bluuebunny/crossref_metadata_2025_split huggingface dataset:
Wing mass estimation algorithm for medium range box wing aircraft
Authors: Jemitola P. O. , Monterzino G. , Fielding J. 
Published: February 2016

AbstractA procedure for defining an empirical formula for the mass estimation of the fore and aft wings field Uof a medium range box wing aircraft is described. The procedure is based upon the work of Howe for estimating the wing mass of conventional cantilever wing aircraft. The paper outlines the procedure used to relate conventional cantilever wings to box wing aircraft wings. Using a vortex lattice tool, finite element methods and regression analysis, the modification performed on the coefficient in Howe’s method to enable its use on a medium range box wing aircraft is outlined. The results show that the fore and aft wings would use the same correction coefficient and that the aft wing would therefore be lighter than the fore wing on a medium range box wing aircraft.
Validation of mathematical models for helicopter flight simulators past, present and future challenges
Authors: Pavel M. D. , White M. , Padfield G. D. , Roth G. , Hamers M. , Taghizad A. 
Published: February 2016

AbstractAt the heart of a flight simulator resides the mathematical representation of aircraft behaviour in response to control inputs, atmospheric disturbances and system inputs including failures and malfunctions. While this mathematical model can never be wholly accurate, its fidelity, in comparison with real world behaviour, underpins the usefulness of the flight simulator. The present paper examines the state of the art achieved in validating mathematical models for helicopter simulators, addressing the strengths and weaknesses of the present European standard for the qualification of helicopter flight simulators, JAR FSTD-H (previously JAR-STD-1H/2H/3H). Essential questions are examined, such as: What is the required model fidelity to guarantee a simulation is sufficiently representative to be fit for purpose? Are the tolerances set in the current standards fine enough that they lead to only minor changes in handling qualities? What is an acceptable tuning process for the simulation? What is the effect of modelling fidelity on the overall pilot control strategy? What is the relationship between the settings of the simulator cueing environment and the behaviour of the pilot? What is the industrial experience on qualification of flight simulators that might usefully inform developments? Many of these questions were addressed in Europe in a previous GARTEUR Action Group (AG) HC/AG-12 the results of which are documented in this paper. Solutions are proposed for improving the current JAR-FSTD standard with respect to validation of mathematical models.
New directions for air traffic control simulators: A discussion to guide the selection and renovation of simulators
Authors: Çetek C. , Aybek F. , Çinar E. , Cavcar A. 
Published: February 2016

AbstractThe authors have recent experience of the specification, procurement, customisation and operation of advanced air traffic control simulator systems. This paper summarises the specification of the software and hardware components of air traffic control simulators within a framework of user needs and advances in simulation technology. Both the required conventional characteristics and the desired innovative features which can accommodate future requirements are defined using taxonomy within this framework. Considering the high investments and challenges involved in the acquisition and operation of modern simulators, the proposed classifications and approaches will provide important benefits to users when selecting or renovating their systems.
High resolution computation of compressible condensing/evaporating moist-air flow for external and internal flows
Authors: Hamidi S. , Kermani M. J. 
Published: February 2016

AbstractIn this paper computation of high resolution, compressible condensing/evaporating moist-air flow for a series of two-dimensional internal and external flows is performed. It has been observed that the inflow wetness content can change the aerothermodynamics of the flow field, e.g. the shock angles. In the case of flow expansion through nozzles, it has also been observed that the content of wetness at nozzles exit in the case of moist-air is more than five times higher than that of pure steam under similar operating conditions. The reason is due to the internal flow of heat from steam portion toward air that accelerates the steam condensation rate. The solver is spatially third- and temporally second-order accurate. Validations of the numerical code are performed versus the experimental data of Moore et al (1973) in the case of very large humidity ratio, ω→∞ (pure steam). Also in the case of low humidity ratio, ω→0 (dry air), the exact solution of shock tube and wedge cases are used to validate our numerical results.
Formation of vortices and spanwise flow on an insect-like flapping wing throughout a flapping half cycle
Authors: Phillips N. , Knowles K. 
Published: February 2016

AbstractThis paper presents an experimental investigation of the evolution of the leading-edge vortex and spanwise flow generated by an insect-like flapping-wing at a Reynolds number relevant to flapping-wing micro air vehicles (FMAVs) (Re = ~15,000). Experiments were accomplished with a first-of-its-kind flapping-wing apparatus. Dense pseudo-volumetric particle image velocimetry (PIV) measurements from 18% – 117% span were taken at 12 azimuthal positions throughout a flapping half cycle. Results revealed the formation of a primary leading-edge vortex (LEV) which saw an increase in size and spanwise flow (towards the tip) through its core as the wing swept from rest to the mid-stroke position where signs of vortex breakdown were observed. Beyond mid-stroke, spanwise flow decreased and the tip vortex grew in size and exhibited a reversal in its axial direction. At the end of the flapping half cycle, the primary LEV was still present over the wing surface, suggesting that the LEV remains attached to the wing throughout the entire flapping half cycle.
Experimental studies on rectangular jets with trapezoidal tabs
Authors: Arokkiaswamy A. , Verma S. B. , Venkateswaran S. 
Published: February 2016

