This paper presents a simulation study of ion thruster plume effects on formation flying spacecraft. Two formation flight applications using micro-ion propulsion are considered: an L2-Halo orbit interferometer formation and a LEO micro-satellite formation. Worst case scenarios in both missions have been investigated. We focus our study on thruster configurations resulting in possible indirect plume impingement on satellites outside of the direct impingement zone. Indirect impingement which cannot be predicted except through plasma simulation or in-flight measurements might expose critical spacecraft elements such as optical sensors to a harsh contamination environment. A high-fidelity electrostatic plasma simulation code for parallel computing platforms was used in the study.

In our study, we found that using miniature scale ion propulsion in both formation missions creates a very low charge-exchange plasma environment which results in tolerable contamination environment for other spacecraft in the formation.

References 1 and 2 discuss the shock position in over-expanded nozzles, and in particular the transition from nozzle flow characterised by a normal shock in the nozzle to nozzle flow characterised by oblique shocks and separation from the wall. As is well known, the shock position and pressure distribution for unseparated overexpanded flow can be adequately explained using one-dimensional fluid mechanics. Ref. 2, while suggesting a criterion for transition to separated flow, maintains that the separation point is not predictable. The criterion suggested by ref. 2 is that whenever the pressure ratio p2/p0, associated with expansion to the nozzle exit plane followed by normal shock compression at the exit Mach number exceeds the nozzle pressure ratio pb/p0, separated flow will occur. Based on this criterion and the double valuedness of the p2/p0 locus, it is argued that at a nozzle pressure ratio of 0·624, a continuous increase of nozzle exit to throat area ratio will provide for transition from unseparated normal shock flow to separated flow and back to unseparated normal shock flow.

A simple closed-form formula for the coefficient of surface pressure, Cp is given in this paper. The formula is based on first and second approximations of the full problem and can be used for the Newtonian flow past two-dimensional thin airfoils at small angles of attack and with attached shock waves. It thus extends Cole's zero order theory to the effects of non-zero (γ – 1) and finite values of Mx where γ is the ratio of the specific heats of the gas and Mx is the free stream Mach number. The analysis presented relies on a previous recent formulation of the hypersonic small disturbance theory which has proved to be more advantageous than the hypersonic small disturbance theory. The results are compared with other approximate methods and the agreement is found to be generally good.

If the motion of an aircraft is restricted, near the ground, to zero bank, comparatively simple solutions to the resulting equations of motion can be obtained, enabling the response to an applied yawing moment to be calculated rapidly. In this paper the analysis is made for three simple forms of applied yawing moment. From the results obtained in a particular case it is possible to judge the suitability of the fin and rudder design chosen on the basis of maximum sideslip reached. A general curve of overswing, in terms of the aerodynamic derivatives of the aircraft, enables the maximum sideslip to be calculated rapidly in a particular case.

Because of typogiaphical difficulties the symbols ^ have been replaced by ° with v, r and t throughout the paper.

If it implied anything in the context of high performance turbine engines, the mutilated quotation from “The Hound of Heaven” might conjure up a vision of afterburning turbojets, or turbo-ramjets, propelling second or third generation supersonic transports at very great heights. This would be a misleading picture, and it will probably emerge that the quotation has little to do with what follows, as this is largely concerned with subsonic flight—on which I believe the economic future of aviation will depend for many years.

The terminology ‘rotorcraft-pilot coupling’ denotes phenomena arising from interaction between pilot and rotorcraft. Among these, the present work deals with ‘pilot-assisted oscillations’ that derive from unintentional pilot actions on controls due to seat vibrations, and are strictly related to rotor-aeroelasticity/airframe-structural-dynamics coupling, with involvement of blade control actuator dynamics. Focusing the attention on helicopters, a comprehensive rotorcraft model is developed and applied, with main rotor unsteady aerodynamics described in state-space form. This makes it particularly suited for stability and frequency-response analysis, as well as control applications. Numerical investigations address two critical rotorcraft-pilot coupling aeroelastic issues: stability of vertical bouncing and gust response in hovering. Results from main rotor unsteady aerodynamics modelling are compared with widely-used quasi-steady aerodynamics predictions. These suggest that, for accurate RPC/PAO phenomena predictions, mathematical modelling should include the three-dimensional, unsteady-flow effects, and that the pilot-in-the-loop passive behaviour produces a beneficial effect on the load factor generated by gust encountering.

