bluuebunny/crossref_metadata_2025 huggingface dataset:
Supersonic boundary-layer interactions with various micro-vortex generator geometries
Authors: Lee S. , Loth E.
Published: February 2016

Various types of micro-vortex generators (μVGs) are investigated for control of a supersonic turbulent boundary layer subject to an oblique shock impingement, which causes flow separation. The micro-vortex generators are embedded in the boundary layer to avoid excessive wave drag while still creating strong streamwise vortices to energise the boundary layer. Several different types of µVGs were considered including micro-ramps and micro-vanes. These were investigated computationally in a supersonic boundary layer at Mach 3 using monotone integrated large eddy simulations (MILES). The results showed that vortices generated from μVGs can partially eliminate shock induced flow separation and can continue to entrain high momentum flux for boundary-layer recovery downstream. The micro-ramps resulted in thinner downstream displacement thickness in comparison to the micro-vanes. However, the strength of the streamwise vorticity for the micro-ramps decayed faster due to dissipation especially after the shock interaction. In addition, the close spanwise distance between each vortex for the ramp geometry causes the vortex cores to move upwards from the wall due to induced upwash effects. Micro-vanes, on the other hand, yielded an increased spanwise spacing of the streamwise vortices at the point of formation. This resulted in streamwise vortices staying closer to the floor with less circulation decay, and the reduction in overall flow separation is attributed to these effects. Two hybrid concepts, named ‘thick-vane’ and ‘split-ramp’, were also studied where the former is a vane with side supports and the latter has a uniform spacing along the centreline of the baseline ramp. These geometries behaved similar to the micro-vanes in terms of the streamwise vorticity and the ability to reduce flow separation, but are more physically robust than the thin vanes.
Towards Integrated design of fluidic flight controls for a flapless aircraft
Authors: Crowther W. J. , Wilde P. I. A. , Gill K. , Michie S. M.
Published: February 2016

Fluidic flight controls enable forces and moments for flight vehicle trim and manoeuvre to be produced without use of conventional moving surface controls. This paper introduces a methodology for the design of Circulation Control (CC) and Fluidic Thrust Vectoring (FTV) as fluidic controls for roll and pitch. Work was undertaken as part of the multidisciplinary FLAVIIR project, with the goal of providing full authority fluidic flight controls sufficient for a fully flapless flight of an 80kg class demonstrator aircraft known as DEMON. The design methodology considers drag, mass, volume and pneumatic power requirements as part of the overall design cost function. It is shown that the fundamental flow physics of both CC and FTV are similar, and hence there are strong similarities to the design approach of each. Flight ready CC and FTV hardware has been designed, manufactured and ground tested. The CC system was successfully wind tunnel demonstrated on an 85% scale half model of the DEMON. The design condition of a control ΔCL of 0·1 was achieved with a blowing coefficient of 0·01, giving a useable control gain of 10. The FTV system was static tested using a micro gas turbine source. The control characteristic was ‘N’ shaped, consisting of an initial high gain response in a negative sense (gain = −30) followed by a low gain response in a positive sense (gain = +3) at higher blowing rate. CC and FTV control hardware directly contributes to around 6% to the overall mass of the flight vehicle, however provision of pneumatic power carries a significant mass penalty unless generated as part of an integrated engine bleed system.
Supersonic aerodynamics of a projectile with slot cavities
Authors: Ibrahim A. , Filippone A.
Published: February 2016

This paper presents the results of experimental and computational investigations on the effect of slot cavities on a supersonic projectile. Experimental work was carried out to show the effects of the slots on the drag at Mach numbers M = 1·36, 1·65, 1·83. The computational analysis was done at M = 1·36. A single configuration of the slot pattern was used with two different slot widths (0·5mm and 2·0mm). Flow features were investigated in the slotted area and at the base. The analysis presented includes the pressure distribution, the supersonic cavity flow and the effect of the slots on the overall aerodynamic drag. Unlike the case of slots at transonic speeds, the suction and blowing mechanism is not found at supersonic Mach numbers. Streamwise cavity slots cause a small base drag reduction. The reduction in total drag is modest when the width is 0·5mm. However, the experiments showed that with the wider slots (2mm) the drag actually increases at Mach numbers from 1·36 to 1·83.
Validation study for prediction of iced aerofoil aerodynamics
Authors: Marques S. , Badcock K. J. , Gooden J. H. M. , Gates S. , Maybury W.
Published: February 2016

