For aircraft that employ distributed propeller propulsion systems, the distributed propeller slipstream increases the analysis complexity. The objective of this paper is to rapidly analyse the influence of propeller slipstream on a wing using a fast prediction approach to perform conceptual design studies. This fast approach is implemented through a panel/viscous vortex particle hybrid method taking into account the air viscosity effect. The parametric studies of propeller streamwise, spanwise, vertical installed position, propeller number and rotational direction are conducted for a rectangular wing platform in two different propeller-wing configurations. The results indicate that the propeller slipstream causes both the augmentations of the wing lift and drag in a traditional tractor propeller layout. For an over-the-wing propeller configuration, however, the obvious lift increase and drag decrease can be obtained. A rear propeller position relative to the wing chord leads to a beneficial increase in lift while a fore propeller location is able to decrease the wing drag. The maximum increment of the lift-to-drag ratio can be achieved by 17.6% when the propeller is located at 30% of the wing chord, which shows a considerable advantage in improving the wing aerodynamic efficiency.

A novel statistical model is presented to quantify situation awareness in the operation of small civilian Unmanned Aircraft Systems (UAS). Today, the vast majority of small Unmanned Aircraft Systems (UAS) operation takes place under Visual Line of Sight (VLOS) of a human operator, who is wholly responsible for the safety of the flight. As operation begins to move to Beyond Visual Line of Sight (BVLOS), it is likely that this responsibility will become shared between operator and the increasingly autonomous UAS itself. Before we seek to quantify the safety of such a system, it is beneficial to analyse the safety of existing Visual Line of Sight (VLOS) operations to provide a target level of safety. Prior to considering any on-board decision making, it is essential to ensure that the artificial situation awareness system of a UAS in Beyond Visual Line of Sight (BVLOS) is at least as good as awareness of a human operator. The paper provides a probabilistic theory and model for the high-level abstractions of situation awareness to guide future assessment of BVLOS operations.

Extensive experiments were conducted to study the effect of various parameters on the surface pressure distribution and transition point of an aerofoil section used in a wind turbine blade. In this paper details of the variation of transition point on the aforementioned aerofoil are presented. The aerofoil spanned the wind-tunnel test section and was oscillated sinusoidally in pitch about the quarter chord. The imposed variables of the experiments were free stream velocity, amplitude of motion, mean angle-of-attack, and oscillation frequency.

The spatial-temporal progressions of the leading-edge transition point and the state of the unsteady boundary-layer were measured using eight closely-spaced, hot-film sensors (HFS). The measurements show that:

(i) Reduced frequency has a pronounced effect on the variations of the transition point.

(ii) There exists a hysteresis loop in the dynamic transition location and its shape varies with the reduced frequency and mean angle-of-attack.

The three-dimensional equations of motion of a flexible towed cable are developed. A closed-form approximation for the equilibrium shape of a cable subjected to arbitrary aerodynamic loading is developed and used in the study of a heavy, vibrating tow cable. Natural frequencies of vibration and cable shapes are computed for typical cables and are shown to be in good agreement with exact, numerically obtained values.

After two years research work in the Universities of this country and elsewhere, there is need for a comprehensive review of the present position on the suppression of aerodynamic noise from jet engines. This report outlines the fundamental understanding of the problem that has been achieved, the basic experimental work done, and the noise suppression techniques which have been suggested.

A study is made of measured noise levels on advanced types of engines and an analysis is made of reductions required and so far achieved with the proposed noise suppression devices. While the basic principles underlying noise formation, both in subsonic and supersonic jet streams are beginning to be understood, the noise reductions so far achieved are still insufficient and warrant an extended programme of research.

The report includes helicopter noise analyses with pressure and pulse jets, and touches on the problem of structural fatigue in the vicinity of a jet stream.