The theory of stress concentration is such a fascinating study that the whole of this lecture could be most happily devoted to discussion of the stress distribution round all sorts of special boundaries. At the same time the most detailed knowledge of elastic stress distribution is of little value unless we can be sure how the stresses computed will be reflected in the actual strength of the part. For that reason this paper is divided into three sections: the first deals with certain moderately simple cases of stress and load concentration and shows how the stress analysis is confirmed by static and fatigue tests: the second describes some investigations of special two–dimensional boundaries and indicates, some fairly general theoretical conclusions which can be drawn: in the third I shall try to review the possible reasons why, in practice, the best found conclusions are not always borne out.

An investigation of active control of the swept shock wave/boundary-layer interaction using ‘smart’ flap actuators is presented. The actuators are manufactured by bonding piezoelectric material to an inert substrate to control the bLeed/suction rate through a plenum chamber. The cavity provides communication of signals across the shock, allowing rapid thickening of the boundary layer approaching the shock. This splits the shock foot into a series of weaker shocks forming a lambda structure, thus reducing wave drag. Active control allows optimisation of the unimorph deflection, hence rate of mass transfer.

In this paper, results of the interaction using pressure sensitive paint (PSP) are emphasised. It is shown that the use of PSP, in conjunction with discrete pressure data, enables the main features of the interaction to be observed when the actuators are subject to different deflections.

Vortex merger is studied within the context of two-dimensional discrete vortex sheets and demonstrated on two equally oriented circular vortices and aircraft tip and flap vortices. It is confirmed that, depending on the wing load distribution, the latter may or may not coalesce into a single counter-rotating pair. The interaction of a vortex with an equally oriented shear layer, governed by the same physical principle, suggests a possible intensification of an aircraft vortex in cross-wind shear.

In recent years some types of aircraft have shown signs of structural trouble due to repeated loading and exploratory thought has therefore been given to the general problems arising. There are two main aspects to be considered: the nature and magnitude of the loads repeatedly occurring in flight and the behaviour and strength of aircraft structures under such loads. In this country a good deal of attention has already been given to the first side of the question, mainly by V-g recorder work, but also by some strain gauge investigations, but the other side of the matter—the resistance of aircraft structures to repeated loading— has so far received less attention. Direct experiments have not in this country gone much beyond exploratory work on a series of Typhoon tailplanes, described by Oaks and Townshend.

The contraction of a wind tunnel should be free from adverse pressure gradients, since this might cause boundary layer separation.

A wall contour has been designed for a circular cross-section contraction using incompressible flow theory. This gave a favourable pressure gradient at the beginning of the contraction where separation is likely to occur.

Appendix I compares the theory with experimental results obtained from a model of a proposed supersonic tunnel of which the contraction is rectangular in cross-section and which has been based on the results obtained in this report.

The Silent Aircraft Initiative was a Cambridge-MIT Institute programme involving a large team of researchers from both the University of Cambridge and the Massachusetts Institute of Technology (MIT). The aim of the project was to produce a concept aircraft design that would be so quiet it would be imperceptible in the urban environments around airports.. This paper gives an overview of how all the sources of engine noise were carefully addressed within the Silent Aircraft design. Novel technologies, a new engine configuration, improved airframe integration, new operational procedures and advanced component design were all required in order to reduce the overall engine noise level to the Silent Aircraft target. The study suggests that in order to dramatically reduce the noise of future aircraft engines a number of major design changes must be combined.

In order to increase the effectiveness of interceptor missile, velocity rendezvous angle control is required in the terminal phase. The purpose of this paper is to obtain an optimal guidance law which can achieve specified velocity rendezvous angle as well as zero terminal miss distance. A biased optimal guidance law based on a simplified mathematical model is deduced for interceptors engaging against invasion aircraft. Different from previous literatures on this issue, the presented guidance law suitable for intercepting high-speed maneuvering target. Another advantage is, under centimeter level miss distance setting, the guidance law needs smaller guidance command near the terminal time, which can successfully avoid command saturation. Simulation results demonstrate the effectiveness of the presented guidance law.

