—The experience from several environmental trials on helicopters is reviewed to show how the life of main and tail rotor blades was reduced by erosion and how effective various modifications have been in protecting these blades. The compressors of the gas turbine power units were also affected by sand ingestion and intake filters were used on later trials to protect the engines. The engine erosion problem is discussed with reference to the position of the intake, the degree and type of filtration and the design of the compressor itself. The design features of the helicopter which affect its ability to operate in dust and salt-laden atmospheres are also considered, along with suggestions for further research and development.

An expansion tube with a free piston driver has been used to generate quasi-steady hypersonic flows in argon and air at flow velocities in excess of 9 km/s. Irregular test flow unsteadiness has limited the performance of previous expansion tubes, and it has been found that this can be avoided by attention to the interaction between the test gas accelerating expansion and the contact surface in the primary shock tube. Test section measurements of pitot pressure, static pressure and flat plate heat transfer are reported. An approximate analytical theory has been developed for predicting the velocities achieved in the unsteady expansion of the ionising or dissociating test gas.

Recent experimental and computational studies have indicated that rearward curvature of a wing can reduce the induced drag factor to values less than that obtained from the unswept elliptical wing considered optimal in classical wing theory. Wake non-planarity associated with a wing with aft swept tips has been suggested as a reason for this behaviour. This paper provides a systematic examination of the wake non-planarities induced by a variety of wing planforms and quantifies the induced drag reductions which may be expected. Results suggest a dependence of induced drag factor upon angle of attack, and therefore upon lift coefficient. The crescent wing is identified as the planform to derive the most benefit from its non-planar wake shape, typically a 4% reduction in induced drag factor at moderate angle of attack.

This Note is concerned with the axisymmetric buckling of orthotropic circular plates. A circular plate fabricated with radial and/or circumferential reinforcements may be idealised as an orthotropic plate of uniform thickness h having different material properties in the radial and circumferential directions.

In this review paper, a recurring theme is the interplay between flying qualities as a technical discipline and flying qualities as an operational attribute. This interplay provides the setting for a presentation on the status of helicopter flying quahties, using maritime helicopter operations as the focus, particularly the recovery phase and the helicopter-ship dynamic interface. Poor weather, inducing high sea states and ship motion, along with complex, invisible and disturbing airflow and degraded visibility, make the dynamic interface a particularly demanding environment for both pilot and helicopter.

Flying qualities are a product of four elements — the aircraft, the pilot, the task and the environment — and the maritime application serves to give this viewpoint its full perspective. Mission-oriented flying quahties engineering is described within the systems framework of Aeronautical Design Standard 33 (ADS-33), utilising concepts like the mission task element, usable cue environment, response type and dynamic response criteria. The innovative constructs introduced by ADS-33 for flying qualities in degraded visual conditions are given special attention. The paper argues that the requirements for what constitutes safe and easy, or Level 1, flying qualities now exist and are well substantiated. New aircraft can now be designed to achieve these performance and safety standards while existing aircraft can be upgraded with integrated flight management systems featuring advanced control/flying qualities technologies.

In this context, the paper also promotes the concept of concurrent requirements capture and preliminary design, to maximise the likelihood that user requirements are achieved and that designs are robust and work first time. The multidisciplinary nature of flying qualities is emphasised, embracing aeromechanics and flight dynamics, controls and displays and human factors. Similarly, the importance of high-fidelity design tools and comprehensive evaluation methods in the concurrent process is stressed.

Handling deficiencies can increase the risk of helicopter accidents, particularly in degraded visual conditions or in emergencies where excursions beyond the operational flight envelope can lead to piloting difficulties. These situations are the new challenges for flying qualities engineers, and two areas are discussed here in some detail. First, flight in severely degraded visual conditions, highlighting the importance of understanding the fundamentals of human visual perception in the development of integrated control and display augmentation. Second, handling qualities following tail rotor failures are discussed and results from current research to develop new advice for aircrew are presented. The author takes the view that much more can and needs to be done to assist the pilot in the management of the tension between performance and safety in helicopter operations, through the provision of improved flying qualities.

By considering the equations of motion it has been shown that the flow in the outer part of a two-dimensional, curved, turbulent wall jet is approximately self-preserving if the ratio of jet thickness to wall radius of curvature is constant along the jet. This condition is satisfied for a jet blowing over a surface of logarithmic spiral profile, for which the radius of curvature R increases linearly with distance s along the wall.

Measurements of velocity profiles and rates of growth of wall jets for surfaces with curvature ratios and 1 are presented. These are compared with solutions obtained using an eddy viscosity theory, and with the flow of jets round circular cylinders. The measured jets are found to be approximately self-preserving in form, and to have rates of growth which are much larger than the jets on circular cylinders with corresponding values of s/R.