AbstractAn experimental investigation was carried out to study the flow development of a jet issuing from a 2:1 rectangular nozzle with mixing tabs using two-component hotwire anemometry. A pair of tabs of trapezoidal configuration (with 2% total blockage area) is placed on the minor-axis side of the rectangular nozzle and tested for two tab inclination angles of 135° and 45°, with respect to the flow direction. Tests were conducted for a nominal jet exit velocity of 20m/sec corresponding to a Reynolds number based on nozzle equivalent diameter of 5·013 × 104. Relative to the plain jet, the jet with tabs show significant reduction in jet-core length (by 67%) followed by a faster decay in jet centreline velocity (U/Ue). This is also accompanied by a significant upstream shift in peak centreline turbulence intensity (u’/Ue). The presence of tabs is observed to inhibit the jet growth along the minor-axis plane thereby introducing large distortion in the jet cross-sectional development that ultimately leads to jet-core bifurcation along its major-axis. While a mushroom-like flow structure develops behind the tab with 135° inclination, the flow structure behind a 45° inclined tab rather takes the shape of the tab itself. The former flow development is seen to enhance the jet growth more along the minor-axis while the latter improves the jet growth more along the major-axis plane. From application point of view, since both tab inclinations result in more or less similar jet characteristics, a 135° inclined tab would be preferable over a 45° inclined tab from the view of improved jet mixing.
A conceptual wing-box weight estimation model for transport aircraft
Authors: Ajaj R. M. , Friswell M. I. , Smith D. , Isikveren A. T. 
Published: February 2016

AbstractThis paper presents an overview of an advanced, conceptual wing-box weight estimation and sizing model for transport aircraft. The model is based on linear thin-walled beam theory, where the wing-box is modelled as a simple, swept tapered multi-element beam. It consists of three coupled modules, namely sizing, aeroelastic analysis, and weight prediction. The sizing module performs generic wing-box sizing using a multi-element strategy. Three design cases are considered for each wing-box element. The aeroelastic analysis module accounts for static aeroelastic requirements and estimates their impact on the wing-box sizing. The weight prediction module estimates the wing-box weight based on the sizing process, including static aeroelastic requirements. The breakdown of the models into modules increases its flexibility for future enhancements to cover complex wing geometries and advanced aerospace materials. The model has been validated using five different transport aircraft. It has shown to be sufficiently robust, yielding an error bandwidth of ±3%, an average error estimate of -0·2%, and a standard error estimate of 1·5%.
Creating a space debris catalogue for an orbital band with suitable candidates for active removal
Authors: Dean D. 
Published: February 2016

AbstractThere are over 16,000 pieces of debris that are larger than 10cm in diameter being tracked which could cause a catastrophic destruction of a satellite if they were to collide. To reduce the likelihood of this happening, debris production is being reduced through debris mitigation techniques; however the debris already in space will need to be removed as well. To design an effective debris removal mission, the first logical step would be to gather as much detail as possible on debris that is known to exist within the orbital band of interest. This report presents a catalogue of debris, created entirely from open source data, for an orbital band within the 900-1,000km altitude range and 82-83° inclination, and offers a possible candidate for deorbit – the Cosmos 3M final-stage rocket body, of which there are 142, each with a mass of 1,435kg, length of 6·5m and diameter of 2·4m.
Pressure measurements over a hemisphere-cylinder body attached with a forward facing spike
Authors: Kalimuthu R. , Mehta R. C. , Rathakrishnan E. 
Published: February 2016

AbstractThe present paper presents oil flow visualisations and pressure measurements over a hemisphere-cylinder body attached with a forward facing spike at Mach 6 and Reynolds number of 1·38 × 108 at 0° and 5° angle-of-attack. The oil flow pictures depict the separation region in the vicinity of the spike on the hemisphere-cylinder body. The oil flow visualisations will help to locate the reattachment shock wave on the hemisphere-cylinder body and also understand the flow field behavior on the blunt-nosed spike configuration. The pressure measurements over the hemisphere-cylinder body depend on the shape and the length of the spike. The pressure distributions over the blunt-nosed body show significant influence of the angle-of-attack. The maximum pressure coefficient on the hemisphere-cylinder body is a function of the spike length, shape of the aerodisk and angle-of-attack. The windward and leeward sides pressure variations show dependence of the geometrical parameters of the spike and shape of the spike. The hemisphere and the flat-faced aerodisk cause considerable reduction of pressure leading to decrease of aerodynamic drag compared to the conical spike.
Optimisation of round bodies for aerodynamic performance and stability at supersonic speeds
Authors: Jiajan W. , Chue R. S. M. , Nguyen T. , Yu S. 
Published: February 2016

