The developments in the use of composite honeycomb in passenger aircraft structures has been growing rapidly over the past ten years. Starting with its use in the semi-structural internal use as a flooring material such as the cross plied UD glass-nomex cored floors of the wide body jet aircraft — 747 — L1011 and DC10, to the external uses for control surfaces, fins, undercarriage doors, engine pods, etc, and the ultimate all-composite fuselage of the LearFan project.

The progress in the light aircraft field by comparison, is almost non-existent, the major use of composites is restricted almost entirely to the manufacture of sailplanes and the majority of these use foam stabilisation rather than honeycomb.

This paper describes a preliminary investigation of the torsional stress concentration factors for circular shafts containing grooves determined using an analogue computer. The range covered by this analysis is for grooves which increase in depth from a minimum of 0·05 of the diameter of the shaft, with radii which vary from 0·5 to 0·05 of the diameter at the minimum section. Results are presented in the form of a data sheet, and a comparison is made with the approximate expression of Neuber. An investigation is also made of the modifications to the stress concentration factor if the flank is not perpendicular to the centre-line of the shaft.

Submerged intakes are widely used on aircraft as a means of supplying internal flow systems with ambient air. It is the purpose of this note to show by analysis of existing experimental data that the boundary layer at the location of the submerged intake can have an appreciable effect on its performance.

Information here relates to the NACA type of divergent submerged intake with a ramp angle, a, of 7° and an aspect ratio, w/d, of 4 because this particular combination of planform and geometry is now almost universally used.

The dynamic analysis of heated structures is of considerable importance. Significant heating of a flight vehicle may take place at high speed causing reduction in its structural stiffness and vibration frequencies. Such changes in frequencies and mode shapes can substantially affect various static and dynamic aeroelastic instabilities.

In general, availability of a similarity law simplifies the solution of a problem. The following presentation studies similarity relations existing between an unheated vibrating cylinder and one with temperature gradient across its thickness.

The high cost of new weapon systems and the loss of tactical capability when vehicles are out of operation for unscheduled repairs are forcing the military services to require a higher order of structural reliability. Fleets are being kept in operation for many years beyond the time originally anticipated during design, resulting in many cases, in the need for costly structural repairs and modification to cure fatigue problems. Consequently, new weapon system procurement contracts are being written with a “fatigue-life” guarantee as a contractual requirement. This situation calls for careful attention to details during the design phase. More testing, both exploratory to aid in detail design decisions, and confirmatory to validate the design and demonstrate compliance with the contractual requirements, is required. Better test methods and more meaningful test results are demanded from the test laboratory. Methods for more accurate predictions of the loads experience during the entire service usage are required of the structural analysts.

Recent developments in the design of rigid airships, as commercial air liners in England and Germany, and for military purposes in America, show a demand for greater lifting capacity, cruising range and speed. The increase in size is mainly responsible for radical changes in the structural arrangement of the transverse frames, introduced by the designers of the British airships, R.100 and R.101.

In an investigation, on the characteristics of the wake from slender cone-cylinders at large angles of attack, the strength and position of the vortices in the wake were determined for a large range of flow conditions. Two types of experiments were conducted, namely schlieren studies and wake traverses.

Schlieren photographs (such as Fig. 1) clearly reveal straight parallel vortex lines in the wake for subcritical values of the cross-flow Reynolds number component and for cross-flow components of the free stream Mach number up to at least 1·4. In ref. 1 numerous such schlieren photographs were analysed using the impulse flow analogy (refs. 2 and 3 for example). According to the analogy the progressive development of the wake along the body when viewed in cross-flow planes is similar to the growth with time of the flow behind a two-dimensional cylinder started impulsively from rest with the same cross-flow conditions.

Mr. Lyall's remarks may be answered from the evidence already present in refs 1 and 2 if these are considered objectively. They establish that the ratio of the shear stresses as given by Preston tubes is correct. Evidence on this point is given by the value obtained for rmt from the Preston tube measurement of shear stress ratio and its agreement with the correlation of rmt as a function of radius ratio given by Leung, Kays and Reynolds.

To write a paper on the work of the Royal Aircraft Establishment is a formidable task. To summarise the programme of work of the Establishment would either provide material for a series of lectures or, if compressed to a single lecture, would lead to a cursory and superficial treatment. I make no apologies, therefore, for not attempting any sort of complete survey. Within the scope of this lecture all I intend to do is to give a brief description of the RAE, what it is, what it does and how it is organised and then to consider a few items drawn from three of the RAE's research programmes, which have been selected to show some of the characteristics of the RAE and the kind of work which goes on there.

The first series of items to be presented fall in the field of the slender wing. The great potentiality of the slender wing for a wide range of applications was apparent from the time the concept was first considered. What I want to stress is the enormous range of detailed researches which are required before such a concept can be used in practical designs.

Despite the fact that there have been few reported instances of the flutter of circular, cylindrical, iso-tropic shells, there still remains a challenge to the analyst to derive a theory which will correlate with the available experimental data.

