The evolution of low Earth orbit satellites under the action of atmospheric drag and J2 oblateness perturbations has been widely investigated. Simple analyses describing the altitude decay of a circular orbit through energy arguments have also been widely available. This paper extends these simple analyses by providing compact analytic expressions for true anomaly and ascending node angle along with a more accurate expression for altitude decay. The resulting solutions then allow a fast and efficient method of investigating the orbital evolution of decaying, precessing circular orbits. Applications include the rapid evaluation of station-keeping budgets for both single platforms and the orbital evolution of satellite constellations. Passive station-keeping by continuous low thrust acceleration is investigated and shown to be naturally unstable.

For a 15° boat-tailed heat shield or payload shroud, the unsteady, compressible Navier-Stokes equations are numerically solved. For a Reynolds number of the order of one million and for four different freestream Mach numbers, the time integration is performed for 64 x 30 grid computational domain. The entire flow-field is analysed to delineate the gross dominant characteristics. It is observed that the flow in the boat-tail region is completely attached for Mach number 2·47 and beyond whereas it is separated for lower Mach numbers. A comparison between the numerical solutions and the available experimental results is provided and, for incipient flow separation conditions, an evaluation of unsteady surface pressure data is presented.

An essay which is to be submitted in competition for the R.38 Memorial Prize may fairly presume a knowledge of airship history in recent years, and I shall here touch only on those circumstances which have influenced the writing of this paper. Between the cessation of airship construction in the autumn of 1921 and its revival in the summer of 1924, all branches of official research were discontinued by order of the Air Council, excepting an inquiry into the theoretical aspects of stress-calculation which had been urged by the Aeronautical Research Committee in its report on the disaster to R.38.

In the Realm of stability and control, we are at present travelling headlong into a region of new problems. Besides tidying up the outstanding items on the types of aircraft we are used to seeing flying about to-day, a not inconsiderable task when one considers the troubles we encounter when we undertake the design of a new so-called “conventional” aeroplane, we have to tackle two major lines of research and development. First, the peculiar problems associated with the stability and control of aircraft of large size, both military and civil, into which difficulties associated with high Mach numbers intrude themselves but little; and secondly, the even more extensive problems associated with the stability and control of aircraft, both large and small, which are to fly at very high Mach numbers, not only at their design cruising and diving speeds, but also in the extremely important slow speed conditions. Mr. M. B. Morgan of the R.A.E. has dealt with some aspects of the second series of problems, so I propose to concentrate on some of the problems involved in the first series, that is those primarily associated with the increase of aircraft size.

The simplest case of asymmetric heating of a fluid flowing between parallel plates is considered. This occurs when heat is transferred to the fluid through one plate, the other being adiathermal. For a theoretical analysis, it is convenient to assume fully developed flow in a parallel channel of infinite aspect ratio, this condition being not widely different to that met with in practical applications.

Wind-tunnel tests of a heavy-class helicopter model were carried out to evaluate the effectiveness of several components optimised for drag reduction by computational fluid dynamics analysis. The optimised components included different hub-cap configurations, a fairing for blade attachments and the sponsons. Moreover, the effects of vortex generators positioned on the back ramp were investigated. The optimisation effect was evaluated by comparison of the drag measurements carried out for both the original and the optimised helicopter configurations. The comprehensive experimental campaign involved the use of different measurement techniques. Indeed, pressure measurements and stereo particle image velocimetry surveys were performed to achieve a physical insight about the results of load measurements. The test activity confirms the achievement of an overall reduction of about 6% of the original model drag at cruise attitude.

One of the most critical aspects of a small satellite is the electrical power system (EPS) since the electrical power is necessary for the satellite to operate correctly during its predefined lifetime. The electrical power system consists mainly of solar cells, batteries, voltage converters and protection circuits. The electrical power system is responsible of providing stable power to the rest of the satellite subsystems.

In satellite electrical power systems, overcurrent protection is now becoming an important function handled by the power distribution module (PDM). This paper proposes a method to evaluate the suggested protection. With the proposed procedure we should be able to verify that every possible failure does not travel through the EPS and cause a fatal degradation of the electrical power system. This will allow a complete evaluation of functionality of the protection hardware.

This paper discusses the design and implementation of the power distribution module (PDM) for the coming generation of small satellites for the Algerian Space Agency (ASAL). The design must provide a reliable protection for the subsystems from the overcurrent associated with a device failure.