In most one-parameter methods of calculating laminar boundary layers it is assumed that the non-dimensional parameters H=δ1/δ2, I=δ2τW/μu1 and L = 2{I−λ(H+2)}, depend only upon the pressure gradient parameter λ=u1δ22/v. In this paper it is shown theoretically that a more accurate, two-parameter representation is

L=F0(λ)−μG0(λ)

I2=F1(λ)−μG1(λ),

where μ=λ2U1U1/(U1)2. Careful examination of the available range of exact solutions of the boundary layer equations has enabled the four functions F0, G0, F1, G1, to be tabulated, and the above functional forms agree with the exact solutions to a remarkable accuracy.

In view of the fact that a reasonable first approximation to L is usually , we write

,

and it is then shown that the momentum integral equation becomes

This equation is easily solved by iteration, setting g=0 in the first approximation, and convergence is extremely rapid.

The method is, in effect, a refinement of that due to Thwaites, which is universally accepted as one of the better of the existing calculation methods. Detailed calculations made by the present method indicate that the errors are only 5% of those given by the Thwaites method.

Heat transfer levels have been investigated behind a rearward-facing step in a superorbital expansion tube. The heat transfer was measured along a flat plate and behind 2 and 3mm steps with the same length to step height ratio. Results were obtained with air as the test gas at speeds of 6·76kms-1 and 9·60kms-1 corresponding to stagnation enthalpies of 26MJ/kg and 48MJ/kg respectively. A laminar boundary layer was established on the flat plate and measured heat transfer levels were consistent with classical empirical correlations. In the case of flow behind a step, the measurements showed a gradual rise in heat transfer from the rear of the step to a plateau several step heights downstream for both flow conditions. Reattachment distance was estimated to be approximately 1·6 step heights downstream of the 2mm step at the low enthalpy condition through the use of flow visualisation.

Earlier papers described a method of calculating three-dimensional turbulent boundary layers based on the use of momentum-integral equations in the streamwise and cross-flow directions. Here the method is applied to a problem which is initially formulated in a coordinate system appropriate to the somewhat complex body geometry. Transformation to a streamline coordinate system is then made before the application of a rapidly converging iterative method of solution. The calculations, which are confined to single Reynolds number and a particular value of the rotation parameter, show the very large increases in drag and torque that accompany early transition.

The high enthalpy and high Mach number effects on flat plate boundary layer flow are examined. In particular, the behaviour of the density profile under high enthalpy, real gas conditions is considered. Mach-Zehnder interferograms of the flow over a flat plate are analysed using a two-dimensional Fourier transform procedure in order to obtain the density field. The comparison between theory and experiment is only fair. It must be pointed out, however, that the real gas effects on the density profiles were found to be small. Estimation of the velocity boundary layer thickness as obtained from a measurement of the thermal boundary layer thickness is seen to compare well with a theoretical prediction.

The factors influencing survival in aircraft accidents can be classified into configurational, environmental, procedural or behavioural. New methodologies which have been developed in order to simulate emergency behaviour are introduced. A series of experimental programmes in which the new behavioural techniques are applied to configurational, environmental and procedural changes in the cabin are described. The results suggest that the techniques provide the behavioural and statistical data required for the assessment of design options or safety procedures for use in aircraft emergencies.

The linear-quadratic-Gaussian (LQG) control synthesis has the advantage of dealing with the uncertain linear systems disturbed by additive white Gaussian noise while having incomplete system state information available for control-loop feedback. This paper hence explores the feasibility of designing and implementing a stability augmentation autopilot for fixed-wing unmanned air vehicles using the LQG approach. The autopilot is composed of two independently designed LQG controllers which control the longitudinal and lateral motions of the aircraft respectively. The corresponding linear models are obtained through a system identification routine which makes use of the combination of two well-established identification methods, namely the subspace method and prediction error method. The two identification methods complement each other well and this paper shows that the proposed system identification scheme is capable of attaining satisfactory state-space models. A complete autopilot design procedure is devised and it is shown that the design process is simple and effective. Resulting longitudinal and lateral controllers are successfully verified in computer simulations and actual flight tests. The flight test results are presented in the paper and they are found to be consistent with the simulation results.

