Analysis of the performance of a 1/4.71 model-scale and full-scale Sikorsky S-76 main rotor in hover is presented using the multi-block computational fluid dynamics (CFD) solver of Glasgow University. For the model-scale blade, three different tip shapes were compared for a range of collective pitch and tip Mach numbers. It was found that the anhedral tip provided the highest Figure of Merit. Rigid and elastic full-scale S-76 rotor blades were investigated using a loosely coupled CFD/Computational Structural Dynamics (CSD) method. Results showed that aeroelastic effects were more significant for high thrust cases. Finally, an acoustic study was performed in the tip-path-plane of both rotors, showing good agreement in the thickness and loading noise with the theory. For the anhedral tip of the model-scale blade, a reduction of 5% of the noise level was predicted. The overall good agreement with the theory and experimental data demonstrated the capability of the present CFD method to predict rotor flows accurately.

The first tandem-wing aircraft I saw appeared in an early 1960s edition of the boys comic—Lion. I dismissed it as a figment of the artist's imagination, but later came across the self-same aircraft in War Planes of World War 2—Fighters—Volume 2 by W. Green. From then on I have been collecting information on tandem-wing aircraft. Some eight years later what started as an interest in aircraft took an academic turn when I did a design study of a tandem-wing aircraft as part of an MSc.

A turbulent boundary layer, subjected to a freestream velocity which changes sinusoidally with time around a spatially uniform mean value has been studied experimentally. The experiments were conducted in a specially designed unsteady-flow water tunnel. A two-component laser Doppler anemometer was used to measure the instantaneous velocities, and hence the turbulent stresses, in the periodic boundary layer. A surface-mounted heat-flux gauge was used to measure the unsteady wall shear stress. The experiments were conducted at relatively high amplitudes of oscillation (≈ 40% of the mean) and over a wide range of reduced frequencies (1.7-28). While the present data on the oscillatory flow properties generally confirm previously known trends the information on the turbulent shear stress and skin friction obtained in the present studies is more complete than previously available. The experimental data, all of which has been archived on magnetic tape, form a comprehensive base for the development and verification of predictive models for unsteady turbulent boundary layers.

The Mamba engine has now been in existence some two–and–a–half years and in that comparatively short period of time it has been developed to the stage of being a type tested engine with over 5,000 hours test bed running and 100 hours flying as a background. The purpose of this paper is to give a brief account of the progress of the development and of the many problems which have been met and overcome during the period. Because of the many aspects of the development work it is not possible to deal in great detail with any particular item but rather to present the over–all picture, from the original design conception to the complete power plant of to–day.

In the present investigation, the results obtained during the ground vibration tests of a closed-loop control system conducted on a full-scale fighter to attenuate vertical fin buffeting response using strain actuation are presented. The experimental results are supported by numerical analyses using a finite element aeroelastic model of the structure. Two groups of actuators consisting of piezoelectric elements distributed over the structure were designed to achieve authority over the first and second modes of the vertical fin. The control laws were synthesised using the linear quadratic Gaussian (LQG) method for a time-invariant two-input, two-output (2x2 MIMO) control system. Three different pairs of sensors including strain gauges and accelerometers at different locations were used to close the feedback loop. The results demonstrated that actual reductions of up to 18% in the root-mean-square (RMS) values of the fin dynamic response measured by the strain transducer at the critical point for fatigue at the root were achieved for the second mode under the most severe buffet condition.

An experimental and numerical study was performed on a Clark Y aerofoil with a 10% chord leading edge Kruger flap to examine its aerodynamic performance at Reynolds numbers of 0·6 × 106, 1 × 106, and 1·6 × 106, to help to identify the forces and moments acting on a basic configuration. A detailed comparison of the numerical and experimental data is presented in this paper. The leading edge flap was effective at high angles of attack with an increase in CL of up to 18% over a conventional no flap configuration and delayed separation by up to 3°. The moments around the Kruger flap rotation point were calculated from the numerical analysis as an initial stage in the design of a UAV passive flap system and they are also presented in the paper.

We develop a simple theory for hot air balloon performance with fixed thermal power and linear heat transfer to the environment, applicable to low-temperature situations such as Titan’s atmosphere. The theory results in a closed-form solution and it is shown that an optimum balloon diameter exists – the maximum payload is achieved when the envelope mass and payload mass are equal. It is also shown simply that the floating mass for a given power has a stronger sensitivity to heat transfer coefficient than to the envelope specific mass. A hot air balloon on Titan with a ~2kW heat source could loft a theoretical maximum payload of ~195kg or ~100kg with appropriate margins.

