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Aircraft Engineering and Aerospace Technology
Journal Prestige (SJR): 0.354
Citation Impact (citeScore): 1
Number of Followers: 213  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 0002-2667 - ISSN (Online) 1748-8842
Published by Emerald Homepage  [356 journals]
  • Thermally induced dynamics of deployable solar panels of nanosatellite
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This study aims to predict the types of thermally induced dynamics (TID) that can occur on deployable solar panels of a small form factor satellite, CubeSat which flies in low Earth orbit (LEO). The TID effect on the CubeSat body is examined. Design/methodology/approach A 3U CubeSat with four short-edge deployable solar panels is considered. Time historic temperature of the solar panels throughout the orbit is obtained using a thermal analysis software. The results are used in numerical simulation to find the structural response of the solar panel. Subsequently, the effect of solar panel motion on pointing the direction of the satellite is examined using inertia relief method. Findings The thermal snap motion could occur during eclipse transitions due to rapid temperature changes in solar panels’ cross-sections. In the case of asymmetric solar panel configuration, noticeable displacement in the pointing direction can be observed during the eclipse transitions. Research limitations/implications This work only examines an LEO mission where the solar cells of the solar panels point to the Sun throughout the daylight period and point to the Earth while in shadow. Simplification is made to the CubeSat structure and some parameters in the space environment. Practical implications The results from this work reveal several practical applications worthy of simplifying the study of TID on satellite appendages. Originality/value This work presents a computational method that fully uses finite element software to analyze TID phenomenon that can occur in LEO on a CubeSat which has commonly used deployable solar panels structure.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-13T02:38:42Z
      DOI: 10.1108/AEAT-07-2018-0185
       
  • Influences of airfoil profile on lateral-directional stability of aircraft
           with flying wing layout
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this study is to analyze influence of airfoil profile on lateral-directional flying quality of flying wing aircraft. The lateral-directional stability is always insufficient for aircraft with the layout due to the absence of vertical stabilizer. A flying wing aircraft with double-swept wing is used as research object in the paper. Design/methodology/approach The 3D model is established for the aircraft with flying wing layout, and parametric modeling is carried out for airfoil mean camber line of the aircraft to analyze lateral-directional stability of the aircraft with different camber line parameters. To increase computational efficiency, vortex lattice method is adopted to calculate aerodynamic coefficients and aerodynamic derivatives of the aircraft. Findings It is found from the research results that roll mode and spiral mode have a little effect on lateral-directional stability of the aircraft but Dutch roll mode is the critical factor affecting flying quality level of such aircraft. Even though changes of airfoil mean line parameters can greatly change assessment parameters of aircraft lateral-directional flying quality, that is kind of change cannot have a fundamental impact on level of flying quality of the aircraft. In case flat shape parameters are determined, the airfoil profile has a limited impact on Dutch roll mode. Originality/value Influences of airfoil profile on lateral-directional flying quality of aircraft with double-swept flying wing layout are revealed in the thesis and some important rules and characteristics are also summarized to lay a theoretical basis for design of airfoil and flight control system of aircraft with the layout.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-12T03:10:22Z
      DOI: 10.1108/AEAT-04-2018-0119
       
  • Spin flight mode identification with OEEMD algorithm
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The paper aims to present an innovative method for identification of flight modes in the spin maneuver, which is highly nonlinear and coupled dynamic. Design/methodology/approach To fix the mode mixing problem which is mostly happen in the EMD algorithm, the authors focused on the proposal of an optimized ensemble empirical mode decomposition (OEEMD) algorithm for processing of the flight complex signals that originate from FDR. There are two improvements with the OEEMD respect to the EEMD. First, this algorithm is able to make a precise reconstruction of the original signal. The second improvement is that the OEEMD performs the task of signal decomposition with fewer iterations and so with less complexity order rather than the competitor approaches. Findings By applying the OEEMD algorithm to the spin flight parameter signals, flight modes extracted, then with using systematic technique, flight modes characteristics are obtained. The results indicate that there are some non-standard modes in the nonlinear region due to couplings between the longitudinal and lateral motions. Practical implications Application of the proposed method to the spin flight test data may result accurate identification of nonlinear dynamics with high coupling in this regime. Originality/value First, to fix the mode mixing problem in EMD, an optimized ensemble empirical mode decomposition algorithm is introduced, which disturbed the original signal with a sort of white Gaussian noise, and by using white noise statistical characteristics the OEEMD fix the mode mixing problem with high precision and fewer calculations. Second, by applying the OEEMD to the flight output signals and with using the systematic method, flight mode characteristics which is very important in the simulation and controller designing are obtained.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-12T03:09:02Z
      DOI: 10.1108/AEAT-12-2017-0280
       
  • Disturbances rejection based on sliding mode control
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper aims to investigate an effective algorithm for different types of disturbances rejection. New dynamics are designed based on disturbance. Observer-based sliding mode control (SMC) technique is used for approximation the disturbances as well as to stabilize the system effectively in presence of uncertainties. Design/methodology/approach This research work investigates the disturbances rejection algorithm for fixed-wing unmanned aerial vehicle. An algorithm based on SMC is introduced for disturbances rejection. Two types of disturbances are considered, the constant disturbance and the sinusoidal disturbance. The comprehensive lateral and longitudinal models of the system are presented. Two types of dynamics, the dynamics without disturbance and the new dynamics with disturbance, are presented. An observer-based algorithm is presented for the estimation of the dynamics with disturbances. Intensive simulations and experiments have been performed; the results not only guarantee the robustness and stability of the system but the effectiveness of the proposed algorithm as well. Findings In previous research work, new dynamics based on disturbances rejection are not investigated in detail; in this research work both the lateral and longitudinal dynamics with different disturbances are investigated. Practical implications As the stability is always important for flight, so the algorithm proposed in this research guarantees the robustness and rejection of disturbances, which plays a vital role in practical life for avoiding any kind of damage. Originality/value In the previous research work, new dynamics based on disturbances rejection are not investigated in detail; in this research work both the lateral and longitudinal dynamics with different disturbances are investigated. An observer-based SMC not only approximates the different disturbances and also these disturbances are rejected in order to guarantee the effectiveness and robustness.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-11T03:35:49Z
      DOI: 10.1108/AEAT-04-2018-0121
       
  • Oil spill remote monitoring by using remotely piloted aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to analyze the application of remotely piloted aircraft (RPA) for remote oil spill sensing. Design/methodology/approach This paper is an analysis of RPA strong points. Findings To increase the accuracy and eliminate potentially false contamination detection, which can be caused by external factors, an oil thickness measurement algorithm is used with the help of the multispectral imaging that provides high accuracy and is versatile for any areas of water and various meteorological and atmospheric conditions. Research limitations/implications SWOT analysis of implementation of RPA for remote sensing of oil spills. Practical implications The use of RPA will improve the remote sensing of oil spills. Social implications The concept of oil spills monitoring needs to be developed for quality data collection, oil pollution control and emergency response. Originality/value The research covers the development of a method and design of a device intended for taking samples and determining the presence of oil contamination in an aquatorium area; the procedure includes taking a sample from the water surface, preparing it for transportation and delivering the sample to a designated location by using the RPA. The objective is to carry out the analysis of remote oil spill sensing using RPA. The RPA provides a reliable sensing of oil pollution with significant advantages over other existing methods. The objective is to analyze the use of RPA employing all of their strong points. In this paper, technical aspects of sensors are analyzed, as well as their advantages and limitations.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-06T11:52:50Z
      DOI: 10.1108/AEAT-12-2017-0273
       