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.

The next generation of civil large aero-engines will employ greater bypass ratios compared with contemporary architectures. This results in higher exchange rates between exhaust performance and specific fuel consumption (SFC). Concurrently, the aerodynamic design of the exhaust is expected to play a key role in the success of future turbofans. This paper presents the development of a computational framework for the aerodynamic design of separate-jet exhaust systems for civil aero-engines. A mathematical approach is synthesised based on class-shape transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts and nozzles. This geometry formulation is coupled with an automated viscous and compressible flow solution method and a cost-effective design space exploration (DSE) approach. The framework is deployed to optimise the performance of a separate-jet exhaust for very-high-bypass ratio (VHBR) turbofan engine. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared with a baseline architecture, through optimum exhaust re-design. The proposed method is able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.

Results from numerical simulations of the three dimensional flow in the intake-isolator of a dual mode scramjet are presented. The FANS calculations have utilised the SST k -ω turbulence model. The effect of cowl length and cowl convergence angle on the inlet mass capture ratio, flow distortion, shock strength and pressure rise are studied in detail. Three cowl lengths and four or five cowl convergence angles for each cowl length are considered. The predicted values of the dimensionless wall static pressure and inlet mass capture ratio are compared with experimental data reported in the literature. The numerical predictions are shown to agree well with the experimental data. In addition, details of the flow field such as shocks, expansion fans and shock boundary interaction are also captured accurately. Inlet unstart is also demonstrated for one case.

From necessity, military aircraft often operate in a highly fatigue damaging environment and history has shown in lost lives and aircraft the consequences of failure to appreciate fully the usage environment. The need for robust and cost effective structural usage monitoring of military aircraft to ensure operations are conducted within acceptable levels of risk is paramount. Furthermore, increased economic pressures require ever-inventive methods to be employed to maximise the lives of military fleets; structural usage monitoring will be a key asset in this drive. A highly cost effective indirect structural health and usage neural network (SHAUNN) monitoring system is proposed. A SHAUNN uses regression relationships determined by artificial neural networks to predict stresses, strains, loads, or fatigue damage from flight parameters. Within this paper the development of a SHAUNN monitoring system is described. Flight parametric data, captured during Operational Loads Measurement of the Royal Air Force Dominie TMk1 aircraft have been used to predict stresses at the key structural location in the wing, using mapping relationships determined by artificial neural networks. A framework for the development of the SHAUNN monitoring system is discussed and the basic architecture of the multilayer perceptron artificial neural network is described. It is concluded that this technology could provide the basis for an accurate, cost-effective structural usage monitoring system and further work to investigate the prediction of ground – based stresses in the wing is recommended.

The use of a submerged inlet is advantageous in modern aircrafts because of its low drag resistance, small radar cross section and ease of maintenance. Although it is well known that the forebody boundary layer deteriorates the aerodynamic performance of a submerged inlet, the level of impact has not yet been fully quantified. To quantify the forebody boundary-layer effect, a submerged diverter was designed to remove a portion of the low-energy boundary flow. The flow pattern and aerodynamic performance of a submerged inlet, with and without the diverter, were investigated by wind-tunnel experimentation and numerical simulations. The effects of mass flow, free stream speed, angle-of-attack and sideslip angle on the aerodynamic characteristics of the inlet with and without the submerged diverter were studied, over an operating envelope of M0 = 0.3 ∼ 0.6, $\alpha$ = –6$^{\circ}$ ∼ 8$^{\circ}$ and $\beta$ = 0$^{\circ}$ ∼ 4$^{\circ}$. The results indicate that both the total pressure loss and the circumferential distortion can be significantly reduced with the removal of the forebody boundary-layer low-energy flow. Meanwhile, the main mechanisms for losses in the submerged inlet were also analysed.

