodynamics described in state-space form. This makes it particularly suited for stability and frequency-response analysis, as well as control applications. Numerical investigations address two critical rotorcraft-pilot coupling aeroelastic issues: stability of vertical bouncing and gust response in hovering. Results from main rotor unsteady aerodynamics modelling are compared with widely-used quasi-steady aerodynamics predictions. These suggest that, for accurate RPC/PAO phenomena predictions, mathematical modelling should include the three-dimensional, unsteady-flow effects, and that the pilot-in-the-loop passive behaviour produces a beneficial effect on the load factor generated by gust encountering.

Computational Fluid Dynamics (CFD) has become an attractive method of choice in the design of many aerospace vehicles because of advances in numerical algorithms and convergence acceleration methods. However, the flow around an advanced fighter aircraft is complicated and usually unsteady due to the presence of vortex-dominated flows. The accuracy and predictability of conventional turbulence models for these applications may be questionable and therefore results obtained from these models must be validated and evaluated on the basis of experimental data from wind tunnels and/or flight tests. This work aims to validate CFD simulations of X-31 wind-tunnel models with and without a belly-mounted sting. The sting setup facilitates forced sinusoidal oscillations in one of three modes of: pitch, yaw, and roll. However, the results show that measured aerodynamic data are altered by the turbulent wake behind the sting, even at small angles of attack. The high angle-of-attack flow around the X-31 is also very complicated and unsteady due to canard and wing vortices. Therefore, validation of CFD models for predicting these complex flows can be a very challenging task. The X-31 wind-tunnel experiments were carried out in the German Dutch low-speed wind tunnel at Braunschweig and include aerodynamic force and moment measurement as well as span-wise pressure distributions at locations of 60% and 70% chord length. This data set is used to validate the Cobalt and Kestrel flow solvers and the results are similar and match quiet well with experiments for small to moderate angles of attack. The main discrepancies between CFD and measurements occur close to the wing tip, where leading-edge flaps are located.

When an aero-engine is operating on the ground the formation of a potent inlet ground vortex can arise which has the ability to ingest foreign object debris. The ingestion of foreign objects can cause notable damage to engine components as well as overall performance degradation. The assessment of foreign object ingestion has been conducted using a combination of computational fluid dynamics, analytical modelling and an Euler-Lagrange uncoupled discrete phase particle tracking method. The flow fields for a full-scale aero-engine have been simulated for a range of ground clearances and intake velocity ratios under crosswind conditions. The sensitivity of the debris ingestion thresholds and characteristics to particle size and material density has been evaluated. The characteristics of the ingestion location within the intake are also considered for a range of operating conditions and particle parameters.

This paper aims to present an integrated multidisciplinary simulation framework, deployed for the comprehensive assessment of combined helicopter–powerplant systems at mission level. Analytical evaluations of existing and conceptual regenerative engine designs are carried out in terms of operational performance and environmental impact. The proposed methodology comprises a wide-range of individual modeling theories applicable to helicopter flight dynamics, gas turbine engine performance as well as a novel, physics-based, stirred reactor model for the rapid estimation of various helicopter emissions species. The overall methodology has been deployed to conduct a preliminary trade-off study for a reference simple cycle and conceptual regenerative twin-engine light helicopter, modeled after the Airbus Helicopters Bo105 configuration, simulated under the representative mission scenarios. Extensive comparisons are carried out and presented for the aforementioned helicopters at both engine and mission level, along with general flight performance charts including the payload-range diagram. The acquired results from the design trade-off study suggest that the conceptual regenerative helicopter can offer significant improvement in the payload-range capability, while simultaneously maintaining the required airworthiness requirements. Furthermore, it has been quantified through the implementation of a representative case study that, while the regenerative configuration can enhance the mission range and payload capabilities of the helicopter, it may have a detrimental effect on the mission emissions inventory, specifically for NOx(Nitrogen Oxides). This may impose a trade-off between the fuel economy and environmental performance of the helicopter. The proposed methodology can effectively be regarded as an enabling technology for the comprehensive assessment of conventional and conceptual helicopter-powerplant systems, in terms of operational performance and environmental impact as well as towards the quantification of their associated trade-offs at mission level.

This paper describes the analysis and the minimum weight optimisation of a fuselage composite stiffened panel made from carbon/epoxy material and stiffened by five omega stringers. The panel investigated inside the European project MAAXIMUS is studied using a fast tool, which relies on a semi-analytical procedure for the analysis and on genetic algorithms for the optimisation. The semi-analytical approach is used to compute the buckling load and to study the post-buckling response. Different design variables are considered during the optimisation, such as the stacking sequences of the skin and the stiffener, the geometry and the cross-section of the stiffener. The comparison between finite element and fast tool results reveals the ability of the formulation to predict the buckling load and the post-buckling response of the panel. The reduced CPU time necessary for the analysis and the optimisation makes the procedure an attractive strategy to improve the effectiveness of the preliminary design phases.

A Repetitively Enhanced Neural Networks (RENN) method is developed and presented for complex and implicit engineering design problems. The enhanced neural networks module constructs an accurate surrogate model and avoids over-fitting during neural networks training from supervised learning data. The optimiser is executed by the enhanced neural networks models to seek a tentative optimum point. It is repetitively added into the supervised learning data set to refine the surfaces until the RENN tolerance is reached. The RENN method demonstrates the effectiveness and feasibility for a 2D highly non-linear numerical example and the structure design of a two-member frame reaching a convergent solution at 10 and 15 iterations at the maximum error of 1% when compared with the exact solution. Then, the RENN method is applied for a long endurance unmanned aerial vehicle (UAV) aerofoil design optimisation. A Class/Shape function transformation (CST) geometry parameterisation method represents an accurate UAV aerofoil with ten geometry design variables. The high-fidelity analysis solver with structured mesh is used for a UAV aerofoil design problem. Using the RENN method, an optimal UAV aerofoil is obtained using 88 high fidelity evaluations at an error of 1·24%. The process reduces the computational time by 81·2% compared with the full high fidelity model. The optimal aerofoil shows a drag reduction of 10·8% in the cruise condition and an improvement in the maximum lift coefficient and stall angle-of-attack when compared with the baseline AG24 aerofoil.