The mixing of single or multiple jets is relevant in several engineering applications ranging from flame stabilisation in combustion to high lift devices for V/STOL aircraft. In some of these applications, expeditious manufacturing and ease of installation may dictate the use of sharp-edged slots in preference to nozzles with contoured upstream shaping.

It has been suggested that cruciform nozzle exit shapes may be suitable for improving mixing and thus thrust augmentation in STOL aircraft. The effects of nozzle geometry on the performance of an axisymmetric thrust augmentor have been investigated experimentally and a very slight improvement in the performance of the axisymmetric thrust augmentor was observed when cruciform nozzles were used after the tests had been performed with circular nozzles.

With an annual growth in travel demand of about 5% globally, managing the environmental impact is a challenge. In 2019, the International Civil Aviation Organisation (ICAO) issued emission reduction targets, including well-to-wake greenhouse gas (GHG) emissions reduced at least 50% from 2005 levels by 2050. This discusses several technologies from an aircraft design perspective that can contribute to achieving these targets. One thing is certain: aircraft will look different in the future. The Transonic Truss-Braced Wing and Flying V configurations are promising significant efficiency improvements over conventional configurations. Electric propulsion, in various architectures, is becoming a feasible option for general aviation and commuter aircraft. It will be a growing field of aviation with zero-emissions flight and opportunities for special missions. Lastly, this paper discusses methods and design processes that include all relevant disciplines to ensure that the aircraft is optimised as a complete system. While empirical methods are essential for initial design, Multidisciplinary Design Optimisation (MDO) incorporates models and simulations integrated in an optimisation environment to capture critical trade-offs. Concurrent design places domain experts in one site to facilitate collaboration, interaction, and joint decision-making, and to ensure all disciplines are equally considered. It is supported by a Collaborative Design Facility (CDF), an information technology facility with connected hardware and software tools for design analysis.

The drag of slant-based three-dimensional bodies such as the upswept rear fuselage of some aircraft (see for instance McCluney and Marshall and Peake) and car shapes (Morel) can be high and it is generally recognised that this high drag is associated with powerful longitudinal vortices in the flow direction. Morel has shown that the formation of these vortices, and therefore the drag, is a function of the slant angle of the base. His experiments were conducted with the body at zero incidence only and it appears that the effect of body incidence on the drag is unrecorded, although Bearman mentions that it has a significant effect.

This note reports tests in a shock tunnel in which a fully integrated scramjet configuration produced net thrust. The experiments not only showed that impulse facilities can be used for assessing thrust performance, but also were a demonstration of the application of a new technique(1) to the measurement of thrust on scramjet configurations in shock tunnels. These two developments are of significance because scramjets are expected to operate at speeds well in excess of 2 km/s, and shock tunnels offer a means of generating high Mach number flows at such speeds.

A theory is presented on the effect of wake turbulence of a leading aircraft on the roll stability of a following aircraft, leading to a simple formula for the safe separation distance between the two aircraft that provides estimates of aircraft separation distances comparable to existing empirical regulations, based on experience. The formula includes the effects of flight and atmospheric conditions, and the characteristics of the leading and following aircraft; it applies to similar or dissimilar aircraft, and it indicates the parameters and conditions leading to increasing or decreasing separation. The formula is applied not only to the three International Civil Aviation Organization (ICAO) categories of aircraft (light, medium and heavy, respectively, Cessna Citation, B737 and B747) but also to ‘special’ aircraft requiring larger separation distance (Boeing 757) and to the world’s largest airliner (Airbus A380).

In tandem with the fast-growing demand for Unmanned Aerial Vehicles (UAVs) for surveillance and reconnaissance, advanced controllers for these critical systems are needed. This paper proposes a flight dynamics controller design that considers various uncertainties for the Hydra Technologies UAS-S4 Ehécatl. In order to be realistic, in addition to flight dynamics nonlinearities, three main sources of uncertainties are considered, as those caused by unknown controller’s parameters, modeling errors, and external disturbances. A Robust adaptive fuzzy logic controller is designed, in charge of nonlinear flight dynamics in presence of a variety of uncertainties. The nonlinear flight dynamics is modeled based on the Takagi-Sugeno method relying on the soft association of local linear models. Since this controller is model-based, an optimal reference model is defined, which is stabilised by the Linear Quadratic Regulator procedure. A fuzzy logic controller is then designed for the nonlinear model. Lastly, with the aim to handle the uncertainties, the gains of the fuzzy controller are reconfigured, and are continuously adjusted by Lyapunov-based robust adaptive laws. The performance of the UAS-S4 Robust adaptive fuzzy logic controller is evaluated in terms of lateral and longitudinal flight dynamics stabilisation, and the reference model state variables tracking under various uncertainties.

