The design, construction, and validation of a new academic wind tunnel is described in detail. The wind tunnel is of a classical, blow-down type and generates a pressure-induced, turbulent separation bubble on a flat test surface by a combination of adverse and favorable pressure gradients. The Reynolds number, based on momentum thickness just upstream of separation, is Reθ ≃ 5,000 at a free-stream velocity of Uref = 25ms−1. The length of the separation bubble is estimated at 0°42 ± 0°02m by three different methods. Results of a numerical simulation demonstrate the absence of flow separation in the wind-tunnel contraction. This results in a turbulence level of about 0·05% in the test section. Oil-film visualisation experiments show that the flow near the wall is strongly three-dimensional in the recirculating region and that the topology of the limiting streamlines is consistent with experiments performed on configurations with fixed separation. Finally, spatial variations of the forward-flow fraction have been documented using a thermal-tuft probe and are shown to compare well with the results of the oil-film visualisation.

This third technical note, on the background of the author's Main Lecture to the Society(1), illustrates the application of the technique given in ref. 2 on a simple example, and develops the basic ideas further, with particular reference to the effect of reinforcing flanges and initial loads, as may arise through differential thermal straining. A more ambitious example demonstrating the improvement achieved by the new triangle appears in ref. 1.

Comparative mixing of a Mach 2 elliptical free jet from a convergent-divergent elliptic nozzle with an aspect ratio of 2:1 in the presence of adverse and marginally favourable pressure gradients has been studied experimentally. It is found that the mixing of the elliptical jet is higher than that for the equivalent circular jet at all the levels of expansion. The decay of elliptic jet is significantly higher than the equivalent circular jet in all three zones of the jet field – the core, characteristic decay and fully developed regions. The reason for the faster decay of the elliptic jet is found to be the continuous variation in the size of the mixing-promoting vortices shed from the nozzle exit owing to its azimuthal asymmetry. The evolution of the jet and its axis-switching phenomenon has been studied using iso-pitot pressure contours taken at different axial locations in the plane normal to the jet axis. As expected, the elliptic jet spreads faster along the minor axis plane than the major axis plane, leading to axis-switching at all the levels of expansion studied. The axis-switching of the elliptic jet shifts upstream with increase in nozzle pressure ratio (NPR) from 4 to 5; from 5 to 7, it shifts downstream. But at marginally under-expanded condition of NPR 8, the axis-switching is found to shift slightly upstream. The occurrence of axis-switching in the elliptic jet indicates enhanced near-field mixing, compared to the equivalent circular jet. The shadowgraph pictures of the jet reveal that the waves prevailing in the elliptic jet are significantly weaker than those in the circular jet.

Experiments were conducted to study the types of flow that occur on the lee surface of delta wings at supersonic speeds. Two sets of flat topped delta wings of different thickness (wedges with 10° and 25° normal angle respectively), each with leading edge sweep angles of 45°, 50°, 60° and 70°, were tested. The measurements, carried out at Mach numbers of 1·4, 1·6, 1·8, 2·0, 2·5 and 3·0, included oil flow visualisations (on both sets of wings) and static pressure distributions (on the thicker wing only). In addition, a 60° sweptback delta wing with a normal angle of 40° was also tested. The tests on this wing included both oil flow visualisations and static pressure measurements. From these and other existing measurements, the leeside flows have been classified into nine distinct types, namely (i) leading edge separated flow with secondary separation, (ii) leading edge separated flow with secondary and tertiary separation, (iii) leading edge separated flow with a shock wave beneath the primary vortex, (iv) leading edge separated flow with shock-induced secondary separation, (v) fully attached flow, (vi) flow attached at the leading edge with inboard shock-induced separation, (vii) mixed type of flow, (viii) flow with a leading edge separation bubble and (ix) leading edge separated flow with a shock wave lying on the lee surface in between the leading edge vortices. These types of flow have been displayed in a plane of Mach number and angle of attack normal to the leading edge. The experimental results indicate that increasing wing thickness has no qualitative effect on the types of flow observed but does shift the boundaries between some of the types of flow.

This paper investigates the effects of the design variables of an aerial deployment mechanism on the robustness of the aerial deployment through a multibody dynamics simulation. The aircraft is modelled as three joined rigid bodies: a right wing, a left wing and a centre body. A spring-loaded hinge is adopted as an actuator for deployment. The design variables are the hinge torque and the deployment timing. The robustness is evaluated using a sigma level method. The margins for the safe deployment conditions are set for the evaluation functions. The dispersive input variables are the initial drop velocity, the surrounding gust velocity, the initial pitch angle and the initial height. The design point with a deployment torque scale value F of 0.7 and a right-wing deployment delay time TSR of 1.0 s can safely deploy in the low-torque deployment condition. This design point is able to accomplish both a safe deployment and a lightweight deployment mechanism.