  • Deflection analysis of the airship structure based on the tapered
           inflatable beam
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to analyse the deflection of the flexible airship structure in a new way which can decrease the calculation amount and improve the calculation speed. Design/methodology/approach Infinitesimal method and tapered inflatable beam theory are combined to study the mechanics characteristics of the airship. Firstly, infinitesimal method is introduced into the airship structure analysis. The airship structure can be divided into several tapered inflatable beam elements. Then, tapered inflatable beam theory is improved and a developed model of the tapered inflatable beam under bending moment is presented. Besides, it is proved that deflection caused by pure load and pure moment can be linearly superimposed. Finally, the deflection of the airship structure is studied by means of tapered inflatable beam theory. Findings This paper improved the tapered inflatable beam theory. Besides, the proposed method for deflection analysis of the flexible airship in this paper can reach the same accuracy with traditional finite element method (FEM). However, the number of beam elements is much less than the one of FEM shell elements, which will decrease the calculation amount much and improve the calculation speed. Practical implications The flexible airship is a new and developing research area in engineering practice. The proposed method in this paper provides one precise and high-speed way to analyse the deformation of the airship. Originality/value The paper draws its value from the contributions to development of inflatable structure and the flexible airship mechanics research.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-01T12:49:56Z
      DOI: 10.1108/AEAT-04-2018-0138
       
  • Flow characteristics of two-dimensional synthetic jets under diaphragm
           resonance excitation
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to experimentally study the external flow characteristic of an isolated two-dimensional synthetic jet actuator undergoing diaphragm resonance. Design/methodology/approach The resonance frequency of the diaphragm (40 Hz) depends on the excitation mechanism in the actuator, whereas it is independent of cavity geometry, excitation waveform and excitation voltage. The velocity response of the synthetic jet is influenced by excitation voltage rather than excitation waveform. Thus, this investigation selected four different voltages (5, 10, 15 and 20 V) under the same sine waveform as experiment parameters. Findings The velocity field along the downstream direction is classified into five regions, which can be obtained by hot-wire measurement. The first region refers to an area in which flow moves from within the cavity to the exit of orifice through the oscillation of the diaphragm, but prior to the formation of the vortex of a synthetic jet. In this region, two characteristic frequencies exist at 20 and 40 Hz in the flow field. The second region refers to the area in which the vortices of a synthetic jet fully develop following their initial formation. In this region, the characteristic frequencies at 20 and 40 Hz still occur in the flow field. The third region refers to the area in which both fully developed vortices continue traveling downstream. It is difficult to obtain the characteristic frequency in this flow field, because the mean center velocities (ū) decay downstream and are proportional to (x/w)−1/2 for the four excitation voltages. The fourth region reveals variations in both vortices as they merge into a single vortex. The mean center velocities (ū) are approximately proportional to (x/w)0 in this region for the four excitation voltages. A fifth region deals with variations in the vortex of a synthetic jet after both vortices merge into one, in which the mean center velocities (ū) are approximately proportional to (x/w)−1 in this region for the four excitation voltages (x/w is the dimensionless streamwise distance). Originality/value Although the flow characteristics of synthetic jets had reported for flow control in some literatures, variations of flow structure for synthetic jets are still not studied under the excitation of diaphragm resonance. This paper showed some novel results that our velocity response results obtained by hot-wire measurement along the downstream direction compared with flow visualization resulted in the classification of five regions under the excitation of diaphragm resonance. In the future, it makes valuable contributions for experimental findings to provide researchers with further development of flow control.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-01T12:49:40Z
      DOI: 10.1108/AEAT-12-2017-0277
       
  • Reliability analysis for a hypersonic aircraft’s wing spar
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a novel structural reliability analysis scheme with considering the structural strength degradation for the wing spar of a generic hypersonic aircraft to guarantee flight safety and structural reliability. Design/methodology/approach A logarithmic model with strength degradation for the wing spar is constructed, and a reliability model of the wing spar is established based on stress-strength interference theory and total probability theorem. Findings It is demonstrated that the proposed reliability analysis scheme can obtain more accurate structural reliability and failure results for the wing spar, and the strength degradation cannot be neglected. Furthermore, the obtained results will provide an important reference for the structural safety of hypersonic aircraft. Research limitations/implications The proposed reliability analysis scheme has not implemented in actual flight, as all the simulations are conducted according to the actual experiment data. Practical implications The proposed reliability analysis scheme can solve the structural life problem of the wing spar for hypersonic aircraft and meet engineering practice requirements, and it also provides an important reference to guarantee the flight safety and structural reliability for hypersonic aircraft. Originality/value To describe the damage evolution more accurately, with consideration of strength degradation, flight dynamics and material characteristics of the hypersonic aircraft, the stress-strength interference method is first applied to analyze the structural reliability of the wing spar for the hypersonic aircraft. The proposed analysis scheme is implemented on the dynamic model of the hypersonic aircraft, and the simulation demonstrates that a more reasonable reliability result can be achieved.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-01T12:41:35Z
      DOI: 10.1108/AEAT-11-2017-0242
       
  • Stress, strain and displacement analysis of geodetic and conventional
           fuselage structure for future passenger aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present the results of calculations that checked how the longerons and frames arrangement affects the stiffness of a conventional structure. The paper focuses only on first stage of research – analysis of small displacement. Main goal was to compare different structures under static loads. These results are also compared with the results obtained for a geodetic structure fuselage model of the same dimensions subjected to the same internal and external loads. Design/methodology/approach The finite element method analysis was carried out for a section of the fuselage with a diameter of 6.3 m and a length equal to 10 m. A conventional and lattice structure – known as geodetic – was used. Findings Finite element analyses of the fuselage model with conventional and geodetic structures showed that with comparable stiffness, the weight of the geodetic fuselage is almost 20 per cent lower than that of the conventional one. Research limitations/implications This analysis is limited to small displacements, as the linear version of finite element method was used. Research and articles planned for the future will focus on nonlinear finite element method (FEM) analysis such as buckling, structure stability and limit cycles. Practical implications The increasing maturity of composite structures manufacturing technology offers great opportunities for aircraft designers. The use of carbon fibers with advanced resin systems and application of the geodetic fuselage concept gives the opportunity to obtain advanced structures with excellent mechanical properties and low weight. Originality/value This paper presents very efficient method of assessing and comparison of the stiffness and weight of geodetic and conventional fuselage structure. Geodetic fuselage design in combination with advanced composite materials yields an additional fuselage weight reduction of approximately 10 per cent. The additional weight reduction is achieved by reducing the number of rivets needed for joining the elements. A fuselage with a geodetic structure compared to the classic fuselage with the same outer diameter has a larger inner diameter, which gives a larger usable space in the cabin. The approach applied in this paper consisting in analyzing of main parameters of geodetic structure (hoop ribs, helical ribs and angle between the helical ribs) on fuselage stiffness and weight is original.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-25T09:34:47Z
      DOI: 10.1108/AEAT-07-2018-0216
       
