This paper presents a methodology for studying the feasibility of stratosphere airship for high altitude long endurance missions and arriving at the baseline specifications of conventional configuration of stratosphere airship, given the performance and operational requirements. Based on this methodology, the AODAP platform (Airship Optimisation Design and Analysis Platform) was developed. Some innovative concepts used in AODAP that are different from previous methods and codes are presented. The shape optimisation of airship was introduced into the design process, and several optimum objectives can be selected including minimum drag, minimum weight and composite objective based on MDO (Multidisciplinary Design Optimisation). The methodology was validated for other design concepts previously developed for similar missions and also was compared to a low altitude vehicle. The baseline specifications of stratosphere airships designed for various shapes using this methodology are presented. The results of sensitivity analyses for a specified airship are discussed, and the sensitivity of airship length with some critical parameters including area density of envelope fabric, area density of solar cell, efficiency of solar cell and efficiency of fuel cell for the specified shape is also provided.

In this note, the experimental results of wind-tunnel measurements of lift coefficient and lift-curve slope for aerofoil NACA 0012 obtained from the VTI-Institute Zarkovo, are presented. The results were obtained from tests and integrations of surface static-pressure data over a model of the NACA 0012 aerofoil section. The data were obtained for a free-stream Mach number range of 0·25-0·8 and a chord Reynolds number range of 2-25MRe. The essential results of these measurements along with the results from other authors are presented and evaluated. The principal factors which influence the accuracy of two-dimensional wind-tunnel test results are analysed. The influences of Reynolds number, Mach number and wall interference with reference to solid and flow blockage as well as the influence of side-wall boundary-layer control are analysed. Interesting results brought to light the Reynolds number effects of the test model versus Reynolds number effects of the facility in subsonic and transonic flow as well as the effects of the side-wall boundary-layer control and wall interference.

The appearance of the Kármán vortex street in the wake of bluff bodies has been well-known for sixty years. But the simple question of whether it appears also in the wakes of two cylinders in tandem at the same speed range, seems to have been unanswered as yet. The main reason for this is presumably the lack of practical interest. The only exception was in the early years of aviation in the era of biplanes. The use of streamlined or cylindrical struts was common and stimulated the investigations of Panell et al 1915, Wong & Yu 1917, Frazer & Simmons 1917, Biermann & Herrnstein 1933, and others.

An experimental investigation was conducted to study the effect of synthetic jet (oscillatory, zero net mass flow jet) blowing near the wing tip, as a means of diffusing the trailing vortex. Velocity measurements were taken, using a Particle Image Velocimetry system, around the tip and in the near wake of a rectangular wing, which was equipped with several blowing slots. The effect of the synthetic jet was compared to that of a continuous jet blowing from the same configurations. The results show that the use of synthetic jet blowing is generally beneficial in diffusing the trailing vortex and comparable to the use of continuous jet. The effect was more pronounced for the highest blowing coefficient used. The driving frequency of the jet did not generally prove to be a significant parameter. Finally, the instantaneous and the phase-locked velocity measurements helped explain the different mechanisms employed by the continuous and synthetic jets in diffusing the trailing vortex.

Jet flap theory for thin aerofoils has been extended to include the effect of jet entrainment on the external flow when the jet is blown over the upper surface of the aerofoil. The effective camber of the aerofoil is increased by the sink effect due to entrainment and the increase of lift at zero incidence is proportional to the square root of the jet momentum coefficient. Formulae and charts are presented to facilitate the determination of the increments of lift and pitching moment due to this effect. The theory is shown to be in first-order agreement with the exact solution for a circular-arc aerofoil of small camber with distributed sinks on the upper surface.

The new theory is compared with four old sets and one new set of experimental data. It greatly improves the accuracy of prediction for cases where the incidence and flap angle are small. The new theory substantiates the usefulness of a small flap in applications of the jet flap principle.

Spanwise velocity correlations of the turbulent vortices shed by a two-dimensional bluff body of D-section are presented for different aspect ratios of the shedding body between end-plates. From these it is shown that the vortices are only shed two-dimensionally when the gap between the end-plates is less than or equal to four times the base height of the shedding body.

This paper considers actuator redundancy management for a redundant multirotor Unmanned Aerial Vehicle (UAV) under actuators failures. Different approaches are proposed: using robust control (passive fault tolerance), and reconfigurable control (active fault tolerance). The robust controller is designed using high-order super-twisting sliding mode techniques, and handles the failures without requiring information from a Fault Detection scheme. The Active Fault-Tolerant Control (AFTC) is achieved through redistributing the control signals among the healthy actuators using reconfigurable multiplexing and pseudo-inverse control allocation. The Fault Detection and Isolation problem is also considered by proposing model-based and model-free modules. The proposed techniques are all implemented on a coaxial octorotor UAV. Different experiments with different scenarios were conducted for the validation of the proposed strategies. Finally, advantages, disadvantages, application considerations and limitations of each method are examined through quantitative and qualitative studies.

