This paper describes a comparative evaluation of two data smoothing algorithms for use in a two step estimation-beforemodelling procedure for aircraft parameter identification. A simple fixed lag smoother is compared with the usual, and more complex, modified-Bryson-Frazier smoother in the first, state estimation, step of the aircraft parameter identification procedure. The comparison is illustrated by application to the analysis of the Dutch Roll motion of the Embraer EMB-312 Tucano. Both algorithms were found to give results of comparable accuracy although the fixed lag smoother is computationally more efficient. It was concluded that the fixed lag smoother algorithm is an acceptable alternative to the modified-Bryson-Frazier algorithm in aircraft parameter identification applications.

This paper deals with the problem of determining an optimal controller which minimises the norm of a given aircraft while guaranteeing that an upper limit of its norm is not exceeded. The problem is tackled by means of a linear matrix inequality formulation, which allows one to constrain the eigenvalues of the model to within a fixed region of the complex plane. Key features of the method are the possibility of simultaneously

• 1. keeping the maximum value of the input demand under control

• 2. taking flying quality requirements into account

• 3. assuring a minimum level of system robustness against uncertainties of the model.

The approach is simple to handle and is well suited to the design of aircraft stability augmentation systems. A discussion of two case studies demonstrates the effectiveness of the procedure.

A computational method has been developed to investigate the aerodynamic interaction between a helicopter rotor and the fuselage in steady-state level forward flight. In order to model a compound helicopter the fuselage is fitted with wings. The model uses deformable vortex rings for the rotor wake, a source panel representation of the fuselage and a lifting surface method for the wings which can handle arbitrary geometries. The wake representation is based on a free wake analysis, relaxing the nodal points on the vortex rings and wing wake elements according to the induced velocities in the flow field after every time step. In this way, neither wake is restricted to planar shape. At each time step the fuselage and wing forces were calculated. The trim attitude of the fuselage was determined by the pressure field surrounding it using a potential flow model for the external flow and a fully developed turbulent boundary layer model on the fuselage surface. Parametric studies were carried out to determine the influence of wing position, wing span, wing angle of incidence and advance ratio on the results. The computer program incorporates a graphical postprocessing output routine allowing the visualisation of the wake shape, the induced velocity and the angles of incidence on the rotor disc to be made, together with the velocities at various wing stations and the lift coefficients along the span.

This paper considers the convection of the dynamic stall vortex. A brief discussion of existing data is given. A comparison of two sets of data from different experimental facilities is then presented, and it is indicated that an important anomaly exists concerning the behaviour of the dynamic stall vortex. It emerges that freestream Mach number is the only parameter that can account for the observed differences. The importance of this parameter is then discussed in the context of vorticity flux from the aerofoil surface.

An experimental study of the effects on the low-speed aerodynamic characteristics of a strake-like fillet is described, modelled on one used on an Airbus A320 variant fitted at the leading edge of a swept wing-plate junction. The wing, swept back at 20°, was of NACA 0015 section and chord 500 mm, both normal to its leading edge. A turbulent boundary layer had developed on the plate well ahead of the junction. The tests were conducted at a unit Reynolds number of 1.56 x 106 m-1.

Surface pressure distributions were measured on the plate in the neighbourhood of the leading edge junction and also on the aerofoil and fillet at wing incidences of 0°, 3°, 6°, 9° and 12°. These were supplemented by surface oil-flow studies.

The mean velocity and turbulence intensity fields around the leading edge were examined for incidences of 0° and 9°, using both a single tube yaw meter developed for the purpose and an X-wire anemometer. The X-wire anemometer was also used downstream of the trailing edge of the swept wing; five of the Reynolds stresses and the mean velocity field were measured.

The sectional lift coefficients on the aerofoil were found to diminish as the junction was approached, slightly more so with the fillet than without it. The sectional drag coefficients due to pressure increased as the junction was approached, the fillet moderating this increase to only a small extent.

However, the addition of the drooped fillet modified the flow considerably. The horseshoe-like vortex was less well defined than without it. At zero incidence, the peak in the turbulence intensity levels was virtually eliminated on what became effectively the compression side of the wing due to the local camber introduced by the asymmetric fillet. The turbulence levels were also reduced by the addition of the fillet at an incidence of 9°. However, the turbulent activity was spread through a larger proportion of the viscous region. The secondary flows and the turbulence activity in the wake are associated with unrecoverable kinetic energy and will be manifest as drag on the surfaces forming the junction.

It is concluded that a carefully designed fillet, optimised for the cruise incidence of an aircraft, can reduce the peak turbulence levels in the junction. It remains unclear whether the total drag associated with the junction flow can be reduced significantly.

However because of its effects on the turbulence, there may be other benefits, for example on the efficiency of downstream elements, such as fuselage-mounted engine intakes or the following stages of an axial flow machine. Junction fillets might also be used to control the scouring of river beds around bridge piers.

