I would like to add yet another note to those already published on the notions of viscous and hysteretic damping in a simple oscillator. In particular, I would like to hammer a nail into the coffin of the so-called “complex stiffness.“

One of the features which hysteretic damping may be defined to possess is the property that the energy loss per cycle (in sinusoidal motion) is independent of the frequency of oscillation, ω. This property is derived from some experimental evidence. It is shown in Reference 1 that this behaviour is reproduced by the particular integral of the equation

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.

The object of this paper is to set down an overall picture of the practical aspects of compressor blade vibration. In doing so it is inevitable that a certain content is effectively a repeat of other people's contributions to the subject. The authors make no apology for this and offer the paper as a "state of the art" as practised at Bristol Siddeley Engines Ltd. In bringing together the many relevant issues of this problem, it is hoped that a useful presentation is obtained.

There is little hesitation in indicating the most obvious and important practical aspect of blade vibration as that of the inherent danger of blade failure and it is known that the axial flow compressor is susceptible to complete failure caused by the fatigue of one blade.

The multi-stage axial flow compressor is a highly complex aerodynamic machine with, its early and late stages required to operate over wide ranges of incidence and Mach number conditions. To some extent, of course, use of variable geometry can reduce this to less taxing proportions. The aerodynamic environment of front stages has broad boundaries somewhat wider than those relating to rear stages, while middle blade rows generally operate under sensibly constant conditions. This diminishing “severity” of operation, front, rear to middle stage blading is not unrelated to the incidence of blade vibration problems which shows a similar pattern.

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.

This paper evaluates a time marching simulation method for flutter which is based on a solution of the Euler equations and a linear modal structural model. Jameson’s pseudo time method is used for the time stepping, allowing sequencing errors to be avoided without incurring additional computational cost. Transfinite interpolation of displacements is used for grid regeneration and a constant volume transformation for inter-grid interpolation. The flow pseudo steady state is calculated using an unfactored implicit method which features a Krylov subspace solution of an approximately linearised system. The spatial discretisation is made using Osher’s approximate Riemann solver with MUSCL interpolation. The method is evaluated against available results for the AGARD 445.6 wing. This wing, which is made of laminated mahogany, was tested at NASA Langley in the 1960s and has been the standard test case for simulation methods ever since. The structural model in the current work was built in NASTRAN using homogeneous plate elements. The comparisons show good agreement for the prediction of flutter boundaries. The solution method allows larger time steps to be taken than other methods.

Steady state aerofoil ground effect is studied both numerically and analytically. Discrete vortex and linear vortex panel methods are applied to a parabolic arc and symmetric Joukowski aerofoil, respectively. A single vortex model for the flow over the parabolic arc aerofoil is developed. The single vortex model and other analytical solutions, valid either near or far from the ground, are compared with the numerical results. The numerical results are used to delineate the influences of angle of attack, camber and thickness. For small values of camber and angle of attack, normalised lift is enhanced near the ground and reduced far from it. For a fixed distance above the ground, normalised lift decreases with increasing angle of attack and camber. Thickness reduces lift at all heights above the ground.

An experiment was made on a turbulent boundary layer with adverse and favourable pressure gradients alternating twice along the stream. The most important findings of the present study are as follows. In the region of the adverse pressure gradient downstream of the strong acceleration, an internal boundary layer develops within the boundary layer. In this internal layer large advection of turbulent energy occurs and the flow is locally non-equilibrium even near the wall. As a result the usual logarithmic law and turbulence similarity in the viscous sublayer break down.

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.

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.

This paper presents a review of the various aeronautical air conditioning systems that are currently available and discusses possible system configurations in the context of the aeronautical environmental control systems. Descriptions of the standard vapor compression cycle and air cycles are provided. The latter includes, simple-cycle, bootstrap-cycle, simple-bootstrap cycle (3-wheel) and condensing cycle (4-wheel). Water separation and air recirculation systems are also explored. A comparison between vapor compression cycles and air cycles is provided, as well as a comparison between different air cycles. Air cycle units are far less efficient than vapor compression cycle units, but they are lighter and more reliable for an equivalent cooling capacity. Details regarding the aircraft conceptual design phase along with general criteria for the selection of an air conditioning system are provided. Additionally, industry trends and technological advances are examined. Conclusions are compiled to guide the systems engineer in the search for the most appropriate design for a particular application.

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.

As propeller-driven aircraft are the best choice for short/middle-haul flights but their acoustic emissions may require improvements to comply with future noise certification standards, this work aims to numerically evaluate the acoustics of different modern propeller designs. Overall sound pressure level and noise spectra of various blade geometries and hub configurations are compared on a surface representing the exterior fuselage of a typical large turboprop aircraft. Interior cabin noise is also evaluated using the transfer function of a Fokker 50 aircraft. A blade design operating at lower RPM and with the span-wise loading moved inboard is shown to be significantly quieter without severe performance penalties. The employed Computational Fluid Dynamics (CFD) method is able to reproduce the tonal content of all blades and its dependence on hub and blade design features.

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.

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.

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.

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.

In the present work, the authors examine two computational approaches that can be used to study flexible flapping systems. For illustration, a fully coupled interaction of a fluid system with a flapping profile performing harmonic flapping kinematics is studied. In one approach, the fluid model is based on the Navier-Stokes equations for viscous incompressible flow, where all spatio-temporal scales are directly resolved by means of Direct Numerical Simulations (DNS). In the other approach, the fluid model is an inviscid, potential flow model, based on the unsteady vortex lattice method (UVLM). In the UVLM model, the focus is on vortex structures and the fluid dynamics is treated as a vortex kinematics problem, whereas with the DNS model, one is able to form a more detailed picture of the flapping physics. The UVLM based approach, although coarse from a modeling standpoint, is computationally inexpensive compared to the DNS based approach. This comparative study is motivated by the hypothesis that flapping related phenomena are primarily determined by vortex interactions and viscous effects play a secondary role, which could mean that a UVLM based approach could be suitable for design purposes and/or used as a predictive tool. In most of the cases studied, the UVLM based approach produces a good approximation. Apart from aerodynamic load comparisons, features of the system dynamics generated by using the two computational approaches are also compared. The authors also discuss limitations of both approaches.

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.

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.