  • Numerical studies of active flow control on wing tip extension
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose In the European project AFLoNext, active flow control (AFC) measures were adopted in the wing tip extension leading edge to suppress flow separation. It is expected that the designed wing tip extension may improve aerodynamic efficiency by about 2 per cent in terms of fuel consumption and emissions. As the leading edge of the wing tip is not protected with high-lift device, flow separation occurs earlier than over the inboard wing in the take-off/landing configuration. The aim of this study is the adoption of AFC to delay wing tip stall and to improve lift-to-drag ratio. Design/methodology/approach Several actuator locations and AFC strategies were tested with computational fluid dynamics. The first approach was “standard” one with physical modeling of the actuators, and the second one was focused on the volume forcing method. The actuators location and the forcing plane close to separation line of the reference configuration were chose to enhance the flow with steady and pulsed jet blowing. Dependence of the lift-to-drag benefit with respect to injected mass flow is investigated. Findings The mechanism of flow separation onset is identified as the interaction of slat-end and wing tip vortices. These vortices moving toward each other with increasing angle of attack (AoA) interact and cause the flow separation. AFC is applied to control the slat-end vortex and the inboard movement of the wing tip vortex to suppress their interaction. The separation onset has been postponed by about 2° of AoA; the value of ift-to-drag (L/D) was improved up to 22 per cent for the most beneficial cases. Practical implications The AFC using the steady or pulsed blowing (PB) was proved to be an effective tool for delaying the flow separation. Although better values of L/D have been reached using steady blowing, it is also shown that PB case with a duty cycle of 0.5 needs only one half of the mass flow. Originality/value Two approaches of different levels of complexity are studied and compared. The first is based on physical modeling of actuator cavities, while the second relies on volume forcing method which does not require detailed actuator modeling. Both approaches give consistent results.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-16T12:27:05Z
      DOI: 10.1108/AEAT-01-2018-0053
       
  • Jet engine degradation prognostic using artificial neural networks
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to propose and develop artificially intelligent methodologies to discover degradation trends through the detection of engine’s status. The objective is to predict these trends by studying their effects on the engine measurable parameters. Design/methodology/approach The method is based on the implementation of an artificial neural network (ANN) trained with well-known cases referred to real conditions, able to recognize degradation because of two main gas turbine engine deterioration effects: erosion and fouling. Three different scenarios are considered: compressor fouling, turbine erosion and presence of both degraded conditions. The work consists of three parts: the first one contains the mathematical model of real jet engine in healthy and degraded conditions, the second step is the optimization of ANN for engine performance prediction and the last part deals with the application of ANN for prediction of engine fault. Findings This study shows that the proposed diagnostic approach has good potential to provide valuable estimation of engine status. Practical implications Knowledge of the true state of the engine is important to assess its performance capability to meet the operational and maintenance requirements and costs. Originality/value The main advantage is that the engine performance data for model validation were obtained from real flight conditions of the engine VIPER 632-43.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-16T12:10:48Z
      DOI: 10.1108/AEAT-01-2018-0054
       
  • Remarks about factors shaping the heliosphere
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of the paper is to give a brief description of the new topic introduced for the first time at the EASN Conferences. Design/methodology/approach The topic concerns the heliosphere, the nearest surrounding of the Sun and thus the nearest vicinity of the Earth. The heliosphere is created due to the interaction between the solar wind and the local interstellar medium. Findings This paper does not include any new information about the heliosphere and only introduces a new topic to this journal. It is briefly shown how heliospheric structures are formed, what factors affect a shape of the heliosphere, what measurements are made by Ulysses, Voyager and IBEX space missions (important for the heliosphere modeling) and how obtained data are used to validate theoretical results. Practical implications To categorize the paper under one of these classifications, research paper, viewpoint, technical paper, conceptual paper, case study, literature review or general review, the authors chose a paper type, general review, as the closest category to this paper. However, it is not a purpose of this paper to provide an extensive review of the community efforts to investigate the physical processes in the vicinity of the heliosphere interface. This is mostly a status report. Originality/value As the new topic in this journal, the article introduces in detail only a small number of aspects connected with heliosphere models. Interplanetary and interstellar magnetic field structures are primarily described. Other factors are only mentioned.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-16T12:10:46Z
      DOI: 10.1108/AEAT-01-2018-0044
       
  • A simulation investigation of helicopter ground resonance phenomenon
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon. Design/methodology/approach The considered physical model of helicopter standing on ground with rotating rotor consists of fuselage and main transmission gear treated as stiff bodies connected by elastic elements. The fuselage is supported on landing gear modeled by spring-damper units. The main rotor blades are treated as set of elastic axes with lumped masses distributed along blade radius. Due to Galerkin method, parameters of blades motion are assumed as a combination of bending and torsion eigen modes. A Runge–Kutta method is applied to solve equations of motions of rotor blades and helicopter fuselage. Findings The presented simulation method may be applied in preliminary stage of helicopter design to avoid ground resonance by proper selection of landing gear units and blade damper characteristics. Practical implications Ground resonance may occur in form of violently increasing mutual oscillations of helicopter fuselage and lead-lag motion of rotor blades. According to changes of stiffness and damping characteristics, simulations show stable behavior or arising oscillations of helicopter. The effects of different blade balance or defect of blade damper are predicted. Originality/value The simulation method may help to determine the envelope of safe operation of helicopter in phase of take-off or landing. The effects of additional disturbances as results of blades pitch control as swashplate deflection are introduced.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-14T12:06:55Z
      DOI: 10.1108/AEAT-11-2017-0256
       
  • Preliminary design of 3D printed fittings for UAV
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Additive manufacturing technology, also commonly called as 3D printing technology, is entering rapidly into the aerospace world and seems to be its future. Many manufacturing processes are replaced by this technology because the ease of use, low costs and new possibilities to make complicated parts. However, there are only few solutions which present manufacturing of structurally critical parts. Design/methodology/approach Complete process of deriving loads, design of fitting geometry, numerical validation, manufacturing and strength testing was presented. The emphasis was made to show specific features of 3D technology in printed fittings for UAV. Findings The research confirms that the technology can be used for the application of fittings manufacturing. Attention needs to be paid, during the design process, to account for specific features of the 3D printing technology, which is described in details. Practical implications Without a doubt, additive manufacturing is useful for manufacturing complicated parts within limited time and with reduction cost. It was also shown that the manufactured parts can be used for highly loaded structures. Originality/value The paper shows how additive manufacturing technology can be used to produce significantly loaded parts of airplanes’ structure. Only few such examples were presented till now.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-14T06:07:23Z
      DOI: 10.1108/AEAT-07-2018-0182
       