Ice accretions can significantly change the aerodynamic performance of wings and rotor blades. Significant performance degradation can occur when ice accreations cause regions of separated flow, to predict this change implies, at a minimum, the solution of the Reynolds-Averaged Navier-Stokes equations. This paper presents validation for two generic cases involving the flow over aerofoil sections with added synthetic ice shapes. Results were obtained for two aerofoils, namely the NACA 23012 and a generic multi-element configuration. These results are compared with force and pressure coefficient measurements obtained in the NASA LTPT wind-tunnel for the NACA 23012, and force, PIV and boundary-layer measurements obtained at DNW for the multi-element case. The level of agreement is assessed in the context of industrial requirements.
Recent extensions and applications of the ‘CST’ universal parametric geometry representation method
Authors: Kulfan B. M.
Published: February 2016

For aerodynamic design optimisation as well as for multidisciplinary design optimisation studies, it is very desirable to limit the number of the geometric design variables. In Ref. 1, a ‘fundamental’ parametric aerofoil geometry representation method was presented. The method included the introduction of a geometric ‘class function/shape function’ transformation technique, CST, such that round nose/sharp aft end geometries as well as other classes of geometries could be represented exactly by analytic well behaved and simple mathematical functions having easily observed physical features. The CST method was shown to describe an essentially limitless design space composed entirely of analytically smooth geometries. In Ref. 2, the CST methodology was extended to more general three dimensional applications such as wing, body, ducts and nacelles. It was shown that any general 3D geometry can be represented by a distribution of fundamental shapes, and that the ‘shape function/class function’ methodology can be used to describe the fundamental shapes as well as the distributions of the fundamental shapes. A number of applications of the ‘CST’ method to nacelles, ducts, wings and bodies were presented to illustrate the versatility of this new methodology. In this paper, the CST method is extended to include geometric warping such as variable camber, simple flap, aeroelastic and flutter deflections. The use of the CST method for geometric morphing of one geometric shape into another is also shown. The use of CST analytic wings in design optimisation will also be discussed.
Climate functions for the use in multi-disciplinary optimisation in the pre-design of supersonic business jet
Authors: Grewe V. , Stenke A. , Plohr M. , Korovkin V. D.
Published: February 2016

Mitigation of climate change is a challenge to science and society. Here, we establish a methodology, applicable in multi-disciplinary optimisation (MDO) during aircraft pre-design, allowing a minimisation of the aircraft’s potential climate impact. In this first step we consider supersonic aircraft flying at a cruise altitude between 45kfeet (~13·5km, 150hPa) and 67kfeet (~20·5km, 50hPa). The methodology is based on climate functions, which give a relationship between 4 parameters representing an aircraft/engine configuration and an expected impact on global mean near surface temperature as an indicator for the impact on climate via changes in the greenhouse gases carbon dioxide, water vapour, ozone and methane. These input parameters are cruise altitude pressure, fuel consumption, fuel flow and Mach number. The climate functions for water vapour and carbon dioxide are independent from the chosen engine, whereas the climate functions for ozone and methane depend on engine parameters describing the nitrogen oxide emissions. Ten engine configurations are taken into account, which were considered in the framework of the EU-project HISAC. An analysis of the reliability of the climate functions with respect to the simplified climate-chemistry model AirClim and a detailed analysis of the climate functions is given.
New methodology and code for Hawker 800XP aircraft stability derivatives calculation from geometrical data
Authors: Anton N. , Botez R. M. , Popescu D.
Published: February 2016

The new FDerivatives code was conceived and developed for calculating static and dynamic stability derivatives of an aircraft in the subsonic regime, based on its geometrical data. The code is robust and it uses geometries and flight conditions to calculate the aircraft’s stability derivatives. FDerivatives contains new algorithms and methods that have been added to DATCOM’s classical method, presented in a USAF Stability and Control DATCOM reference. The new code was written using MATLAB and has a complex structure which contains a graphical interface to facilitate the work of potential users. Results obtained with the new code were evaluated and validated with flight test data provided by CAE Inc. for the Hawker 800XP business aircraft.
Investigation of rotorcraft-pilot couplings under flight-path constraint below the minimum-power speed
Authors: Lu L. , Padfield G.D. , Jump M.
Published: February 2016

In some situations, closed loop control by the pilot can result in the combined pilot-aircraft system becoming marginally stable or even unstable. This can happen whether the pilot is controlling attitude or flight path. In this paper, an investigation into helicopter stability under flight-path constraint below the minimum-power speed is reported. The work provides a theoretical basis for flight path handling qualities criteria particularly for flight on the, so-called, back-side of the power curve. The research uses the theory of weakly coupled systems by partitioning the helicopter longitudinal dynamics to investigate three interacting subsystems – classically the surge mode, the phugoid mode and the heave mode. Under certain conditions, strong control of flight path or vertical speed is shown to drive the aircraft-pilot system unstable and a conflict is shown to exist between feedback gain values to guarantee stability of both the surge and the flight path motions. This conflict constitutes a potential source of adverse rotorcraft-pilot couplings. The problems are exacerbated in cases when the use of collective control is restricted. The phenomenon is explored in both ground based simulation and flight test to provide a verification of the theory.
Modelling ion propulsion plume interactions with spacecraft in formation flight
Authors: Kafafy R. , Cao Y.
Published: February 2016

