This work presents a study of acoustic oscillations generated and the wave structure associated with supersonic flow past wall mounted 3D open cavities of varying length-to-width (L/W) ratio. Experiments were conducted to investigate the acoustic signature generated by the cavities at freestream Mach number of 1·7. The effect of L/W ratio of the cavity on the dominant modes of the acoustic signature registered on different walls of the cavities is investigated for an L/W range of 0·83-4. Shift in the dominant acoustic mode is observed as L/W ratio changes from 3 to 4. Statistical analysis of pressure data showed existence of acoustic waves and spreading of acoustic energy over different modes with change in cavity width. Time averaged schlieren visualisation indicated variation of shock and shear layer structure in the mainstream for the different cavities. Acoustic waves generated by the presence of the cavity and the dynamic behaviour of the shear layer were observed during instantaneous shadowgraph visualisation. Numerical simulation was done to make a prior assessment of the flow structure and the results are in good agreement with those from experiments. Ratio of mass exchange between cavity and mainflow and the cavity volume was observed to have profound effect on the magnitude of pressure oscillations generated by the cavities.

This paper describes and presents results from static wind tunnel tests conducted on a 60° delta wing at a root chord Reynolds number of 2·7 x 106. In these tests, the wing was instrumented with 192 miniature pressure transducers which, in conjunction with a powerful multi-channel data-logging system, allowed the distribution of time-varying surface pressures to be measured at high temporal resolution. Analysis indicates that the distribution of root mean square pressure on the leeward surface of the wing can provide considerable insight into the behaviour of both the primary and secondary vortex structures. In addition, it has been established that the frequency content of pressure signals measured in the vicinity of these vortex structures is sensitive to the vortex state. It is suggested that these data features can be directly attributed to previously observed behavioral characteristics of the vortex breakdown process.

This note assesses the benefits of active aeroelastic structures (AAS) in enhancing flight performance and control authority. A representative AAS concept, whose torsional stiffness and shear centre position can be altered depending on the instantaneous flight condition, is employed in the wing of a medium altitude long endurance (MALE) UAV. A multidisciplinary design optimisation (MDO) suite is used in this study. It turns out that AAS can be very effective when used for enhancing control authority of the vehicle but have limited benefits in terms of flight performance (lift to drag).

Stratospheric airships are lighter-than-air vehicles that work at an altitude of 20km in the lower calm portion of the stratosphere. They can be used as real-time surveillance platforms for environment monitoring and civil communication. Solar energy is the ideal power choice for long-endurance stratospheric airships. Attitude control is important for airships so that they can point at a target for observation or adjust the attitude to improve the output performance of solar panels. Stratospheric airships have a large volume and semi-flexible structure. The typical actuators used are aerodynamic surfaces, vectored thrust and ballonets. However, not all these actuators can work well under special working conditions, such as low density and low speed. In this study, moving-mass control is introduced to stratospheric airships because its control efficiency is independent of airspeed and atmospheric density. A nonlinear feedback controller based on generalised inverse with a nonlinear mapping module is designed to implement moving-mass control. Such a new station keeping scheme with moving masses is proposed for airships with different working situations.

Flying a light aeroplane involves a combination of pilot and aeroplane performing a set task, within a specific environment. The pilot is continuously sampling and selecting available sensory cues, interpreting those cues, making decisions and manipulating the primary controls (stick and rudder) to safely achieve flying objectives. The ‘feel’ of an aeroplane (a flying quality) is directly associated with the stick and rudder forces and how the aeroplane responds to control inputs. Classical theory has been applied to estimate the apparent (as felt by the pilot) longitudinal stick-free static stability (change of stick force with airspeed) of a typical, two-seat, high-wing light aeroplane. The theory has been extended to consider the effects of tail downwash and flap deflection. The results are compared with actual flight tests and show that the method may be used for the initial assessment of longitudinal stick-free static stability and more importantly, tendencies towards neutral or negative stability affecting flight safety.

Modern low-altitude unmanned aircraft (UA) detection and surveillance systems mostly adopt the multi-sensor fusion technology scheme of radar, visible light, infrared, acoustic and radio detection. Firstly, this paper summarises the latest research progress of UA and bird target detection and recognition technology based on radar, and provides an effective way of detection and recognition from the aspects of echo modeling and micro motion characteristic cognition, manoeuver feature enhancement and extraction, motion trajectory difference, deep learning intelligent classification, etc. Furthermore, this paper also analyses the target feature extraction and recognition algorithms represented by deep learning for other kinds of sensor data. Finally, after a comparison of the detection ability of various detection technologies, a technical scheme for low-altitude UA surveillance system based on four types of sensors is proposed, with a detailed description of its main performance indicators.