A numerical investigation of generic open-wheel racing car wing and wheel geometry has been conducted, using original sub-scale experimental data for validation. It was determined that there are three main interactions that may occur, identifiable by the path that the main and secondary wing vortices take around the wheel. Interaction ‘A’ occurs when the main and secondary wing vortices both travel outboard of the wheel; interaction ‘B’ is obtained when only the main wing vortex passes inboard of the wheel; while interaction ‘C’ sees both wing vortices travel inboard of the wheel. The different interactions are achieved when geometric changes to the wing affect the pressure distribution about the endplate, either by altering the magnitude of suction generated by the wing or by changing the locations of peak suction and vortices relative to the wheel’s stagnation regions. As a result, the influence that the wing and wheel have on each other – in comparison to the same bodies in isolation – varies, resulting in significant consequences for downforce and drag.

The unsteady flowfield of a series of mixed-compression hypersonic inlets with different bleeding rates were numerically simulated. Firstly unstart oscillatory flow of hypersonic inlets caused by downstream massflow choking was discussed. Then the effects of boundary layer bleeding on the averaged performance parameter of hypersonic inlets, and on the dominant amplitude and frequency of unstart oscillatory flow of hypersonic inlets were presented. The reasons why the boundary-layer bleeding can suppress unstart oscillatory flow of hypersonic inlets were analysed. In conclusion, the averaged performance parameter of hypersonic inlets during a big buzz is improved greatly, and the dominant frequency of unstart oscillatory flow of hypersonic inlets is reduced in contrast with no bleeding, and all these are benefit to the design and operation of hypersonic inlets.

King Abdul Aziz University of Saudi Arabia (KAU) and Tokai University in Japan have collaborated to design and manufacture a solar powered unmanned air vehicle (UAV), Sun Falcon, which has capability of continuous one day fight with intended design extension towards night flight. The project is a student-driven endeavour involving some 30 students. Both universities are equally involved in the actual design studies of the performance characteristics, aerodynamic design, propulsion and structural analysis. Tokai University is in charge of the actual on-site supervision and examination of on-going manufacturing processes and ultimate fabrication of the prototype model.

The conceptual design of the Sun Falcon was meticulously worked out in consideration of the operational mission, which included such flight characteristics as the cruising velocity, flight altitude, payload, flight time, rate of climb, power requirements and so on. The weather condition patterns in Saudi Arabia, which remained fairly supportive of the solar cell performance, were also deemed crucial in the design process. However, the design of a solar plane had other challenges in terms of power unit accommodation and payload consideration in comparison to other conventional UAVs. In this paper, an outline of the design features of the Sun Falcon is presented and other notable design features particular to solar UAVs are discussed. It was learnt in this exercise that the selection of the base aerofoil is perhaps one of the most important design items, as the Reynolds number for such UAVs understandably drops notably lower than conventional aeroplanes and such features as the camber curvature and wing area must cater for the installation of solar panels whose size, strength and quantity must respect local weather conditions. For the Sun Falcon, the actual design process examined two candidate aerofoils FX74-CL5-140 (FX74) and SD7037-092-88(SD7037) both of which were abundantly suitable for furnishing the required aerodynamic characteristics. SD7037 was ultimately chosen as it provided the best geometry and camber line in terms of accommodation and placement of the solar panels. Further scrutiny demonstrated that this latter aerofoil provided better take-off performance and superior L/D behaviour under cruise conditions. In order to check out the aerodynamic performance in general and overall stability and control characteristics, a preflight test under battery power (a 2,500 mAh li-po 4-cell 14·8v) was achieved on 4 June 2013. Other solar powered based tests are currently under way at present.

I have chosen to describe the methods employed by the Research Division of the United States Navy Bureau of Aeronautics with the hope that others engaged in a similar line of endeavour may be able to make profitable use of this information in the solution of their specific problems. The development of the material to be discussed arose directly from the need for quick, reasonably accurate and simple means for studying a broad series of aircraft designs leading toward fulfilling an overall military requirement. This process is carried out in the Bureau of Aeronautics by the Research Division prior to the issuing of a type specification with a request for proposal to the various manufacturers. Before these specifications are decided upon, many designs must be investigated and reported to those initially responsible for the issuing of military requirements. By this process intelligent compromises are arrived at and the best possible aeroplane chosen. These compromises reflect the performance, the ability of operation from carriers, and the effectiveness of the design as a military weapon.

The established control anticipation parameter longitudinal handling qualities criterion is based on the assumption that the short term dynamic response of the aeroplane is classical, or second-order-like. Modern fly-by-wire aircraft often have longitudinal short term dynamics which are not second-order-like and to which it is difficult to apply the criterion. This paper presents a proposed generic control anticipation parameter which is a modified version of the control anticipation parameter and which may be applicable to both unaugmented and augmented aircraft of all types. The appropriateness of the modified criterion was illustrated by designing a number of command and stability augmentation control laws, constrained to meet the criterion, for application to a medium weight fly-by-wire civil transport aircraft. The handling characteristics conferred by the control laws were assessed in a series of flight simulator trials, the results of which are briefly presented.

In 1931 Peters showed that flow entry conditions had a marked effect on diffuser performance. Later, Winternitz and Ramsay indicated that the pressure recovery coefficient Cp of a conical diffuser might be regarded as a linear function of β(θ/d)i for values of β(θ/d)i up to 0·1. This note draws attention to the fact that diffusers with much larger values of this parameter are of interest and presents some experimental and theoretical results. Work is continuing along the lines indicated.