This paper presents a simulation study of ion thruster plume effects on formation flying spacecraft. Two formation flight applications using micro-ion propulsion are considered: an L2-Halo orbit interferometer formation and a LEO micro-satellite formation. Worst case scenarios in both missions have been investigated. We focus our study on thruster configurations resulting in possible indirect plume impingement on satellites outside of the direct impingement zone. Indirect impingement which cannot be predicted except through plasma simulation or in-flight measurements might expose critical spacecraft elements such as optical sensors to a harsh contamination environment. A high-fidelity electrostatic plasma simulation code for parallel computing platforms was used in the study.

In our study, we found that using miniature scale ion propulsion in both formation missions creates a very low charge-exchange plasma environment which results in tolerable contamination environment for other spacecraft in the formation.
Nonlinear transient analysis of moderately thick rectangular composite plates
Authors: Tanriöver H. , Şenoca E.
Published: February 2016

This paper presents an analytical-numerical methodology for the geometrically nonlinear analysis of laminated composite plates under dynamic loading. The methodology employs Galerkin technique, in which suitable polynomials are chosen as trial functions. In the solution process, Newmark’s scheme for time integration, and modified Newton-Raphson method for the solution of resulting nonlinear equations are used. In the formulation, first order shear deformation theory based on Mindlin’s hypothesis and von Kármán type geometric nonlinearity are considered. The results are compared to that of finite strips, and Chebyshev series published elsewhere. The method is found to determine closely both the displacements and the stresses. A finite element analysis has also been carried out for the validation of the results. The present method can be efficiently and easily applied for the nonlinear transient analysis of laminated composite plates with various boundary conditions.
Measurements of velocity Field in the wake of an oscillating wind turbine blade
Authors: Soltani M. R. , Mahmoudi M.
Published: February 2016

A series of tests were carried out to study the unsteady wake behaviour behind an aerofoil which is a section of a wind-turbine blade. The model is oscillated in pitch about its quarter chord axis at various reduced frequencies, amplitudes, and mean angles-of-attack. Instantaneous and mean velocity profiles were obtained using total and static pressure at 35 vertically aligned points behind the aerofoil via two similar rakes. The rakes were located at a distance of 1·5 chord length behind the model. An estimation of the real time and average variations of the linear momentum deficit during the oscillation cycle is obtained and has been compared with the corresponding static data. The results show that oscillation amplitude and reduced frequency have pronounced effects on the wake characteristics.
Blended wing body structural concept development
Authors: Velicki A. , Thrash P.
Published: February 2016

A lightweight robust airframe design is one of the key technological advancements necessary for the successful launch of a blended wing body aircraft. The non-circular pressure cabin dictates that substantial improvements beyond current state-of-the-art aluminium and composite structures is needed, and that improvements of this magnitude will require radically new airframe design and manufacturing practices. Such an approach is described in this paper. It is a highly integrated structural concept that is tailored and optimised to fully exploit the orthotropic nature and unique processing advantages inherent in dry carbon fibres, while also employing stitching to enable a unique damage-arrest design approach.
Improved test capabilities for cost-effective performance evaluation of airborne electronic warfare systems
Authors: Pywell M. , Midgley-Davies M.
Published: February 2016

State of the art test capabilities are described that can enable optimised electronic warfare (EW) system development programmes for military air platforms. EW systems are key enablers of air platform survivability and the above is important as defence ministries and industry world-wide wrestle with affordability, technical and industrial challenges. The military end user needs improved capability quickly, with high availability; defence ministries want this at lowest possible cost and with minimum risk. Satisfying these requirements, while remaining in business against a background of fierce international competition and world financial crisis, is indeed challenging. Improvements are described that can enable improved technical performance and lower cost, time and risk programmes than hitherto, with emphasis on RF EW systems and the recent multi-£M upgrade to BAE Systems’ EW Test Facility, with its aircraft-sized anechoic chamber. This upgrade assures a further decade of world-class UK capability for testing manned and unmanned air platforms.
Evolution of generic flight vehicle design synthesis
Authors: Chudoba B. , Heinze W.
Published: February 2016

