The program VICONOPT is used to find the optimum (least mass) dimensions of a range of stiffened wing panels which are subject to buckling and material strength constraints and are loaded in axial compression with a sinusoidal manufacturing imperfection. Design plots are presented to show the effects that various rib spacings and stiffener types have on optimum design mass. A simplified model of a complete wing box is used to illustrate the design of a full wing panel and plots of optimum values of design variables at various stations along the wing have been obtained. The results were chosen to illustrate the practicality of optimisation with reference to manufacture of a full wing panel and to show the effect of changing the sophistication of modelling and theory used for the range of panels considered. The important aspects of the choice of design variables and design concepts are highlighted and percentage savings in mass, compared with an optimum metal panel design, are given for the various (global) optima found along with some examples of (rejected) local optima.

Advanced cooling techniques involving internal enhanced heat transfer and external film cooling and thermal barrier coatings (TBCs) are employed for gas turbine hot components to reduce metal temperatures and to extend their lifetime. A deeper understanding of the interaction mechanism of these thermal protection methods and the conjugate thermal behaviours of the turbine parts provides valuable guideline for the design stage. In this study, a conjugate heat transfer model of a turbine vane endwall with internal impingement and external film cooling is constructed to document the effects of TBCs on the overall cooling effectiveness using numerical simulations. Experiments on the same model with no TBCs are performed to validate the computational methods. Round and crater holes due to the inclusion of TBCs are investigated as well to address how film-cooling configurations affect the aero-thermal performance of the endwall. Results show that the TBCs have a profound effect in reducing the endwall metal temperatures for both cases. The TBC thermal protection for the endwall is shown to be more significant than the effect of increasing coolant mass flow rate. Although the crater holes have better film cooling performance than the traditional round holes, a slight decrement of overall cooling effectiveness is found for the crater configuration due to more endwall metal surfaces directly exposed to external mainstream flows. Energy loss coefficients at the vane passage exit show a relevant negative impact of adding TBCs on the cascade aerodynamic performance, particularly for the round hole case.

Shock-wave/turbulent boundary-layer interactions (SWTBLIs) with separation are known to be inherently unsteady but their physical mechanisms are still not totally understood. An experimental investigation has been performed in a supersonic wind tunnel at a freestream flow Mach number of 2·42. The interaction between a shock wave created by a shock generator (α = 3°, α = 9°, α = 13° and α = 15° deflection angles) and a turbulent boundary layer with thickness δ = 5mm has been studied. High-speed Schlieren visualisations have been obtained and used to measure shock wave unsteadiness by means of digital image processing. In the interactions with separation, the reflected shock’s unsteadiness has been in the order of 102Hz. High-speed wall pressure measurements have also been obtained with fast-response micro-transducers along the interactions. Most of the energy of the incoming turbulent boundary layer is broadband and at high frequencies (>104Hz). An addition of low-frequency (<104Hz) fluctuation energy is found at separation. Along the interaction region, the shock impingement results in an amplification of fluctuation energy due to the increase in pressure. Under the main recirculation region core there is only an increase in high frequency energy (>104Hz). Amplification of lower frequency fluctuation energy (>103Hz) is also observed close to the separation and reattachment regions.

Fuel outgassing (oxygen evolution) within aircraft fuel tanks presents a serious flammability hazard. Time constants representing oxygen transfer rate, from the fuel into a tank’s ullage, are used to model the effect of outgassing on tank flammability. These time constants are specific to a single aircraft type and flight envelope and may not accurately represent fuel outgassing behaviour for other aircraft types with differing fuel tank configurations and flight envelopes. To improve current modelling practice for more accurate flammability analysis dimensional modelling has been used to determine the rate of oxygen evolution from Jet A-1 fuel in an aircraft fuel tank. Measurements of oxygen evolution rate, made on a dimensionally similar model, have been projected to an A320 aircraft. The evolution of oxygen from the fuel was found to increase monotonically with time. Fitting the test data with an inverse-exponential function enabled oxygen release rate and its associated time constant (τ) to be determined. Dimensional modelling of aviation fuel outgassing using model fuel tanks will enable oxygen evolution rate from aviation fuel to be determined for a wide range of aircraft fuel tank configurations and environments without the need for flight testing. In turn the accuracy of flammability assessment of aircraft fuel tanks will be improved and significant cost savings made.

