An experimental investigation of the viscous decay of a steady, three-dimensional, helical vortex is presented. The vortex was generated by a rectangular, symmetrical half aerofoil (N.A.C.A. 0009) of 10·5 in. span, 3 in. chord, cantilevered from the wall of the circular test section of a low speed wind tunnel. Only the local total head and flow directions were measured with a five-hole pressure sensitive probe at one to eight chord lengths behind the trailing edge of the aerofoil, and the radial, tangential and axial velocity distributions in the trailing vortex were derived. The static pressure variation in the vortex flow and the difficulties in its accurate measurement are discussed. All tests were performed at a Reynolds number based on the free-stream velocity and chord length of the aerofoil of 10,000.

The various experimental velocity distributions in the vortex were in good agreement with a linearised, axi-symmetric, laminar, incompressible, viscous vortex due to Newman. The decay of velocity and geometry parameters, however, were found to be approximately eight to ten times faster than predicted by the analysis based on a laminar kinematic viscosity. It was also observed that the amount of circulation which actually rolled up in the trailing vortex was only about 58 per cent of that expected around the (N.A.C.A. 0009) aerofoil used for this investigation.

This paper presents highlights from research conducted at the University of Liverpool to determine suitable fidelity criteria and guidelines for the modelling and simulation of the helicopter-ship dynamic interface environment. The paper begins by describing the characteristics of the helicopter-ship dynamic interface, explaining the motivation behind the research and reviewing the state-of-the-art in dynamic interface simulation. The development of a dynamic interface research environment based on an existing research simulator operated by the University of Liverpool is then described, before key results from a number of piloted simulation experiments are presented. These experiments were specifically designed to address fidelity sensitivity issues, such as, are unsteady airwake models necessary, or can a steady airwake model induce appropriate levels of pilot workload? What influence does the modelled ship geometry, or choice of atmospheric wind conditions have on the airwake model and on pilot workload? Finally, the paper concludes by briefly describing the relevance of these research findings to current and future industry programmes.

Wire gauzes are used in air duct systems and wind tunnels for many purposes. Methods of estimating the pressure loss through a gauze currently available are often found in practice to be unsatisfactory. In this paper data obtained by several experimenters are correlated. It is shown that the effect of Reynolds number on pressure drop is similar for a large range of gauze porosities. An empirical method of estimation of the pressure drop is given which can be used with confidence for gauzes of porosity (defined as the proportion of area not blocked by the wires) from 0·27 to 1·00, and over a range of Reynolds number, based on approach velocity and wire diameter, from 20 to 600.

An Euler optimisation for a BWB configuration with winglets incorporating an array of three-dimensional shock control bumps is carried out by employing an efficient adjoint-based optimisation methodology. A high fidelity multi-block grid with over two million grid points is generated to resolve the shape of the 3D shock control bumps, the winglet as well as the overall BWB shape, which are parameterised by over 650 design variables. In order to perform such a large aerodynamic optimisation problem feasibly, the optimisation tools such as the flow solver and the adjoint solver have to be parallelised with a good parallel efficiency. This paper reports the parallel implementation efforts on the adjoint solver; especially on the calculation of the sensitivity derivatives, which has to be looped over the total number of design variables. Results from the optimisation of the wing master sections, winglet aerofoil sections and the three dimensional bumps indicate a significant improvement regarding the aerodynamic performance against the baseline geometry for the given planform layout of the aircraft.

This paper examines the dynamic stalling of a finite wing of aspect ratio 3·0 when subject to constant pitch motions up to and beyond stall. In particular, unsteady surface pressure data were obtained at 192 locations on the wing surface and these were then analysed to provide information on the nature and phasing of dynamic stall events in both the chordwise and spanwise directions. It was also possible to obtain sectional force and moment coefficients by integration of the pressures measured on specific chordal arrays. This provided valuable insight into the load distribution on the wing throughout the range of motion. On this basis, it was established that the wing loading distribution was consistent with conventional understanding of steady wing loading up to the incidence at which the dynamic stall vortex was initiated. Beyond this point, the formation and subsequent convection of the vortex structure was found to be strongly three-dimensional but, nevertheless, exhibited many of the features of two-dimensional dynamic stall.

Experiments are described in which the lift on a rectangular element of a two-dimensional wing and on a finite aspect ratio wing has been measured in grid turbulence. By measuring the spectrum of the lift and the spectrum of the turbulence upwash component, an experimental value for the turbulent admittance may be found. This is compared with a calculated value based upon linearised theory.