When defining a new product like an aircraft, space access vehicle or space mission, the Advanced Projects Group evaluates the available design space and compares it with the design space required to accomplish the specified mission. As with any product development process, the general life-cycle characteristics are established first during the conceptual design (CD) phase, clearly before a design proposal can be released to the follow-on design phases such as preliminary design (PD), detail design (DD), flight test (FT), and finally operation and disposal. As a rule of thumb, it can be assumed that around 80% of the flight vehicle configuration and mission tandem are determined during the CD phase alone, which is the key phase where the initial brainstorming has to take place. Clearly, it is the responsibility of the CD team to simulate the entire life-cycle of the project from ‘cradle to grave’ where the focus is on correctness rather accuracy in order to identify the design space and offer an overall proof of design convergence. Currently, the important primary aerospace vehicle and mission design decisions at CD level are still made using extremely simple analysis and heuristics. A reason for this scenario is the difficulty in synthesising the range of individual design disciplines for both, classical and novel aerospace vehicle conceptual designs, in more than an ad hoc fashion. Although the CD segment is seen as the most important step in the product development phase due to its pre-defining function, it is the least well understood part of the entire product evolution process due to its level of abstraction. This paper presents the roadmap towards the next generation of aerospace life-cycle synthesis systems, a software and management process capable to immediately calculate cost and time implications while simultaneously linking design, manufacturing, testing, and operation. A historical review of how design has been accomplished until today is presented. The design approaches are categorised and the characteristics of today’s state-of-the-art design synthesis systems are discussed. A specification for the new class of intelligent generic design synthesis systems is presented capable of satisfying the demands imposed by the new breed of high-performance aircraft, space access vehicles, space missions, and others. Finally, the development status of the next generation aerospace vehicle design synthesis (AVDS-PrADO) simulation-based acquisition environment is presented.
Wake vortex encounters in a flight simulator
Authors: Allerton D. J. , Spence G. T.
Published: February 2016

A real-time model of a wake vortex has been developed for an engineering flight simulator, which enables studies to be undertaken into the effects of wake vortex encounters with civil transport aircraft. A brief summary of the methods used to create a vortex model and integrate it with a real-time flight simulator is presented. Tools have been developed to determine the structural loads on an aircraft during a simulated wake vortex encounter and to visualise the airframe loads. Four wake vortex encounters were flown with different pilot recovery strategies and automatic flight controls. It is clear from analysis of these results that it is difficult for a pilot to appreciate the margin of tail loading during an encounter and a prototype training display is outlined to enable pilots to fully understand the effects of rudder inputs in recovery from these events.
The effect of gravity on a shaped sonic boom
Authors: Cain T.
Published: February 2016

Previously it was demonstrated that gravity is directly responsible for the attenuation of sonic booms as they propagate towards the ground, while the gradient in acoustic impedance has no effect on the wave strength. This was a significant departure from the well accepted acoustic theory and naturally led to questions concerning comparison with experiment and the implications for low boom design. This paper presents a comparison of measured ground signatures of the Northrup-Grumman Shaped Sonic Boom Demonstrator (SSBD) with a method of characteristics extrapolation from a close proximity flight measurement of static pressure under the aircraft. Comparison is also made with Whitham’s theory, extended to include gravity and the ambient temperature variation, as presented in the previous paper. The calculations are both in good agreement with the experimental ground signature. The usual equations that define the acoustic propagation (gravitational body force neglected) are transformed so that the computational algorithms of the extended Whitham theory are applicable. The transformation simplifies the acoustic model and reveals the conditions under which the errors in its prediction will be large. At the Mach number of 1.4, corresponding to the SSBD flight, the error is relatively small and the acoustic prediction is also in close agreement with the ground signature.
Mechanics and aerodynamics of perching manoeuvres in a large bird of prey
Authors: Carruthers A. C. , Thomas A. L. R. , Walker S. M. , Taylor G. K.
Published: February 2016

This paper reviews recent results on the mechanics and aerodynamics of perching in a large bird of prey, the Steppe Eagle Aquila nipalensis. Data collected using onboard and high-speed video cameras are used to examine gross morphing of the wing planform by the flight muscles, and smaller-scale morphing of the wing profile by aeroelastic deflection of the feathers, Carruthers et al. High-resolution still images are used to reconstruct the shape of the wing using multi-station photogrammetry, and the performance of the measured wing profile is analysed using a panel code, Carruthers et al. In bringing these lines of research together, we examine the role of aeroelastic feather deflection, and show that the key to perching in birds lies not in high-lift aerodynamics, but in the way in which the wings and tail morph to allow the bird to transition quickly from a steady glide into a deep stall.
Design support for metallic machined ribs using knowledge based engineering techniques
Authors: van der Laan T. , van Dalen F. , Vermeulen B.
Published: February 2016

In recent years increasingly pressure has been applied on aircraft component suppliers to reduce design lead-time and design cost of aircraft components. To achieve this reduction in lead-time and cost, advanced automation tools that automate part of the engineering process can be used. Knowledge based engineering (KBE) tools are one such automation tool type and automate part of the engineering process based on existing knowledge within a company. In this paper the development process and use of a KBE application to design machined ribs in an industrial setting is discussed. The development of the KBE tool has resulted in the standardisation of the design methodology. Furthermore it has reduced machined rib design lead-time in a project to develop a bussines jet empennage by 40%.
Modelling the flight dynamics of the hang glider
Authors: Cook M. V. , Spottiswoode M.
Published: February 2016

The development of the non-linear equations of motion for the hang glider from first principles is described, including the complex geometry of control by pilot ‘weight shift’. By making appropriate assumptions the linearised small perturbation equations are derived for the purposes of stability and control analysis. The mathematical development shows that control is effected not by pilot weight shift, but by centre of gravity shift and that lateral-directional control by this means is weak, and is accompanied by significant instantaneous adverse response.