The authors have recent experience of the specification, procurement, customisation and operation of advanced air traffic control simulator systems. This paper summarises the specification of the software and hardware components of air traffic control simulators within a framework of user needs and advances in simulation technology. Both the required conventional characteristics and the desired innovative features which can accommodate future requirements are defined using taxonomy within this framework. Considering the high investments and challenges involved in the acquisition and operation of modern simulators, the proposed classifications and approaches will provide important benefits to users when selecting or renovating their systems.

The work reported here extends the work of Nikuradse on the categorisation of surface roughness. Nikuradse’s work considered pipe flow, whilst in this paper boundary layer flow is investigated. Important and fundamental differences between these two types of flow are shown. An experiment upon turbulent boundary layers over sandpaper-type roughnesses enabled a fluid-mechanic categorisation of these surfaces to be obtained for this flow. This is compared with some shape and size categorisations of the roughnesses from detailed measurements.

This paper considers the problem of a three-axis flexible satellite attitude stabilisation subject to the vibration of flexible appendages and external environmental disturbances, which affect the rigid body motion. To solve this problem, a disturbance observer is proposed to estimate and thereby reject the flexible appendage vibration. Based on the H∞ and Linear Matrix Inequality (LMI) approach, a controller for spacecraft with flexible appendages is proposed to ensure robustness as well as attitude stability with high precision. Stability analysis of the overall closed-loop system is provided via the Lyapunov method. The simulation results of three-axis flexible spacecraft demonstrate the robustness and effectiveness of the proposed method.

The Inaugural Lecture of the Air Law Group of the Society “Crime in Aircraft” by the Hon. Mr. Justice Wilberforce was given in the Lecture Theatre, 4 Hamilton Place, London W.l. on 31st October 1962. Mr. L. A. Wingfield, M.C., D.F.C., A.R.Ae.S., Solicitor to the Society, presided. Introducing the lecturer Mr. Wingfield said that the Hon. Mr. Justice Wilberforce was a Judge of the Supreme Court (Chancery Division). He had also been the leader for the United Kingdom at no less than four diplomatic Conferences on Air Law—Geneva in 1948, Rome in 1952, The Hague in 1955 and Guadalajara in 1961—he had also been the Chief United Kingdom member of the Legal Committee of I.C.A.O.; to find a more distinguished contribution to air law than Sir Richard's would be difficult.

The Society's Air Law Group arose indirectly from a lecture to the Society in January 1961 by Mr. H. Caplan on “The Law versus Science in Aeronautics”. The Group was formed in September of that year, and since its inception Sir Richard had acted as Chairman. One of the first tasks tackled by the Group had been the study of the draft convention adopted by I.C.A.O. at Munich in September 1959 on offences occurring on board aircraft. At the request of the Ministry of Aviation the Steering Committee of the Group, voicing an independent view, reported on the draft convention.

Not only was this the Group's inaugural lecture, it was also the first occasion in the Society's 96 years that it had been addressed by one of Her Majesty's Judges.

Before beginning his lecture the Hon. Mr. Justice Wilberforce congratulated the Society on its initiative in forming the Air Law Group and paid tribute to the Secretary of the Society, Dr. A. M. Ballantyne, T.D., B.Sc, F.R.Ae.S., F.A.I.A.A., Hon.F.C.A.S.I., for the work he had done and the assistance he had given in organising the Group.

The following report is taken from the tape recording of the lecture.

Large deflection problem of a spring loaded hinged nonuniform cantilever beam subjected to a rotational distributed loading is formulated by means of a second-order non-linear integro-differential equation. The problem is examined by considering the beam of rectangular cross-section with linear depth taper subjected to a uniform rotational distributed load. The elastic curves of a beam for this special case are presented.

The chief aim of the vibration engineer is to prevent the occurrence of structural failure due to excessive vibration. A secondary aim of only slightly less importance is to ensure comfort for personnel by the elimination of physiologically unpleasant vibrations; this consideration is a vital factor in the effective design of operational aircraft. The technique of vibration research is largely a matter of detecting and avoiding resonant conditions, and any method of predicting natural frequencies by calculation is welcome if only because it offers the possibility of reducing the practical testing programme.