Estimating the wing structural weight of an extremely manoeuvrable Unmanned Aerial Vehicle (UAV) during conceptual design has proven to be a significant challenge due to its high load factor (the ratio of an aircraft lift to its weight). The traditional empirical method relies on existing statistical data of previously built aircraft, then is inadequate for the innovative UAV structure design which can endure extremely manoeuvrable load (load factor is greater than 9g). In this paper, the finite element model for wing structure of an extremely manoeuvrable UAV with foreplane was built, and the structural weight was estimated by static aeroelastic optimisation considering structural strength and buckling constraints. The methodology developed here is only consisted of three components, which is much less than that for existing method, thus the procedure developed here sacrificed some accuracy, but it’s faster and more suitable for aircraft conceptual design. It was validated by the overlap between the weights given by the methodology, and the results from empirical equations when the load factors are less than 9g. Through the analysis procedure developed, the wing structural weights of the extremely manoeuvrable UAV were given under different load cases (load factor changes from 5g to 12g).

A numerical study of the flow topologies over the 60° delta wing of the AGARD-B model at Mach 0·80 has revealed that vortex bursting occurs between 13°-15° angle-of-attack, while vortex separation occurs above 18°. These aerodynamic features have been identified as additional comparison criteria which need to be replicated for facilities using the model for calibration or inter-tunnel comparison purposes. The numerical simulations were performed using ANSYS Fluent V13, a structured mesh with near wall treatment and the Spalart-Allmaras and κ-ω SST turbulence models, and validated experimentally in a 5′ × 5′ transonic facility. Other aspects not previously identified or studied are firstly a recovery shock between the primary and secondary vortex that exists only when vortex bursting occurs, and secondly the lack of a shock between the wing and vortex when the flow topology corresponds to the centreline shock region as observed in other studies.

The design and building of accurately scaled aeroelastic wind-tunnel models is difficult, time consuming and very costly. With the increasing usefulness of computational methods for predicting aeroelastic phenomena, few complex models have been built in recent years. New fighter aircraft projects are also becoming more and more scarce, and transport aircraft have essentially the same configuration since half a decade. This also significantly reduces the need for aeroelastic wind-tunnel models. However, there still is an interest in the results from aeroelastic testing. In some cases new and radical configurations may warrant wind-tunnel testing and in other cases complex phenomena arising in fight testing may need carefully performed experiments to resolve problems. However, there is definitely a trend towards building models and performing testing in the support of the development of computational methods.

The developments in computer technology do not only improve the computational methods for aeroelasticity. Modern Computer Aided Design and Manufacturing techniques can significantly improve the quality and efficiency of the design and build process for aeroelastic models. There have also been some recent improvements in measurement techniques which have proven very useful for testing of aeroelastic wind-tunnel models. The paper will present some new design and build techniques developed for the manufacturing of a large scale wind-tunnel model of a canard delta wing fighter aircraft configuration. In the build process fiber-reinforced composites will be used, hence, challenges and possible solutions concerning the ability to produce a model with well defined material properties and fiber angles will be discussed. Further challenges arise when both measurement equipment and adjustable control surfaces should be attached inside the model using techniques that are possible to describe with computational methods. In addition, equipment, such as pressure taps, and control surface mechanics need to fit and function in a flexible structure. As a result, the above requirements will lead to necessary compromises in the design, hence, the paper will present the choices taken during the build process and for which reasons. The use of an optical positioning measurement system will also be discussed for both the validation of model properties and non-contact measurement of model deformations during wind-tunnel testing.

In this article, the effect of on-surface flow visualisation (SVF) techniques on measured loads over an airfoil are explored. Titanium dioxide based mixture effects on the lift and drag coefficient are experimentally quantified at low Reynolds numbers by recording the time history as the patterns evolve and freeze. With statistical comparison based on Student’s t-distribution method, it was determined that the effect on the drag coefficient was minimal but the lift coefficient was slightly attenuated. Additionally, it was observed that at high angles-of-attack the temporal history of the flow as the wind tunnel ramps up may alter the steady-state flow field in the presence of a SFV mixture.

