Airframe–propulsion integration concepts that use boundary-layer ingestion (BLI) have the potential to reduce aircraft fuel burn. One concept that has been recently explored is NASA’s STARC-ABL aircraft configuration, which offers the potential for fuel burn reduction by using a turboelectric propulsion system with an aft-mounted electrically driven BLI propulsor. So far, attempts to quantify this potential fuel burn reduction have not considered the full coupling between the aerodynamic and propulsive performance. To address the need for a more careful quantification of the aeropropulsive benefit of the STARC-ABL concept, we run a series of design optimisations based on a fully coupled aeropropulsive model. A 1D thermodynamic cycle analysis is coupled to a Reynolds-averaged Navier–Stokes simulation to model the aft propulsor at a cruise condition and the effects variation in propulsor design on overall performance. A series of design optimisation studies are performed to minimise the required cruise power, assuming different relative sizes of the BLI propulsor. The design variables consist of the fan pressure ratio, static pressure at the fan face, and 311 variables that control the shape of both the nacelle and the fuselage. The power required by the BLI propulsor is compared with a podded configuration. The results show that the BLI configuration offers 6–9% reduction in required power at cruise, depending on assumptions made about the efficiency of power transmission system between the under-wing engines and the aft propulsor. Additionally, the results indicate that the power transmission efficiency directly affects the relative size of the under-wing engines and the aft propulsor. This design optimisation, based on computational fluid dynamics, is shown to be essential to evaluate current BLI concepts and provides a powerful tool for the design of future concepts.

The Navier-Stokes equations are solved iteratively on a small digital computer for the class of flows generated within a rectangular “cavity” by a surface passing over its open end. Solutions are presented for depth/breadth ratios ƛ=0.5 (shallow), 10 (square), 20 (deep) and Reynolds number 100. Flow photographs ore obtained which largely confirm the predicted flows. The theoretical velocity profiles and pressure distributions through the centre of the vortex in the square cavity are calculated.

In an appendix an improved finite difference formula is given for the vorticity generated at a moving boundary.

Since Thorn began his pioneering work some thirty-five years ago the number of numerical solutions which have been obtained for the equations of incompressible viscous fluid motion remains small (see bibliographies of Thom and Apelt, Fromm). The known solutions are principally for steady streaming flows, although two methods have now been used with success for non-steady flows (Payne jets and Fromm flow past obstacles). By contrast this paper is concerned with the class of closed flows generated in a rectangular region of varying depth/breadth ratio by a surface passing over an open end. This problem has been considered for a number of reasons.

This is the fifth occasion on which I have had the honour to address our Society and the third time I have chosen to discuss the same subject: “The stalling of aeroplanes.” I should apologise for thus harping upon a single string were it not that no subject could be more appropriate to an address in memory of Wilbur Wright, and were it not that until the stall is thoroughly understood and its attendant dangers pushed out of the way, aviation will not be free to play the great rôle which it is undoubtedly destined to play in the civilisation of the world.

Wind tunnels are far older than powered flight. Including the aeroplane of the Wright Brothers, the introduction of new types of aircraft has almost invariably been preceded by extensive model investigation in wind tunnels. These investigations can be classified into two broad areas. One is development testing, in which models are tested and modified with the purpose of producing a specific aircraft configuration. The other covers the research and general experimentation which is necessary for the determination of the new rules governing entry into a new régime of flight. Thus, wind tunnels may conveniently be referred to either as “development” or as “research” wind tunnels, according to the type of work which is done in them.

Detailed measurements in the wake flow behind blunt based two-dimensional and three-dimensional bodies have been made at M = 2. The results throw some light on the behaviour of separated flows and indicate the importance of flow reversal. The results for the two-dimensional body are in agreement with the Crocco and Lees’ theory on separated shear flows. A preliminary study of the effect of air injection at the base shows that the base pressures increase significantly.

Although columns built almost 5,000 years ago still stand in Egypt, the first theoretical calculation of the load-carrying capacity of columns was published by Euler only 209 years ago. The classical theory established in this publication is still the foundation of all stability analyses and is likely to remain so in the future. It was derived for perfectly elastic columns and for specific end conditions. These end conditions are not realised in the ordinary column test nor do they correspond exactly to the conditions prevailing in an aeroplane wing when it is subjected to a gust or to the shock of landing. They are not the same as the end conditions of a column in a building when it collapses during a tornado or an earthquake.

Even at small angles of incidence, the flow separates from the sharp leading edges of a slender wing. These flow separations usually lead to the formation of two free vortex layers, joined to the leading edges of the wing and rolling up to form spiral-shaped vortex sheets above the upper surface of the wing. This vortex formation is illustrated schematically in Fig. 1. The streamlines on the vortex sheet follow helical paths. Smoke injected near the wing apex for flow visualisation remains concentrated close to the axis of the core of the vortex sheet.