Structural Health Monitoring (SHM) is the component of damage prognosis systems responsible for interrogating a structure to detect, locate, and identify any damage present. Guided wave (GW) testing methods are attractive for this application due to the GW ability to travel over long distances with little attenuation and their sensitivity to different damage types. The Composite Long-range Variable-direction Emitting Radar (CLoVER) transducer is introduced as an alternative concept for efficient damage interrogation in GW SHM systems. This transducer has an overall ring geometry, but is composed of individual wedge-shaped anisotropic piezocomposite sectors that can be individually excited to interrogate the structure in a particular direction. The transducer is shown to produce actuation amplitudes larger than those of a similarly sized ring configuration for the same electric current input. The electrode pattern design used allows each sector to act as an independent actuator and sensor element, decreasing the number of separate transducers needed for inspection. The fabrication and characterisation procedures of these transducers are described, and their performance is shown to be similar to that of conventional piezocomposite transducers. Experimental studies of damage detection demonstrating the proposed interrogation approach are also presented for simulated structural defects.

An ultrasonic guided wave technique based on time delay comb transducers is introduced for aircraft inspection. It is demonstrated that for isotropic plate structures the time delay comb transducers with appropriate excitation are capable of performing discontinuity detections without knowing the precise isotropic material properties of the objects being inspected. Fibre-reinforced composite plates are also considered. The wave skew effects are investigated using both the slowness curve and the Poynting vector. A composite inspection technique that takes advantage of the skew effects is proposed. Using time delay comb transducers to excite the guided wave modes with different skew angles, the proposed technique is capable of scanning the composite plate in different directions without moving or rotating the transducers. By contrast with the applications in the isotropic cases, knowledge of the material properties and other necessary information that is needed to produce dispersion curves is generally required. Experimental results are provided as a validation of the proposed techniques.

This paper studies hovering capability of flapping two-dimensional tandem wing sections inspired by a real dragonfly wing configuration and kinematics. Based on unsteady numerical simulations, the dragonfly corrugated wings have been benchmarked against a flat wing in terms of the aerodynamic forces and flow structures generated during a flapping cycle. The timing of rotation at each stroke is studied by pitch rotation at three different rates, i.e., 80%, 60% and 40% of a flapping period. The results suggest that the longer time pitch rotation with the period of 80% of the overall flapping period is closer to the force calculations obtained of a balanced flight, that is, the mean vertical force supports the dragonfly weight of 0.754 g with a small difference of 0.92% and the mean horizontal force indicates negligible thrust. However, the corrugated wing performs aerodynamically differently from the flat plate with differences in and in by ±2.06% for the corrugated shape. The vorticity flow field for both wings have been recorded at some instants of flapping motions which give more explanation of such dissimilarity.

When designing flight control laws using linearisations of an aircraft model about different flight conditions, some form of scheduling of the resultant gains will often be required to implement the controller over wide operating regions. In practice, the controller gains are often scheduled against relatively slowly-varying system states such as altitude or velocity. However, it may also be desirable to schedule gains against rapidly-varying states such as angle-of-attack, thereby generating a cyclic dependence through hidden coupling terms. Previous published work at Bristol has developed a numerical method of accounting for this dependence when scheduling state feedback gains against coupled states. The resulting ‘dynamic gain schedule’ is shown to significantly improve the transient response of the aircraft model during rapid manoeuvring and to reduce the chances of control surface actuator position limit saturation. In this paper, the novel design process, using eigenstructure assignment, is applied to a mathematical second-order longitudinal aircraft model which represents an approximate BAe Hawk wind-tunnel model. The dynamic gain scheduled controller is shown to work extremely well in practice when applied to the closed-loop experimental rig. Despite the highly nonlinear characteristics of the model aerodynamics and tailplane actuation system, as well as unmodelled high turbulence levels, dynamic gain scheduling demonstrates stable closed loop control even in regions where the nonlinearities are such that conventional gain scheduling fails to produce a stable response.