The classic problem of stability and transition to turbulence of an incompressible flow over a flat plate subjected to uniform suction through the surface is considered. This is an important “model” problem for laminar flow control and the results have direct application to the design of low drag aerofoils. Detailed mean boundary layer calculations are presented and the flow stability is determined by obtaining numerical solutions of the Orr-Somerfield equation. Transition conditions are estimated by using the semi-empirical en method. Accurate values have been obtained for the levels of suction necessary to delay instability and transition indefinitely. Consideration has also been given to the effect of surface suction upon the plate drag.

The class of aircraft known as gyroplanes (or autogyros) helped to pave the way for the development of the helicopter. However, they have found no application in contemporary commercial or military aviation. It is in recreational or sport flying that the gyroplane has proved popular. Most if not all designs are however homebuilts, and as a consequence little analysis of any significance has been conducted on the flight mechanics of these aircraft. This Paper presents an analysis of the longitudinal stability of gyroplanes. Simple consideration of the basic forces and moments that might influence stability, in the context of elementary rotor behaviour, is used to assess the fundamental nature of gyroplane longitudinal stability. This is quantified by the application of a sophisticated generic rotorcraft mathematical model to the gyroplane problem. It is concluded that the basic configuration can be longitudinally stable with appropriate design, and that the rotorspeed degree of freedom must also be included when modelling the aircraft.

Inverse simulation is becoming a more widely used technique in flight mechanics studies. The ability to predict control and response time histories for a predefined manoeuvre or task has obvious benefits in certain applications (handling qualities studies related to the ADS-33 requirements for helicopters, for example). The main criticisms of the technique have always been concerned with the numerical problems usually encountered when solving the equations of motion in the inverse manner. In this paper these problems are highlighted and a simple technique which overcomes them is presented. This technique provides a robust mechanism whereby reliable and consistent inverse simulations are possible.

The effects of flap deflection, configuration incidence and leading edge bluntness on boundary layer separation and transition have been studied experimentally. A quasi two-dimensional flat plate equipped with a full span trailing edge control flap has been employed for these tests.

The studies have been carried out in a hypersonic gun tunnel at M∞ = 8·2 and Re∞/cm = 9·0 × 104. The flap deflection angles studied were in the range 0° ≤ β ≤ 30°. The incidence range was from zero to α = 10° (positive α is nose down). Leading edge bluntness effects were simulated by the introduction of a hemi-cylindrical leading edge.

The flow structure was studied using high speed Schlieren photography as well as surface pressure and heat transfer measurements. Liquid crystals were employed to study the threedimensionality of the flow structure for selected configurations. Analytical theories have been developed to estimate the flap pressure and heat transfer levels for the sharp and blunt configurations. These are compared against experimental measurements.

The paper presents a method for the prediction of mean flow data (i.e. skin friction, velocity profile and shape parameter) for adiabatic two-dimensional compressible turbulent boundary layers at zero pressure gradient. The transformed law of the wall, law of the wake, the van Driest model for the complete inner region and a correlation between the Reynolds number based on the boundary layer integral length scale (ReΔ*.) and the Reynolds number based on the boundary layer momentum thickness (Reθw) were used to predict the mean flow quantities. The results for skin friction coefficient show good agreement with a number of existing theories including the van Driest II and the Huang et al. Comparison with a large number of experimental data suggests that at least for transonic and supersonic flows, the velocity profile as described by van Driest and Coles is Reynolds number dependent and should not be presumed universal. Extra information or perhaps a better physical approach to the formulation of the mean structure of compressible turbulent boundary layers even in zero pressure gradient and adiabatic condition, is required in order to achieve complete (physical and mathematical) convergence when it is applied in any prediction methods.

Flow transition in the developing region of a plane wall jet is studied experimentally by means of hot-wire measurements. The Reynolds number, based on the jet exit velocity and nozzle exit height, varies from 3 x 102 to 3 x 104 in the investigations. The results indicate that the most important factor causing wall jet flow to change from the initial exit state to the final turbulent state is the rapid increase of turbulent intensity from the formation of vortical structures and their interactions with the wall. Data also shows that the transition process of the wall jet depends on the Reynolds number in the operating range. When the Reynolds number is greater than 2000, the mean velocity distribution of the wall jet changes directly from a top-hat profile at the nozzle exit to a turbulent profile in the turbulent developed region. When the Reynolds number is smaller than 2000, Glauert's laminar velocity profile is found between the nozzle exit and the turbulent region. It is also found that the transition will be delayed and the transition region prolonged with the decrease of the Reynolds number.

This paper develops the theory of turning flight in a moving air mass to explain some of the problems encountered in making downwind turns, particularly in light aircraft. Although the basic mechanics of flight and aerodynamics of the aircraft are well understood, the methods used by the pilot to control the aircraft in turning flight are less well established and can be critical in maintaining a safe flight path. In this paper the effect of the control strategy likely to be employed by a pilot flying close to the ground is used to show how the energy loss during a turn can lead to an unsafe flight path.