Global air traffic demand has shown rapid growth for the last three decades. This growth led to more delays and congestion within terminal manoeuvring areas (TMAs) around major airports. The efficient use of airport capacities through the careful planning of air traffic flows is imperative to overcome these problems. In this study, a mixed-integer nonlinear programming (MINLP) model with a multi-objective approach was developed to solve the aircraft sequencing and scheduling problem for mixed runway operations within the TMAs. The model contains fuel cost functions based on airspeed, altitude, bank angle, and the aerodynamic characteristics of the aircraft. The optimisation problem was solved by using the $\varepsilon$-constraint method where total delay and total fuel functions were simultaneously optimised. We tested the model with different scenarios generated based on the real traffic data of Istanbul Sabiha Gökçen Airport. The results revealed that the average total delay and average total fuel were reduced by 26.4% and 6.7%, respectively.

Aircraft refuelling is a major cause of flight delays because it is a slow process. Further, if it does not begin as soon as the aircraft is available for ground handlers, there is an increasing risk of it being terminated after the final passenger has boarded. Usually, the process only begins after information regarding the required quantity of fuel is passed through the flight dispatcher, and this information typically requires a certain time to reach the ground handlers. Therefore, it is intended to test a new scenario: to begin refuelling with a minimum level and, if necessary, fill up the remainder with the final fuel figures when received. The aim of this paper is to analyse the application of Six Sigma in this process through Student’s t-test and statistical process control. The collected data in this case study include the amount of fuel supplied and flight delays (which are mainly caused by refuelling). The results demonstrate that the new process is favourable, and that the average length of flight delays is reduced from 14 to 6 min, which is an improvement of 57%. It is concluded that the application of Six Sigma in the aircraft refuelling process saves time and improves on-time performance levels, which is relevant to the scientific literature, thereby aiding in mitigating the risk of fines and penalties.

Accurate prediction of supersonic and hypersonic turbulent flows is essential to the design of high-speed aerospace vehicles. Such flows are mainly predicted using the Reynolds-Averaged Navier–Stokes (RANS) approach in general, and in particular turbulence models using the effective viscosity approximation. Several terms involving the turbulent kinetic energy (k) appear explicitly in the RANS equations through the modelling of the Reynolds stresses in such approach, and similar terms appear in the mean total energy equation. Some of these terms are often ignored in low, or even supersonic, speed simulations with zero-equation models, as well as some one- or two-equation models. The omission of these terms may not be appropriate under hypersonic conditions. Nevertheless, this is a widespread practice, even for very high-speed turbulent flow simulations, because many software packages still make such approximations. To quantify the impact of ignoring these terms in the RANS equations, two linear two-equation models and one non-linear two-equation model are applied to the computation of five supersonic and hypersonic benchmark cases, one 2D zero-pressure gradient hypersonic flat plate case and four shock wave boundary layer interaction (SWBLI) cases. The surface friction coefficients and velocity profiles predicted with different combinations of the turbulent kinetic energy terms present in the transport equations show little sensitivity to the presence of these terms in the zero-pressure gradient case. In the SWBLI cases, however, comparisons show that inclusion of k in the mean flow equations can have a strong effect on the prediction of flow separation. Therefore, it is highly recommended to include all the turbulent kinetic energy terms in the mean flow equations when dealing with simulations of supersonic and hypersonic turbulent flows, especially for flows with SWBLIs. As a further consequence, since k may not be obtained explicitly in zero-equation, or certain one-equation, models, it is debatable whether these models are suitable for simulations of supersonic and hypersonic turbulent flows with SWBLIs.

A method is described whereby, at any point in an infinite parallel annulus, the approximate axial velocity due to a single row of high aspect ratio blades may be calculated from a knowledge of the conditions of flow adjacent to the blades. The method is based on the assumption of a simplified expression for the radial velocity, being the product of an unknown function of the radius and an exponential term independent of the radius containing an undetermined constant; the function and the undetermined constant are calculated by reference to the conditions of flow in the plane of the row considered. The flow due to any number of rows is then obtained by summing the radial velocity fields due to each row and obtaining the axial velocities by integration of the equation of continuity.

The solution of the problem with infinitely many rows is shown to have a simple form by virtue of the fact that the flow (provided that the velocities remain finite) settles down to a pattern which is periodic by one stage pitch.

It has been found that the addition of a whirl velocity component to an otherwise axial flow in a conical diffuser can lead to improvements in static pressure recovery and in efficiency, according to a restricted definition. Diffusers of total conical angle 10, 20 and 30 deg. were tested in detail; improvements were evident only in the last two, which would otherwise have suffered from flow separation problems caused by their relatively large divergence angles. The agency for improvement seems to be a beneficial redistribution of axial velocities.