  • Identification of a degradation of aerodynamic characteristics of a
           paraglider due to its flexibility from flight test
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Aerodynamics of paragliders is very complicated aeroelastic phenomena. The purpose of this work is to quantify the amount of aerodynamic drag related to the flexible nature of a paraglider wing. Design/methodology/approach The laboratory testing on scaled models can be very difficult because of problems in the elastic similitude of such a structure. Testing of full-scale models in a large facility with a large full-scale test section is very expensive. The degradation of aerodynamic characteristics is evaluated from flight tests of the paraglider speed polar. All aspects of the identification such as pilot and suspension lines drag and aerodynamics of spanwise chambered wings are discussed. The drag of a pilot in a harness was estimated by means of wind tunnel testing, computational fluid dynamics (CFD) solver was used to estimating smooth wing lift and drag characteristics. Findings The drag related to the flexible nature of the modern paraglider wing is within the range of 4-30 per cent of the total aerodynamic drag depending on the flight speed. From the results, it is evident that considering only the cell opening effect is sufficient at a low-speed flight. The stagnation point moves forwards towards the nose during the high speed flight. This causes more pronounced deformation of the leading edge and thus increased drag. Practical implications This paper deals with a detailed analysis of specific paraglider wing. Although the results are limited to the specific geometry, the findings help in the better understanding of the paraglider aerodynamics generally. Originality/value The data obtained in this paper are not affected by any scaling problems. There are only few experimental results in the field of paragliders on scaled models. Those results were made on simplified models at very low Reynolds number. The aerodynamic drag characteristics of the pilot in the harness with variable angles of incidence and Reynolds numbers have not yet been published.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-09T01:37:57Z
      DOI: 10.1108/AEAT-06-2018-0162
       
  • Earth-to-space and high-speed “air” transportation: an
           aerospaceplane design
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed. Design/methodology/approach The LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile. Findings It was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth. Practical implications The proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before. Originality/value This paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-09T01:36:35Z
      DOI: 10.1108/AEAT-08-2017-0196
       
  • Experimental study of the solubility and diffusivity of CO2 and O2 in RP-3
           jet fuel
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This study aims to get the essential data of the solubility and diffusion coefficient of gas in jet fuel for appropriately designing a kind of on-board inert gas generation system. Design/methodology/approach A test apparatus based on pressure–decay method was constructed to measure solubility and diffusion coefficient of gas in liquid. The test apparatus and method were verified via measurement of solubility and diffusion of CO2 in the pure water. Findings The solubility of CO2 and O2 in RP-3 jet fuel with the temperature from 253 to 313 K under three various pressures were measured and compared with theoretical value calculated by a relative density method provided in the standard of ASTM D2780-92, and the deviation is within 10 per cent. The diffusion coefficients of CO2 and O2 in RP-3 jet fuel are determined by monitoring the gas pressure in a hermetic cell versus time with the temperature from 253 to 333 K. The measured diffusivity-temperature relation can be well fitted through the Arrhenius equation for engineering applications. The obtained correlation can be used to predict the diffusion coefficient of CO2 and O2 in RP-3 jet fuel under a wide temperature range. Practical implications The semi-empirical correlation of solubility and diffusion coefficient in RP-3 jet fuel obtained from the experimental data could be used to support the design of an inert gas generation system. Originality/value There are no essential data of solubility and diffusion of CO2 and O2 in RP-3 jet fuel; therefore, it is fatal if the quantity and rate of mass transfer of CO2 and O2 in RP-3 jet fuel must be assessed, e.g. during the design of green on-board inert gas generation system.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-08T08:31:33Z
      DOI: 10.1108/AEAT-05-2017-0133
       
  • LCA of the maintenance of a piston-prop engine
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This study aims to introduce an approach to evaluate environmental impact of a piston-prop engine from the view point of life cycle assessment (LCA). Design/methodology/approach In the aviation industry, safety is an important issue. For reliable and safe flights, the maintenance of aerial vehicles and engines is mandatory. Additionally, regular and correct maintenance plays a key role in keeping efficiency at a high level. With this in mind, a LCA of a regular 50 hourly maintenance process of Cessna type training aircraft is conducted. During the assessment, the starting of the engine before maintenance, replacement of the oil filter, test procedure of the spark plugs, a compressor test, engine cleaning and engine starting following maintenance are taken into account. Findings At the end of the study, normalization and characterization values for the maintenance, electricity consumption during maintenance and used fuel are obtained. Practical implications Regarding the number of this type aircraft worldwide, the current study offers a valuable contribution to the literature. The authors also intend to introduce an approach which may be useful for the assessment of large body aircraft still in service. Originality/value The present paper is a pioneer for future applications of LCA methodology to piston-prop engines and training aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-07T01:36:39Z
      DOI: 10.1108/AEAT-05-2017-0116
       
  • Hybrid energy systems in unmanned aerial vehicles
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The presented research is carried out in reaction to the soaring costs of fuel and tight control over environmental issues such as carbon dioxide emissions and noise. The purpose of this paper is to study the feasibility of applying the environmental-friendly energy source in an unmanned aerial vehicles (UAVs) propulsion system. Design/methodology/approach Currently, the majority of UAVs are still powered by conventional combustion engines. An electric propulsion system is most commonly found in civilian micro and mini UAVs. The UAV classification is reviewed in this study. This paper focuses mainly on application of electric propulsion systems in UAVs. Investigated hybrid energy systems consist of fuel cells, Li-ion batteries, super-capacitors and photovoltaic (PV) modules. Current applications of fuel cell systems in UAVs are also presented. Findings The conducted research shows that hybridization allows for better energy management and operation of every energy source onboard the UAV within its limits. The hybrid energy system design should be created to maximize system efficiency without compromising the performance of the aircraft. Practical implications The presented study highlights the reduction of the energy consumption, necessary to perform the mission and maximizing of the endurance with simultaneous decrease in emissions and noise level. Originality/value The conducted research studies the feasibility of implementing the environmental-friendly hybrid electric propulsion systems in UAVs that offers high efficiency, reliability, controllability, lack of thermal and noise signature, thus, providing quiet and clean drive with low vibration levels. This paper highlights the main challenges and current research on fuel cell in aviation and draws attention to fuel cell – electric system modeling, hybridization and energy management.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-04T03:00:25Z
      DOI: 10.1108/AEAT-08-2018-0218
       
  • Fibre Bragg grating sensor applications for structural health monitoring
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Structural health monitoring (SHM) has become an attractive subject in aerospace engineering field considering the opportunity to avoid catastrophic failures by detecting damage in advance and to reduce maintenance costs. Fibre Bragg Grating (FBG) sensors are denoted as one of the most promising sensors for SHM applications as they are lightweight, immune to electromagnetic effects and able to be embedded between the layers of composite structures. The purpose of this paper is to research on and demonstrate the feasibility of FBG sensors for SHM of composite structures. Design/methodology/approach Applications on thin composite beams intended for SHM studies are presented. The sensor system, which includes FBG sensors and related interrogator system, and manufacturing of the beams with embedded sensors, are detailed. Static tension and torsion tests are conducted to verify the effectiveness of the system. Strain analysis results obtained from the tests are compared with the ones obtained from the finite element analyses conducted using ABAQUS® software. In addition, the comparison between the data obtained from the FBG sensors and from the strain gauges is made by also considering the noise content. Finally, fatigue test under torsion load is conducted to observe the durability of FBG sensors. Findings The results demonstrated that FBG sensors are feasible for SHM of composite structures as the strain data are accurate and less noisy compared to that obtained from the strain gauges. Furthermore, the convenience of obtaining reliable data between the layers of a composite structure using embedded FBG sensors is observed. Practical implications Observing the advantages of the FBG sensors for strain measurement will promote using FBG sensors for damage detection related to the SHM applications. Originality/value This paper presents applications of FBG sensors on thin composite beams, which reveal the suitability of FBG sensors for SHM of lightweight composite structures.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-04T02:58:23Z
      DOI: 10.1108/AEAT-11-2017-0255
       
