This paper develops a 3D homogenisation based continuum damage mechanics (HCDM) model for fibre-reinforced composites undergoing micromechanical damage under cyclic loading. Micromechanical damage in a representative volume element (RVE) of the material occurs by fibre-matrix interfacial debonding, which is incorporated in the model through a hysteretic bilinear cohesive zone model. The proposed model expresses a damage evolution surface in the strain space in the principal damage co-ordinate system or PDCS. PDCS enables the model to account for the effect of non-proportional load history. The material constitutive law involves a fourth order orthotropic tensor with stiffness characterised as a macroscopic internal variable. Cyclic damage parameters are introduced in the monotonic HCDM model to describe the material degradation due to fatigue. Three dimensional damage in composites is accounted for through functional forms of the fourth order damage tensor in terms of components of macroscopic strain and elastic stiffness tensor. The HCDM model parameters are calibrated from homogenisation of micromechanical solutions of the RVE for a few representative cyclic strain histories. The proposed model is validated by comparing results of the HCDM model with pure micromechanical analysis results followed by homogenisation. Finally, the potential of cyclic HCDM model as a design tool is demonstrated through macro-micro analysis of cyclic damage progression in composite structures.

This paper provides a method that helps the aircraft designer to develop a performance constraint chart (PCC) for vertical/short takeoff and landing (V/STOL) aircraft that produce hybrid lift (static lift in combination with aerodynamic lift). The PPC provides a first estimation for the thrust-to-weight ratio (T0/MTOW) and wing loading (MTOW/S). The method is applicable to concepts, where static lift and main-engine thrust are coupled (e.g. the F-35B system, turbojet and lift-fan, coupled by a shaft) and to concepts, where static lift is produced by separate devices (e.g. lift-engines or other concepts for static lift). It includes thrust vectoring of the main engine.

The method includes the evaluation of certain flight stages, or segments, as short take-off and landing (STOL), vertical landing (VL), one engine inoperative (OEI) climb (for civil aircraft concept applications) and cruise, all these have to be considered. For each of these five segments, standard flight mechanics equations are extended by a static lift component, an augmentation ratio (a factor that describes the dependency of the thrust and the static lift, if coupled) and the thrust vectoring angle. Hence these equations are modified in a way that the aircraft designer can directly calculate T0/MTOW and MTOW/S for the performance requirements of each segment. Thus, an optimum design point can be selected. Inputs are aerodynamic coefficients, maximum lift coefficient of the wing, mass fraction from take-off to landing, additional static lift during take-off and landing, number of engines, augmentation ratio of the propulsion system and required take-off and landing field length.

A performance constraint chart for a JSF F-35B type aircraft is modelled to show the application of the method. As an application to civil aircraft a PCC and a parameter optimisation for the civil regional jet ‘HyLiner’ is presented.

The present paper reports a way of using an artificial neural network (ANN) for modelling methane-air jet diffusion turbulent flame characteristics, such as temperature and chemical species mass fractions in a gas turbine combustion chamber. Since the neural network needs sets of examples to adapt its synaptic weights in the training phase, we used pre-assumed probability density function (PDF) method and considered chemical equilibrium chemistry model to compute the flame characteristics for generating the examples of input-output data sets. In this approach, flow and mixing field results are presented with a non-linear first order k-ε model. The turbulence model is applied in combination with preassumed β-PDF modelling for turbulence-chemistry interaction. The training algorithm for the neural network is based on a back-propagation supervised learning procedure, and the feed-forward multilayer network is incorporated as neural network architecture. The ability of ANN model to represent a highly non-linear system, such as a turbulent non-premixed flame is illustrated, and it can be summarized that the results of modelling of the combustion characteristics using ANN model are satisfactory, and the CPU-time and memory savings encouraging.

Ultrasonic nondestructive evaluation methods are routinely used to detect and size fatigue cracks near fastener holes in aircraft structures as a part of scheduled maintenance. In contrast, statistical crack propagation models provide an estimate of the expected fatigue life assuming a known crack size and future fatigue loadings. Here an integrated approach for in situ diagnosis and prognosis of fastener hole fatigue cracks is proposed and implemented that incorporates both ultrasonic monitoring and crack growth laws. The sensing method is an ultrasonic angle beam technique, and cracks are automatically detected from the ultrasonic response. An extended Kalman filter is applied to combine ultrasonically estimated crack sizes with a crack growth law, effectively using the time history of the ultrasonic results rather than only the most recent measurement. A natural extension of this method is fatigue life prognosis. Results from fatigue tests on 7075-T651 aluminium coupons show improved crack size estimates as compared to those obtained from ultrasonic measurements alone, and also demonstrate the capability of predicting the remaining life. This approach for fatigue crack detection, sizing and prognosis is an example of a general strategy for in situ monitoring of structural damage whereby improved results are achieved from the integration of noisy measurements with imperfect crack growth models.