An account is given of a numerical method for calculating transonic flow past an aerofoil with an allowance for viscous effects, providing that the boundary layer remains fully attached over the aerofoil surface. The method has been developed by combining, in an iterative manner, calculations of the inviscid flow with calculations of the compressible boundary layer and wake. The solution for the inviscid flow is obtained by an iterative scheme, originally established by Garabedian & Korn, which has been modified to give a more realistic representation of shock waves. The boundary-layer development is treated as laminar initially; at a certain transition position a turbulent boundary layer is assumed to develop, and this is determined by the lag-entrainment method of Green et al. Comparisons of the results from the numerical scheme with some experimental measurements are shown for various examples in which shock waves of moderate strength are present. The method predicts, with reasonable accuracy, both the detailed pressure distribution and the variation of drag coefficient with lift coefficient.

The discrete element displacement method is used to analyse the instability of flat rectangular plates subjected to arbitrary systems of in-plane loading. Critical load intensities calculated in a range of applications, some of which include discrete reinforcing members, are compared with the predictions of past investigators and demonstrate the accuracy of the procedure.

It is generally accepted that the displacement patterns chosen to represent the deformation of a finite element should be such that there is complete compatibility with adjacent elements. Although some writers have doubted the necessity for the complete satisfaction of compatibility away from the reference nodes, provided the displacement pattern is capable of assuming certain uniform strain states, the desirability of complete compatibility is not in question. The displacement fields suggested in this note satisfy complete compatibility.

This paper presents results of an investigation of the wake flow of a series of two-dimensional bluff bodies with different boundary-layer separation angles. These results are compared with the results of an earlier investigation of a similar series of axisymmetric bodies. It is found that when the same length scales characterising the wakes that were proposed for the axisymmetric case are used in the current investigation, then there is good agreement between the two cases in the variation of the length scales with boundary-layer separation angle. The universal Strouhal number previously proposed for axisymmetric bodies is derived for the two-dimensional bodies under investigation and is found to remain a constant for the different bodies, but at a different numerical value from that found for the axisymmetric bodies.

After a brief discussion of alternative ways in which the drag of an aircraft wing can be derived theoretically, attention is focused on a technique whereby the separate ‘far-field’ components of drag — viscous, trailing-vortex (induced) and wave — are calculated separately. In particular, a new approximate method is described for estimating the wave drag. Based on an exact two-dimensional analysis involving the flow conditions just upstream of the shock wave, a simple formula is derived which, to first order, involves only a knowledge of the Mach number distribution on the surface of the aerofoil ahead of the shock and of the surface geometry at its foot. The accuracy of this formula is assessed for aerofoils by comparison with more exact theoretical results and with experiment. A ‘strip theory’ extension to swept wings is proposed and illustrated by applying it to a particular transport-type wing body combination. Using experimental pressure measurements as input, all three components of drag are estimated theoretically, and by adding their sum to separate balance measurements of the body drag comparisons between ‘theory’ and experiment for the overall drag can be made. These show a satisfactory standard of accuracy, the error varying between —5% and — 1% of the total drag over a wide range of Mach number and lift coefficient.

The effect of multigrid acceleration implemented within an upwind-biased Euler method for hovering rotor flows is presented. The requirement to capture the vortical wake development over several turns means a long numerical integration time is required for hovering rotors, and the solution (wake) away from the blade is significant. Furthermore, the flow in the region near the blade root is effectively incompressible. Hence, the solution evolution and convergence is different to a fixed wing case where convergence depends primarily on propagating errors away from the surface as quickly as possible, and multigrid acceleration is shown to be less effective for hovering rotor flows. It is found that a simple V-cycle is the most effective, smoothing in the decreasing mesh density direction only, with a relaxed trilinear prolongation operator. Results are presented for multigrid computations with 2, 3, 4, and 5 mesh levels, and a CPU reduction of approximately 80% is demonstrated for five mesh levels.

Most service failures in fatigue start on a free surface, as in many cases the most severe stress conditions exist there and are of a bi-axial stress nature, such as a drive shaft under combined bending and twisting. This particular case of multi-axial stress fatigue, because of its importance in design, has seen considerable investigation, although several aspects of the stress situation complicate the analysis of the test results with regard to the development of a general criterion of fatigue failure under multi-axial stress. The principal stress axes rotate with variations in the ratio of bending to twisting and the shaft is subjected to a stress gradient. Only the extreme portions of the shaft in the plane of bending are under the maximum bending stress, whereas the complete surface of the shaft is subjected to the maximum torsion stress.