Multicriteria trajectory optimisation is expected to increase aviation safety, efficiency and environmental compatibility, although neither the theoretical calculation of such optimised trajectories nor their implementation into today’s already safe and efficient air traffic flow management reaches a satisfying level of fidelity. The calibration of the underlying objective functions leading to the virtually best available solution is complicated and hard to identify, since the participating stakeholders are very competitive. Furthermore, operational uncertainties hamper the robust identification of an optimised trajectory. These uncertainties may arise from severe weather conditions or operational changes in the airport management. In this study, the impact of multicriteria optimised free route trajectories on the air traffic flow management is analysed and compared with a validated reference scenario which consists of real flown trajectories during a peak hour of Europe’s complete air traffic in the upper airspace. Therefore, the TOolchain for Multicriteria Aircraft Trajectory Optimisation (TOMATO) is used for both the multicriteria optimisation of txrajectories and the calculation of the reference scenario. First, this paper gives evidence for the validity of the simulation environment TOMATO, by comparison of the integrated reference results with those of the commercial fast-time air traffic optimiser (AirTOp). Second, TOMATO is used for the multicriteria trajectory optimisation, the assessment of the trajectories and the calculation of their integrated impact on the air traffic flow management, which in turn is compared with the reference scenario. Thereby, significant differences between the reference scenario and the optimised scenario can be identified, especially considering the taskload due to frequent altitude changes and rescinded constraints given by waypoints in the reference scenario. The latter and the strong impact of wind direction and wind speed cause wide differences in the patterns of the lateral trajectories in the airspace with significant influence on the airspace capacity and controller’s taskload. With this study, the possibility of a successful 4D free route implementation into Europe’s upper airspace is proven even over central Europe during peak hours, when capacity constraints are already reaching their limits.

This paper presents a new flight trajectory optimisation method, based on genetic algorithms, where the selected optimisation criterion is the minimisation of the total cost. The candidate flight trajectories evaluated in the optimisation process are defined as flight plans with two components: a lateral flight plan (the set of geographic points that define the flight trajectory track segments) and a vertical flight plan (the set of data that define the altitude and speed profiles, as well as the points where the altitude and/or speed changes occur). The lateral components of the candidate flight plans are constructed by selecting a set of adjacent nodes from a routing grid. The routing grid nodes are generated based on the orthodromic route between the flight trajectory’s initial and final points, a selected maximum lateral deviation from the orthodromic route and a selected grid node step size along and across the orthodromic route. Two strategies are investigated to handle invalid flight plans (relative to the aircraft’s flight envelope) and to compute their flight performance parameters. A first strategy is to assign a large penalty total cost to invalid flight profiles. The second strategy is to adjust the invalid flight plan parameters (altitude and/or speed) to the nearest limit of the flight envelope, with priority being given to maintaining the planned altitude. The tests performed in this study show that the second strategy is computationally expensive (requiring more than twice the execution time relative to the first strategy) and yields less optimal solutions. The performance of the optimal profiles identified by the proposed optimisation method, using the two strategies regarding invalid flight profile performance evaluation, were compared with the performance data of a reference flight profile, using identical input data: initial aircraft weight, initial and final aircraft geographic positions, altitudes and speed, cost index, and atmospheric data. The initial and final aircraft geographic positions, and the reference flight profile data, were retrieved from the FlightAware web site. This data corresponds to a real flight performed with the aircraft model used in this study. Tests were performed for six Cost Index values. Given the randomness of the genetic algorithms, the convergence to a global optimal solution is not guaranteed (the solution may be non-optimal or a local optima). For a better evaluation of the performance of the proposed method, ten test runs were performed for each Cost Index value. The total cost reduction for the optimal flight plans obtained using the proposed method, relative to the reference flight plan, was between 0.822% and 3.042% for the cases when the invalid flight profiles were corrected, and between 1.598% and 3.97% for the cases where the invalid profiles were assigned a penalty total cost.

In this paper we study large angle rotational manoeuvres of spacecraft. The problem of attitude control is formulated and solved through potential functions, thus simplifying frequent reorientation manoeuvres. A combination of gas jet and FEEP thrusters is employed in order to obtain short settling times while meeting stringent pointing requirements. In this way, spacecraft reorientation is achieved in two integrated steps by means of a discrete/continuous control law. Autonomous avoidance of undesired space orientations is obtained and constraints due to the solar array pointing requirements are satisfied. A case study is discussed where the methodology is applied to a large space telescope.