The aeroelastic modeling and flutter characteristics of a shear deformable wing/stores configuration under pull-up angular velocity is investigated. An isotropic non-uniform wing, which structural model incorporates flexibility in transverse shear and warping effects, is considered. The aeroelastic governing equations and boundary conditions are determined via Hamilton’s variational principle. In order to exactly consider the span wise location and properties of the attached stores the generalised function theory is used. The partial differential equations are transformed into a set of eigenvalue equations through the extended Galerkin’s approach. Numerical simulation highlighting the effects of the pull-up angular velocity and store parameters and configurations, such as mass ratio and their attachment locations, on the flutter speed are presented. The results of flutter analyses are validated with the published results and good agreement is observed. Furthermore, the procedure for an optimal deployment of stores is obtained for the case of the wing with four stores.

Subsonic wind tunnel experiments were conducted to study the effect of forced transition on the pressure distribution in the concave portion of contraction. Further more, the effect of early transition on the turbulence level in the test section of the wind tunnel is studied. Measurements were performed by installing several trip strips at two different positions in the concave portion of the contraction. The results show that installation of the trip strips, have significant effects on both turbulence intensity and on the pressure distribution. The reduction in the free stream turbulence as well as the wall static pressure distribution varied significantly with the location of the trip strip. The results confirm the significant impact of the tripped boundary layer on the control of adverse pressure gradient. The trip strip at X/L = 0.115 improves pressure distribution in contraction and reduces turbulence intensity in the test section, considerably.

This paper aims to demonstrate the viability of using the large eddy simulation (LES) CFD methodology to model a representative, complete STOVL aircraft geometry at touch down. The flowfield beneath such a jet-borne vertical landing aircraft is inherently unsteady. Hence, it is argued in the present work that the LES technique is the most suitable tool to predict both the mean flow and unsteady fluctuations, and, with further development and validation testing, this approach could be a replacement, and certainly a complementary aid, to expensive rig programmes. The numerical method uses a compressible solver on a mixed element unstructured mesh. Examination of instantaneous flowfield predictions from these LES calculations indicate close similarity with many flow features identified from ground effect flow visualisations, which are well known to be difficult to model using RANS-based CFD. Whilst significant further work needs to be carried out, these calculations show that LES could be a practical tool to model, for example, Hot Gas Ingestion for the Joint Strike Fighter aircraft.

This paper addresses the development of multiple UAV deployment simulation models that include representative aerodynamic cross-coupling effects. Applications may include simulations of autonomous aerial refuelling and formation flying scenarios. A novel wake vortex model has been developed and successfully integrated within a Matlab/Simulink simulation environment. The wake vortex model is both sufficiently representative to support studies of aerodynamic interaction between multiple air vehicles, and straightforward enough to be used within real time or near real time air-to-air simulations. The model integration process is described, and computational results of a two-vehicle-formation flight are presented.

This paper describes the market for air vehicles that offer Very Short Take Off and Landing (VSTOL) capabilities, also known as VSTOL aircraft. A commercial VSTOL aircraft will only succeed if it provides profits to both manufacturers and airlines. Being able to operate out of shorter fields offers airlines greater flexibility and therefore provides a benefit to the airline. However, adding VSTOL performance also adds cost. Successful VSTOL aircraft designs must balance the added cost against the additional benefits VSTOL performance will provide.

The first problem in any new VSTOL design is determining what field length provides the maximum benefit relative to cost. In order to be able to determine this one needs to know not only what it will cost to design and build the new aircraft, but its value to the airlines, as reflected by the demand for it and the price they are willing to pay. Only then is it possible to design the aeroplane the market wants at a sustainable price and in sufficient quantities to optimise the manufacturer’s return on investment.

This paper is a comprehensive parametric analysis of helicopters and aeroplanes that shows how much the market values short field performance by determining how much airline customers have been paying for it and how the market reacts to changes in its price. The analysis simultaneously considers over 230 currently produced models of fixed wing and rotary commercial aircraft, ranging in size from a Cessna 172 to an Airbus A380 and with balanced field lengths ranging from less than 150 to over 10,500ft.

The data show that a single, statistically significant demand curve exists for all of these vehicles, with prices that range across several orders of magnitude. This curve consists of several segments or sub-markets including those for regional and rotary wing aircraft. Both of those markets offer vehicles over a wide range of prices. The analysis shows that VSTOL price is a function of the balanced field length provided by the vehicles within these market segments. Comparing demand and sustained price against cost reveals particular field lengths for which it will be possible to make a profit building VSTOL aircraft.