The various methods which axe available for controlling the rolling motions of a model suspended in an electro-magnet system are discussed. Data on the moment capability and damping achieved with some of the possible control techniques are presented, together with details of the roll control system adopted for the six-component magnetic balance which has been developed at Southampton University for an 8 in × 6 in supersonic wind tunnel.

The Chairman, in introducing the Author, described the subject of his paper as one of paramount importance to all who were concerned in the design and operation of helicopters Mr Rogers was an indentured aeronautical engineer apprentice at the Fairey Aviation Company from 1940 to 1946, during which time he obtained the Ordinary and Higher National Certificates in Mechanical and Aeronautical Engineering He was awarded an S B A C scholarship to the College of Aeronautics in 1946 and obtained a diploma with distinction in Aircraft Design

A method is developed for the calculation of the aerodynamic forces acting on a “swallow tail” wing (Figs. l(a) and l(b)) of small aspect ratio. For a given incidence, lift and induced drag coefficient are, within the limits of the theory, proportional to aspect ratio and independent of Mach number. The exact solution is derived for the limiting case of small sweepback of the trailing edge. A numerical method is indicated for the more general case.

A theory is developed for the prediction of the pressure at the reattachment point in subsonic two-dimensional steady base flow. Theoretically estimated values of the reattachment pressure seem to agree well with experimental ones.

With the aviation industry facing increasing environmental and energy challenges, there has been a growing demand for sustainable aviation fuel (SAF). Previous studies have shown the role of ice accretion, release and blockage in aviation-related incidents and accidents with conventional jet fuel. However, there is no available data that establishes the magnitude of influence new fuel compositions will pose on ice formation and accretion in aircraft fuel systems. A recirculating fuel test rig capable of cooling fuel from ambient to −30°C within 4h was built by Airbus to simulate conditions in an aircraft wing tank and allow characterisation of ice accretion. The key characteristic was the pressure drop across an inline fuel strainer for the different SAF explored but visual analysis of ice accretion on the strainer mesh (filters used in protecting fuel feed pumps) was also performed for individual experimental runs for comparison. Measurements revealed that 100% conventional fuel exhibited a higher propensity to strainer blockage compared to the SAF tested. However, all SAF blends behaved differently as the blending ratio with Jet A-1 fuel had an impact on the pressure differential at different temperatures. Data from this work are essential to establish confidence in the safe operation of future aircraft fuel systems that will potentially be compatible with 100 % SAF.

An aircraft and landing gear loads model was developed to assess the Margin of Safety (MS) in main landing gear components such as the main fitting, sliding tube and shock absorber upper diaphragm tube. Using a technique of Bayesian sensitivity analysis, a number of flight parameters were varied in the aircraft and landing gear loads model to gain an understanding of the sensitivity of the MS of the main landing gear components to the individual flight parameters in symmetric two-point landings. The significant flight parameters to the main fitting MS, sliding tube bending moment MS and shock absorber upper diaphragm tube MS include: longitudinal tyre-runway friction coefficient, aircraft vertical descent velocity, aircraft Euler pitch angle and aircraft mass. It was also shown that shock absorber servicing state and tyre pressure do not contribute significantly to the MS.

A method for the generation of assumed modes in beams with arbitrary boundary conditions is described. The method is applied to two cantilevered beams using a modified Rayleigh-Ritz vibration analysis. The results obtained show excellent agreement with those obtained by exact methods of analysis.

Military aircrafts are often subjected to severe flight maneuvers with high Angles-of -Attack (AOA) and Angles of Sideslip (AOSS). These flight attitudes induce non-uniform in flow conditions to their gas turbine engines which may include distortion of inlet total pressure and total temperature at the Aerodynamic Interface Plane (AIP). Operation of the downstream compression system with distorted inflow typically results in reduced aerodynamic performance, reduced stall margin, and increased blade stress levels. In the present study the steady state total pressure distortion induced to the Aerodynamic Interface Plane due to the aircraft’s flight attitude have been estimated in terms of distortion descriptors. The distorted conditions at the interface between the intake and the engine have been predicted by using Computational Fluid Dynamics (CFD), where 33 different aircraft flight attitudes have been tested. Based on the obtained results the effect of Angle-of-Attack (AOA) and Angle of Side Slip (AOSS) on the distortion descriptors have been studied. The results showed that the distortion effect becomes more pronounced whenever this specific airframe configuration is exposed to incoming flow with an AOSS. Among the tested cases, the greatest total pressure defect at the AIP in terms of difference from the average value and of circumferential extent was calculated for the flight attitudes of 0·35M flight with 0° AOA and 8° AOSS and 0·35M fight with 16° AOA and 16° AOSS.

The paper describes in detail an older method than Relaxation of approximating to the solution of equations of the Laplace and Poisson type. The corresponding fourth order equations are discussed briefly, and a method of dealing with certain non-linear equations is indicated. A description is also given of the propagation of errors in the fields due to various causes.