During a recent open week at the National Gas Turbine Establishment at Pyestock, an experiment involving just over 1,000 subjects was made to establish some measure of the disturbance value of such rhythmic or amplitude modulated noises as are generated by helicopter rotors. The results were (i) that, as may be implied from Niese's results with modulated tones, the subjective assessment of such a noise depends upon its peak sound pressure level (i.e. 82·5 db. in Fig. 3) rather than upon its r.m.s. value, (ii) that the effect of the modulation frequency is negligible in the range from 4 to 12 cycles/sec. and (iii) that these results are invariant with peak sound pressure level in the range from 85 to 95 db. re 2 x 10–4 dynes/cm2.

A computational study has been carried out to analyse the shock wave turbulent boundary layer interaction in a two-dimensional compression ramp flow for a free stream Mach number of 2·94. Ramp angles ranging from 8° to 24° are used to produce the full range of possible flow fields, including flows with minor separation, moderate separation, and significant amount of separation. The model has been analysed using 2D numerical simulations based on a commercially available Computational Fluid Dynamics (CFD) Code, Fluent, that employs k-ω Shear Stress Transport (SST) turbulence model. The computed data for surface pressure distribution indicated a good agreement with experiment. Numerical results obtained through the present series of computations indicate an increased extent of separated zone, and thus show increased upstream influence when compared to experiment. Total pressure loss has shown to increase downstream of separation location and increase when corner angle increases. Secondary separation has been observed for higher angles.

The design of a vehicle with cruise capability which can be packaged in a volume limited space, such as a cylindrical tube, centres around the problem of folding the lifting and control surfaces in such a manner that their deployment is simple and reliable. Several folding schemes are illustrated in Fig. 1. A wing fabricated of a flexible material and suspended by lines from the fuselage (parawing, Fig. 1a) can be folded to occupy only a small volume. These wings, however, are difficult to deploy at high forward speeds and have characteristically poor lift/ drag ratios unsuitable for high performance cruise capability. Another concept which allows storing the wing in a volume-limited space is the “scissor” wing (Fig. 1b) which, when deployed, swings out from a cavity in the fuselage.

A three-dimensional unsteady Navier-Stokes solver has been used to simulate stall inception in a single row ten passage segment of a transonic fan, the NASA rotor-67. At subsonic flow conditions, the 3D results illustrate a rotating stall inception with short scale part-span cells rotating at around 80% rotor speed, similar to that observed in some low speed experiments. However, at a supersonic relative inflow condition, the results show that an isolated blade row tends to stall in a one-dimensional breakdown pattern without first experiencing rotating stall. At near-stall conditions, significant self-excited unsteadiness is generated by the interaction between the tip-leakage vortex and the passage shock wave. Further computations for two-dimensional configurations indicate that it is possible to have a rotating pattern of instability in transonic blade rows associated with circumferential synchronised shock oscillation.

I intend in this lecture to discuss tlie ways in which I think analogue computing can be of assistance to aeronautics in the future. I shall therefore be less concerned with how they are built, than with how they are used, and in fact, what circumstances justify using them at all. Up to a few years ago the only large computors were analogue ones, so the user had no choice, but when the automatic digital computor became established as a practical machine (and even before) there were many people to suggest that the analogue was obsolete. This view is now less widely held, even by makers and designers of digital machines, and reviews of aeronautical computing aids have usually found some place for the analogue computor, without establishing any general principles on which the choice has been made, except, perhaps, to suggest that the accuracy required is the criterion.

I shall start by describing three of the types of computor most commonly used in aeronautics, the “ response simulator,” the “ electrolytic tank” and the “network analyser” and use these to illustrate the particular properties of analogue computing.

The results of a series of shear buckling tests on shallow, flat-sided, corrugated webs are presented, and a comparison made with a theoretical analysis in which the web is assumed to buckle either as an orthotropic plate or in a local mode. An expression is derived for the efficiency of a corrugated web in shear, and used to compare the four shapes of cross-section tested.

The recent period has been filled with exceptionally interesting developments and advances, resulting in high-performance conventional and non-conventional manned and unmanned aircraft. Although those vehicles seem to comply well with specific mission performance requirements, one is still confronted with an apparent weakness to reliably stabilise and control throughout the flight envelope. Since the provision of satisfactory stability and control characteristics invariably compromises flight performance, it becomes essential to identify and integrate performance-optimal stability and control design solutions early during the flight vehicle definition phase. In particular, the conceptual design of integrated control effectors for advanced aircraft is far from being trivial. Never before have we been presented with such tremendous wealth of specialised data and information suitable for detail design of controls. In contrast, never before has it been necessary to approach any one of the primary design disciplines still as entirely ad hoc and inconsistent as in the case of designing controls during the conceptual design phase. This need initiated the development of a configuration independent (generic) stability and control methodology capable of sizing primary control effectors of fixed wing subsonic to hypersonic designs of conventional and unconventional, symmetric and asymmetric configuration layouts. This paper summarises the methodology concept and demonstrates its versatility and validity by analyzing selected stability and control characteristics of the Northrop YB-49 flying wing.