  • Stability analysis of the experimental airplane powered by a pulsejet
           engine
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present stability analysis of a small pulsejet-powered airplane. This analysis is a part of a student project dedicated to designing an airplane to test valved pulsejet engine in flight conditions. Design/methodology/approach The panel method was chosen to compute the airplane’s aerodynamic coefficients and derivatives for various geometry configurations, as it provides accurate results in a short computational time. Also, the program (PANUKL) that was used allows frequent and easy changes of the geometry. The evaluation of dynamic stability was done using another program (SDSA) equipped with means to formulate and solve eigenvalue problem for various flight speeds. Findings As a result of calculations, some geometry corrections were established, such as an increase of the vertical stabilizer’s size and a new wing position. Resulting geometry provides satisfactory dynamic and static stability characteristics for all flight speeds. This conclusion was based on criteria given by MIL-F-8785C specifications. This paper presents the results of the first and the final configuration. Practical implications The results shown in this paper are necessary for the continuation of the project. The aircraft’s structure was being designed in the same time as the calculations described in this paper proceeded. With a few modifications to make up for the changes of external geometry, the structure will be ready to be built. Originality/value The idea to design an airplane specifically to test a pulsejet in flight is a unique one. Most RC pulsejet-powered constructions that can be heard of are modified versions of already existing models. What adds more to the value of the project is that it is being developed only by students. This allows them to learn various aspects of aircraft design and construction on a soon-to-be real object.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-04T02:53:44Z
      DOI: 10.1108/AEAT-07-2018-0184
       
  • Influence of Gurney flaps on aerodynamic characteristics of a
           canard-configuration aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to mount Gurney flaps at the trailing edges of the canards and investigate their influence on aerodynamic characteristics of a simplified canard-configuration aircraft model. Design/methodology/approach A force measurement experiment was conducted in a low-speed wind tunnel. Hence, the height and shape effects of the Gurney flaps on the canards were investigated. Findings Gurney flaps can increase the lift and pitching-up moment for the aircraft model tested, thereby increasing the lift when trimming the aircraft. The dominant parameter to influence aerodynamic characteristics is the height of Gurney flaps. When the flap heights are the same, the aerodynamic efficiency of the triangular Gurney flaps is higher than that of the rectangular ones. Moreover, the canard deflection efficiency will be reduced with Gurney flaps equipped, but the total aerodynamic increment is considerable. Practical implications This paper helps to solve the key technical problem of increasing take-off and landing lift coefficients, thus improving the aerodynamic performance of the canard-configuration aircraft. Originality/value This paper recommends to adopt triangular Gurney flaps with the height of 3 per cent chord length of the canard root (c) for engineering application.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-03T09:55:45Z
      DOI: 10.1108/AEAT-08-2017-0181
       
  • Strategic approach to managing human factors risk in aircraft maintenance
           organization: risk mapping
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Aviation has multi-cultural business environment in all aspects as operational and management. Managing aviation requires high awareness on human factor risk which includes organizational behavior-related topics. The greatest risk to an enterprise’s ability to achieve its strategic goals and objectives is the human factor. Both organizational behavior and corporate culture behavior with social psychology are the most vital aspects of management and strategy in terms of human resources. Related risks, including organizational behavior and culture, have the potential to directly impact on both business performance and corporate sustainability. Therefore, in this paper, the most prominent risks were determined in accordance with social psychology, and after identification of human factor-based risks, these have prioritized and prepared risk mapping with fresh approach. For this reason, this study aims to develop risk mapping model for human factors that takes into account interrelations among risk factors three dimensional based new approach. This approach includes both identification of human factor based risks, prioritization them and setting risk mapping according to corporate based qualifications via tailoring risk list. Developed risk map in this paper will help to manage corporate risks to achieve improved performance and sustainability. Design/methodology/approach This new organizational behavior- and culture-focused risk mapping model developed in this study has the potential to make significant contribution to the management of the human factor for modern management and strategy. In enterprise risk management system, risk mapping is both strong and effective strategic methodology to manage ergonomics issue with strategic approach. Human factor is both determinative and also strategic element to both continuity and performance of business operations with safely and sound. In view of management and strategy, vitally, the human factor determines the outcome in both every business and every decision-making. Findings It is assumed that, if managers manage human risk you may get advantages to achieving corporate strategies in timely manner. Aviation is sensitive sector for its ingredients: airports, airlines, air traffic management, aircraft maintenance, pilotage and ground handling. Aim of this paper is to present risk management approach to optimize human performance while minimizing both failures and errors by aircraft maintenance technician (AMT). This model may apply all human factors in other departments of aviation such as pilots and traffic controllers. AMT is key component of aircraft maintenance. Thus, errors made by AMTs will cause aircraft accidents or incidents or near miss incidents. In this study, new taxonomy model for human risk factors in aircraft maintenance organizations has been designed, and also new qualitative risk assessment as three dimensions is carried out by considering the factors affecting the AMT’s error obtained from extensive literature review and expert opinions in the field of aviation. Human error risks are first categorized into two main groups and sub three groups and then prioritized using the risk matrix via triple dimension as probability, severity and interrelations ratio between risks. Practical implications Risk mapping is established to decide which risk management option they will apply for managers when they will look at this map. Managers may use risk map to both identify their managerial priorities and share sources to managing risks, and make decisions on risk handling options. This new model may be a useful new tool to manage ergonomic human factor-based risks in developing strategy in aviation business management. In addition, this paper will contribute to department of management and strategy and related literature. Originality/value This study has originality via new modeling of risk matrix. In this study, dimension of risk analysis has been improved as three dimensions. This study has new approach and new assessment of risk with likelihood (probability), impact (severity) and interrelations ratio. This new model may be a useful new tool to both assess and prioritize mapping of ergonomic-based risks in business management. In addition, this research will contribute to aviation management and strategy literature and also enterprise risk management literature.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-03T09:53:37Z
      DOI: 10.1108/AEAT-06-2018-0160
       
  • Reachable domain of spacecraft with a single normal impulse
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to obtain the reachable domain (RD) for spacecraft with a single normal impulse while considering both time and impulse constraints. Design/methodology/approach The problem of RD is addressed in an analytical approach by analyzing for either the initial maneuver point or the impulse magnitude being arbitrary. The trajectories are considered lying in the intersection of a plane and an ellipsoid of revolution, whose family can be determined analytically. Moreover, the impulse and time constraints are considered while formulating the problem. The upper bound of impulse magnitude, “high consumption areas” and the change of semi-major axis and eccentricity are discussed. Findings The equations of RD with a single normal impulse are analytically obtained. The equations of three scenarios are obtained. If normal impulse is too large, the RD cannot be obtained. The change of the semi-major axis and eccentricity with large normal impulse is more obvious. For long-term missions, the change of semi-major axis and eccentricity leaded by multiple normal impulses should be considered. Practical implications The RD gives the pre-defined region (all positions accessible) for a spacecraft under a given initial orbit and a normal impulse with certain magnitude. Originality/value The RD for spacecraft with normal impulse can be used for non-coplanar orbital transfers, emergency evacuation after failure of rendezvous and docking and collision avoidance.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-03T09:42:37Z
      DOI: 10.1108/AEAT-03-2017-0079
       