The purpose of a joint is to transfer load between the two parts being joined. As a result of this load transfer there will be a stress variation in the components in the joint region, as well as stresses in the joining medium (fasteners or adhesive). In fhjs work, the stress results of a two-dimensional finite element (FE) analysis are used to understand failure modes of a bolted joint in low-temperature cure carbon fibre reinforced plastic (CFRP) woven laminates loaded in tension, and to predict the bearing strength. Good agreement with experimentally observed damage modes and strengths is achieved with, in some cases, a difference of less than 10%.

Piezoelectric wafer active sensors (PWAS) are lightweight and inexpensive transducers that enable a large class of structural health monitoring (SHM) applications such as: (a) embedded guided wave ultrasonics, i.e., pitch-catch, pulse-echo, phased arrays; (b) high-frequency modal sensing, i.e., electro-mechanical impedance method; and (c) passive detection. The focus of this paper is on the challenges posed by using PWAS transducers in the composite laminate structures as different from the metallic structures on which this methodology was initially developed. After a brief introduction, the paper reviews the PWAS-based SHM principles. It follows with a discussion of guided wave propagation in composites and PWAS tuning effects. Then, the mechanical effect is discussed on the integration of piezoelectric wafer inside the laminate using a compression after impact. Experiments were performed on a glass fibre laminate, employing PWAS to measure the attenuation coefficient. Finally, the paper presents some experimental and multi-physics finite element method (MP-FEM) results on guided wave propagation in composite laminate specimens.

Within the framework of a major European co-operative research project on rotorcraft aerodynamics and acoustics (Helishape), a parametric model rotor test was conducted in the open test section of the DNW using DLR's MWM test rig, and a highly instrumented model of a fully articulated ECF rotor equipped with blades of advanced design with two exchangeable blade tips. One set of blade tips (7A) was rectangular, while the other (7AD1) was a swept-back parabolic-anhedral shape. The objectives of this experimental research were to evaluate noise reduction techniques (conceptionally by variation of rotor speed, dedicated tip shapes and advanced aerofoils, as well as operationally, by identifying low noise, BVI-minimising, descent procedures) and to validate the partners’ aerodynamic and aeroacoustic codes. A comprehensive set of simultaneous acoustic and aerodynamic blade surface pressure data and blade dynamic and performance data was measured. In addition, valuable information on the tip-vortex geometry and blade-vortex miss distance was obtained by LLS flow visualisation. The experimental equipment, the test procedures, and the test matrix are briefly described. A survey of the main results is presented and the trends of the most important parameter variations for both rotors are discussed.

An improved variational optimization approach is established to optimize and analyse the propulsion efficiency of the high-altitude contra-rotating propellers for high-altitude airships based on the Vortex Lattice Lifting Line Method. The optimum radial circulation distribution, chord and pitch distribution are optimized under the maximum lift-to-drag ratio of aerofoils. To consider the effects of the actual Reynolds number and the Mach number of each aerofoil section, aerodynamics such as lift coefficient, drag coefficient and lift-to-ratio are obtained by interpolating a CFD database, which is established by numerical simulations under different Reynolds number, Mach number and angles-of-attack. The improved method is verified by validation cases on a high-altitude CRP using the three-dimensional steady Reynolds-averaged Navier-Stokes solver and moving reference frames technique. The optimization results of thrust, torque and efficiency for both the individual front/rear propeller and CRP are shown to agree reasonably well with the CFD results. Using the improved approach, the influence of blade numbers, diameter, rotation speeds, axial distance and torque ratio on the optimum efficiency of CRPs is illustrated in detail by conducting parametric studies.

The inflation of a parachute encompasses problems of aerodynamics, dynamics, and elasticity. The most noticeable effects of the interaction of the events are the opening force and the filling time. Attempts to analyse the opening process and to predict the opening force date back to 1927. Since then a considerable number of methods have been published. Some of them are primarily analytical, others more empirical, and several deal with isolated problems. The earliest and most analytically oriented studies are by Scheube and O'Hara.

