A simple but effective method is presented to analyse the effects of asymmetric external stores on flutter of an aircraft in this paper. The method is a kind of ‘direction derivative’ technique, and can be easily applied to predict whether the asymmetric external stores increase or decrease the flutter speed. By using the present technique the varying tendency of flutter speed can be determined preliminarily. This can be very useful and valuable to the designers at the beginning stage of flutter analysis. The effectiveness of the proposed method is proved analytically.

An incompressible method is presented to predict the upwash corrections associated with vortical flow as a result of wind-tunnel side wall effects. An image system is used to simulate the tunnel side walls which are assumed to be solid. An integral expression is formulated, representing the average upwash induced over the wing by the image system. Wall effects may be determined for flows with and without vortex breakdown. Comparisons of the results with upwash predictions from a Navier-Stokes study show close accord. The upwash expression also displayed the ability to successfully predict corrections for flows involving vortex breakdown.

The requirement for an aircraft to operate in the thin atmosphere at high altitudes for long periods provides a particular challenge to all areas of aeronautical engineering. This is reflected in a range of difficulties encountered when attempting to design aircraft for this type of operation.

This paper first considers the reasons for adopting an unmanned solution for high altitude long endurance (Hale) aircraft. It then indicates how the demands of Hale operations lead to problems in applying the design approach taken for more common aircraft types. Some of the work performed at the College of Aeronautics (CoA), in attempting to address the problem areas for the design process and produce initial designs for a range of Hale unmanned aircraft (UMA) types, is then reviewed. The lack of data/methods to allow the prediction of Hale UMA structure mass, engine performance at high altitudes and aerodynamic parameters for low Reynolds number operation and high aspect ratio configurations are identified as particular problems. However, initial design, based on standard approaches, has been shown to be possible, provided caution is exercised.

A theoretical study is presented for the investigation of a potential-flow model for enhancing lift over a flat-plate aerofoil experiencing thin aerofoil stall. Rather than suppressing the leading-edge separation, flow is assumed to separate tangentially at the leading edge and made to reattach smoothly at the tip of a forward-facing fence joining the plate tangentially on its upper surface to avoid any unnecessary stagnated flow. The length of the fence and its location from the leading edge form two geometrical parameters. At any positive angle of attack, the resulting bounding streamline emanating from the leading edge and terminating at the tip of the fence is simulated by using suitable mathematical singularities subject to boundary conditions such as attaining a finite velocity at each critical point of the conformal mapping involved, and the condition of finite pressure gradient at reattachment, when applicable. Computational results from varying these two geometrical parameters indicate that the lift from each model is enhanced, as compared with the attached flow model around a simple flat plate and the original separated flow model by Kirchhoff.

Modelling the response of solid rocket motors to bullet and fragment impacts is a high priority among the military services from standpoints of both safety and mission effectiveness. Considerable effort is being devoted to characterising the bullet and fragment vulnerability of solid rocket motors, and to developing solid rocket motor case technologies for preventing or lessening the violent responses of rocket motors to these impact loadings. Because full-scale tests are costly, fast-running analytical methods are required to characterise the response of solid rocket motors to ballistic impact hazards. In this study, a theoretical first-principles-based model is developed to determine the partitioning of the kinetic energy of an impacting projectile among various solid rocket motor failure modes. Failure modes considered in the analyses include case perforation, case delamination, and fragmentation of the propellant simulant material. Energies involved in material fragmentation are calculated using a fragmentation scheme based on a procedure developed in a previous impact study utilising propellant simulant material. The model is found to be capable of predicting a variety of response characteristics for analogue solid rocket motors under high speed projectile impact that are consistent with observed response characteristics. Suggestions are made for improving the model and extending its applicability to a wider class of impact scenarios.

The Beddoes near wake model, developed for high resolution blade vortex interaction computations, enables efficient numerical evaluation of the downwash due to trailed vorticity in the near wake of a helicopter rotor. The model is, however, limited by the assumption that the near wake lies in the plane of the rotor and, in some cases, by its inability to accurately evaluate the induced velocity contribution from vorticity trailed from inboard blade sections. In this paper, modifications to the method are proposed which address these issues and allow it to be used with confidence over a wider range of rotor flows.

The effects of the model surface to free stream adiabatic temperature ratio (Tw/Tad) on periodic transonic flow over a 14% thick biconvex aerofoil are evaluated using a computational fluid dynamic approach. The analysis is based on the thin layer Navier Stokes equations with Baldwin-Lomax turbulence model. The results of computations showed that on biconves aerofoils there is a large effect of heat transfer on instantaneous pressure distributions and periodic buffet excitation level confirming some of the available experimental data. The effects observed have an implication in wind tunnel measurement of buffet associated periodic transonic flows.