When the Royal Aeronautical Society did me the honour several weeks ago of asking me to give a lecture on sound-proofing of aircraft, I hoped to be able to prepare a formal paper and read it to you. I did not realise at that time, however, that my temporary duties as consulting engineer to the civil aircraft division of the French Air Ministry would take practically all my time. So I arrived in London with only the scantiest notes and shall be obliged to talk more or less extemporaneously on the broader subject of comfort in air travel. I hope that I may have the opportunity to supplement this lecture by a more elaborate paper at a later time.

It is twenty years ago almost to the day since the world was saddened by the news of the death of one of its heroes—Wilbur Wright. The inspiring story of the successful attack made by the two brothers, Wilbur and Orville, on the age-long problem of human flight is too well known to be told at length here. But it is remarkable that the first flight by a motor driven aircraft on 17th December, 1903, was made only two years after Chanute, in a speech before the Western Society of Engineers, had felt compelled to use such cautious words as: “There is some hope that, for some limited purposes at least, man will eventually be able to fly through the air.”

This paper looks at how the aerospace industry can achieve the Acare goal of greener manufacturing, maintenance and disposal. It looks further than merely reducing waste and eliminating hazardous materials and processes and suggests that the organisational structure of the industry will play an important role in facilitating a move towards such a goal. Greater co-operation or integration within the industry at all stages of the product life cycle chain is a fundamental requirement as individual companies run a risk of increasing the total environmental burdens if they concentrate solely on reducing their own impacts without considering the effect a change they make may have on other companies. The use of comprehensive environmental supply chain management systems and end of life plans can smooth the implementation of extended product responsibility and accelerate the benefits of greener manufacturing, maintenance and disposal.

Computational Fluid Dynamics (CFD) has become an attractive method of choice in the design of many aerospace vehicles because of advances in numerical algorithms and convergence acceleration methods. However, the flow around an advanced fighter aircraft is complicated and usually unsteady due to the presence of vortex-dominated flows. The accuracy and predictability of conventional turbulence models for these applications may be questionable and therefore results obtained from these models must be validated and evaluated on the basis of experimental data from wind tunnels and/or flight tests. This work aims to validate CFD simulations of X-31 wind-tunnel models with and without a belly-mounted sting. The sting setup facilitates forced sinusoidal oscillations in one of three modes of: pitch, yaw, and roll. However, the results show that measured aerodynamic data are altered by the turbulent wake behind the sting, even at small angles of attack. The high angle-of-attack flow around the X-31 is also very complicated and unsteady due to canard and wing vortices. Therefore, validation of CFD models for predicting these complex flows can be a very challenging task. The X-31 wind-tunnel experiments were carried out in the German Dutch low-speed wind tunnel at Braunschweig and include aerodynamic force and moment measurement as well as span-wise pressure distributions at locations of 60% and 70% chord length. This data set is used to validate the Cobalt and Kestrel flow solvers and the results are similar and match quiet well with experiments for small to moderate angles of attack. The main discrepancies between CFD and measurements occur close to the wing tip, where leading-edge flaps are located.

At a time when a great deal of research on UCAVs is aimed at maximising their autonomy it should not be forgotten that human operators will, ultimately, remain ‘in control’ to some degree. The decision-sharing relationship between the operator and the UCAV depends on political constraints as well as the intelligence of the UCAV system. This in turn dictates the amount and type of information to be exchanged and the way in which it is communicated with the operator.

Operational flexibility is a key military driver and in order to achieve it, a variable autonomy command interface, combined with information fusion and intelligent decision-support systems, will be required. To be effective the operator will need to work in concert with the UCAV system rather than act simply as a command source and an information sink. The implications of this ‘partnership’ for command and monitoring requirements and in particular for weapons release authorisation, are discussed in this paper.

The exact solution to the stability of a simply supported rectangular plate loaded along two opposite edges by several equidistant point loads is presented. Criteria are derived which, if met, permit the true stress state throughout the plate to be radically simplified.