The development of a comprehensive semi-empirical mathematical model of the flexible wing aerodynamics is described. In particular, the modelling attempts to quantify camber and twist dependencies. The performance of the model is shown to compare satisfactorily with measured hang glider wing data obtained in earlier full scale experiments. The mathematical aerodynamic model is then used to estimate the hang glider stability and control derivatives over the speed envelope for substitution into the linearised equations of motion.
Comparison of a grid-based CFD method and vortex dynamics predictions of low Reynolds number cylinder flows
Authors: Baranyi L. , Lewis R. I.
Published: February 2016

Computational fluid dynamics models range from the finite difference type grid-based method to the Lagrangian style vortex cloud simulation technique for solving the Navier-Stokes equations. This paper undertakes a comparison of these two methods for the classical datum bluff body case of flow past a stationary circular cylinder at low Reynolds numbers in the range 10 to 220. Comparisons include time-history, time-mean and root-mean-square values of oscillating drag and lift coefficients, frequency of vortex shedding and related vortex street wake flow patterns. Particularly close agreement was obtained for Strouhal number versus Reynolds number, and good agreement for time-mean value of drag coefficients; comparison was also made with experimental results. Attempts are also made to calculate the skin friction and surface pressure components of the cylinder drag, revealing the significance of skin friction drag within this range and its relative insignificance above a Reynolds number of 220.
High performance computing and computational aerodynamics in the UK
Authors: Emerson D. R. , Sunderland A. J. , Ashworth M. , Badcock K. J.
Published: February 2016

The establishment of the UK Applied Aerodynamics Consortium in 2004 brought together many of the UK’s leading research groups to tackle challenging aerodynamic problems on the national computing facility, HPCx. This paper provides a brief history of some early pioneers of numerical simulation and highlights some key contributions to development in parallel processing that laid the foundations for today’s researchers. The transition from vector to massively parallel processing is discussed from a UK viewpoint along with technological barriers that could have a significant impact on future systems. Solutions to these barriers are already being sought and the paper discussed some of the novel technologies that may be deployed in the future. In its short history, the consortium has made substantial progress and this is briefly discussed with several highlights that illustrate the scientific output. Although a number of challenges are identified, particularly with respect to developing a comprehensive visualisation capability, the consortium is well placed to build upon its initial success.
An unsteady, moving mesh CFD simulation for Harrier hot-gas ingestion control analysis
Authors: Richardson G. A. , Dawes W. N. , Savill A. M.
Published: February 2016

Hot gas ingestion (HGI) can be a problematic feature of short take-off vertical landing (STOVL) aircraft during the descent phase of landing, or while on the ground. The hot exhaust gases from the downwards pointing nozzles can be re-ingested into the engine intakes, causing power degradation or reduced engine surge margin. The flow-fields that characterise this phenomenon are complex, with supersonic impinging jets and cross-flows creating large ground vortices and fountain up-wash flows.

A flow solver has been developed to include a suitable linear mesh deformation technique for the descending aircraft configuration. The code has been applied to predict the occurrence of HGI, by simulating experimental results from a 1/15th scale model of a descending Harrier. This has enabled an understanding of the aerodynamic mechanisms that govern HGI, in terms of the near-field and far-field effects and their impact on the magnitude of temperatures at the engine intake.

This paper presents three sets of CFD results. First a validation exercise shows predicted results from the twin-jet with intake in crossflow test-case. This is an unsteady Reynolds averaged Navier Stokes (URANS) solution for a static geometry (there is no moving mesh). This allows comparison with experiment. Secondly, a full descent phase URANS Spalart-Allmaras (SA) turbulence model calculation is done on an 8·5m cell mesh for half the flow domain of the Harrier model and test-rig without dams/strakes. This shows how the HGI flow mechanisms affect the engine intake temperature profiles, for the case where there are no flow control methods on the underside of the aircraft. Thirdly, the full descent phase URANS SA turbulence model calculation is done on a 22·4m cell mesh for the full flow domain of the Harrier model and test-rig, with the dam/strake geometry included in the structured mesh region.
Loss reduction on ultra high lift low-pressure turbine blades using selective roughness and wake unsteadiness
Authors: Howell R. J. , Roman K. M.
Published: February 2016