The flow near the trailing edge of a circulation-controlled aerofoil is studied. The circulation is controlled by tangential blowing around the bluff trailing edge of the aerofoil.

Velocity profiles in the blowing jet on the trailing edge of a circulation-controlled aerofoil have been measured and a calculation method has been developed for this type of circulation control. The development of the blowing jet is calculated by means of a velocity-profile equation, an angular-momentum integral equation, and an entrainment equation. Improved criteria are suggested for establishing the starting conditions for the development calculation. A realistic method is presented for taking into account the effects of the boundary layer which develops upstream of the blowing slot. A separation criterion for turbulent wall jets and a curvature correction for the entrainment rate are proposed. The calculated blowing-jet developments are found to be in good agreement with experiment. There is fair agreement between the calculated and measured blowing rates required to produce given lift coefficients on the aerofoil.

The paper presents a solution for the buckling under shear of a rectangular plate which is clamped along one edge and simply-supported along the other edges. The authors have also re-examined the case of one pair of opposite edges clamped and the other pair simply-supported.

As part of a series of investigations of the swirl in S-ducts, the present paper gives the results of tests on a duct with both horizontal and vertical offsets, for comparison with earlier work on ducts with horizontal offset only. Results obtained bear a similarity to those of Reference 2, for which the duct had the same cross-sectional shapes and areas, but the addition of a vertical offset magnifies the resultant swirl considerably. As before, the swirl is found to be sensitive to entry conditions, in that a solid spoiler, blocking off 15% of entry width from the inside wall, reverses the sense of the final swirl. Taken overall, the results throw additional light on the nature of the swirl problem. A particularly interesting and potentially important correlation is obtained between the final swirl (in magnitude and sense of rotation) and a pressure difference between the top and bottom walls of the duct, taken in the end plane of the second bend.

This paper describes the work done to achieve high parallel performance for an unstructured, unsteady turbomachinery computational fluid dynamics (CFD) code. The aim of the work described here is to be able to scale problems to the thousands of processors that current and future machine architectures will provide. The CFD code is in design use in industry and is also used as a research tool at a number of universities. High parallel scalability has been achieved for a range of turbomachinery test cases, from steady-state hexahedral mesh cases to fully unsteady unstructured mesh cases. This has been achieved by a combination of code modification and consideration of the parallel partitioning strategy and resulting load balancing. A sliding plane option is necessary to run fully unsteady multistage turbomachinery test cases and this has been implemented within the CFD code. Sample CFD calculations of a full turbine including parts of the internal air system are presented.

A description is given of a supersonic pressure-tube wind tunnel which has been constructed at A.R.D.E. This is a blow-down tunnel which uses as a reservoir a long tube filled with gas under pressure. A quasi-steady supersonic flow is achieved by expanding in a convergent-divergent nozzle the subsonic flow behind rarefaction waves which propagate down the tube when a diaphragm at the nozzle exit is burst. The theory of the operation of the tunnel is given and calculations are made of the boundary-layer growth along the tube. Pressure-time records were obtained in the tube, and a high speed camera was used to obtain pictures of the flow round a model. Measurements also included a pitot-tube traverse of the nozzle exit, and the Mach number distribution was determined from the ratio of the pitot to the stagnation pressure. Tests showed that, as predicted, a constant stagnation pressure was obtained ahead of the nozzle, and it is considered that a tunnel of this type would be a cheap and simple way of obtaining an intermittent tunnel with adequate running time for many types of test, and capable of operating at a Reynolds number of more than 107 per inch at a Mach number of about 3·5.

This paper reviews recent results on the mechanics and aerodynamics of perching in a large bird of prey, the Steppe Eagle Aquila nipalensis. Data collected using onboard and high-speed video cameras are used to examine gross morphing of the wing planform by the flight muscles, and smaller-scale morphing of the wing profile by aeroelastic deflection of the feathers, Carruthers et al. High-resolution still images are used to reconstruct the shape of the wing using multi-station photogrammetry, and the performance of the measured wing profile is analysed using a panel code, Carruthers et al. In bringing these lines of research together, we examine the role of aeroelastic feather deflection, and show that the key to perching in birds lies not in high-lift aerodynamics, but in the way in which the wings and tail morph to allow the bird to transition quickly from a steady glide into a deep stall.

Increasingly stringent operational demands on military aircraft, together with the widening flight envelopes made possible by advances in aerodynamic, structural and propulsion technology, have forced the complication of stability and control augmentation systems (S/CAS) upon the aircraft designer. The early SAS configurations were conceived as devices to enhance the natural response characteristics of the aircraft; later, CAS evolved as a way of making the desired response more independent of the basic aircraft characteristics. Many recent examples of S/CAS are very complex, with equations of twentieth or higher order commonly being used for linear handling qualities analysis. In view of the fact that aircraft of this complexity have been part of the handling qualities picture for ten years or more, it is surprising that no widely accepted criteria for these aircraft have evolved.