A new wing-tip concept with morphing upper surface and interchangeable conventional and morphing ailerons was designed, manufactured, bench and wind-tunnel tested. The development of this wing-tip model was performed in the frame of an international CRIAQ project, and the purpose was to demonstrate the wing upper surface and aileron morphing capabilities in improving the wing-tip aerodynamic performances. During numerical optimisation with ‘in-house’ genetic algorithm software, and during wind-tunnel experimental tests, it was demonstrated that the air-flow laminarity over the wing skin was promoted, and the laminar flow was extended with up to 9% of the chord. Drag coefficient reduction of up to 9% was obtained when the morphing aileron was introduced.

The forces on an aerofoil have been measured at various angles of incidence for 0.5 < R < 2,000.

Below R = 2 the lift is found to be given approximately by :—

L, = (3/2) µlV sin 2x

and the drag at all angles of incidence tends to:

D = µlV.

An approximate theoretical estimate is also given of the effect of spacing, on the drag of a series of flat plates at R → o.

The challenges of helicopter simulation are being tackled across a broad front as technology is developed to meet the needs of Industry. Traditionally, the strongest impetus has come from the training community and this is likely to continue for some time as simulation technology advances at increasing pace, raising fidelity standards. The development of PC-based simulation technologies is providing a significant spur in this development and lowering the cost, making complete simulation systems of reasonably high fidelity available to smaller organisations. This paper describes the first year of operation with such a system at the University of Liverpool – HELIFLIGHT. With its full motion, wide field-of-view visuals, programmable force-feel system and the comprehensive FLIGHTLAB modelling environment, we describe the HELIFLIGHT system as high fidelity and the first year of utilisation saw extensive use in a variety of handling qualities and pilot-vehicle technology research and teaching. Some of these are described in the paper and from a year of many highlights, the EU-funded programme to develop handling qualities for a civil tilt rotor aircraft is selected to demonstrate the versatility of the HELIFLIGHT system.

The unsteady air flow over and around the helicopter landing deck of a naval vessel is known to cause high pilot workload and to limit the helicopter's operational envelope for launch and recovery. Previous research has suggested that modifications to the ship's hangar edges can beneficially modify the flow over the deck. This paper examines the effectiveness of five hangar-edge modifications using computational fluid dynamics–generated airwakes and flight mechanics modelling, as well as piloted flight trials in a motion-base simulator. Results are presented, in terms of unsteady helicopter loads and pilot workload ratings, for modifications to the windward vertical rear edge of the hangar and with an oblique wind. The results demonstrate that while the airwake can be altered by superstructure modifications, the integrated effect of the altered airwake on the entire helicopter does not necessarily give the desired result; indeed of the five modifications tested, two were seen to be beneficial while three caused an increase in pilot workload compared with the unmodified hangar. Overall, the paper shows that the airwake can be modified by superstructure design changes, and that the effect on the helicopter can be measured through modelling and simulation. It is also demonstrated that making judgements on the severity of the airwake based on the aerodynamic flow field alone can be misleading. The benefit of these simulation tools is that they can be used during the ship design process to evaluate the effect of the superstructure aerodynamics, rather than wait until after the ship is built.

The Clean Sky Joint Undertaking is currently managing two large-scale research and innovation programmes under FP7 and Horizon 2020 to contribute to the strengthening of the European aeronautical sector ensuring global leadership and competitiveness. This paper describes the research and demonstration programmes in Clean Sky (2008–2017) and Clean Sky 2 (2014–2024) related to propulsion technologies for the next-generation aircraft. The bulk of this work is addressed in Clean Sky 1 under the “Sustainable And Green Engines” (SAGE) programme and under the “ENGINES” programme in Clean Sky 2. The High-Level Objectives are described for each engine architecture as well as the targets in terms of CO2 and noise reduction versus a year 2000 reference unless stated otherwise. An overview of the new engine concepts that would satisfy the ACARE objectives is presented, including the main technologies which are to be developed to ensure the successful demonstration of each of those new engine concepts.

A lecture called “Some aspects of buffeting” was given to the Society in 1961 by my colleague Mr. R. Fail. In that lecture the emphasis was primarily on the prediction of buffet onset in flight from model tests in wind tunnels, and on the prevention of buffeting. In the subsequent 11 years there has been much more interest in the prediction of the severity of buffeting above the buffet boundary, for both transport and fighter type aircraft, if buffeting cannot be prevented or alleviated. I have tried to reflect this interest and I hope that the title of this paper will also suggest a cautious examination of the physical processes at work above the buffet boundary, when the boundary layer has separated. We still really know very little about these processes but I hope that the paper may stimulate further research and questioning.

