For the safety problems caused by the limited landing space of the deck during the arresting process of the carrier-based aircraft, a dynamic model of the carrier-based aircraft’s landing and arresting is built. Based on the batch simulation method, the lateral dynamics safety envelope of the aircraft during the arresting was defined, and the dynamic response of the key points in the envelope during the arresting process was investigated. Subsequently, the influence of engine thrust and aircraft quality on the arresting safety envelope was studied based on reasonable safety evaluation indicators, and the safety status envelope of the deck arresting was given. Then, the particular Hamilton-Jacobi partial differential equation is used to obtain the lateral dynamics safety envelope of the carrier-based aircraft in the process of landing and arresting by backward inversion. Results indicate that engine thrust and landing quality have little effect on the yaw angle in the arresting safety boundary during the arresting. Additionally, with the engine thrust and landing quality increase, the maximum safe off-centre distance gradually decreases, and the safety boundary decreases accordingly. During the phase of landing glide, the engine thrust and quality have little effect on the maximum safe eccentric distance. When the engine thrust is increased by 40%, the maximum safe yaw angle is reduced from 0.3°, and the safety boundary is reduced by 4.2%. When the aircraftquality increases by 40%, the maximum safe yaw angle is reduced by 0.4°, and the safety boundary is reduced by 2.8%. The findings of this paper can provide framework for the research on theaircraft-to-carrier dynamic matching characteristics of the carrier-based system, and is of great significance to the research on improving the safety of the carrier-based aircraft landing arresting.

This research paper presents an application of the integrated process and product design (IPPD) approach for selecting the best joint configuration for dissimilar material joining in the early product design phase. The proposed methodology integrates the multi-criteria decision making (MCDM) approach with quality function deployment (QFD) to identify the key criteria for joint selection, including load-carrying capacity, size, cost per joint, ease of manufacturing, time consumption and deformation. Three types of joints (rivet, weld and adhesive) and two hybrid joints (adhesive-weld and adhesive-rivet) are considered for three dissimilar material configurations: carbon fiber-reinforced plastic (CFRP) aluminum, CFRP steel, and aluminum-steel. QFD is utilised to transform job requirements into design criteria, and in the second phase, the Technique for Order Preference by Similarity to Ideal Solution (TOPSIS) is employed to choose the optimal joint configuration based on the weighted criteria acquired in the previous phase. The selected joint configuration is then validated through experimental study. The findings demonstrate that the proposed IPPD approach with QFD-TOPSIS techniques is highly effective for selecting mechanical joints for dissimilar material joining in the early design phase. The study concludes that the adhesive-rivet hybrid joint is the optimal solution among all alternatives. The proposed methodology can ultimately lead to improved product reliability and performance, as well as reduced development time and cost.

In the present work, a folding wing system (FWS) was developed for guided ammunitions, so that the swept-back angle could be adjusted during both gliding and diving phases. Unlike previous designs, the FWS does not have any fixing mechanisms or brake elements, and it provides folding functionality to reduce the drag force during the terminal phase. We conducted mechanism design, manufactured the FWS, performed system identification and designed various controllers including linear quadratic regulator (LQR), linear quadratic integrator (LQI), sliding mode control (SMC) and second-order sliding mode control (SOSMC) to adjust and hold the desired swept-back angles. Then, the performance of the FWS was tested experimentally under two different flight scenarios, with and without aerodynamic loads. While all controllers operated with almost zero steady-state error (SSE) in the absence of aerodynamic loads, the SOSMC was the most effective controller under aerodynamic loads, considering SSE, delay, chattering, and energy consumption.

In the first part of the paper a brief review is made of major items for research in the helicopter field, the subjects touched on including rotor aerodynamics, stability and control, evaluation of configurations, vibrations and fatigue, and operational aspects An outline is given, in the section on stability and control, of work being done at A & A E E in connection with the assessment of the longitudinal handling characteristics of helicopters The second part of the paper contains an account of recent work at A & A E E on the low speed and take-off performance of a helicopter An empirical method of low speed performance estimation is described, and the variation with wind speed of the ground effect on a rotor is discussed theoretically and on the basis of experimental results A theoretical analysis of the forward take-off motion of a helicopter is briefly presented, and the final section is concerned with the performance of a multi-engine helicopter in the event of failure of one engine, with particular reference to the possibility of safe operation from the type of site proposed for civil use

This paper concerns the free vibration of linear systems which are subject to damping. The general theory for a system having n degrees of freedom is recalled and illustrated by means of a simple example. Using an analytical method, the approximate theory for small damping is derived, showing the nature of the entrainment of the principal modes; the results are found to confirm those of Southwell (who uses a physical argument) and they are illustrated by an example. The approximate theory is shown to need modification when there are two nearly-equal natural frequencies and the appropriate corrections are found. The analysis is made somewhat complicated by the existence of two criteria of “ smallness ” instead of one; these are the difference between the close natural frequencies and the amount of the damping.

This paper considers the problem of a reinforced circular hole in a bent or twisted flat elastic plate having small deflections. The reinforcement is chosen so that the ratio of its radial width to the diameter of the hole is small; this implies that there is no radial stress variation across sections of the reinforcement. Using the classical theory for bending, expressions are obtained for the boundary conditions at the junction of the reinforcement and plate. The expressions are general in so far as they are also applicable to plates under the action of lateral loading. The following particular problems are considered: —(a) uniform bending about one or two axes, (b) uniform twisting about two axes. For the particular case of uniform bending about one axis, data has been obtained for the stress concentration around holes for plates and reinforcements having a reasonably wide range of practical dimensions.