Numerical analysis of the flow in weapon bays modelled as open rectangular cavities of length-to-depth (L/D) ratio of 5 and width-to-depth (W/D) ratio of 1 with doors-on and doors-off is presented. Flow conditions correspond to Mach and Reynolds numbers (based on cavity length) of 0·85 and 6·783m respectively. Results from unsteady Reynolds-averaged Navier-Stokes (URANS), large-eddy simulation (LES) and detached-eddy simulation (DES) are compared with the simulation methods demonstrating the best prediction of this complex flow. It was found that URANS was not able to predict the change of flow characteristics between the doors-on and doors-off configurations. In addition, the energy content of the cavity flow modes was much better resolved with DES and LES. Further, the DES was found to be quite capable for this problem giving accurate results (within 3dB of) experiments and appears to be a promising alternative to LES for modelling massively separated flows.

In this study, a new aero-propulsive model (APM) was derived from the flight manual data of a transport aircraft using Genetic Algorithms (GAs) to perform accurate trajectory predictions. This new GA-based APM provided several improvements to the existing models. The use of GAs enhanced the accuracy of both propulsive and aerodynamic modelling. The effect of compressible drag rise above the critical Mach number, which was not included in previous models, was considered along with the effects of compressibility and profile camber in the aerodynamic model. Consideration of the thrust dependency with respect to Mach number and the altitude in the propulsive model expression was observed to be a more practical approach. The proposed GA model successfully predicted the trajectory for the descent phase, as well, which was not possible in previous models. Close agreement was observed when comparing the time to climb and time to descent values obtained from the model with the flight manual data.

In commercially available gas turbine performance simulation tools, individual engine components are typically represented with non-dimensional maps of experimental or default data. In those cases where actual component characteristics are not available and default characteristics are used instead, conventional tools can deviate substantially at off-design and transient conditions. Similarly, when real component characteristics are available, conventional engine cycle simulation tools can not predict the performance of the engine at other than nominal conditions satisfactorily, or account for the impact of changes in component geometry.

This study looked into the full integration of two-dimensional streamline curvature component models with a low fidelity cycle program. Firstly, the obtained engine performance was compared against the one calculated based on default component characteristics. As a second case study, a range of flight Mach numbers and angles of attack were examined together with the effect of three different intake lip geometries on the performance of a notional, two-spool, low-bypass ratio, military engine. Two-dimensional models were used in the engine cycle analysis to provide a more accurate, physics- and geometry-based estimate of intake and fan performances.

The analysis carried out by this study demonstrated relative changes in the predicted engine performance larger than 1%. For briefness, representative results are presented and discussed in this paper for one flight Mach number and angle of attack setting. More importantly, this research effort established the necessary methodology and technology required towards a full, two-dimensional engine cycle analysis at an affordable computational resource in the very short term.

The motion of liquid in the containers of a vehicle may have severe influence upon the stability of the vehicle. For this reason many authors have investigated the behaviour of liquid with a free surface due to various modes of excitation. In the earlier development of small vehicles, the frequencies of the elastic structures were relatively high compared to the propellant frequencies and the effect of their interaction could be neglected. With the increase of modern space vehicles this interaction of the propellant with the elastic structure is becoming a pronounced influence upon the design of the vehicle and its control system. This has some influence upon the stability and response behaviour of a large space vehicle, since the close grouping of propellant sloshing and structural frequencies around the control frequency of the vehicle creates continuous interaction during flight.

In this paper a non-linear theory of the supersonic bangs from an accelerating body is worked out, using an extension of Whitham’s method. The strength of the bow shock is determined, and it is shown that the effect of acceleration on the pressure rise is appreciable at large distances from the nose for Mach numbers near unity. The results are compared with those of linear theory; the estimate for the pressure rise is found to be about a half of the linear value. The results also include an expression for the curvature of an attached shock at the nose, and a formula for the decay of a shock after it is detached from a decelerating body.

This paper deals with the force coefficients on the leeward of two bluff bodies, in particular with the origin and nature of the lift force. Initially a general description of the characteristics of the force coefficients, plus the effect of parameters such as free-stream turbulence and surface roughness in the form of stranding, is given.

A resume of all the known suggestions for the origin of the lift force is listed and a series of experiments, attempting to test these suggestions using simple methods, is described. These methods consist mainly of force measurement, involving force measurement on the leeward conductor using different shapes for the leeward and windward bodies. However, static pressure tapping and flow visualisation are also employed to a small degree.

It is shown that none of the suggestions gives a satisfactory explanation for the origin of the lift force. However, several interesting points were noted; for instance the dependence of the lift profiles for the leeward body on its own characteristics in the wake and not particularly on the wake characteristics. Hence it is shown that the use of wake parameters, measured without the presence of the leeward body when attempting to assess the forces thereon, is an over-simplification of the situation as far as lift is concerned while appearing to work quite well for drag.