Bird strike is a major consideration when designing fans for civil aero engines. Current methods rely on impact tests and structural optimisation but it is highly desirable to have predictive numerical models to assess the aerodynamic and aeroelastic stability of bird-damaged fan assemblies. A detailed feasibility study towards such a prediction capability, consisting of a CFD solver coupled to a finite element representation of the structure, is reported in this paper. A whole-assembly model was used for both the fluid and the structural domains, an approach necessitated by the loss of cyclic symmetry due to one or more blades undergoing plastic deformation under the effect of the bird impact. It was assumed that two consecutive blades had suffered unequal amount of bird damage, the so-called heavy and medium-damaged blades.

Prior to his departure from the Royal Aircraft Establishment in 1938, Roxbee Cox published an impressive range of specialist papers relevant to the design and analysis of airships and aircraft.

His earliest published paper (Refs are on p 227), on the external forces on an airship structure, was based on his work on the R101 at the Royal Airship Works, Cardington. This comprehensive monograph, which was awarded the R38 Memorial Prize by the Royal Aeronautical Society, addresses, in turn, the range of loading conditions that needs to be considered, safety factors to be used in design, calculation of individual load components, and estimation of stresses within the structure (this consisted of longitudinal and transverse members braced with wires).

The aerodynamic performance of aerofoils performing unsteady motions is important for the design of helicopter rotors. In this respect the study of aerofoils undergoing in-plane oscillations (translation along the horizontal axis) provides useful insight into the flow physics associated with the advancing blade in forward flight. In this paper a numerical method is developed in which the unsteady thin layer Navier-Stokes equations are solved for aerofoils performing rigid body motions. The method has been applied to the calculation of the flowfield around a NACA 0012 aerofoil performing in-plane motions representative of high-speed forward flight. Comparison of computed pressure data with experimental measurements is generally found to be good. The quantitative differences observed between computations and experiment are thought to have arisen mainly as a consequence of the low aspect ratio of the model rotor employed in the windtunnel tests.

An analysis of paraglider performance has revealed that wing section drag is the most significant contribution to overall drag. Wind tunnel measurements performed on two-dimensional hollow models indicate that intake drag is less significant than previously thought. An experimental investigation into the characteristics of a ‘quasi ’ -two-dimensional flexible model consisting of solid ribs covered with a fabric skin was performed at realistic Reynolds numbers. The main cause of performance deterioration was found to be a significant reduction in section lift coefficient when compared to a similar solid wing section. This is believed to be mainly due to two factors: a large trailing edge separation and the deformation of the wing between ribs. The deformation was measured and it was shown that the deformed shape is less capable of generating high lift coefficients than the design section. It is thought that the extent of the trailing edge separation is increased due to the presence of streamwise grooves caused by the shape deformation of the wing. The shape of the separated region was found to be strongly three-dimensional with the separation point being about half a chord-length further upstream along the ribs. A small separation bubble was also observed immediately behind the lip of the intake, due to the fabric ‘flaring’ open. Based on the observations presented a number of suggestions for improved wings have been made.

The classical criteria of the 1920s for the avoidance of flexure-torsion flutter were based on parameters such as stiffness ratio and the position across the typical chord of the aerodynamic, elastic and mass centres.

In more recent times, and in the context of aeroelastic tailoring of lifting surfaces constructed with composite materials, a new terminology has emerged — as evident from technical papers produced in the USA over the past decade. The prevention of flexure-torsion flutter, or the raising of the critical speed, is now achieved by providing more wash-in or less wash-out, where these terms do not have their established aeronautical meanings — and for this reason (and other, non-semantic ones) are to be deprecated.

By recourse to a typical section model (with quasi-steady and unsteady compressible aerodynamics), the writer argues that, for conventionally constructed wings, the new criteria are ‘fuzzy’ and incomplete versions of the earlier criteria in respect of the positioning of the various ‘centres’ (i.e., elastic, mass and aerodynamic) across the typical chord, and therefore that the new terminology is redundant in this context.

For laminated composite wing structures, even when the construction is uniform and the fiexural axis straight, it is established that the inclination of this axis with respect to the CG and aerodynamic axes may be so large that the fiexural and shear centres at the tip could be several chords forward or aft of the mid-chord axis; the criteria of the 1920s are then irrelevant. The wash-in/wash-out criteria may then be said to ‘come into their own’, albeit that it is shown herein that they remain fuzzy and incomplete — even for quasisteady binary problems. A crude modal binary model of a rudimentary laminated composite wing is included to illustrate this and other features. By recourse to a higher-order flutter formulation, the writer demonstrates that the wash-in/wash-out criteria are, in certain respects, unreliable.

An experimental investigation of high frequency combustion instability in a liquid rocket engine of 3kN thrust was conducted. The diagnostic method of detecting combustion instability was based on the measurement of the dynamic pressure following an artificial disturbance in the combustion chamber. Test data spectral analysis was performed by using the stochastic vibration data processing method. Although the engine has demonstrated an absence of tendency to spontaneous instability, insufficient stability of the original engine design was evident during tests with artificial triggers. A Helmholtz type resonator in form of partitioned cavities tuned to a few different, close frequencies was designed and installed in the combustion chamber wall to avoid spontaneous or triggered combustion instability. The experimental results from hot tests with artificial triggers confirmed the high efficiency of the applied acoustic resonator.