  • Experimental investigation of plasma vortex generator in flow control
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to compare the effects of two different configurations of plasma streamwise vortex generators (PSVG), including comb-type and mesh-type in controlling flow. This is demonstrated on the NACA 0012 airfoil. Design/methodology/approach The investigations have been done experimentally at the various electric and aerodynamic conditions. The surface oil flow visualization method has been used to the better understanding of the flow physics and the interaction of the oncoming flow passing over the airfoil and the vortex generated by comb-type PSVG. Findings This paper demonstrates the potential capabilities of the mesh-type and comb-type PSVGs in controlling flow in unsteady operation. It was found that the vortex generated by the mesh-type PSVG in unsteady operation was an order of magnitude stronger than comb-type PSVG. The flow visualisation technic proved that only a part of the plasma actuator is effective in the condition that the actuator is installed only on a portion of the upper surface of the airfoil. Originality/value This paper experimentally confirms the capabilities of the mesh-type PSVG unsteady operation in compare with comb-type PSVG in controlling flow, whereby recommends using mesh-type PSVG in the leading edge in front of comb-type PSVG on the entire wingspan to prevent the stall.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-03T09:37:03Z
      DOI: 10.1108/AEAT-07-2018-0194
       
  • Performance improvement of helicopter rotors through blade redesigning
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to determine dependencies between a rotor-blade shape and a rotor performance as well as to search for optimal shapes of blades dedicated for helicopter main and tail rotors. Design/methodology/approach The research is conducted based on computational methodology, using the parametric-design approach. The developed parametric model takes into account several typical blade-shape parameters. The rotor aerodynamic characteristics are evaluated using the unsteady Reynolds-averaged Navier–Stokes solver. Flow effects caused by rotating blades are modelled based on both simplified approach and truly 3D simulations. Findings The computational studies have shown that the helicopter-rotor performance may be significantly improved even through relatively simple aerodynamic redesigning of its blades. The research results confirm high potential of the developed methodology of rotor-blade optimisation. Developed families of helicopter-rotor-blade airfoils are competitive compared to the best airfoils cited in literature. The finally designed rotors, compared to the baselines, for the same driving power, are characterised by 5 and 32% higher thrust, in case of main and tail rotor, respectively. Practical implications The developed and implemented methodology of parametric design and optimisation of helicopter-rotor blades may be used in future studies on performance improvement of rotorcraft rotors. Some of presented results concern the redesigning of main and tail rotors of existing helicopters. These results may be used directly in modernisation processes of these helicopters. Originality/value The presented study is original in relation to the developed methodology of optimisation of helicopter-rotor blades, families of modern helicopter airfoils and innovative solutions in rotor-blade-design area.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-01-02T01:37:52Z
      DOI: 10.1108/AEAT-01-2018-0009
       
  • Unsteady aerodynamic characteristics of a morphing wing
    • Pages: 1 - 9
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 1-9, January 2019.
      Purpose The purpose of this paper is to investigate the unsteady aerodynamic characteristics in the deflection process of a morphing wing with flexible trailing edge, which is based on time-accurate solutions. The dynamic effect of deflection process on the aerodynamics of morphing wing was studied. Design/methodology/approach The computational fluid dynamic method and dynamic mesh combined with user-defined functions were used to simulate the continuous morphing of the flexible trailing edge. The steady aerodynamic characteristics of the morphing deflection and the conventional deflection were studied first. Then, the unsteady aerodynamic characteristics of the morphing wing were investigated as the trailing edge deflects at different rates. Findings The numerical results show that the transient lift coefficient in the deflection process is higher than that of the static case one in large angle of attack. The larger the deflection frequency is, the higher the transient lift coefficient will become. However, the situations are contrary in a small angle of attack. The periodic morphing of the trailing edge with small amplitude and high frequency can increase the lift coefficient after the stall angle. Practical implications The investigation can afford accurate aerodynamic information for the design of aircraft with the morphing wing technology, which has significant advantages in aerodynamic efficiency and control performance. Originality/value The dynamic effects of the deflection process of the morphing trailing edge on aerodynamics were studied. Furthermore, time-accurate solutions can fully explore the unsteady aerodynamics and pressure distribution of the morphing wing.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-02T12:13:23Z
      DOI: 10.1108/AEAT-04-2017-0101
       
  • Design improvements and flap deflection evaluations with considering
           centrifugal load on active trailing edge flap
    • Pages: 10 - 19
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 10-19, January 2019.
      Purpose The purpose of this paper is to present the design, analysis and experiments of the active trailing-edge flap (Seoul National University Flap, SNUF) for vibration reduction in the helicopter rotor prior to the small-scale blades planned to test in a whirl tower. Design/methodology/approach The predictions of the hinge moment in both steady and unsteady flows were obtained through computational fluid dynamics calculations. When compared with the results originated from analytical formulations, the proposed method showed improved prediction capabilities. To validate the deflection of the flap under the centrifugal load by rotating, static analysis was conducted using both contact and rotating condition of MSC NASTRAN. The corresponding experiment also was performed using the vertical frame for simulating the effect of the centrifugal force. Findings The hinge moment of the flap is predicted through unsteady analysis in the actuation frequency of 3/rev. The material of the guide in the flap mechanism was selected through static analysis under both contact and rotating condition. Finally, reduction of the deflection occurred because of the load in the axial direction of the hinge like the centrifugal load. Practical implications The important aspects, such as design, analysis, and experiments for the active trailing-edge flap were shown. Originality/value This paper showed the relationship of the displacement, block force and voltage of the piezo-actuator, combined with the hinge moment predicted. The methodology and the experiment were presented for simulating the centrifugal force acting on the flap.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-13T07:46:52Z
      DOI: 10.1108/AEAT-10-2016-0165
       
  • Guidance law to control impact time constraining the seeker’s field
           of view'
    • Pages: 20 - 29
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 20-29, January 2019.
      Purpose The purpose of this paper is to present a novel guidance law that is able to control the impact time while the seeker’s field of view (FOV) is constrained. Design/methodology/approach The new guidance law is derived from the framework of Lyapunov stability theory to ensure interception at the desired impact time. A time-varying guidance gain scheme is proposed based on the analysis of the convergence time of impact time error, where finite-time stability theory is used. The circular trajectory assumption is adopted for the derivation of accurate analytical estimation of time-to-go. The seeker’s FOV constraint, along with missile acceleration constraint, is considered during guidance law design, and a switching strategy to satisfy it is designed. Findings The proposed guidance law can drive missile to intercept stationary target at the desired impact time, as well as satisfies seeker’s FOV and missile acceleration constraints during engagement. Simulation results show that the proposed guidance law could provide robustness against different engagement scenarios and autopilot lag. Practical implications The presented guidance law lays a foundation for using cooperative strategies, such as simultaneous attack. Originality/value This paper presents further study on the impact time control problem considering the seeker’s FOV constraint, which conforms better to reality.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-09-27T02:51:19Z
      DOI: 10.1108/AEAT-06-2017-0151
       