Drag measurements were made on a blunt-based body of revolution with the addition of several alternative boat-tailed afterbodies. The best of these afterbodies, with a length of only 60 per cent of the maximum diameter, gave almost as much reduction of drag as a conventional streamline tail. The pressure distribution on this body was measured for various lengths. It does not seem to be possible at present to predict quantitatively the drag of this type of body but some qualitative design principles have been suggested.

Experiments to determine the overall buckling and faceplate wrinkling loads of sandwich panels with carbon fibre reinforced plastic faces and honeycomb cores when loaded under uniaxial compression are described. The results are compared with theoretical predictions which assume that the faceplates are orthotropic. The effect on the experimental loads of certain coupling terms, which occur in the constitutive equations of those faceplates which are not orthotropic, is then considered.

The equations of motion for an isolated laminar viscous vortex at moderate to large Reynolds numbers are linearised, by assuming that both the rotational velocity and the deficit of longitudinal velocity are small compared with that in the free stream. The rotational motion and the longitudinal motion may then be superimposed and solutions are readily obtained for each. If the vortex is generated by a body with profile drag it is predicted that the deficit of longitudinal velocity will be positive, which is in agreement with experimental observation. Further details of the solution and its relation to the flow in real vortices are discussed; and the theory is compared with some measurements in a turbulent vortex.

The purpose of the research presented here is to exploit actuation via smart materials to perform shape control of an aerofoil on a small aircraft and to determine the feasibility and advantages of smooth control surface deformations. A type of piezoceramic composite actuator known as Macro-Fiber Composite (MFC) is used for changing the camber of the wings. The MFC actuators were implemented on a 30° swept wing, 0·76m wingspan aircraft. The experimental vehicle was flown using two MFC patches in an elevator/aileron (elevon) configuration. Preliminary flight and wind-tunnel testing has demonstrated the stability and control of the concept. Flight tests were performed to quantify roll control using the MFC actuators. Lift and drag coefficients along with pitch and roll moment coefficients were measured in a low-speed, open-section wind tunnel. A vortex-lattice analysis complemented the database of aerodynamic derivatives used to analyse control response. The research, for the first time, successfully demonstrated that piezoceramic devices requiring high voltages can be effectively employed in small air vehicles without compromising the weight of the overall system.

This paper is concerned with two-dimensional, incompressible fluid motion generated within a square cavity (a) by an outer stream and (b) by the action of a flat surface passing over one of its sides. This type of motion (“cavity flow”) is considered to consist of a boundary layer surrounding an inviscid “core.” A solution of a linearised form of von Mises’ equation is obtained for steady flow in the boundary layer for constant pressure and that of a small periodic variation around the walls of the cavity. From this analysis the vorticity imparted to the core is obtained. The motion in the core is then determined on the basis of the persistence of this value of the vorticity. Experimental results are given for (a) and (b) and compared with the theory. A method of solution of the non-linear boundary layer problem is indicated.

The knowledge of the flow through a rotor of a rotary-wing aircraft is essential for the study of many helicopter problems The behaviour of a helicopter with respect to stability, region of roughness, etc, depends in many cases on the airflow through the disc Most existing stability or performance theories fail when they concern low forward speed and vertical descent because of some simplifying assumptions

An attempt has been made to find a rather simple expression to describe the variations in axial velocity through a rotor as a function of the azimuth angle

Further, a chart is produced giving the average axial flow through the disc for all possible combinations of forward and vertical velocities, valid for all kinds of one-rotor aircraft

The results are used to improve the theory of some problems as mentioned above and are applied, where possible, to the Sikorsky S51 helicopter, to compare the theory with experimental data

The effects of slanting the base of a slender axisymmetric cylinder (length/diameter ratio of 9), aligned with the flow, was studied experimentally. The body was equipped with interchangeable rear ends covering a range of slant angles between 0° (vertical) and 70°. It was found that the base slant has a very dramatic effect on body drag, particularly in a relatively narrow range of slant angles where the drag coefficient exhibits a large local maximum (over-shoot). Detailed study of the flow showed that the drag overshoot is related to the existence of two very different Separation patterns on the slanted base. One pattern is similar to that found behind axisymmetric bodies with no base slant, and its main feature is the presence of a closed Separation region adjacent to the base. The other pattern is highly three-dimensional with two streamwise vortices forming along the sides of the slanted base. This pattern sets in very abruptly at a “critical” slant angle α ∼ 47°. Drag force measurements showed that, at first, the drag coefficient slowly increases with the slant angle, but then jumps suddenly upwards to more than double its baseline value (from CD = 0.24 to CD = 0.625) at the critical angle. At angles higher than that CD decreases again, and at 70° it is about equal to the baseline value. Further effects of the slant angle are the generation of a large side force and a significant increase in near-wake flow periodicity.

Competitive pressures and economic constraints are driving aircraft manufacturers towards an ever-increasing exploitation of CFD for design, optimisation and prediction of off-design conditions. Such exploitation is favoured by rapid advances in meshing technology, numerical algorithms, visualisation tools and computer hardware. In contrast, the predictive capabilities of mathematical models of turbulence are limited — indeed, are often poor in regions of complex strain — and improve only slowly. The intuitive nature of turbulence modelling, its strong reliance on calibration and validation and the extreme sensitivity of model performance to seemingly minor variations in modelling details and flow conditions all conspire to make turbulence modelling an especially challenging component of CFD, but one that is crucially important for the correct prediction of complex flows.