Estimation of the lift of an aerofoil is one of the fundamental measurements of fluid mechanics. Lift is commonly measured using a load cell or a force balance. Non-intrusive methods to measure lift are usually pressure based. Aerofoils may be pressure tapped where small surface orifices are connected via tubing to a pressure measurement system, either a multi-tube manometre or an electronic system. Both measurement options add cost and complication, especially in an educational setting. Pressure tapping small aerofoils can also be difficult, especially if the models are rapid prototyped (RP). Low model surface resolution (from RP manufacture) and confined geometry complicate model assembly and finishing. Boundary-layer transition caused by poorly implemented tappings (too large a diametre or poorly aligned, i.e. straight aft) can also alter results. Wall pressure tappings may also be used and have the benefit of being non-intrusive. To implement, the test section roof and floor is tapped with a streamwise row of ports that facilitate measurement of the wall pressure signature. Integration of the pressure differential then relates to the lift produced. This measurement methodology still requires a multi-channel pressure acquisition system and modification of the wind tunnel. In Refs 4,5 methods are presented that facilitate calculation of the instantaneous forces acting on a body through flow field measurements determined using particle image velocimetry. However, the required flow field measurements encompass those surrounding the body, and are not a simple point measurement. In Ref. 6 a method is presented to estimate the lift of an aerofoil using two Pitot-static tubes that are used to measure the velocity above and below the aerofoil’s quarter chord. Wall corrections are required to yield an accurate lift estimate.

The impact point dispersion of mortar missiles can be drastically reduced with a control ring of lateral pulse jets located around the centre of gravity and a homing head to measure or to derive estimates of appropriate missile-target states, e.g. the line-of-sight rates. A simulation model including a missile as a six-degree-of-freedom vehicle, a jet pulse flight control system, and a relative missile-to-target motion was used for the comparison of four guidance laws i.e. proportional navigation guidance, augmented proportional navigation, augmented proportional navigation with rendezvous, and adaptive sliding-mode guidance. This paper focuses on the efficiency of pulse jet control on miss distance, and thus makes the assumption that sensor measurements and the guidance states required to apply each of the guidance laws are perfectly known. Proportional navigation and the adaptive sliding mode guidance exhibit a large miss distance due to limited control authority. Augmented proportional navigation is slightly better than augmented proportional navigation with rendezvous for the same design parameters and they both give small miss distances with limited control authority, but they both require a free gyro. A proper selection of the design parameters — the number of pulse jets and the magnitude of the individual pulse jet thrust for a particular dispersion of flight parameters and the instant of the guidance start — is required to achieve optimum dispersion reduction. The minimum intensity of the individual pulse jet impulse required for the ‘pin point’ accuracy (the circular error probable smaller than 1m) of the mortar missile was determined for all presented guidance laws considering dispersion from the nominal trajectory.

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.

This paper demonstrates the application of an integrated rotorcraft multidisciplinary design and optimisation framework, deployed for the purpose of preliminary design and assessment of optimum regenerative powerplant configurations for rotorcraft applications. The proposed approach comprises a wide-range of individual modelling theories applicable to rotorcraft flight dynamics, gas turbine engine performance and weight estimation as well as a physics-based stirred reactor model, for the rapid estimation of various gas turbine gaseous emissions. A single-objective Particle Swarm Optimiser is coupled with the aforementioned rotorcraft design framework. The overall methodology is deployed for the design and optimisation of a reference multipurpose Twin-Engine-Light civil rotorcraft, modelled after the Bo105 helicopter, which employs two Rolls-Royce Allison 250-C20B turboshaft engines. Through the implementation of a single-objective optimisation strategy, notionally based optimum engine design configurations are acquired in terms of engine weight, mission fuel burn and mission gaseous emissions inventory at constant technology level.

The acquired optimum regenerative engine configurations are subsequently deployed for the design of conceptual rotorcraft regenerative engine configurations, targeting improved mission fuel economy, enhanced payload-range capability as well as overall environmental impact, while maintaining the respective rotorcraft airworthiness requirements. The proposed methodology essentially constitutes as an enabler for designing rotorcraft powerplants within realistic, three-dimensional operations and towards realising their associated design trade-offs at mission level.