This paper presents the design and optimisation of an aerofoil with active continuous trailing-edge flap (CTEF) investigated as a potential rotorcraft active control device. Several structural cross-section models are developed: high-fidelity NASA STRucture ANalysis (NASTRAN) and University of Michigan/Variational Asymptotic Beam Section Code (UM/VABS) models and a reduced-order analysis model. The validation of the reduced-order model is established by comparing its predictions of CTEF deformations with those of NASTRAN and UM/VABS analyses, which both show good agreement. The 2D aerodynamic characteristics of the CTEF aerofoil are evaluated using XFOIL and Computational Fluid Dynamics (CFD) analyses: FUN3D and Overset Transonic Unsteady Rotor Navier-Stokes (OVERTURNS). XFOIL, coupled with the reduced-order structure model, is adopted for optimisation study. The accuracy of XFOIL in predicting the aerodynamic pressure of the CTEF aerofoil is verified using CFD simulations, which shows sufficient fidelity. The predicted variations of aerodynamic coefficients with a CTEF angle are compared among the aerodynamic analyses. The optimisation process is developed and applied to two bimorph bender configurations: a Macro-Fibre Composite (MFC) solid bender and an MFC stack bender. The solid bender is used to confirm the functioning of the optimisation procedure and to use its optimal layout as a reference to the stack design, the primary design object. A linear tapered shape is found to be the optimum for a MFC solid bender, which generates an average of 63% more CTEF angles than those of an optimal rectangular bender. An optimised MFC stack bender is shown to resemble the shape of the solid bender. A four-ply bimorph is considered the best choice among the stack layouts because of its large output of CTEF angles and relatively less plies required. The CTEF angle produced by the four-ply optimal layout ranges from 7.6° to 5.3° with speeds from 0 to 200m/s at an angle of attack (AoA) of 6°. The reduction in the CTEF angle with AoA is less steep than that with speed, ranging from 6.5° to 5.8° with AoA from 0 to 8° at speed of 166m/s. An average of 14% increase in CTEF angles is achieved through optimisation for the four-ply bimorph.

A new method for system stability analysis, the weight functions method, is applied to estimate the longitudinal and lateral stability of a Hawker 800XP aircraft. This paper assesses the application of the weight functions method to a real aircraft and a method validation with an eigenvalues stability analysis of the linear small-perturbation equations. The method consists of finding the weight functions that are equal to the number of differential equations required for system modelling. The aircraft’s stability is determined from the sign of the total weight function – the sign should be negative for a stable model. Aerodynamic coefficients and stability derivatives of the mid-size twin-engine corporate aircraft Hawker 800XP are obtained using the in-house FDerivatives code, recently developed at our laboratory of applied research in active controls, avionics and aeroservoelasticity LARCASE. The results are validated with the flight test data supplied by CAE Inc. for all considered flight cases. This aircraft model was chosen because it was part of a research project for FDerivatives code and continued with weight function method for stability analysis in order to develop a design tool, based only on the aircraft geometrical parameters for subsonic regime. The following flight cases are considered: Mach numbers = 0·4 and 0·5, altitudes = 3,000m, 5,000m, 8,000m and 10,000m, and angles-of-attack α = –5° to 20°.

This paper presents an improved algorithm for foreign object debris (FOD) detection on the runway with several innovative techniques. The detection scheme incorporates four steps of geometric adjustment, background subtraction, clutter suppression and camouflage elimination. After geometric adjustment, the background model is built for each pixel with a set of RGB colour values taken in the past at the same location or in the neighborhood in the step of background subtraction. The background model samples are substituted randomly with an unfixed update period. Furthermore, the steps of clutter suppression and camouflage elimination are added to modify the segmentation map after background subtraction in order to increase the detection probability and decrease the false alarm rate. The overall algorithm is applied to the test data and real data on the runway. The results show that the RGB-based algorithm performs better than the classical gray-based techniques.

A new methodology is described to identify aircraft dynamics and extract the corresponding aerodynamic coefficients. The proposed approach makes use of fuzzy modelling for the identification process where input/output data are first classified by means of the concept of fuzzy clustering and then the linguistic rules are extracted from the fuzzy clusters. The fuzzy rule-based models are in the form of affine Takagi-Sugeno models, that are able to approximate a large class of nonlinear systems. A comparative study is performed with existing techniques based on the employment of neural networks, showing interesting advantages of the proposed methodology both for the physical insight of the identified model and the simplicity to obtain accurate results with fewer parameters to be properly tuned.

The present paper continues the work of Zhu et al. The closed-form expressions for the evaluation of forces on a body in compressible, viscous and rotational flow derived in the previous paper have been extended to different forms. The expressions require only a knowledge of the velocity field (and its derivatives) in a finite and arbitrarily chosen region enclosing the body. The equations are implemented on three-dimensional inviscid flows over wings and wing/body combinations. Further implementation on three-dimensional viscous flows over wings has also been investigated.