AbstractAn optimisation procedure coupled with computational fluid dynamics (CFD) is proposed to minimise the aerodynamic drag and to improve the static and dynamic stabilities of generic rounds at supersonic speeds (Mach 1·5 to 4). First, the Active-set algorithm, Sequential Quadratic Programming (SQP) is used as the optimisation method for drag minimisation. The objective function is the zero-lift drag computed from a semi-empirical solution. The constraints are based on the geometric restrictions of the body. CFD is then employed to validate the accuracy of the drag prediction from the semi-empirical solution and to incorporate the stability requirements into the optimisation process. A supersonic round body is considered as an example application. The optimised body provides up to 15% drag reduction and 46% increase in gyroscopic stability while remaining dynamically stable over the whole range of the operating Mach numbers.
Energy absorption of sandwiched honeycombs with facesheets under in-plane crushing
Authors: Atli-Veltin B. , Gandhi F. 
Published: February 2016

AbstractThe in-plane crushing and energy absorption of sandwiched honeycomb cores with facesheets are examined through finite element simulations. Assuming no debonding between the facesheet and honeycomb core (which would be the case if manufacturing techniques such as brazing are used to produce very strong bonds between the facesheeet and the core), intracellular buckling mode for thin facesheets, and wrinkling mode for thick facesheets are observed. In the dimpling mode, deformation is governed by the core, honeycomb vertical cell walls do not deform, and the inclined wall deformation does not vary through the cell depth. In the wrinkling mode, deformation is governed by the facesheet, the vertical cell walls deform significantly, and the inclined cell wall deformation varies through the cell depth. Increasing cell angle increased Specific Energy Absorption (SEA) for honeycombs with thin facesheets. Decreasing vertical cell wall length increased SEA for honeycombs with thick facesheets. Increasing wall thickness and decreasing core depth increased SEA for honeycombs with thin and thick facesheets. With geometric changes, SEA increased ~3 times over the baseline configurations. For a given keel beam dimensions, using fewer rows of larger cells reduces the effective non-dimensional core-depth, thereby increasing the effect of the facesheet and the SEA significantly. The SEA of sandwiched honeycombs with facesheets in in-plane crushing appears to be competitive with, or better than, SEA honeycombs in out-of-plane crushing.
Simulations of a low-boom, axisymmetric, external compression inlet
Authors: Coyne T. , Loth E. , Koncsek J. , Davis D. , Conners T. , Howe D. 
Published: February 2016

AbstractComputational simulations have been used to study a new low-boom, axisymmetric, external compression supersonic inlet with an on-design Mach number of 1·7. The inlet incorporates a relaxed compression surface with a near-zero cowl angle to help reduce external oblique shock waves due to spillage and cowl geometry. To reduce mechanical complexity the inlet is designed with zero-bleed. To understand the impact on performance and shock overpressure caused by the inlet itself, several throat and diffuser designs were simulated. The computations utilised a Reynolds-averaged Navier-Stokes code. Inflow properties were held consistent with the operational characteristics of the NASA GRC 8′ × 6′ Supersonic Wind Tunnel (SWT) for experimental testing of the inlet of Mach 1·67 at a Reynolds number of 5·4 × 106. Stagnation pressure recovery performance for the baseline condition exceeded 94% at design mass flow rates, and reduced only slightly with increases in Mach number (consistent with theoretical predictions) and extension of cowl position. The simulations also showed that the relaxed compression surface combined with the near-zero cowl angle helps to significantly reduce external oblique shocks This is partly due to the reduced inlet spillage in combination with a reduced overall turning angle placed on the free-stream flow relative to the cowl shape.
Density measurements for rectangular free jets using background-oriented schlieren
Authors: Tipnis T. J. , Finnis M. V. , Knowles K. , Bray D. 
Published: February 2016

AbstractAn experimental study incorporating the use of the Background-Oriented Schlieren (BOS) technique was performed to measure the density field of a rectangular supersonic jet. This technique is easier to set up than conventional schlieren since the optical alignment involving the various mirrors, lenses and knife-edge is replaced by a background pattern and a single digital camera. The acquired images which contain information of density gradients in the flow are solved as a Poisson equation and further processed using deconvolution and tomographic algorithms to generate a 3D domain which contains information about the actual density. 2D slices can then be extracted to quantitatively visualise the density along any required planes. The results from supersonic axisymmetric jets are used for validation of the code; these show excellent agreement with pre-validated CFD data. The results for a rectangular supersonic jet are then obtained. These show good agreement with the CFD data, in terms of shock-cell spacing and overall structure of the jet. The technique has proved useful for investigating axis-switching, a phenomenon generally associated with non-axisymmetric jets.
The influence of simulated missile warhead fragment damage on the aerodynamic characteristics of two-dimensional wings
Authors: Irwin A. J. , Render P. M. 
Published: February 2016