  • A numerical study on the effects of design parameters on the acoustics
           noise of a high efficiency propeller
    • Pages: 30 - 37
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 30-37, January 2019.
      Purpose A numerical study on the aerodynamic noise generation of a high efficiency propeller is carried out. Design/methodology/approach Three-dimensional numerical simulation based on Reynolds averaged N-S model is performed to obtain the aerodynamic performance of the propeller. Then, the result of the aerodynamic analysis is given as input of the acoustic calculation. The sound is calculated using the Farassat 1A which was derived from Ffowcs Williams–Hawkings equation and is compared with the measurements. Findings Moreover, the fan is modified for noise reduction by changing its geometrical parameters such as span, chord length and torsion angle. Originality/value The variation trend of aerodynamic and acoustic are compared and discussed for different modification tasks. Some meaningful conclusions are drawn on the noise reduction of propeller.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T07:11:02Z
      DOI: 10.1108/AEAT-08-2017-0183
       
  • High AOA short landing robust control for an aircraft
    • Pages: 38 - 49
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 38-49, January 2019.
      Purpose The purpose of this paper is to propose a high angle of attack short landing model for switched polytopic systems as well as to derive an equation for fluidic thrust vector deflection angle based on pressure to reduce the velocity during the landing phase of flight. Design/methodology/approach In this paper, robust control algorithm is proposed for a non-linear high angle of attack aircraft under the effects of non-linearities, tottering hysteresis, irregular and wing rock atmosphere. High angle of attack short landing flight under asynchronous switching is attained by using the robust controller method. Lyapunov function and the average dwell time scheme is used for obtaining the switched polytopic scheme. The asynchronous switching and loss of data are controlled asymptotically. The velocity of aircraft has been lucratively reduced during the landing phase of flight by using the robust controller technique. Findings The proposed algorithm based on robust controller including the effects of non-linearities guarantee the successful reduction of velocity for high angle of attack switched polytopic systems. Practical implications As the landing phase of an aircraft is one of the complicated stage, this algorithm plays a vital role in stable and short landing under the condition of high angle of attack (AOA). Originality/value In this paper, not only the velocity of flight has been reduced, but also the high angle of attack has been attained during the landing phase, because of which the duration of landing has been reduced as well, while in most of the previous research, it is based on low angle of attack and long landing duration.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T02:31:37Z
      DOI: 10.1108/AEAT-05-2017-0134
       
  • Novel development of dynamic behavior of carbon fiber reinforced polymer
           sandwich panels with stepwise graded adhesive layer
    • Pages: 50 - 59
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 50-59, January 2019.
      Purpose The purpose of this study is to focus on the developments of carbon fiber reinforced polymer (CFRP) panels with stepwise graded properties on adhesive layer. The various arranges of the graded properties of the adhesive layer have been checked according to experimental results of the literatures and based on applicability. Design/methodology/approach The finite element (FE) models and experimental modal tests of the manufactured CFRP sandwich panel specimens have been investigated. The core thickness, core density and orientation of the fiber direction of the sandwich panel face – sheets have been parametrically checked based on modal behavior. Two fully free and fully clamped boundary conditions (BC) have been checked in stepwise graded adhesive zone (SGAZ) cases and first five non-zero natural frequencies (NF) have been compared. Dynamic response of the SGAZ includes modal analysis and transient dynamic loading have been performed numerically with ABAQUS 6.12 well-known FE code. Findings The first non-zero NF of SGAZ Case 4 was 11.69 per cent higher than homogenous Case 2 and 7.06 per cent lower than Case 1 in fully free boundary conditions. A total of 26.38 per cent is the greatest discrepancy between fist five non-zero NFs of all cases with two BCs (Case 1 vs Case 2 in fully clamped BC). Maximum structural damping behavior and minimum stress picks have been studied during transient dynamic loading analysis of CFRP panel with SGAZ. SGAZ Case 3 (middle adhesive with lower modulus) has increased the maximum structural damping while reducing the minimum out of plain tip displacements during transient dynamic loading by 111.26 per cent in comparison with homogenous Case 2. Also, Case 3 has reduced the Mises stress picks on the adhesive region by 605.68 per cent. Practical implications Making a stepwise graded adhesive region (without any added mass) has been shown that it is a novel and useful way to achieve a wide range of stiffness on CFRP panels. Originality/value Development of the sandwich panels with various stiffness and damping properties.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-02T12:38:24Z
      DOI: 10.1108/AEAT-08-2017-0179
       
  • Autonomous planetary rover navigation via active SLAM
    • Pages: 60 - 68
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 60-68, January 2019.
      Purpose This paper aims to introduce an efficient active-simultaneous localization and mapping (SLAM) approach for rover navigation, future planetary rover exploration mission requires the rover to automatically localize itself with high accuracy. Design/methodology/approach A three-dimensional (3D) feature detection method is first proposed to extract salient features from the observed point cloud, after that, the salient features are employed as the candidate destinations for re-visiting under SLAM structure, followed by a path planning algorithm integrated with SLAM, wherein the path length and map utility are leveraged to reduce the growth rate of state estimation uncertainty. Findings The proposed approach is able to extract distinguishable 3D landmarks for feature re-visiting, and can be naturally integrated with any SLAM algorithms in an efficient manner to improve the navigation accuracy. Originality/value This paper proposes a novel active-SLAM structure for planetary rover exploration mission, the salient feature extraction method and active revisit patch planning method are validated to improve the accuracy of pose estimation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-13T07:54:58Z
      DOI: 10.1108/AEAT-12-2016-0239
       
  • Cold flow studies in a vortex thrust chamber
    • Pages: 69 - 77
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 69-77, January 2019.
      Purpose The purpose of this paper is to estimate the chamber pressure and flow behaviour in a vortex thrust chamber (VTC) during the cold flow with hydrogen and oxygen as propellants. Design/methodology/approach Experiments are carried out in a VTC with a different mixture ratio of hydrogen and oxygen. The pressures developed inside the VTC are measured. Numerical simulations are carried out to understand the flow patterns of fuel and oxidizer inside the VTC. Findings The chamber pressure is influenced by the type of injection of propellant and mixture ratio. Tangential injection of propellant is the key parameter for an increase of the chamber pressure of the VTC. Research limitations/implications The pressure measurements are carried out in cold flow conditions without combustion happening in the VTC. Practical implications The practical implication is that when the combustion in the VTC ceases, the thrust generated due to the propellants in cold flow conditions can be assessed. Originality/value The VTC with the tangential injection of propellant generates higher chamber pressure.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-06T11:32:19Z
      DOI: 10.1108/AEAT-07-2017-0167
       