Numerical investigations are performed to study the features of symmetry breaking phenomena in a symmetric supersonic combustor. By changing the fuel equivalence ratio, the symmetry breaking phenomena are numerically observed. Further, the hysteresis behaviours are observed along with the phenomena of symmetry breaking by increasing and then decreasing the the fuel equivalence ratio. Furthermore, more complex hysteresis phenomenon, i.e., multi-loop hysteresis, is observed numerically. The multi-loop hysteresis consists of both major and minor loops, and is associated with the interaction between multiple flow structures. Due to the fact that different flow structures would result in different flow conditions, the findings regarding the existence of hysteresis behaviours can significantly affect the combustion process and even the performance of the engine. It is hoped that the observation of the (multi-loop) hysteresis phenomenon may bring it to the attention of those investigators whose experiments maybe affected by its occurrence.

This paper focuses on maintenance test flying as pertaining to commercial jet aircraft. Aircraft maintenance manuals or regulatory prescriptions require aircraft to be test flown prior to being released to service, following specific maintenance, repair, overhaul or modification events. Further, such flights might be conducted to accept or return aircraft as demonstration of their serviceability. Similarly, test flights to verify aircraft performance are at times required.

Such maintenance test flights are mostly conducted by pilots rated to fly those aircraft primarily in commercial line operations, typically with no or little training on such specific maintenance or performance test flight procedures or techniques. The international regulatory environment remains in flux with discussion ongoing.

In conducting such flights a pilot is exposed to activities outside the normal aircraft standard operating procedures up to the edge of the operational flight envelope. Whilst not intentional, abnormal aircraft behaviour can nevertheless result in inadvertent flight outside the envelope with a consequential potential loss of control. History has shown that the predominant cause of fatal accidents during maintenance test flying result from complex loss of control scenarios not recovered or not recoverable. This raises the question of adequacy of pilot training.

Maintenance test flights are a necessary component for the industry to maintain its exceptional safety standards and minimising loss of life. But such flights in themselves remain demonstrably one or two orders of magnitude more risky than commercial flights.

As concern for the environment increases, so does the desire to reduce emissions and improve fuel efficiency. One avenue being investigated in the aerospace industry is the reduction in losses in a gas turbine engine through endwall contouring, which has shown some promise in improving efficiency. The current investigation was aimed at investigating the unsteady effects downstream of a rotor with a non-axisymmetric endwall contour originally designed for a cascade.

Previous investigations (numerical and experimental) have shown that the endwall contour improved the turbine efficiency by reducing the cross passage secondary flows. The contour however had a detrimental effect on the tip flow. The current investigation found that there were negligible unsteady differences between the annular and contoured rotor, with the average comparing well to the steady state results. It was also found that the effects at the tip were due to the experimental blades untwisting during operation.

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.

One of the aerospace design engineer’s goals aims to reduce drag for increased aircraft performance, in terms of range, endurance, or speed in the various flight regimes. To accomplish this, the designer must have rapid and accurate techniques for computing drag. At subsonic Mach numbers drag is primarily a sum of lift-induced drag and zero-lift drag. While lift-induced drag is easily and efficiently determined by a far field method, using the Trefftz plane analysis, the same cannot be said of zero-lift drag. Zero-lift drag (CD,0) usually requires consideration of the Navier-Stokes equations, the solution of which is as yet unknown except by using approximate numerical techniques with computational fluid dynamics (CFD). The approximate calculation of zero-lift drag from CFD is normally computed with so-called near-field techniques, which can be inaccurate and too time consuming for consideration in the design environment. This paper presents a technique to calculate zero-lift and boundary-layer drag in the subsonic regime that includes aeroelastic effects and is suitable for the design environment. The technique loosely couples a two-dimensional aerofoil boundary-layer model with a 3D aeroelastic solver to compute zero-lift drag. We show results for a rectangular wing (baseline), a swept wing, and a tapered wing. Then compare with a rectangular wing with variable thickness and camber, thinning out from the root to tip (spanwise direction), thus demonstrating the practicality of the technique and its utility for rapid conceptual design.