This paper describes the development of a generic aircraft systems dispatch reliability design methodology (ASDRDM) that has been developed for use during early phases of the aircraft systems design process.

The methodology incorporates prediction of both reliability and maintainability through the aircraft design hierarchy, down to component level. It can be applied at the early design stage, but can also be used for advanced design phases and can use generic or actual failure rate and mean time to repair data. It allows designers to modify system architectures and component reliability and maintainability characteristics. The paper shows the validation that has been performed, and its use is demonstrated by a case-study.

This paper presents a relative fuel burn evaluation of the transonic strut-braced-wing configuration for the regional aircraft class in comparison to an equivalent conventional tube-and-wing aircraft. This is accomplished through multipoint aerodynamic shape optimisation based on the Reynolds-averaged Navier-Stokes equations. Aircraft concepts are first developed through low-order multidisciplinary design optimisation based on the design missions and top-level aircraft requirements of the Embraer E190-E2. High-fidelity aerodynamic shape shape optimisation is then applied to wing–body–tail models of each aircraft, with the objective of minimising the weighted-average cruise drag over a five-point operating envelope that includes the nominal design point, design points at $\pm 10\%$ nominal $C_L$ at Mach 0.78, and two high-speed cruise points at Mach 0.81. Design variables include angle-of-attack, wing (and strut) twist and section shape degrees of freedom, and horizontal tail incidence, while nonlinear constraints include constant lift, zero pitching moment, minimum wing and strut volume, and minimum maximum thickness-to-chord ratios. Results show that the multipoint optimised strut-braced wing maintains similar features to those of the single-point optimum, and compromises on-design performance by only two drag counts to achieve up to 11.6% reductions in drag at the off-design conditions. Introducing low-order estimates for approximating full aircraft performance, results indicate that the multipoint optimised strut-braced-wing regional jet offers a 13.1% improvement in cruise lift-to-drag ratio and a 7.8% reduction in block fuel over a 500nmi nominal mission when compared to the similarly optimised Embraer E190-E2-like conventional tube-and-wing aircraft.

Quantitative and qualitative results of a series of experiments conducted on a rotor in ground effect at low forward speeds are presented. The velocity over a wide area of the ground effect wake was measured using particle image velocimetry, and the evolution of the flow is described as the forward speed increases. The formation of a dust cloud leading to so-called helicopter brown-out was simulated through a series of flow visualisation experiments. The technique involved sprinkling a fine dust on the ground below and ahead of the rotor. Larger dust clouds were observed at lower forward speed, and the dust cloud penetrated into the areas of the flow including those where vorticity levels were of low magnitude and occasional velocity fluctuations from the mean were large.

A new method of analysing the response of linear systems to sinusoidal excitation, which avoids the assumptions of present methods, is developed in this paper. It enables one to obtain the eigenvalues and righthand eigenvectors of a system with viscous damping from the response of the system. If the inertia matrix of the system is known, the damping and stiffness matrices can also be obtained. No assumptions are made about the form of these matrices, and the method is not restricted to systems with only small amounts of damping.

When J. A. Brinell, at the end of the last century, interested himself in the property hardness of metals, he could have had no idea of the importance to be attained later by the new line of investigation he was launching nor of the interesting machines and their applications for which his work was to provide the inspiration.

In the aircraft industry this form of test is not solely used to determine the property hardness as such, but is used to give some idea of the maximum tensile strength of the material. This use is illustrated by some of the specifications controlling the quality of aircraft material, in which hardness tests are required on a specified percentage of bars to ensure that the bars which are riot tensile tested will give a maximum tensile strength of not less than the minimum nor greater than the maximum specification values.

A simple formulation is established for the determination of vibration characteristics of pre-twisted beams by the matrix displacement method. An important point in the analysis is the assumption of complete polynomial displacement distribution in the principal directions. In particular, two twisted elements which employ quintic polynomials are introduced and used in the calculation of natural frequencies of pre-twisted beams for several boundary conditions. The results demonstrate rapid convergence.