Because the sustained price for a VSTOL aircraft is a function of its field length performance, and the quantity of a model sold is a function of its price, it follows that it is possible to predict the profit potential for a given VSTOL design in advance of its development. A thorough study of potential configurations therefore offers manufacturers the ability to design VSTOL aircraft that provide the maximum possible profit while fully satisfying the requirements of their airline customers.

The long-term performance of aeronautical structures is typically discussed in terms of concepts such as structural integrity, durability, damage tolerance, fracture toughness, etc. These familiar concepts are usually addressed by considering balance equations, crack growth relationships, and constitutive equations with constant material properties, and constant or cyclically applied load conditions. Loading histories are represented by changing stress (or strain) states, only. But for many situations, especially associated with high performance aircraft, the local state of the material may also change during service, so that the properties used in those equations are functions of time and history of applied conditions. For example, local values of stiffness, strength, and conductivity are altered by material degradation to create ‘property fields’ that replace the global constants, and introduce time and history into the governing equations. The present paper will examine a small set of such problems and offer a construct for using related solutions to estimate future performance based on history of use and current material state, a concept typically called prognosis.

The aim of the present work is the investigation about the use of blade trailing-edge flaps for the reduction of vibratory loads arising at the hub of helicopter main rotors in forward flight. The alleviation of these loads is achieved through multicyclic higher harmonic actuation of the blade flaps, which is related to measured vibratory loads amplitude. The feedback control law is obtained by an optimal control process based on the minimisation of a cost function, under the constraint of compatibility with the nonlinear equations governing blade aeroelasticity. In the numerical investigation concerning a four-bladed rotor in level flight conditions, a computationally efficient local controller methodology is applied, with the attention focused on the effectiveness of the control algorithm, along with its robustness with respect to differences (existing in real applications) between the aeroelastic models used for control law synthesis and validation.

A methodology is described that determines the statistical confidence bounds on the results from flight flutter tests: modal parameter estimates, flutter margin values and flutter speed estimates, without the need for Monte-Carlo simulation. The approach is based on least squares statistics and eigenvalue perturbation theory applied to the various stages of the analysis process, starting with frequency and damping estimation through to the flutter margin calculations. The technique is demonstrated upon a number of data sets from aeroelastic simulations of flight flutter tests.

This work looks to harness atmospheric energy through thermal soaring to optimise the flight persistence of Micro Air Vehicles (MAVs) and Unmanned Air Vehicles (UAVs). There are two key challenges when considering thermal soaring, the first being the locating of thermals and the second being the extraction of the maximum potential energy from the thermals. Thermal location is by no means an exact science with experienced glider pilots needing to consider many factors to improve the probability of encountering a thermal. As thermals are caused by the uneven heating of the Earth’s surface it is, however, possible to predict likely thermal locations. With the application of a suitable guidance algorithm which considers these ‘hot spots’ it is possible to increase the likelihood of encountering a thermal. Once a thermal is found it is important to attempt to extract the maximum energy from the thermal. To do this the vehicle needs to move quickly to the centre of the thermal. There are many potential techniques for thermal centring, some of which appear to be entirely contradictory to each other, the crucial factor determining the success of such a technique has been found to be the response time of the onboard sensors. This paper considers the many aspects of thermal soaring such as locating thermals, thermal detection and thermal centring. Five different thermal models are presented which are used to demonstrate the thermal centring techniques. Finally a commercial glider simulation package is used to demonstrate the control architecture and simulate a fully autonomous flight.

The aim of this paper is to determine the mathematical relationship (model) between control deflections and structural deflections of the F/A-18 modified aircraft in the active aeroelastic wing technology program. Five sets of signals from flight flutter tests corresponding to the excited sources were measured by NASA Dryden Flight Research Center. These excitation inputs are: differential ailerons, collective ailerons, collective stabilisers, differential stabilisers, and rudders. The signals to be used by the model are of two types: control deflection time histories and corresponding structural deflections on the wing and trailing-edge flaps. We choose to use the subspace identification method based on reconstructing the observability matrix in order to identify the nonlinear multi-input, linear-in-the-states, multi-output system. We identify models (input/output characteristics) by applying this method for a number of sixteen flight conditions for which the Mach number varies from 0·85 to 1·30 and the altitudes vary from 5,000ft to 25,000ft. Very good results are obtained with a fit between the estimated and the measured signals and a correlation coefficient higher than 90%.