It is well known that the main purpose of the by-pass principle is to improve the propulsive efficiency of a simple jet engine by removing some of the energy left in the jet gases and using this to compress an extra quantity of air, known as the by-pass air, this air being ejected rearwards with the jet gases. In this way a greater mass of air is ejected rearwards at a lower jet velocity and thus a better propulsive efficiency is obtained. This is an extremely simplified view of the advantages of the by-pass engine, however, since an equivalent result of obtaining a lower jet velocity can be obtained by designing the jet engine for a lower combustion temperature. The by-pass principle is of advantage because it enables a higher propulsive efficiency to be obtained at the same time as employing a high combustion temperature and therefore a high basic cycle efficiency. If the component efficiencies of a gas turbine were 100 per cent, cycle efficiency would not depend upon combustion temperature at all, and there would thus be no advantage in principle in using the by-pass engine. In practice there would probably be some residual advantage left in that for a given thrust a lower engine weight could be obtained.

Classical gas dynamics is by definition the dynamics of gases that are in local thermodynamic equilibrium everywhere in the flow field so that their thermodynamic behaviour is completely specified by two independent variables of state, for example by pressure p and density p. The fundamental assumption of local thermodynamic equilibrium is not valid in many flow situations of practical importance, and this fact has given rise to a revision and generalisation of classical gas dynamics. This generalisation is usually called “relaxation gas dynamics”. Ten years ago J. Ackeret mentioned this subject briefly at the end of the Daniel and Florence Guggenheim Memorial Lecture on The Role of Entropy in the Aerospace Sciences. He gave that lecture at the 2nd International Congress of the Aeronautical Sciences in Zürich (12th September 1960) and he said on that occcasion: “Non-equilibrium states are of growing importance, for strictly speaking, exact (thermodynamic) equilibrium occurs only relatively seldom in gas dynamics. However, a rational theory of nonequilibrium states must necessarily be extremely complicated….

Some of the uses of ground based flight simulation as a research tool to aid the design of new aircraft and their equipment are described. The function of the simulator is to provide a method for investigating humon flying tasks in the laboratory, so that the relationship between the pilot's capabilities and the equipment's characteristics can be systematically studied. The paper is presented in two parts describing recent work on two research simulators at the RAE.

Part I deals with the use of simulation for studying aircraft stability and control characteristics. The equipment used at RAE for this work is described, with particular emphasis on methods of presenting to the pilot a simulated view of the outside world, and for reproducing some of the motion cues which he experiences in flight. Experimental evidence of the importance of these simulation cues when making aircraft control assessments is also presented. Several examples of simulation studies into the control of conventional and VTOL aircraft are given, to illustrate the type of research problems in this field which may be tackled and the techniques involved in solving them.

Optimisation of aircraft ground operations to reduce airport emissions can reduce resultant local air quality impacts. Single engine taxiing (SET), where only half of the installed number of engines are used for the majority of the taxi duration, offers the opportunity to reduce fuel consumption, and emissions of NOX, CO and HC. Using 3510 flight data records, this paper develops a model for SET operations and presents a case study of London Heathrow, where we show that SET is regularly implemented during taxi-in. The model predicts fuel consumption and pollutant emissions with greater accuracy than previous studies that used simplistic assumptions. Without SET during taxi-in, fuel consumption and pollutant emissions would increase by up to 50%. Reducing the time before SET is initiated to the 25th percentile of recorded values would reduce fuel consumption and pollutant emissions by 7–14%, respectively, relative to current operations. Future research should investigate the practicalities of reducing the time before SET initialisation so that additional benefits of reduced fuel loadings, which would decrease fuel consumption across the whole flight, can be achieved.

In this paper a simple approximation method is presented for rapidly computing the lift distributions of arbitrary aerofoils. The numerical results are compared with those obtained by an exact method and for many purposes show a satisfactory degree of accuracy. The latter, for all practically occurring cases, can be estimated at the start of the computation work with the aid of the comparison examples given.

A detailed and accurate three-dimensional finite element stress analysis was conducted on countersunk rivet holes in a plate subjected to tension loading. The analysis included a wide range of countersunk depths, plate thicknesses, countersunk angles and plate widths. The study confirmed some of the previous results, addressed their differences, provided many new results, and investigated countersunk angle and width effects. Using the detailed FE results and the limiting conditions, a design equation for stress concentration was developed and verified.

The “post-stall” manoeuvre capabilities of current and projected combat aircraft have prompted a resurgence of interest in the dynamic response to wing motion of the vortex burst at high angles of attack. A large number of experimental studies have been reported, but these have tended to be restricted to a limited range of motion profiles and a single specific wing configuration. Presumably due to the complexity of the phenomenon, no attempt appears to have been made to undertake a systematic study of the burst response for a range of wing planforms and motion profiles. Hence, a survey was made of the published data and a semi-empirical correlation derived using a system identification methodology. For preliminary performance studies, the dynamic response of the vortex burst is modelled adequately by a second order lag transfer function. The physical significance of this system model and the implications for the design of a systematic experimental study are discussed.