How long would it take to evacuate a blended wing body (BWB) aircraft with around 1,000 passengers and crew? How long would it take an external post-crash fire to develop non-survivable conditions within the cabin of a BWB aircraft? Is it possible for all the passengers to safely evacuate from a BWB cabin subjected to a post-crash fire? These questions are explored in this paper through computer simulation. As part of project NACRE, the airEXODUS evacuation model was used to explore evacuation issues associated with BWB aircraft and to investigate fire issues, the CFD fire simulation software SMARTFIRE was used. The fire and evacuation simulations were then coupled to investigate how the evacuation would proceed under the conditions produced by a post-crash fire. In conjunction with this work, a large-scale evacuation experiment was conducted in February 2008 to verify evacuation model predictions. This paper presents some of the results produced from this analysis.

Experimental measurements of the separated flows around a range of tapered flat plates normal to a uniform stream are presented. Both the surface pressures and velocities in the wake are discussed. It turns out that even slight taper can modify the wake region of the flow significantly, and so produce a surprising degree of three-dimensionality into the overall flow structure. These experimental results point to the dangers in applying simple strip theories to highly separated flows.

Aerodynamic theory has evolved over the years to describe the flow fields around streamlined aircraft and their components, so as to provide, amongst other quantities, accurate estimates of the aerodynamic forces and moments.

A joint meeting of the Royal Aeronautical Society and the Royal Meteorological Society was held on 14th December 1950 at the Institution of Civil Engineers, Great George Street, London, S.W.I, at which the 814th Lecture to be delivered before the Royal Aeronautical Society was given by Dr. G. S. Hislop. Major G. P. Bulman, C.B.E., F.R.Ae.S., presided at the meeting and after welcoming the members of the Royal Meteorological Society and especially their Secretary, Professor Sheppard, and expressing his regret at the absence in America of the President of the Royal Meteorological Society, Sir Robert Watson Watt, he introduced the Lecturer, G. S. Hislop, Ph.D., B.Sc, A.R.T.C, A.F.R.Ae.S., M.I.Mech.E., Senior Assistant to the Controller of Research and Special Development, British European Airways Corporation.

This paper deals with the design and implementation of the nonlinear control algorithm for the attitude tracking of a four-rotor helicopter known as quadrotor. The choice of this algorithm, which is based on the Backstepping sliding mode technique with adaptive gain, was justified by the fact that it ensures robustness with respect to modelling errors and external disturbances while reducing the chattering phenomenon caused by the sign function in the first order sliding mode based controllers with fixed gains. In order to show the effectiveness of the controller, experimental tests were carried out on a quadrotor. The obtained results show good performance of the proposed controller in terms of stabilisation, tracking and robustness.

A flight simulator study has been carried out to evaluate the performance of modern flight control systems encountering curved wake vortices. During the decay process the shape of wake vortices alters significantly which has an influence on the encounter characteristics and thus on the encounter hazard. To analyse most realistic wake encounters, flow fields of matured vortices have been generated with large-eddy-simulations. These were used for the determination of induced force and moment histories during the encounter. The force and moment histories have been implemented into the equations of motions of a 6 DoF flight simulation. For the sake of comparison encounters have also been simulated with straight vortices as they were mostly used for many other investigations. The most important goal of the study was to analyse the difference between these ideally straight vortices and vortices with a more realistic shape regarding encounter acceptance. The simulator study was conducted in an A330 motion-based full-flight simulator with pilots-in-the-loop. The analysed scenario was a wake encounter during final approach. The encounter conditions corresponded to a heavy-behind-heavy situation for Instrument Flight Rules (IFR) operations. The aircraft was flown either manually (in normal law) or with autopilot engaged. Altogether 93 encounters have been simulated, 38 with straight and 55 encounters with curved vortices. For encounters under manual control the simulator study revealed a potential risk of pilot induced oscillations (PIO) during encounters with curved vortices. With autopilot engaged not even one encounter with curved vortices was classified to be unacceptable. Although significant aircraft response was experienced the autopilot never disengaged automatically in any encounter. Altogether about 12 percent of the encounters were not accepted by the pilots. This is indeed a significant number, especially as the analysed scenarios can be regarded to be realistic situations which can occur in reality even if the applicable separation distances are applied.

One hundred and two years ago, after the Wright brothers had just attempted another unsuccessful flight, they predicted that it would be another 50 years before manned flight was achieved. Only two years later and 100 years ago this year, Orville Wright achieved the first powered flight at Kitty Hawk in North Carolina, illustrating just how difficult it is to predict the future and at the same time launching the pioneering age of aviation.

In recent years a number of authors have analysed the bending of clamped skew plates under uniform pressure. Many of these solutions are very approximate. Favre and Dorman adopted the Ritz method and their expressions for displacement are restrictive. Based on the generalised method postulated by Lardy, Mirsky has presented a solution which involves considerable algebraic and arithmetic work. Komatsu's paper contains a solution using conformal mapping. Quinlan has given a powerful method to solve skew plate problems involving different boundary conditions and various types of loadings, but no numerical work has been done.

The calculation of the inviscid, incompressible flow past non-lifting three-dimensional bodies has been greatly facilitated by A. M. O. Smith and his collaborators, using distributions of elementary singularities over the surface of the body. There is considerable current interest in the extension of this work to lifting bodies, involving the representation of the real viscous wake by a vortex sheet of zero thickness in the inviscid model of the flow. In view of this interest, it is appropriate to consider the way in which such a vortex sheet behaves at the trailing edge of a wing with non-zero trailing-edge angle. Is the sheet tangential to the upper surface of the wing, to the lower surface, or to neither?