A series of mixed-compression hypersonic inlets at different bleeding rates were simulated at different freestream conditions in this paper. The unstart/restart characteristics of hypersonic inlets were analysed and the reasons why the unstart/restart phenomenon is in existence is presented. The unstart/restart characteristics of hypersonic inlets at different bleeding rates were given. The effects of boundary-layer bleeding on the performance parameter (mass-captured coefficient, total-pressure recovery coefficient), starting and restarting Mach number of hypersonic inlets were discussed. In conclusion, boundary-layer bleeding can improve the performance parameter of hypersonic inlets, and can reduce the starting and restarting Mach number, and can broad the operation range of the hypersonic inlet.

Current sonic boom theory is based on linear midfield solutions coupled with acoustic propagation models. Approximate corrections are made within the theory to account for non-linearities, in particular for the coalescence of compression waves and the formation of weak shocks. A very large adjustment is made to account for the increasing acoustic impedance that the waves encounter as they propagate from the low density air at cruise altitude to the high density air at sea level. Typically this correction reduces the calculated over pressure levels by a factor of three. Here the method of characteristics (MOC) is used to prove that the density gradient within a hydrostatic atmosphere has no direct effect on the propagation or intensity of the wave. However gravity and ambient temperature both affect the wave propagation and the combined pressure level attenuation is not dissimilar to that previously attributed to acoustic impedance. Although the flawed acoustic theory has given reasonable predictions of measured sonic booms, the omission of gravity from the equation of motion and the inclusion of a false impedance modification, makes the model unreliable for prediction of future designs, particularly those focused on boom minimisation. As an aid to quiet supersonic aircraft design, Whitham’s theory is extended to include gravity and ambient temperature variation and shown to be in good agreement with a MOC solution for the real atmosphere.

The purpose of the research presented here is to exploit actuation via smart materials to perform shape control of an aerofoil on a small aircraft and to determine the feasibility and advantages of smooth control surface deformations. A type of piezoceramic composite actuator known as Macro-Fiber Composite (MFC) is used for changing the camber of the wings. The MFC actuators were implemented on a 30° swept wing, 0·76m wingspan aircraft. The experimental vehicle was flown using two MFC patches in an elevator/aileron (elevon) configuration. Preliminary flight and wind-tunnel testing has demonstrated the stability and control of the concept. Flight tests were performed to quantify roll control using the MFC actuators. Lift and drag coefficients along with pitch and roll moment coefficients were measured in a low-speed, open-section wind tunnel. A vortex-lattice analysis complemented the database of aerodynamic derivatives used to analyse control response. The research, for the first time, successfully demonstrated that piezoceramic devices requiring high voltages can be effectively employed in small air vehicles without compromising the weight of the overall system.

The aerodynamics of insect-like flapping are dominated by the production of a large, stable, and lift-enhancing leading-edge vortex (LEV) above the wing. In this paper the phenomenology behind the LEV is explored, the reasons for its stability are investigated, and the effects on the LEV of changing Reynolds number or angle-of-attack are studied. A predominantly-computational method has been used, validated against both existing and new experimental data. It is concluded that the LEV is stable over the entire range of Reynolds numbers investigated here and that changes in angle-of-attack do not affect the LEV’s stability. The primary motivation of the current work is to ascertain whether insect-like flapping can be successfully ‘scaled up’ to produce a flapping-wing micro air vehicle (FMAV) and the results presented here suggest that this should be the case.

Numerical models used to simulate the lofting mechanisms of runway stones were developed to assess the threat to aircraft structures from runway debris impacts. An inflated aircraft tyre model, which was validated by comparison with experimental indentation tests, showed that over-rolling of stones under typical take-off conditions led to only modest vertical loft velocities of less than 5 m/s. Experiments using a drop weight impactor simulated a section of aircraft tyre descending upon stones. These tests demonstrated that lofting was achieved for impacts with low rubber thickness. However, for impacts with greater rubber thickness, lofting was suppressed. Using more realistic tyre geometries resulted in launches with backspin, but only horizontally along the ground in the direction of the tyre axis. The speed at which launches occurred was proportional to the rate of descent of the tyre section and would consequently determine the loft speeds due to potential asperity lofting.

The Silent Aircraft Initiative was a Cambridge-MIT Institute programme involving a large team of researchers from both the University of Cambridge and the Massachusetts Institute of Technology (MIT). The aim of the project was to produce a concept aircraft design that would be so quiet it would be imperceptible in the urban environments around airports.. This paper gives an overview of how all the sources of engine noise were carefully addressed within the Silent Aircraft design. Novel technologies, a new engine configuration, improved airframe integration, new operational procedures and advanced component design were all required in order to reduce the overall engine noise level to the Silent Aircraft target. The study suggests that in order to dramatically reduce the noise of future aircraft engines a number of major design changes must be combined.