A detailed discussion on the difference between an initial value line and a line characterised by a boundary condition has been presented. Two types of boundaries are described and illustrated. To examine each boundary, several different calculations have been performed for a straight pipe. The results of the numerical calculations are compared with an analytical solution. It is shown that known pressure and velocity at the pipe ends give the most accurate results. Comparisons are also made between several practical types of calculations which give similar findings. The use of time-dependent boundaries can lead to errors as large as 40 per cent in derived results. It is shown that good accuracy can be restored by converting the boundaries into initial value lines. It is concluded that in general no more than one time-dependent boundary should be used in any calculation. Finally it is demonstrated that errors are not revealed by means of pressure diagrams alone.

Closed form solutions for low thrust orbit raising are well known for mission analysis and vehicle performance prediction. This paper provides extended solutions which include nodal regression through the effect of the J2 oblateness term in the geopotential. Since regression is a strong function of orbit radius, the resulting change in ascending node angle is coupled to the in-plane orbital motion. However, it is demonstrated that the resulting dynamical equations may be separated and solved sequentially. Nodal regression during orbit raising is of importance for solar electric vehicles due to the resulting variation of sun aspect angle and eclipse conditions.

The aerodynamic derivatives for a two-dimensional aerofoil with an oscillating control surface in a high subcritical flow are calculated using the method of Nixon. The aerofoil is assumed to be symmetric and at zero incidence with a zero mean control surface deflection angle. With the exception of the control surface stiffness derivative the magnitudes of the aerodynamic derivatives are greater than the corresponding linear values by between 15–35 per cent; the greatest difference occurs for low values of the frequency parameter. The magnitude of the control surface stiffness derivative is less than the corresponding linear value by the order 10–15 per cent.

The small deflection/initial buckling behaviour of elastic, homogeneous plates is characterised by the absence of coupling between in-plane and out-of-plane quantities, i.e. stress couples, shear strains etc. By contrast, heterogeneous plates, of which generally laminated plates are an example, exhibit ‘shell-type’ behaviour in as much as in-plane and out-of-plane quantities are coupled. Reissner and Stavsky were the first to correctly formulate the field equations for such plates. Since that time a large number of papers on the influence of coupling on the behaviour of many different types of laminated structural elements, such as beams, plates and shells have appeared in the literature. However, it appears that until quite recently, the analysis of these structural elements in the presence of a general (Winkler/Pasternak) elastic foundation has not been undertaken.

The natural frequencies of a plate or a shell can be changed by altering its stiffness. This can be achieved by appropriately varying the mass or curvature, by introducing initial stresses or by attaching stiffeners to the system. The last method has obvious advantages over the other methods. The natural frequencies are thus raised, lowered or remain unchanged depending on (i) the dimensions of the stiffener relative to the plate or shell dimensions, and (ii) its position.

I Have the great honour to give the Fifth Memorial Lecture in tribute to Juan de la Cierva, a great man and pioneer in rotating wing development, whose work was predominantly done in England.

His fundamental practical achievement, resulting from a deep knowledge of aeronautical science, was the demonstration of the fact that a big airscrew with suitable pitch could serve as a lifting surface without the danger of stall which is inherent in fixed-wing aircraft.

The importance of the effects of compressibility on the performance and behaviour of high speed aircraft has been appreciated for a great many years, but in this country it was not until 1937 that serious attention was given to the design of a wind tunnel that would enable the study of compressibility to be undertaken by tests on models of reasonable scale. Prior to 1937 the work which had gone on had established the characteristics of a few aerofoils and simple shapes; at Farnborough the work mainly related to the testing of high tip speed propellers, while at the National Physical Laboratory tests had been made of a two-dimensional character on aerofoils, these tests being made in a small wind tunnel driven on the injector principle, by air discharged from a high pressure reservoir.

A Method is described for providing warning of the onset of undesirable effects produced by flow separation on an aircraft wing. It is based on static pressure measurements near the trailing edge and appears to have advantages over alternative methods.

Recent work has suggested that certain effects of flow separation such as buffeting, aileron buzz and undesirable changes in loading, occur when a “ bubble ” of separated flow originating at the leading edge of a wing, or at a shock wave on its surface, first becomes sufficiently large to affect the flow at the trailing edge.

In a recent paper the suggestion is made that “it would be sensible when discussing the flow to use a Mach number based on the actual propagation velocity in the given circumstances.”

This directs attention to the deeper question of whether or not a Mach number however defined is a rational parameter in terms of which to express effects of flow.

In certain cases the quantity γp/ρ occurs quite naturally, regardless of any physical interpretation that can be given to it.

The advent of a new generation of turbojet powered transport aircraft with engines mounted on the rear fuselage has complicated the task of the aerodynamicist, particularly when an attempt is made to determine the individual contributions to the stability and aerodynamic loading of the aircraft. This paper gives some idea of the additional complexity which is directly attributable to the presence of these engine nacelles and their effects on manoeuvring tail loads. Where rational methods are available for determining these effects they are referred to or included.

In this paper we are restricting our attention to those cases of circumferential inlet distortion where the distortion of the velocity profile is too large to be described as a small perturbation. Thus any theory describing the flow must use a mathematical model which is non-linear.