This paper provides a solution to the active vibration control of a microsatellite with two solar panels. At first, the microsatellite is processed as a finite element model containing a rigid body and two flexible bodies, according to the principles of mechanics, and that the dynamic characteristics are solved by modal analysis. Secondly, the equation involving vibration control is established according to the finite element calculation results. There are several actuators composed of macro fibre composite on the two solar panels for outputting control force. Furthermore, the control voltage for driving actuator is calculated by using fuzzy algorithm. It is clear that the smart structure consists of the flexible bodies and actuators. Finally, the closed-loop control simulation for suppressing harmful vibration is established. The simulation results illustrate that the responses to the external excitation are decreased significantly after adopting fuzzy control.

Civil aircraft that fly long ranges consume a large fraction of civil aviation fuel, injecting an important amount of aviation carbon into the atmosphere. Decarbonising solutions must consider this sector. A philosophical-analytical feasibility of an airliner family to assist in the elimination of carbon dioxide emissions from civil aviation is proposed. It comprises four models based on the integration of the body of a large two-deck airliner with the engines, wings and flight surfaces of a long-range twin widebody jet. The objective of the investigation presented here is to evaluate the impact of liquid hydrogen tank technology in terms of gravimetric efficiency. A range of hydrogen storage gravimetric efficiencies was evaluated; from a pessimistic value of 0.30 to a futuristic value of 0.85. This parameter has a profound influence on the overall fuel system weight and an impact on the integrated performance. The resulting impact is relatively small for the short-range aircraft; it increases with range and is important for the longer-range aircraft. For shorter-range aircraft variants, the tanks needed to store the hydrogen are relatively small, so the impact of tank weight is not significant. Longer range aircraft are weight constrained and the influence of tank weight is important. In the case of the longest range, the deliverable distance increases from slightly over 4,000 nautical miles, with a gravimetric efficiency of 0.3, to nearly 7,000 with a gravimetric efficiency of 0.85.

This paper presents performance analyses of the model-scale ERICA and TILTAERO tiltrotors and of the full-scale XV-15 rotor with high-fidelity computational fluids dynamics. For the ERICA tiltrotor, the overall effect of the blades on the fuselage was well captured, as demonstrated by analysing surface pressure measurements. However, there was no available experimental data for the blade aerodynamic loads. A comparison of computed rotor loads with experiments was instead possible for the XV-15 rotor, where CFD results predicted the FoM within 1.05%. The method was also able to capture the differences in performance between hover and propeller modes. Good agreement was also found for the TILTAERO loads. The overall agreement with the experimental data and theory for the considered cases demonstrates the capability of the present CFD method to accurately predict tiltrotor flows. In a second part of this work, the validated method is used for blade shape optimisation.

Multidisciplinary design and innovative highly automated manufacturing methods are increasingly important to today’s aircraft industry: multidis-ciplinary design because it reduces lead-time and results in a better design, and automated manufacturing methods because they are more capable and reduce manufacturing cost. In this paper a cost estimation model is presented that integrates the manufacturing cost of friction stir welded connections within a multidisciplinary design decision tool. Due to the fact that friction stir welding is a new manufacturing method, the cost estimation model is based on the actual process physics, meaning what the process looks like in terms of processing speeds and characteristics. As an integral part of a multidisciplinary design framework, the developed cost estimation model contributes to a design support tool that assesses not only manufacturing but also structural and aerodynamic issues. It is shown that the cost model developed can be integrated into this more holistic design process support architecture. The predicted costs are accurate to the historical data and allow tradeoff of manufacturing and economic considerations within the context of the multidisciplinary design tool. The tradeoff capability is highlighted through a presented case study that compares the friction stir welding process as an alternative solution to more tradition riveting. Most importantly, this results in a quantitative tradeoff between two processes that shows the manufacturing cycle time of friction stir welding to be reduced by 60% and the recurring assembly cost by 20%.