A series of experiments has been performed in an open-jet wind tunnel on a generic jet-lift STOVL aircraft in transition out of ground effect. In order to obtain information about the forces acting on the aircraft, static pressure measurements were made on the wing and intake lips. The model tested was equipped with a single vectored lift-jet and ‘powered’ intakes. Tests were conducted at a variety of nozzle pressure ratios and effective velocity ratios.

The experiments conclude that a mutual interference exists between the jet and intake flows which generates non-linearly-additive loads on the airframe. This questions the validity of testing jets and intakes separately on a STOVL aircraft of this type.

An Euler correction method is developed for unsteady, transonic inviscid flows. The strategy of this method is to treat the flow-field behind the shock as rotational flow and elsewhere as irrotational flow. The solution for the irrotational flow is obtained by solving the unsteady full-potential equation using Jameson's rotated time-marching finite-difference scheme. Clebsch's representation of velocity is followed for rotational flow. In this representation the velocities are decomposed into a potential part and a rotational part written in terms of scalar functions. The potential part is computed from the unsteady full potential equation with appropriate modification based on Clebsch's representation of velocity. The rotational part is obtained analytically from the unsteady momentum equation written in terms of Clebsch variables. This method is applied to compute the unsteady flow-field characteristics for an oscillating NACA 64A010 aerofoil. The results of the present calculation are found to be in good agreement with both Euler solution and experimental results.

This investigation discusses some proposed techniques to attenuate rotary wing vibration using smart materials. Advantages and disadvantages of these solutions are analysed in greater detail. Numerical results from feasibility studies are summarised to present the accomplishments already achieved and the goals to be pursued in the near future to develop the next generation of the so-called ‘smart rotors’.

Pressure distributions over a dual cavity cascade were studied at supersonic speeds of Mach 1·5 and 2·5. The study was performed through numerical modelling and results compared with model measurements. The Reynolds-averaged Navier-Stokes equations were solved using a finite-volume algorithm in which the inviscid cell interface fluxes were estimated using Roe's approximate Riemann solver with a second-order extension, and turbulence was modelled using a two-equation k-m model with compressibility corrections. Two test configurations were selected: (1) a length-to-depth ratio L/D = 1 cavity followed by another L/D = 1 cavity, and (2) an L/D = 3 cavity followed by an L/D = 1 cavity. The prediction was compared with that of a single cavity of the same L/D. It was found that the pressure field around the L/D = 1 cavity was substantially modified by a preceding L/D = 3 cavity. Changes in the pressure and pressure drag coefficient were observed. The study clarified some earlier observations of unsteady modes over a dual cavity cascade, and confirmed model measurements of the pressure fluctuation under a number of flow and geometry conditions.

A new approach to deriving analytical approximations for the free modes of motion of an aircraft is presented. Based on an asymptotic analysis, it eliminates the need for a priori assumptions regarding the characteristic motion of each mode, and is thus particularly suitable for an introductory exposition. In its basic form, it yields expressions for natural frequencies, damping factors and time constants which agree, to leading order, with well established results. At a more advanced level, it allows us to assess the validity of these results. In particular, we find that the two degree-of-freedom approximations to the short period oscillation and Dutch roll are asymptotically incorrect. This is unlikely to have a significant effect on the accuracy of the former, but is the reason for the poor damping predictions of the latter.

A finite-state aerodynamics methodology is proposed for the analysis of the forces generated by a gust. To illustrate and assess the methodology, gust-response and gust-alleviation applications are included. Finite-state aerodynamics denotes a technique to approximate aerodynamic loads so as to yield an aircraft model of the type ẋ = Ax + Bu (state-space formulation). In this paper, a finite-state formulation is proposed to include the presence of a gust. The aerodynamic loads to be approximated are evaluated here by using a frequency-domain boundary-element formulation; the flow is assumed to be irrotational except for a zero-thickness vortex layer (wake). The gust-alleviation application consists of determining a control law for reducing the response to a vertical gust disturbance, as measured by the centre of mass acceleration. Two optimal-control approaches are considered for the synthesis of the control law: one uses the classical linear-quadratic regulator (LQR), whereas the second includes the additional feed-forward of the gust velocity ahead of the aircraft. Deflections of ailerons and elevators are assumed to be the control variables. Numerical results deal with responses to both a deterministic ‘1 – cosine’ gust distribution and a stochastic von Kármán spectrum. They indicate that the finite-state aerodynamic model proposed is capable of approximating, with a high level of accuracy, both the aerodynamic loads induced by the aircraft kinematics variables and those induced by the control variables, over a wide frequency range.