This paper describes how it is possible to reduce the profile losses on ultra high lift low pressure (LP) turbine blade profiles with the application of selected surface roughness and wake unsteadiness. Over the past several years, an understanding of wake interactions with the suction surface boundary layer on LP turbines has allowed the design of blades with ever increasing levels of lift. Under steady flow conditions, ultra high lift profiles would have large (and possibly open) separation bubbles present on the suction side which result from the very high diffusion levels. The separation bubble losses produced by it are reduced when unsteady wake flows are present. However, LP turbine blades have now reached a level of loading and diffusion where profile losses can no longer be controlled by wake unsteadiness alone.

The ultra high lift profiles investigated here were created by attaching a flap to the trailing edge of another blade in a linear cascade — the so called flap-test technique. The experimental set-up used in this investigation allows for the simulation of upstream wakes by using a moving bar system. Hotwire and hotfilm measurements were used to obtain information about the boundary-layer state on the suction surface of the blade as it evolved in time. Measurements were taken at a Reynolds numbers ranging between 100,000 and 210,000.

Two types of ultra high lift profile were investigated; ultra high lift and extended ultra high lift, where the latter has 25% greater back surface diffusion as well as a 12% increase in lift compared to the former. Results revealed that distributed roughness reduced the size of the separation bubble with steady flow. When wakes were present, the distributed roughness amplified disturbances in the boundary layer allowing for more rapid wake induced transition to take place, which tended to eliminate the separation bubble under the wake. The extended ultra high lift profile generated only slightly higher losses than the original ultra high lift profile, but more importantly it generated 12% greater lift.
On different parameterisation methods to analyse spacecraft attitude manoeuvres in the presence of attitude constraints
Authors: Radice G. , Casasco M.
Published: February 2016

This paper analyses and compares two different attitude representations, using quaternions and modified Rodrigues parameters, in the context of the potential function method applied to autonomously control constrained attitude slew manoeuvres. This method hinges on the definition of novel Lyapunov potential functions in terms of the attitude parameters representing the current attitude, the goal attitude and any pointing constraints, which may be present. It proves to be successful in forcing the satellite to achieve the desired attitude while at the same time avoiding the pointing constraints. A linearised version of the modified Rodrigues parameterisation is also introduced and analysed. Finally advantages and drawbacks of all attitude representations are discussed.
A static compressible flow model of synthetic jet actuators
Authors: Tang H. , Zhong S.
Published: February 2016

In this paper, a simple static compressible flow model for circular synthetic jet actuators is described. It is used to undertake a systematic computational investigation of the effect of changing actuator geometrical and operating parameters on the magnitude of peak jet velocity at the orifice exit of an actuator whose diaphragm displacement and frequency are allowed to vary independently. It is found that, depending on the flow conditions inside the orifice duct, the actuator may operate in two distinct regimes, i.e. the Helmholtz resonance regime and the viscous flow regime. In the Helmholtz resonance regime, the resultant synthetic jet is generated by the mass physically displaced by the oscillating diaphragm coupled with the Helmholtz resonance in the actuator. In the viscous flow regime, the Helmholtz resonance is completely damped by viscous effect such that the jet is produced by the diaphragm oscillation alone. The relationship between actuator geometrical and operating parameters at the optimum condition which yields the maximum peak jet velocity at a given diaphragm displacement is also established for these two regimes. Finally, a preliminary procedure for designing synthetic jet actuators for flow separation control on an aircraft wing is proposed.
The representation of GT component maps using Mach numbers
Authors: Kyritsis V. E. , Pilidis P. , Ramsden K.
Published: February 2016

Component maps are produced under certain environmental conditions using air as the working fluid during static ground operation. Any changes of the component characteristics when operating under different temperature conditions and/or with different working fluid are partially taken into account, because of the existence of the gas constant and the ratio of the specific heats in the non-dimensional mass flow and rotational speed. This provides a second order correction for the component characteristics, which may be adequate for the initial modeling of engines. However, for rigorous performance calculations correction factors are applied to the non-dimensional mass flow, rotational speed and pressure ratio distributions of a map, when deviations from the reference conditions under which it was extracted, are experienced. In the current study, a different approach is considered in order to eliminate the inaccuracies caused by the varying temperature and chemical composition. It makes direct use of inlet and circumferential Mach numbers based upon stagnation temperature in conjunction with dimensionless enthalpy variation. A sensitivity analysis against gas property variations is conducted to quantify the benefits gained in precision. Generally, the well-known relationships correlating the Mach number with total and static properties are based on the assumption of perfect gas and constant gas properties. Introducing dependency on temperature and/or chemical composition for the caloric properties of the semi-perfect gas, proper mean values are defined and some theoretical corrections are provided for the well-known equations. The mass flow compatibility equation is then based on the ‘corrected’ expression correlating dimensionless mass flow and Mach number and takes full account of gas property variations.
The Victoria University of Manchester’s contributions to the development of aeronautics
Authors: Ackroyd J. A. D.
Published: February 2016