The flight test phase of the Nasa-Army UH-60A Airloads Program was completed in February 1994. The objective of the project was to measure the airloads in flight on an UH-60A Black Hawk using 221 blade pressure transducers installed in nine radial arrays. In addition, 21 transducers were installed along the blade leading edge to better identify blade-vortex interaction phenomena. This paper briefly describes the program's background and history, instrumentation, and data acquisition. Airloads data are presented for two test conditions with different levels of complexity: steady, level flight and a diving turn. These flight data provide useful information of relatively simple and complex nonlinear flow for the validation of theoretical analyses. Evidence of shock waves, reverse flow, dynamic stall and blade vortex interactions are presented.

Measurements were made in the Royal Aircraft Establishment High Speed Tunnel of the drag of a wing with various grades of surface roughness in the form of camouflage paint. The measurements were made for a range, of Reynolds number from about 4.5 × 106 to about 15 × 106 at a constant Mach number of 0.2, and for a range of Mach numbers from 0.2 to about 0.7 at a constant Reynolds number of 4.5 × 106 and to about 0.6 at a constant Reynolds number of 6.0 × 106. It was found that for the range of Mach numbers tested compressibility had no appreciable effect on the drag increase due to roughness (Figs. 2 and 3). Further, the drag effect of each roughness tested was such that an equivalent size of sand roughness Ke, of the type tested by Nikuradse, could be readily associated with it. For each finish tested a number of roughness records were taken by means of a roughness gauge developed by Dr. Tomlinson, of the National Physical Laboratory (Fig. 5).

In a Technical Note in the June Journal (p. 354), Dr. A. L. M. Grzedzielski discusses some aspects of the Matrix Force Method of Structural Analysis and in particular has raised objections to the concept of initial strains and its use by the present writers as an artifice in the treatment of cut-outs and modifications in structures. Dr. Grzedzielski asserts that the idea of initial strains is not valid if the flexibility matrix of the unassembled structural elements is other than purely diagonal. Unfortunately this assertion is just not true.

This paper describes the application of the theory of pulse responses to a linearised Euler scheme on a moving grid. The impulse responses of such a linear system are memory functions or temporal representations of the manner in which, and the time over which a perturbation remains active in the response of the system. The exact response of the linear system to an arbitrary input can then be predicted via convolution. The linearised Euler scheme is derived from a full nonlinear Euler method based on Jameson’s cell-centred scheme which has been modified to make it time-accurate and has the necessary terms to account for grid motion. The discrete form of the full Euler equations are linearised about the corresponding nonlinear steady mean flow assuming the unsteadiness in the flow and grid are small. The resulting set of differential equations for the flow perturbations are solved at each time step using a dual-time scheme. These equations can be easily solved to obtain pulse responses, which capture the entire frequency content of the linear system. These pulse responses then provide a simple and efficient method for calculating the system response to many different inputs via convolution. A further use of the pulse responses is their potential to allow a reduced-order linear model to be constructed, which preserves some of the accuracyof the original system. Preliminary work carried out in this area is presented. Results of calculations for heave, pitch and ramp test cases are presented.

A differential equation which contains an eigenvalue is considered as a pair of simultaneous differential equations by augmenting the main equation with the equation λ′=0. This ensures the constancy of the eigenvalue λ. The differential eigenvalue problem is thus reduced to a pair of simultaneous non-linear differential equations with two-point boundary conditions. An iterative method for the solution of the two-point boundary value problem is described.

To demonstrate the method, the normal modes and frequencies in flapping of a helicopter rotor blade are calculated. In the case considered the stiffness and mass/(unit length) of the blade have points of discontinuity. The method may also be applied when the blade parameters are given in the form of experimental data.

It is only within recent years that the helicopter has emerged from the experimental aircraft category to a position in which it can be considered as an effective operational and commercial means of aerial transport. The development of the helicopter to its present stage is due to extensive work in many countries, particularly in America, and it is unfortunate, but unavoidable, that Britain was unable to take a larger share in this work.