A study was conducted on a GA(W)-1 wing in order to investigate the effect of testing inverted wings in ground effect at low Reynolds numbers. The wing was tested at a range of ground clearances and Reynolds numbers and results showed that the wing’s performance was dependent on both these parameters. Surface flow-visualisation and numerical simulation results highlighted the existence of a laminar separation bubble on the wing’s suction surface. The results also indicated that both the bubble’s length and the onset of separation were sensitive to ground clearance and Reynolds number. Attempts were made to minimise the wing’s Reynolds number dependency by using transition strips on the suction surface. The transition strip results highlighted the influence that a laminar separation bubble has on the overall performance of the wing and how its presence alters the force enhancement and reduction mechanisms on an inverted wing in ground effect.

The non-linear theory of supersonic bangs, obtained in Part 1 for a body accelerating along a straight path, is extended to include curved paths. The basic theory remains the same. The important parameter, which appears in the theory, is the acceleration component along the ray, the rays being lines drawn from points on the flight path at an angle cos-1 (1/M) with the direction of motion. It is found that the only essential effect of the curvature of the path is in the modification of this acceleration component to include a term due to the transverse acceleration. With this modification the main results are formally the same as in Part I.

The strength of the bow shock is obtained, and it is found that the effect of the curvature of the path is more pronounced at points on the inside of the curve, and in general it becomes greater as the distance from the body increases. A simple asymptotic formula is obtained which predicts the strength of the shock with an error of less than five per cent, at distances of the order of a hundred body-lengths. Finally, the theory is compared and contrasted with the recent work by Warren.

The “reference enthalpy” and local flat plate concepts are used to derive some simple expressions for the turbulent heat transfer rate distribution over an arbitrary body at supersonic and hypersonic speeds. The connection between pressure and heat transfer rate is established and tested against a number of experimental data.

The method of parameter differentiation is applied to the solution of the non-linear, two-point, ordinary differential equations resulting from an analysis of the finite bending of a thin-walled tube. Starting from a given set of solutions of the differential equations for a particular value of the curvature parameter α, solutions for a range of values of α can be obtained by this method non-iteratively. Very close agreement was obtained with solutions using iterative numerical methods.

Exact solutions for the steady circulatory flow about a circular cylinder with suction applied at its surface have been obtained for the following cases: —

• (a) the cylinder at rest and a circulation at infinity,

• (b) the cylinder rotating and zero circulation at infinity,

• (c) the cylinder rotating in either direction with a circulation at infinity,

provided that, in general, the suction velocity is greater than a certain limiting value.

There is a particular case of (c) when the circulation at the cylinder equals that at infinity in magnitude and direction, for which the circulation is constant at all radii, for all values of R—presumably for blowing as well as suction.

The addition of a uniform stream parallel to the cylinder axis does not affect the circulatory motion, and the solution for the combined flow, which is exact, has some practical interest in that it demonstrates that the boundary layer on a long, circular, rotating boss can be kept thin and that the circulatory motion outside the boundary layer can be prevented by the suitable application of suction before starting the rotation.

The flow (a) is important in connection with the withdrawal of the Thwaites flap. The existence of steady state solutions, coupled with the arguments of Section 5.2, make it appear highly probable that a wing employing this device will retain its lift indefinitely, when the flap is withdrawn.

For high rates of suction flow, the flows exhibit the characteristics of boundary layer flows, in that the vorticity is confined to a narrow annulus enveloping the cylinder and the velocity distribution near the wall tends to the well-known asymptotic distribution of Griffith and Meredith.

It would be interesting to investigate these flows experimentally and to examine the stability of the laminar flow under suction conditions.

The analysis throughout is simple and expressions are obtained for the velocity distribution, the stresses, the vorticity and the torque, and so on.

The Unmanned Aerial Vehicles (UAVs) become more and more popular due to various potential application fields. This paper studies the distributed leader-follower formation flight control problem of multiple UAVs with uncertain parameters for both the leader and followers. This problem has not been addressed in the literature. Most of the existing literature considers the leader-follower formation control strategy with parametric uncertainty for the followers. However, they do not take the leader parametric uncertainty into account. Meanwhile, the distributed control strategy depends on less information interactions and is more likely to avoid information conflict. The dynamic model of the UAVs is established based on the aerodynamic parameters. The establishment of the topology structure between a collection of UAVs is based on the algebraic graph theory. To handle the parametric uncertainty of the UAVs dynamics, a multivariable model reference adaptive control (MRAC) method is addressed to design the control law, which enables follower UAVs to track the leader UAV. The stability of the formation flight control system is proved by the Lyapunov theory. Simulation results show that the proposed distributed adaptive leader-following formation flight control system has stronger robustness and adaptivity than the fixed control system, as well as the existing adaptive control system.