This issue of the Aeronautical Journal celebrates the 50th anniversary of the foundation of the Honours Degree Course in Aeronautical Engineering at the Victoria University of Manchester. The following article therefore describes the aeronautical research and teaching activities of that university up to its recent amalgamation with the University of Manchester Institute of Science and Technology (UMIST) to form the present-day University of Manchester. This juncture provides a further justification for recording the Victoria University’s achievements.

Both the Victoria University and UMIST had their roots in the nineteenth century although, apart from the relatively brief period of the First World War, neither of them was particularly involved in aeronautics until after the Second World War. However, as Sections 6.0-10.0 seek to demonstrate, thereafter the Victoria University’s involvement became considerable. The preceding Sections describe the origins of the Victoria University and UMIST and, in the case of the former institution, the subsequent activities of its staff and graduates in engineering and mathematics which, although not always specifically aeronautical in content, nonetheless had a profound influence on the development of the aeronautical sciences.
On the benefits of lower Mach number aircraft cruise
Authors: Filippone A.
Published: February 2016

The paper reviews the issue of cruise Mach number and addresses the benefits of operating subsonic commercial aircraft at speeds below the long-range cruise speed. The case considered is the flight of transport aircraft for flight segments up to 1,000nm. It is shown that the fuel burned is decreased by as much as 1·8% on a nominal 1,000nm stage length for operation around the long-range cruise Mach number, or below. This is achieved at a cost of a marginal delay on each flight segment (less than three minutes). The longer flight time is likely not to affect the daily operation of the aircraft. The fuel saving is compounded, because the gross take-off weight (GTOW) is recalculated to take into account the reduced fuel consumption at each flight segment. The analysis into the environmental benefits includes the reduction in,andemissions, and the heat released in the high atmosphere. Sensitivity analyses are carried out on the take-off weight, on the aerodynamic coefficients, on the transonic drag rise and the weight uncertainty. It is predicted that the optimal operation of the example aircraft over a nominal 1,000nm route can reduce the fuel consumption by as much as 150,000kg per year in comparison with an operation at the long-range Mach number. The aircraft model has a maximum take-off weight of 170,000kg and is powered by two GE CF6-80C2 engines.
Estimation of lateral-directional parameters using neural networks based modified delta method
Authors: Singh S. , Ghosh A. K.
Published: February 2016

The aim of the study described herein was to develop and verify an efficient neural network based method for extracting aircraft stability and control derivatives from real flight data using feed-forward neural networks. The proposed method (Modified Delta method) draws its inspiration from feed forward neural network based the Delta method for estimating stability and control derivatives. The neural network is trained using differential variation of aircraft motion/control variables and coefficients as the network inputs and outputs respectively. For the purpose of parameter estimation, the trained neural network is presented with a suitably modified input file and the corresponding predicted output file of aerodynamic coefficients is obtained. An appropriate interpretation and manipulation of such input-output files yields the estimates of the parameter. The method is validated first on the simulated flight data using various combinations and types of real-flight control inputs and then on real flight data. A new technique is also proposed for validating the estimated parameters using feed-forward neural networks.
Large-eddy simulation of separated flow over a swept wing with approximate near-wall modelling
Authors: Li N. , Leschziner M. A.
Published: February 2016

The paper investigates, by means of a simulation methodology, the flow separating from a 40 degrees backward-swept wing at 9 degrees incidence and Reynolds number of 210,000, based on the wing-root chord length. The Simulation corresponds to LDA, PIV and suction-side-topology measurements for the same geometry, conducted by other investigators specifically to provide validation data. The finest block-structured mesh contains 23·6 million nodes and is organised in 256 blocks to maximise mesh quality and facilitate parallel solution on multi-processor machines. The near-wall layer is resolved, to a thickness of about y+ = 20, by means of parabolised URANS equations that include an algebraic eddy-viscosity model and from which the wall-shear stress is extracted to provide an unsteady boundary condition for the simulation. The numerical solution is in good agreement with the experimental behaviour over the 50-70% inboard portion of the span, but the simulation fails to resolve some complex features close to the wing tip, due to a premature leading-edge vortex breakdown and loss in vortex coherence. The comparisons and their discussion provide useful insight into various physical characteristics of this complex separated wing flow.
Prediction of drag and lift of wings from velocity and vorticity fields
Authors: Zhu G. , Bearman P. W. , Graham J. M. R.
Published: February 2016