This note attempts to clarify some basic ideas, and correct some common misunderstandings, on wing flexure-torsion flutter. Topics include:

• (i) an understanding of the physics of wing flexure-torsion flutter through a number of specific examples,

• (ii) an examination of the energy interpretation of the physics of flutter, indicating its limited usefulness,

• (iii) a discussion of the significant differences between wing flexure-torsion flutter and the Duncan flutter engine,

• (iv) a graphical representation to assess the contribution of various parameter to flutter onset,

• (v) subcritical response below the critical flutter speed,

• (vi) some practical wing flutter considerations.

Measurements of momentum in regions of turbulent mixing show that an analogy exists between the processes of turbulent and molecular balance which enables a rational application of the differential equation of thermal conduction to the propagation of turbulent momentum to be made. This fact forms the basis of a new phenomenological theory of free turbulent flows.

An important prerequisite for the design, assessment and certification of aircraft and their associated control systems is a quantitative specification of the environment in which the aircraft is intended to operate, for example, atmospheric gust. Gust loads on aircraft may induce detrimental influences such as increased aerodynamic and structural loads, structural deformation and decreased flight dynamic performance. This paper presents a systematic and comprehensive overview of important concepts and applications of gust loads on aircraft. This overview includes a brief research background, concepts, research techniques, influences and load alleviation measures of gust. Finally, we summarise some potential improvements in the future work. It is also recommended to learn from previous experiences to avoid aviation accidents due to flight through atmospheric gusts and turbulence.

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 buffet characteristics of a supercritical aerofoil were investigated in the High Reynolds Number Two-Dimensional Test Facility of the National Aeronautical Establishment (NAE). The test was performed quite deep into the buffet régime. The buffet onset boundary was determined from the divergence of the unsteady balance normal force. Steady and unsteady pressure measurements were obtained for shock-induced separation with reattachment as well as fully separated flows. For flow conditions where discrete shock wave oscillations occurred, ensemble-averaging of the unsteady pressures determined the propagation velocity of the pressure wave induced by periodic shock motion. A model of the self-sustained oscillatory shock motion due to shock-boundary layer interaction was formulated. Broad-band cross-correlations of the pressure field determined the downstream convection velocities of the turbulent eddies of separated flows.

A study of shock tube operation with free-piston compression of the driver gas is presented. It is shown theoretically and experimentally that shock Mach numbers exceeding 30 in air may be obtained with an apparatus of moderate size. The operating limitations imposed by the maximum pressures allowed in the apparatus, by the ratio of the volume of the shock tube to the volume of the compression tube, and by the maximum pressure available for driving the piston are all explored theoretically. The compression process is studied experimentally, and results for three different compression tubes are used to indicate empirical rules for maximising the efficiency of the compression process. Streak photographs of the movement of the luminous gas behind the shock wave in a 1 in diameter shock tube are used to demonstrate the quality of the shock-wave motion produced.

The aim of the study described herein was to develop and verify an efficient neural network based method for extracting aircraft stability and control derivatives from real flight data using feed-forward neural networks. The proposed method (Modified Delta method) draws its inspiration from feed forward neural network based the Delta method for estimating stability and control derivatives. The neural network is trained using differential variation of aircraft motion/control variables and coefficients as the network inputs and outputs respectively. For the purpose of parameter estimation, the trained neural network is presented with a suitably modified input file and the corresponding predicted output file of aerodynamic coefficients is obtained. An appropriate interpretation and manipulation of such input-output files yields the estimates of the parameter. The method is validated first on the simulated flight data using various combinations and types of real-flight control inputs and then on real flight data. A new technique is also proposed for validating the estimated parameters using feed-forward neural networks.

Extensions to Maskell’s original correction method, developed over several years, are consolidated and designated ‘Maskell III’. The procedures were applied to dedicated tests on a family of flat-plate wing models in a small, low-speed wind tunnel at NRC. Test conditions included angles of attack from -10° to 110° and models of up to 16% of tunnel area. Off-centre tests were included with model-to-wall distances down to 0.72 chords. Corrected lift and drag data correlated well between models of different sizes and at different offsets from the tunnel centreline. Comparisons are made with corrections using the pressure-signature and two-variable methods, emphasising post-stall conditions. These showed that the ‘Maskell III’ procedures, which require minimal input, correlated as well as the other methods for most model sizes and positions in the tunnel.