AbstractThe paper describes a method of representing damage on a wing due to multiple warhead fragments, and investigates two of the key variables: fragment impact density and hole diameter. The aerodynamic effects of the damage were quantified by wind-tunnel tests on a two-dimensional wing at a Reynolds number of 5 × 105. The wing was of hollow construction with leading and trailing-edge spars. In all of the cases tested, simulated fragment damage resulted in significant lift losses, drag increases and pitching moment changes. Increasing fragment density or hole size resulted in greater effects. To a first order approximation, both lift and drag increments at a given incidence were related to the percentage wing area removed. Surface flow visualisation showed that low fragment densities and small damage sizes resulted in a complex flow structure on the surface of the wing. This was made up of boundary-layer growth between the damage holes, attached wakes from the forward damage holes and separated surface flow over the rear of the wing. For these cases, individual hole patterns showed similar flow mechanisms to those seen for larger scale gunfire damage cases. Increased fragment density and hole size resulted in upper surface flow separation at the first row of holes. Behind this separation, the flow was attached and consisted of the combined wakes from the forward damage holes. Investigations into the influence of internal model structure indicated that trends in coefficient changes were similar for both hollow and solid wings. However, the magnitudes of the effects were found to be smaller for hollow wings than for solid wings.
Influence of weight modelling on the outcome of wing design using multidisciplinary design optimisation techniques
Authors: Mariens J. , Elham A. , van Tooren M. J. L. 
Published: February 2016

AbstractWeight estimation methods are categorised in different classes based on their level of fidelity. The lower class methods are based on statistical data, while higher class methods use physics based calculations. Statistical weight estimation methods are usually utilised in early design stages when the knowledge of designers about the new aircraft is limited. Higher class methods are applied in later design steps when the design is mature enough. Lower class methods are sometimes preferred in later design stages, even though the designers have enough knowledge about the design to use higher class methods. In high level multidisciplinary design optimisation (MDO) fidelity is often sacrificed to obtain models with shorter computation times. There is always a compromise required to select the proper weight estimation method for an MDO project.An investigation has been performed to study the effect of using different weight estimation methods, with low and medium levels of fidelity, on the results of a wing design using multidisciplinary design optimisation techniques. An MDO problem was formulated to design the wing planform of a typical turboprop and a turbofan passenger aircraft. The aircraft maximum take-off weight was selected as the objective function. A quasi-three-dimensional aerodynamic solver was developed to calculate the wing aerodynamic characteristics. Five various statistical methods and a quasi-analytical method are used to estimate the wing structural weight. These methods are compared to each other by analysing their accuracy and sensitivity to different design variables. The results of the optimisations showed that the optimum wing shape is affected by the method used to estimate the wing weight. Using different weight estimation methods also strongly affects the optimisation convergence history and computational time.
One-dimensional model of a pulsed plasma thruster
Authors: Nada T. R. 
Published: February 2016

AbstractThis paper presents a modified one dimensional model for the pulsed plasma thruster. The model takes into account the impact of thruster geometry and capacitor initial energy on plasma inductance and resistance, as well as the ablated mass per pulse. These three factors have significant impact on the performance of the thruster and have not been considered concurrently before. The model is verified by comparing the estimated specific impulse, impulse bit and thrust-to-power ratio with experimental measurements for space qualified thrusters. The maximum error is estimated to be about 8% when using the modified model, while the conventional snowplow model produces error up to about 70%. The modified model is then employed to generate design charts that would help in selecting the proper combination of thruster geometry and energy level in the preliminary design phase of a pulsed plasma thruster for different space missions.
Buoyancy estimation of a Montgolfière in the atmosphere of Titan
Authors: Dorrington G. E. 
Published: February 2016

AbstractPrediction of the buoyancy of a thermal balloon, or Montgolfière, intended to float in the atmosphere of Titan is discussed. A laboratory based experiment designed to measure the buoyant lift of an electrically-heated, single-wall, natural-shape balloon is described. The experimental results presented closely match an analytical model employing established heat transfer correlations with selected parameter values. When the model is extrapolated to the cryogenic conditions representative of Titan’s lower troposphere, using the same correlations and parameters, the estimated buoyancy is found to be substantially higher than has been previously predicted. To account for the buoyancy difference, it is suggested that the internal free convection heat transfer rate is significantly lower than has been assumed in previous studies. To substantiate this result, it is recommended that further experiments should be performed at higher Rayleigh numbers.