The present paper continues the work of Zhuet 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.
Reducing environmental impacts of aviation with innovative air traffic management technologies
Authors: Williams V. , Noland R. B. , Majumdar A. , Toumi R. , Ochieng W. , Molloy J.
Published: February 2016

Commercially-driven air traffic management (ATM) innovations typically aim to increase air space capacity and/or reduce delays. Here, their potential application for environmental mitigation is discussed. Both carbon dioxide (CO2) and non-CO2 climate impacts are considered, as are noise and air quality issues. We outline the technological, scientific and political barriers to an integrated approach to applying ATM technologies to environmental mitigation. These issues highlight the need to improve comparison and prioritisation of the emissions and effects of aviation.
A numerical study of high pressure turbine forced response in the presence of damaged nozzle guide vanes
Authors: Mare L.di , Imregun M. , Smith A. D. , Elliott R.
Published: February 2016

This paper reports results from numerical computations of low engine order and blade-passing forced response on the rotor of a high pressure turbine due to severe damage to a single nozzle guide vane. The computations are performed using a time-domain, nonlinear viscous compressible flow simulation code. The flow and the levels of forcing for a few selected modes are compared for the undamaged and the damaged configurations. The results show that the response in various modes is affected to a different extent by the damage. The main blade-passing response was found to be largely unaffected, if not marginally reduced. On the other hand, the vibration levels for some modes were seen to be up to eight times higher because of the low-order excitation harmonics created by the damaged passage.
Vortex interaction and breakdown over
double-delta wings
Authors: Gai S. L. , Roberts M. , Barker A. , Kleczaj C. , Riley A. J.
Published: February 2016

Modern high-speed aircraft, especially military, are
very often equipped with single or compound delta
wings. When such aircraft operate at high
angles-of-attack, the major portion of the lift is
sustained by streamwise vortices generated at the
leading edges of the wing. This vortex-dominated
flow field can breakdown, leading not only to loss
of lift but also to adverse interactions with other
airframe components such as the fin or horizontal
tail. The wind tunnel and water studies described
herein attempt to clarify the fluid mechanics of
interaction between the strake and wing vortices of
a generic 76°/40° double-delta wing leading to
vortex breakdown. Some studies of passive control
using fences at the apex and kink region are also
described. Various diagnostic methods-laser sheet
flow visualisation, fluorescent dyes, and pressure
sensitive paints have been used.
Validation of a rotorcraft mathematical
model in autorotation by use of gyroplane flight
tests
Authors: Spathopoulos V. M.
Published: February 2016

Aircraft handling qualities in autorotation are
critical in determining the level of safety of
rotorcraft. For helicopters suffering from an engine
failure, transcending from powered to autorotative
flight occurs rapidly and requires immediate and
accurate pilot reaction. Although it is important
for the handling qualities in this flight state to
be predicted correctly, obvious difficulties will
exist in using flight tests as a means of validation
when autorotation constitutes an abnormal mode of
operation. In the research work presented in this
paper an alternative approach is applied, of
configuring a generic rotorcraft model as a
gyroplane, a type of vehicle for which its main
rotor is constantly in autorotation. Flight tests
are used for the validation purposes both for steady
state and dynamic response cases. Results are
produced to complement those already existing for a
dissimilar gyroplane type thus increasing the level
of confidence obtained. It is concluded that
important handling qualities indicators such as the
steady state trends are correctly predicted although
limitations are imposed due to rotor speed
discrepancy.
Airspace safety in New Zealand: A causal analysis of controller caused airspace incidents between 1994-2002
Authors: Majumdar A. , Ochieng W. Y. , Nalder P.
Published: February 2016

The New Zealand Government takes airspace safety very seriously. The level of safety in New Zealand airspace is measured by the number of recorded airspace incidents. An airspace incident can be thought of as a failure in the chain of operations in the air traffic system when it is provided with an air traffic service (ATS). Some of these incidents result in a loss of separation between aircraft, varying from slight to a very serious loss with a significant risk of collision, known as a near collision. New Zealand’s Civil Aviation Authority (CAA) identifies the causal factors for all airspace incidents using the Reason model of human error, divided into three areas; active failures committed by individuals involved in the incident, local factors relating to the task and the ATS environment and organisational factors originating in the managerial and organisation spheres of the ATS provider. Based upon the CAA’s database, this paper analyses trends in controller caused incidents in the NZ airspace during the period 1994 to 2002 in six controller caused categories. The results indicate that for controller caused incidents, execution errors dominate the active failures category, while in the case of local factors, poor concentration/lack of attention, inadequate checking and controller workload are the dominant factors. For the organisation category, poor planning and inadequate control and monitoring dominate. These results should form the basis of a robust and transparent framework for intervention mechanisms by the New Zealand Civil Aviation Authority for enhanced airspace safety.