The present paper describes the design of a hybrid actuation control concept, a fuzzy logic proportional-integral-derivative plus a conventional on-off controller, for a new morphing mechanism using smart materials as actuators, which were made from shape memory alloys (SMA). The research work described here was developed for the open loop phase of a morphing wing system, whose primary goal was to reduce the wing drag by delaying the transition (from laminar to fully turbulent flows) position toward the wing trailing edge. The designed controller drives the actuation system equipped with SMA actuators to modify the flexible upper wing skin surface. The designed controller was also included, as an internal loop, in the closed loop architecture of the morphing wing system, based on the pressure information received from the flexible skin mounted pressure sensors and on the estimation of the transition location.

The controller’s purposes were established following a comprehensive presentation of the morphing wing system architecture and requirements. The strong nonlinearities of the SMA actuators’ characteristics and the system requirements led to the choice of a hybrid controller architecture as a combination of a bi-positional on-off controller and a fuzzy logic controller (FLC). In the chosen architecture, the controller would behave as a switch between the SMA cooling and heating phases, situations where the output current is 0A or is controlled by the FLC.

In the design phase, a proportional-integral-derivative scheme was chosen for the FLC. The input-output mapping of the fuzzy model was designed, taking account of the system’s error and its change in error, and a final architecture for the hybrid controller was obtained. The shapes chosen for the inputs’ membership functions were s -function, π-function, and z -function, and product fuzzy inference and the center average defuzzifier were applied (Sugeno).

An inviscid model, originally proposed by Bryson to describe the formation of laterally-symmetric vortices above bodies of revolution at large angles of incidence, is extended to describe flows without lateral symmetry. Each vortex is represented by a single line-vortex and conditions of cross-flow stagnation are applied at the postulated separation lines. The transverse component of the force acting on each vortex is balanced by an equal and opposite force acting on the cut connecting it to the adjacent separation line. It is shown that this model allows asymmetrical vortices to form in the flow over slender circular cones even when the separation lines are disposed symmetrically.

Over the subsonic speed range, the pressure field attached to a ducted rotor decays rapidly with distance away from the rotor. This exponential decay, coupled with the low radiation efficiency of a subsonically rotating pressure pattern, is generally sufficient to bring the radiated sound power well below that produced by interaction of the rotor with disturbances in the flow. This is particularly true for multibladed rotors as used in aircraft-engine fans and compressors, and accounts for the importance of rotor-stator interaction as a source of noise at subsonic tip speeds.

It was while lecturing at Cranfield in the late 1970s and early 80s on the aerodynamic design of aircraft wings that the ideas discussed in this paper began to develop. My work on wing tip sails had led me to look in some detail at the flight of birds and their ability to vary wing camber. I realised that this was an attractive way to overcome the need to compromise between the design requirements associated with the different regimes of flight necessary with fixed geometry.

To be economically attractive, civil aircraft have to be capable of operating over a wide range of flight conditions. For example, an executive jet may be required to fly fast and low at light weight over a short stage length, such as London to Paris, on one day, and fly heavily loaded a much longer stage length the next, when it is desirable for the aircraft to fly high and at an economical speed.

This paper presents results from a joint Lockheed Martin/NASA Glenn effort to design and verify an ultra-compact, highly-survivable engine inlet subsonic duct based on the emerging technology of active inlet flow control (AIFC). In the AIFC concept, micro-scale actuation (∼mm in size) is used in an approach denoted ‘secondary flow control’ to intelligently alter a serpentine duct's inherent secondary flow characteristics with the goal of simultaneously improving the critical system-level performance metrics of total pressure recovery, spatial distortion, and RMS turbulence. In this approach, separation control is a secondary benefit, not a design requirement. The baseline concept for this study was a 4:1 aspect ratio ultra-compact (LID= 2·5) serpentine duct that fully obscured line-of-sight view of the engine face. At relevant flow conditions, this type of duct exhibits excessive pressure loss and distortion because of extreme wall curvature. Two sets of flow control effectors were designed with the intent of establishing high performance levels to the baseline duct. The first set used two arrays of 36 co-rotating microvane vortex generators (VGs); the second set used two arrays of 36 micro air-jet (microjet) VGs, which were designed to produce the same ‘vorticity signature’ as the microvanes. Optimisation of the microvane array was accomplished using a design of experiments (DOE) methodology to guide selection of parameters used in multiple Computational Fluid Dynamics (CFD) flow solutions. A verification test conducted in the NASA Glenn W1B test facility indicated low pressure recovery and high distortion for the baseline duct without flow control. With microvane flow control, at a throat Mach number of 0·60, pressure recovery was increased 5%, and both spatial distortion and turbulence were decreased approximately 50%. Microjet effectors also provided significantly improved performance over the baseline configuration.