Targeting higher efficiencies and lower fuel consumption of turbomachines, heat transfer and profile loss are research topics of particular interest. In contrast to that, the interaction of both was, so far, rarely investigated, but gains in importance in recent research activities. The profile loss of engine components can be characterised by the airfoil wakes at the blade rows utilising established measurement and evaluation methods for which an adiabatic flow is typically supposed. To enable the investigation of the influence of heat transfer at the blade on the loss characteristics, a novel evaluation procedure was set up. In addition to the pneumatic data, the total temperature in the airfoil wake at a linear cascade was measured by means of a five-hole probe with an integrated thermocouple. For the evaluation and analysis of these data, different definitions of the loss coefficient were investigated and, finally, extended to account for thermal aspects. Furthermore, established techniques to average the local wake data were applied and compared with special focus to their suitability for non-adiabatic cases. Moreover, an extended version of the mixed-out average as defined by Amecke was utilised applying not only a far-reaching consideration of a temperature gradient but also the inclusion of the third spatial dimension to enable the evaluation of field traverses in addition to single wake traverses. These techniques were applied to wake measurement data from a linear compressor cascade gained in a special test set-up in the high-speed cascade wind tunnel for different operating points and different blade temperatures. The suitability of the new methods could be proven, and initial steps of the aerodynamic analysis of the resulting data are presented. Thereby, the acquired techniques turned out as powerful methods for the evaluation of wake traverses on compressor and turbine cascades under non-adiabatic conditions.

Attention is confined to civil aviation and the present scope of investigating procedures is outlined on a world-wide basis. Deficiencies in these procedures are noted and some suggestions are made for possible future developments of technique. In particular it is suggested that there can be little progress in air safety without the assimilation of “ human engineering data” by investigators and the Industry alike. Opinions expressed are entirely personal to the author and are not associated in any way with the official views of The British Aviation Insurance Co. Ltd.

Landing gear are exposed to cyclic loads from the ground manoeuvres that aircraft perform in-service. Variability is observed in the loading magnitude associated with ground manoeuvres, along with the per-flight variability in ground manoeuvre occurrence and sequencing. Whilst loading magnitude variability has been widely characterised, significant assumptions are required regarding manoeuvre occurrence and sequencing when constructing landing gear load spectra for fatigue design. These assumptions are required due to the limited availability of data concerning ground manoeuvre occurrence and sequencing relating to aircraft in-service and require validation to facilitate the design of more efficient components. ‘Big-Data’ approaches, employing Automatic Dependent Surveillance-Broadcast (ADS-B) transponder data, enable aircraft ground tracks to be identified. This paper presents a methodology to characterise the variability in ground manoeuvre occurrence and sequencing using ADS-B data sourced from Flightradar24® for a wide-body aircraft fleet. Using statistics generated for the fleet, it was identified that significant variability exists in the occurrence and sequencing of turning and braking manoeuvres. The statistics also validate existing assumptions, including that the proportional share of left and right turning manoeuvres is equal. Finally, this paper discusses the utility of ADS-B datasets for constructing landing gear load spectra and monitoring of landing gear in-service.

During the past few years there has been a surprising amount of discussion about the possibility of reviving airships. A number of schemes for new designs of rigid airship have been canvassed, all of them postulating important advances in lighter-than-air technology as compared with the ‘state-of-the-art’ at the time rigid airship development was abandoned just before the Second World War. Consideration of the factors responsible for the airship's failure may throw useful light on the validity of some of the assumptions made in recent proposals.

The phrase “Air-Cushioned Vehicle” describes the complete range of vehicles which obtain some or all of their support from a free pressurised cushion of air contained between the vehicle and the ground.

Vickers’ interest in air cushion vehicles stems directly from the basic work carried out on the Hovercraft—one type of air cushion vehicle—by Mr. C. S. Cockerell and the initiative and encouragement displayed by the N.R.D.C. through its subsidiary Hovercraft Development Ltd.

The problem of determining the distribution of stress over the various components of a structure such as an aircraft frame is one which has always presented considerable difficulty. For a large proportion of the components calculations of the stresses can only be carried out on the basis of assumptions and approximations, the effects of which cannot be estimated with any degree of accuracy and, since considerations of size and weight impose limitations upon the employment of a large safety factor, there has always existed a need for some practical means of checking the results of calculations or of obtaining more reliable estimates of the stresses.

During 1961 preliminary consideration was given by the Weapons Research Establishment to the possibility of measuring the structure of the upper atmosphere by pressure probes on sounding rockets. A similar technique had been used with success in the USA. It was proposed that velocity be measured and that an incidence meter be used to measure pitot and static pressures, from which atmospheric pressure and density could be calculated. There was available at the time comprehensive calibration information on hemispherical-headed incidence meters at Mach numbers below 3. Since the sounding rocket would be flying within the Mach number range 3 to 6 there was a need to extend the existing calibrations to higher Mach numbers. At first it seemed that the results of Baer would do this, but a closer examination showed that his results did not agree with the incidence meter calibrations.