Carbon fibres with high tensile strength are being increasingly utilised in the manufacture of advanced composite aerostructures. A Modified Compact Tension (MCT) specimen is often deployed to measure the longitudinal intralaminar fracture toughness but a high tensile strength often leads to premature damage away from the crack tip. We present an approach whereby the MCT specimen is supported by external fixtures to prevent premature damage. In addition, we have developed a novel measurement technique, based on the fibre failure strain and C-scanning, to determine the crack length in the presence of surface sublaminate delamination which masks the crack tip location. A set of cross-ply specimens, with a ((90/0)s)4 layup, were manufactured from an IMS60/epoxy composite system Two different data reduction schemes, compliance calibration and the area method, are used to determine the fibre-dominated initiation and propagation intralaminar fracture toughness values. Propagation values of fracture toughness were measured at 774.9 ± 5.2% kJ/m2 and 768.5 ± 4.1% kJ/m2, when using the compliance calibration method and the area method, respectively. Scanning Electron Microscopy (SEM) is carried out on the fracture surface to obtain insight into the damage mechanism of high-tensile-strength fibre-reinforced unidirectional composites. The measured tensile fracture toughness value is used in a fully validated computational model to simulate the physical test.

The finite element method is extended to the problem of finite plane strain of elastic solids. A highly elastic body subjected to two-dimensional deformations is represented by an assembly of triangular elements of finite dimension. The displacement fields within each element are approximated by linear functions of the local coordinates. Non-linear stiffness relations involving generalised node forces and displacements are derived from energy considerations. For demonstration purposes, the non-linear stiffness equations are applied to the problems of finite simple shear and generalised shear. For finite simple shear, it is shown that these relations are in exact agreement with finite elasticity theory. Convergence rates of finite element representations of these problems are briefly examined.

The computation of slowly moving Shockwaves is considered. A numerical error is introduced into the solution during an explicit computation, which results in the generation of spurious entropy. The generation of this spurious entropy and its effect on the pressure wave in Euler flows is examined, in one dimension by considering a moving shock in a constant area duct, and in two dimensions by considering an unsteady aerofoil flow. An efficient temporal discretisation is presented, whose advantages over the explicit scheme are two-fold. First, the numerical entropy generation due to shock motion is eliminated, and second, the CPU requirements for a two-dimensional unsteady computation are reduced by a factor of five from that required for the explicit scheme, for no loss of accuracy.

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.

The problem of designing a three-dimensional contraction arose recently in the course of the projected construction of a new low-speed (200 ft. /sec), closed-return wind-tunnel at the Northampton College of Advanced Technology. A three-dimensional contraction was required of contraction ratio 7·4:1 to provide smooth low turbulence conversion to the working section speed within an axial length of 126 inches. The inlet and outlet (working-section) dimensions are shown in Fig. 1 and the sections are similar. It was decided to design the contraction on the basis of an equivalent axi-symmetric contraction whose inlet and outlet circular areas are equal to those of the corresponding irregular octagons and then adapt the derived contraction wall profile to the actual sectional contours. The resulting equivalent inlet and outlet radii are then 55·4 in. and 20·4 in. respectively. The problem was thus reduced to the design of an axi-symmetric contraction of given area ratio within a specified axial length, with favourable velocity characteristics for low turbulence flow in the working section. This latter specification implies the absence of adverse pressure gradients on the walls near the working section and a small percentual difference between axial and wall velocities at the inlet to the working section.

This work presents a method for selecting the gain parameters of a C* control law for an aircraft’s longitudinal motion. The design incorporates various handling quality requirements involving modal, time- and frequency-domain criteria that were fixed by the aircraft manufacturer. After necessary model order-reductions, the design proceeds in essentially two-steps: stability augmentation system (SAS) loop design and control augmentation system (CAS) loop design. The approach partly relies on the use of guardian maps to characterise, in each case, the set of gain parameters for which desired handling quality requirements are satisfied. The approach is applied throughout the full flight envelope of a business jet aircraft and yields satisfactory results.

The value of computational fluid dynamics, CFD, delivered to date has mainly been related to its application to high-speed cruise design. To increase its applicability CFD must apply to the full flight envelope frequently characterised by large regions of separated flows. These flows are encountered by transport aircraft at low speed with deployed high lift devices, at their structural design loads conditions, or subjected to in-flight upsets that expose them to speed and/or angle-of-attack conditions outside the envelope of normal flight conditions to name a few. Such flows can only be characterised by the Navier-Stokes equations. This paper will report the progress toward CFD for full flight envelope. The CFD methods in use at Boeing will be described. Examples presented will address high-lift, loads and stability and control concerns including Reynolds scaling from wind tunnel to flight, vortex generator simulation, spoiler and horizontal tail effectiveness. In general, results shown are in ‘good enough’ agreement with experimental data. Deficiencies and the need for further algorithm and process improvement are noted. The need for automation to enable the large scale use of CFD will also be discussed.