The aim of this note is to compare the plane turbulent free jet diffusing in an infinite stagnant ambient environment with the corresponding plane turbulent wall jet on a smooth boundary regarding the three important characteristics, namely the velocity distribution and the variation of the velocity and length scales, in the region of developed flow.

The use of rule-based fuzzy logic system for estimating the stability and control derivatives for the longitudinal aircraft motion is proposed. The capabilities of the fuzzy logic system in estimating both the short-period and the phugoid mode of motions are explored. The flight data used in the estimation process were generated using the three nonlinear longitudinal equation of motion for a small remotely piloted vehicle with all the aerodynamic coefficients obtained from the wind-tunnel tests. The preferred method of perturbation of the aircraft elevator for data collection is also highlighted. The stability and control derivatives are estimated as the change in the aerodynamic force or moment due to small variation in one of the motion or control variables about its nominal value when the rest of the variables are held constant at their respective nominal values. The changes in the aerodynamic force and moment are predicted using the fuzzy logic system. The results show that the fuzzy logic system has a good potential as alternative tools for parameter estimation from flight data. The proposed method does not require any guesses of the initial values of the flight parameters.

The authors were listed incorrectly in the article by Zanotti(1). The authors should have appeared in the following order:

G. Gibertini, A. Zanotti, G. Droandi, D. Grassi, G. Campanardi, F. Auteri, A. Aceti and A. Le Pape

The article has now been updated with all authors listed in the correct order. This change has been made to the online and print PDF and HTML copies.

The in-plane crushing and energy absorption of sandwiched honeycomb cores with facesheets are examined through finite element simulations. Assuming no debonding between the facesheet and honeycomb core (which would be the case if manufacturing techniques such as brazing are used to produce very strong bonds between the facesheeet and the core), intracellular buckling mode for thin facesheets, and wrinkling mode for thick facesheets are observed. In the dimpling mode, deformation is governed by the core, honeycomb vertical cell walls do not deform, and the inclined wall deformation does not vary through the cell depth. In the wrinkling mode, deformation is governed by the facesheet, the vertical cell walls deform significantly, and the inclined cell wall deformation varies through the cell depth. Increasing cell angle increased Specific Energy Absorption (SEA) for honeycombs with thin facesheets. Decreasing vertical cell wall length increased SEA for honeycombs with thick facesheets. Increasing wall thickness and decreasing core depth increased SEA for honeycombs with thin and thick facesheets. With geometric changes, SEA increased ~3 times over the baseline configurations. For a given keel beam dimensions, using fewer rows of larger cells reduces the effective non-dimensional core-depth, thereby increasing the effect of the facesheet and the SEA significantly. The SEA of sandwiched honeycombs with facesheets in in-plane crushing appears to be competitive with, or better than, SEA honeycombs in out-of-plane crushing.

This paper reviews some of the work performed on the ground effects rig at BAE Systems in Warton, the now unique facility for modelling dynamic vertical landings and measuring hot gas ingestion. The paper discusses the flow diagnostic tools which have been used in recent times, up to the complexity of 3D PIV, and uses examples from the F-35 JSF programme to illustrate these. The paper hopefully demonstrates the difficulties of using these tools, as well as the benefits they can bring to a development programme.

The ground effects rig is a facility designed to measure temperature rise in aircraft intakes during vertical manoeuvres close to the ground. It is a complex facility that comes as close as possible to accurately modelling the flowfields around an aircraft moving both vertically and horizontally near the ground, with the ability to model dynamic pitch and roll at the same time.

Standard instrumentation for the models consists of rapid response thermocouples mounted in a rake at the engine face. 45 thermocouples of 0·05mm diameter with a time constant of about 10ms are used.

Although, with its standard instrumentation, the rig can measure how much hot gas gets to the engine face, it doesn’t show how it got there, or where it came from, which is the knowledge required to improve the design. Hence there is a need for flow diagnostics.

The pay-off, in the final analysis needs to be a better air transport system for the traveller. This will, in turn, generate more traffic for the airline and hence more business for the aircraft manufacturer. To achieve a substantial and profitable share of the available market for new aircraft the manufacturer's product must be produced at the right time, meet the requirements of the airlines, and be competitive in terms of price and capability with the products, both new and existing, of his competitors. Advanced technology, applied in a cost effective manner, can be one of the ways of achieving this competitiveness and of improving the air transport system.

On the 15th May 1941 a small aircraft landed at Cranwell after a flight lasting 17 minutes. The pilot, Fl. Lt. P. E. G. Sayer, got out of the aircraft and was immediately surrounded by an excited group wanting to know how the machine had performed. He answered questions standing by the aircraft and reading from the notes on his knee pad, which he copied on to the paper as shown in Fig. 1. This flight, the first of a Whittle turbojet propelled aircraft, started a new era in aviation. It also marked the beginning of the rapid development of a new heat engine, the fifth since Newcomen’s steam engine.

As you can see from the pilot’s notes, the performance of the E28/39 aircraft in that first flight was modest. An Indicated Air Speed (IAS) of 240 mph was reached. That it was powered by a turbojet can be seen from the readings of rpm, jet pipe temperature, etc. This first engine developed about 850 lb take-off thrust and weighed some 700 lb.