In section 8 the author's apparatus for the study of skin friction, based on direct application of the momentum theorem, has been described (Fig. 99). In a paper published in 1914, this idea was further developed and it was shown that it is possible to calculate approximately in this way the skin friction and the distribution of air velocities in different régime.

Conventional wisdom states that coordinate complete experimental response data are required to effect a successful model update. This paper examines the influence measured rotational degrees of freedom have on the behaviour of the frequency response function (FRF) sensitivity updating procedure. Three finite element (FE) model-based case studies are undertaken which suggest that very little improvement is afforded by the inclusion of measured rotations in the updating process.

Error-location strategies in the past have yielded little qualitative information about the optimum selection of updating parameters in the model updating procedure. Two further cases are investigated; these suggest that no amount of coordinate completeness will rectify the debilitating effects of an inadequate p-value choice. A method of selecting the number and location of updating parameters — based on the FRF sensitivity method — is presented. This method is shown to yield promising results.

Once the linearised lift problem for a supersonic wing has been solved and the disturbance velocity potential or the pressure is known at every point on the surface, the complete flow round the wing and the downwash behind the wing may be found by the following procedure.

This paper offers a critical perspective on the changing organisational structure of the Western commercial aircraft industry. The role of systems integration based on risk-sharing partnerships for new aircraft programmes is explored. We find that build-to-print subcontracting relationships are being replaced by internationally devolved design and engineering tasks for airframe development, signaling a profound change in the geography of commercial aircraft production. While sensible from a financial standpoint, the international outsourcing of design-intensive production entails substantial amounts of technology transfer–including the delivery of proprietary knowledge to risk-sharing partners. For several of the advanced market economies, including Canada, France, Germany, the UK, and the US, the long-range strategic downside is that foreign risk-sharing partners could eventually become competitors. Systems integration on a risk-sharing basis also implies home-country joblosses among skilled workers with expertise in design, engineering, and R&D.

A simple yet physically comprehensive and accurate method for the estimation of the cruise fuel burn rate of turbofan powered transport aircraft operating in a general atmosphere was developed in part 1. The method is built on previously published work showing that suitable normalisation reduces the governing relations to a set of near-universal curves. However, to apply the method to a specific aircraft, values must be assigned to six independent parameters and the more accurate these values are the more accurate the estimates will be. Unfortunately, some of these parameters rarely appear in the public domain. Consequently, a scheme for their estimation is developed herein using basic aerodynamic theory and data correlations. In addition, the basic method is extended to provide estimates for cruise lift-to-drag ratio, engine thrust and engine overall efficiency. This step requires the introduction of two more independent parameters, increasing the total number from six to eight. An error estimate and sensitivity analysis indicates that, in the aircraft’s normal operating range and using the present results, estimates of fuel burn rate are expected to be in error by no more than 5% in the majority of cases. Initial estimates of the characteristic parameters have been generated for 53 aircraft types and engine combinations and a table is provided.

The use of finite element analysis in the modelling of an unloaded countersunk hole revealed that there is a consistent relationship between the maximum stress concentration factor and the design parameters of the hole — countersunk depth, edge distance, t/d ratio and countersunk angle. It has been shown that countersunk depth and edge distance have a strong influence on the Kt value while the t/d ratio has a lesser effect. Variation of countersunk angles between 90° and 110° was found to have a negligible influence on the maximum Kt values.

It has also been shown that normalising the stress concentration factors of a countersunk hole by dividing the Ktvalues at various countersunk depths by the Ktvalues for a plain hole of similar configuration, enables the stress concentration factors in a countersunk hole to be determined from values in a plain hole. This is very useful since extensive work had been done on stress concentration factors for a plain hole and data are therefore more commonly available.

Aeronautic Pitot probes (PPs) are extremely important for airspeed and altitude measurements in aviation. Failure of the instrument due to clogging caused by ice formation can lead to dangerous situations. In this work, a commercial aeronautic PP was characterised experimentally regarding its inner composition, material properties and its thermal performance in a climatic wind tunnel. Performance runs were taken out in order to analyse the thermal response of the PP under various operating conditions with a particular emphasis on the cooling process in the case of a heating element failure. Data for the thermal conductivity, diffusivity and specific heat for each material forming the PP were obtained. A numerical model to simulate the thermal behaviour of the PP was created using Comsol Multiphysics (CM). Experimental data were compared with their numerical counterparts for model validation purposes. After the model was validated, the operation of the PP in flight conditions was simulated. The failure of the conventional heating system was analysed to obtain the time until the PP reaches a tip temperature where ice formation can be expected. The tip temperature undercut the zero degrees Celsius mark 165 seconds after the heating element was switched off. The data collected in this work can be used to implement and validate mathematical models in order to predict the thermal performance of Pitot probes in flight conditions.