The use of active trailing edge flaps on rotors may lead to performance benefits as well as noise and vibration reduction. In this work, computational fluid dynamics, using the HMB2 solver, is used to assess the effect of the trailing edge flaps on the whole flight domain of a modern main rotor. Starting from a baseline blade design, multiple techniques are demonstrated. The flap is first assessed using 2D pitching aerofoil simulations, followed by dMdt simulations, that account for the simultaneous variations of pitch and Mach around the azimuth. It was shown that enhanced lift was obtained while inspection of the moment coefficient showed negative damping for the flap for a limited set of conditions. Due to the 2D formulation, dMdt computations are fast to perform and can be used to inform codes predicting the rotor performance. The flap was then assessed in hover, and only allowed for limited improvement in blade performance at high thrust. In forward flight, the flap was actuated at a frequency of 1 per revolution, and was found to have a strong effect on the loads on the retreating side. The effect on the moments was even stronger. The flight envelope of the blade was explored, and clean and flapped cases were compared. The most noticeable changes occur at high and medium thrust. The CFD method was found to be efficient and robust, without any substantial penalties in CPU time, due to the flap modelling, over the tested conditions.

Advanced design offices have traditionally applied conceptual design techniques based on semi-empirical methods in an attempt to develop an accurate prediction of aircraft designs at the end of the development process. Continuing advances in computer capability and rotorcraft analysis software present an opportunity to re-think conceptual design to include the greater use of physics-based analyses. A roadmap for developing this capability is outlined, taking into account techniques and ideas from Model-Based Systems Engineering, Design Thinking and Multidisciplinary Optimisation. Recent activities that demonstrate some of these desired capabilities are briefly described along with lessons learned.

The need for autonomous location estimation in the form of optical navigation is an essential requirement for forthcoming deep space missions. While crater-based navigation might work well with larger bodies littered with craters, small sub-kilometer bodies do not necessarily have them. We have developed a new pose estimation method for absolute navigation based on photometric local feature extraction techniques thus making it suitable for missions that cannot rely on craters. The algorithm can be used by a navigation filter in conjunction with relative pose estimation such as visual odometry for additional robustness and accuracy. To estimate the position and orientation of the spacecraft in the asteroid-fixed coordinate frame, it uses navigation camera images in combination with other readily available information, such as orientation relative to the stars and the current time for an initial estimate of the asteroid rotation state. Evaluation of the algorithm when using different feature extractors is performed, on one hand, using Monte Carlo simulations and, on the other hand, using actual images taken by the Rosetta spacecraft orbiting the comet 67P/Churyumov–Gerasimenko. Our analysis, where four different feature extraction methods (AKAZE, ORB, SIFT, SURF) were compared, showed that AKAZE is most promising in terms of stability and accuracy.

The problem considered is that of a circular plate which is supported along an internal concentric circle and subjected to uniform pressure over the area bounded by the support line; the outer edge of the plate is free. This form of loading and support is used in the testing of glass panels, where the rigidity of the panel is determined from a measurement of the central deflexion. However, the formula in current use does not allow for the stiffening effect of the overhanging outer annulus.

At 14h42 UTC (3.42pm BST), on Tuesday, 25 July 2000, Concorde F-BTSC, operated by Air France, started its take-off roll on runway 26 Right at Paris Charles de Gaulle, to undertake a charter flight to New York with nine crew and 100 passengers on board. Shortly before rotation, the front right tyre of the left main landing gear ran over a strip of metal, which had fallen from another aircraft, and was damaged. Debris was thrown against the bottom wing surface leading to a rupture of tank five. A major fire, fuelled by the massive leak, broke out almost immediately under the left wing. Problems appeared shortly afterwards on engine two and for a brief period on engine one. The aircraft took off. The crew shut down engine two following an engine fire alarm. The landing gear would not retract. The aircraft flew for about one minute at a speed of 200kt and an altitude of 200ft, but was unable to gain height or speed. Engine one then lost thrust, the aircraft’s angle of attack and bank increased sharply. The thrust on engines three and four fell suddenly. The aircraft crashed onto a hotel at Gonnesse. Everyone on board and four people in the hotel were killed.

The mutual interaction between the main rotor and tail rotor wakes is central to some of the most problematic dynamic phenomena experienced by helicopters. Yet achieving the ability to model the growth and propagation of helicopter rotor wakes with sufficient realism to capture the details of this interaction has been a significant challenge to rotorcraft aerodynamicists for many decades. A novel computational fluid dynamics code tailored specifically for rotorcraft applications, the vorticity transport model, has been used to simulate the interaction of the rotors of a helicopter with a single main rotor and tail rotor in both hover and low-speed quartering flight, and with the tail rotor rotating both top-forward and top-aft. The simulations indicate a significant level of unsteadiness in the performance of both main and tail rotors, especially in quartering flight, and a sensitivity to the direction of rotation of the tail rotor. Although the model thus captures behaviour that is similar to that observed in practice, the challenge still remains to integrate the information from high fidelity simulations such as these into routine calculations of the flight dynamics of helicopters.