It is widely accepted that the effectiveness of cannon in the air-to-air combat scenarío depends on the rate of fire. Present firing rates for 20-30 mm aircraft cannon range from 600 to 1500 rounds per minute. New developments are aimed towards even higher rates of fire. With these higher rates of fire more energy is likely to be transferred to the aircraft fuselage, resulting in possible structural fatigue damage. To absorb the recoil forces several types of recoil system have been developed and are presently in use. Conventional recoil systems are not, however, ideal for very high rates of fire and hence alternatives must be investigated.

With this work the use of an anti-resonant device to reduce the transmission of forces to the aircraft fuselage for a high rate of fire is investigated. Antiresonant isolators operate on the principle that a condition of little or no motion may be enforced at specific points of interest on a system at particular frequencies, through suitable tuning of the system. This is essentially achieved by designing the isolator system in such a way that inertial forces are used to react spring forces under stationary conditions.

Experiments are performed on the effects of a turbulent boundary layer on surface pressure and heat transfer, using a cylindrical model fitted with a variety of flares at Mach 5. The flow over compression corners of up to 20° (below the expected incipient separation value) is examined. Liquid crystal thermography indicates an unexpected peak in heat flux close to the corner. It is suggested that this is due to a small scale separation occurring earlier, i.e. at smaller corner angles, than previously observed. An estimate of the size of such a separation bubble is made and it is suggested that this is a predominantly laminar phenomenon.

Progress in aeroelasticity of fixed and rotary wing aircraft is compared by means of a selective review. Discussion of the role of concurrent engineering leads into modelling and validation, these being dealt with under the headings of aerodynamics, structural and integrated models. Formulation and solution of equations of motion are considered, which relate mainly to rotorcraft, and then aeroelastic performance improvement is covered. This includes optimisation, aeroservoelasticity and passive devices. Basic insight providing models are reviewed and the paper concludes with, further comments on insight, and also future progress in aeroelasticity.

This paper examines the changing role aerodynamic technology plays in the design and development of modern combat aircraft. It reviews several aspects of aerodynamics which contribute to aircraft performance and handling characteristics. It considers the importance of a range of technologies, the contribution each makes to the final integrated solution and comments on the compromises necessary through the design cycle to optimise the overall weapon system, sometimes to the cost of one of the component technologies. The technologies discussed are in no way exhaustive but attempt to encapsulate the breadth of the subject, to illustrate its diversity and to point the way for the future development if aerodynamic technology is to continue to make an important contribution to the design of combat aircraft.

Each of the component technologies are discussed in terms of the contribution it makes, the tools and techniques used to predict, analyse and interpret the technology contribution, and the requirements for the direction of the development of the technology for the future.

Before considering the technologies themselves it is important to understand the environment in which they will be used, which increasingly conditions the manner in which they are employed and suggests the direction for their development.

This paper begins, therefore, by briefly reviewing the changes to the environment in which the fighter pilot operates and the manner by which his requirements, and therefore the specification of the aircraft, are defined before going on to look at the overall contribution each technology makes in the design process.

A novel automatic grid generation method for complex CFD applications has been developed, driven by the need of the industrial CFD practitioner to reduce the turnaround time from an initial CAD geometry to obtaining a flow simulation. The method uses a subdivision strategy, allowing hanging nodes in a Cartesian unstructured mesh environment and is combined with a non-overlapping, semi-structured surface grid. Unusually, the surface grid is the end rather than the starting point of the present approach. This reversal of gridding strategy allows usability to be enhanced, in that the grid generation is reduced to one automatic step. The capabilities of the method are demonstrated by reference to selected inviscid and viscous flow solutions using the Fluent code Rampant v4.0 as a flow solver.

A rapid quasi-analytical method — low order perturbation panel method — is presented for calculating aerodynamic sensitivity derivatives for subsonic wings. The method is based on the low order panel method with the internal Dirichlet problem formulation and analytical differentiations cascaded and inverted. In terms of doublet strength sensitivity to configuration geometry, the computing cost of the present method for the partial derivative matrix calculation is less than one order of magnitude than the existing high order perturbation panel method, but the accuracy is comparable. Furthermore, by applying a physical interpolation instead of a weighted geometrical interpolation, the sensitivity derivatives of pressure, lift and pitching moment coefficient with respect to configuration geometry are derived and calculated in the present paper. Such sensitivity derivatives are lacking in high order perturbation panel method. A few calculative examples are given.

The start condition of a second-throat ejector-diffuser (STED) system has been studied by solving the axisymmetric Navier-Stokes equations with a high order conservative supra-characteristics method (CSCM). According to the conventional concept, when a STED is started, the flow passing through the second-throat contraction should be supersonic so that the normal shock swallowing condition applying for a supersonic windtunnel holds. The present numerical results, however, suggest that a STED can be started with a normal shock forming ahead of the second-throat contraction because the high speed flow in the region close to the wall of the diffuser serves to release the mass blocked by the normal shock, leading to a decrease in chamber pressure as well as the start of the system. This not only contradicts the conventional concept of the start of a STED, but also provides an explanation for the experimental discrepancy between the start conditions of a STED and a supersonic windtunnel.