  • Prioritisation of factors contributing to human error for airworthiness
           management strategy with ANP
    • Pages: 78 - 93
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 78-93, January 2019.
      Purpose Airline business management is set on airworthy strategy. Airline sustainability depends upon corporate-based airworthy strategy as airworthiness is the base to any airline business management and strategy. An airline can ensize its corporate sustainability if it has airworthiness strategy and risk management. The main condition to survive in the airline business is to maintain airworthiness with the fleet, maintenance and corporate-risk management. Aircraft maintenance technician (AMT) has a dual role in aircraft maintenance system as the source of failure in maintenance process via his volatility and unmanageable qualifications and secondly source of manager of maintain airworthiness of the aircrafts in airline. Situational awareness of managers about both limitations and qualifications of human factors is vital determinant to the decision-making process in aviation. Although continuously improving in related literature, one of the biggest weaknesses of the current methods of AMT error or performance is that the ability to model the reciprocal effects of the factors affecting the fault is limited. For this reason, this study aims to develop an analytic network process (ANP) model that takes into account the effects of mutual dependences among factors. Design/methodology/approach Firstly, with the help of experts and extensive literature, 67 factors that contributed to AMT error are identified and grouped. Then, the factors identified as eligible criteria and sub-criteria that contributed to the AMT errors are determined. In this study, the weights of identified criteria that have influence on AMT error try to determine by using ANP method. ANP is the common method to solve multi-criteria decision-making problems and is used to calculate priorities of factors. Criteria determined in this study are classified into three main clusters: “individual-related criteria”, “working environment-related criteria” and “organisational-related criteria”. These main clusters include 15 sub criteria such as communication, documentation (quality/updating/availability) and peer pressure. Findings The result of this study shows that time pressure, organisational culture, safety culture and supervision are the most important criteria that contributed to AMT error. Their weights are 0.207, 0.172, 0.102 and 0.094, respectively. Originality/value There are many difficulties and limitations in measuring the factors that have an influence on AMT errors. For this reason, the weights of criteria and sub-criteria necessary are determined using ANP, and in this manner, it is possible to make better decisions in this process as ANP is a multi-criteria decision-making technique that considers qualitative factors in decision-making problems. The factors’ taxonomy determined as a result of the expert opinions and the extensive literature and the ANP model developed taking into account the dependencies between the factors will contribute to the literature.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T02:32:38Z
      DOI: 10.1108/AEAT-11-2017-0245
       
  • Combustion chamber design and reaction modeling for aero turbo-shaft
           engine
    • Pages: 94 - 111
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 94-111, January 2019.
      Purpose The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance. Design/methodology/approach The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance. Findings The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments. Research limitations/implications Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities. Practical implications This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development. Originality/value The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-06T11:31:40Z
      DOI: 10.1108/AEAT-10-2017-0217
       
  • Design of parallel adaptive extended Kalman filter for online estimation
           of noise covariance
    • Pages: 112 - 123
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 112-123, January 2019.
      Purpose The successful use of the standard extended Kalman filter (EKF) is restricted by the requirement on the statistics information of the measurement noise. The covariance of the measurement noise may deviate from its nominal value in practical environment, and the filtering performance may decline because of the statistical uncertainty. Although the adaptive EKF (AEKF) is available for recursive covariance estimation, it is often less accurate than the EKF with accurate noise statistics. Design/methodology/approach Aiming at this problem, this paper develops a parallel adaptive EKF (PAEKF) by combining the EKF and the AEKF with an adaptive law, such that the final state estimate is dominated by the EKF when the prior noise covariance is accurate, while the AEKF is activated when the actual noise covariance deviates from its nominal value. Findings The PAEKF can reduce the sensitivity of the algorithm to the model uncertainty and ensure the estimation accuracy in the normal case. The simulation results demonstrate that the PAEKF has the advantage of both the AEKF and the EKF. Practical implications The presented algorithm is applicable for spacecraft relative attitude and position estimation. Originality/value The PAEKF is presented for a kind of nonlinear uncertain systems. Stability analysis is provided to show that the error of the estimator is bounded under certain assumptions.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-06T11:30:40Z
      DOI: 10.1108/AEAT-01-2018-0066
       
  • A new jig-shape optimization method for the high aspect ratio wing
    • Pages: 124 - 133
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 124-133, January 2019.
      Purpose Computational efficiency is always the major concern in aircraft design. The purpose of this research is to investigate an efficient jig-shape optimization design method. A new jig-shape optimization method is presented in the current study and its application on the high aspect ratio wing is discussed. Design/methodology/approach First, the effects of bending and torsion on aerodynamic distribution were discussed. The effect of bending deformation was equivalent to the change of attack angle through a new equivalent method. The equivalent attack angle showed a linear dependence on the quadratic function of bending. Then, a new jig-shape optimization method taking integrated structural deformation into account was proposed. The method was realized by four substeps: object decomposition, optimization design, inversion and evaluation. Findings After the new jig-shape optimization design, both aerodynamic distribution and structural configuration have satisfactory results. Meanwhile, the method takes both bending and torsion deformation into account. Practical implications The new jig-shape optimization method can be well used for the high aspect ratio wing. Originality/value The new method is an innovation based on the traditional single parameter design method. It is suitable for engineering application.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T12:50:17Z
      DOI: 10.1108/AEAT-01-2018-0073
       
  • Aerodynamic analysis of nonuniform trailing edge blowing
    • Pages: 134 - 144
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 134-144, January 2019.
      Purpose Numerical and experimental results for different oncoming base-flow conditions indicate that nonuniform trailing edge blowing (NTEB) can expand the performance range of compressors and reduce the thrust on the rotor, while the efficiency of the compressor can be improved by more than 2 per cent. Design/methodology/approach Relevant aerodynamic parameters, such as total pressure, ratio of efficiency and axial thrust, are calculated and analyzed under conditions with and without NTEB. Measurements are performed downstream of two adjacent stator blades, at seven equidistantly spaced reference locations. The experimental measurement of the interstage flow field used a dynamic four-hole probe with phase lock technique. Findings An axial low-speed single-stage compressor was established with flow field measurement system and nonuniform blowing system. NTEB was studied by means of numerical simulations and experiments, and it is found that the efficiency of the tested compressor can be improved by more than 2 per cent. Originality/value Unlike most of the previous research studies which mainly focused on the rotor/stator interaction and trailing edge uniform blowing, the research results summarized in the current paper on the stator/rotor interaction used inlet guide vanes for steady and unsteady calculations. An active control of the interstage flow field in a low-speed compressor was used to widen the working range and improve the performance of the compressor.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T07:10:23Z
      DOI: 10.1108/AEAT-04-2018-0115
       
  • Design of a double parabolic supersonic nozzle and performance evaluation
           by experimental and numerical methods
    • Pages: 145 - 156
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 145-156, January 2019.
      Purpose The purpose of this paper is to design a double parabolic nozzle and to compare the performance with conventional nozzle designs. Design/methodology/approach The throat diameter and divergent length for Conical, Bell and Double Parabolic nozzles were kept same for the sake of comparison. The double parabolic nozzle has been designed in such a way that the maximum slope of the divergent curve is taken as one-third of the Prandtl Meyer (PM) angle. The studies were carried out at Nozzle Pressure Ratio (NPR) of 5 and also at design conditions (NPR = 3.7). Experimental measurements were carried out for all the three nozzle configurations and the performance parameters compared. Numerical simulations were also carried out in a two-dimensional computational domain incorporating density-based solver with RANS equations and SST k-ω turbulence model. Findings The numerical predictions were found to be in reasonable agreement with the measured experimental values. An enhancement in thrust was observed for double parabolic nozzle when compared with that of conical and bell nozzles. Research limitations/implications Even though the present numerical simulations were capable of predicting shock cell parameters reasonably well, shock oscillations were not captured. Practical implications The double parabolic nozzle design has enormous practical importance as a small increase in thrust can result in a significant gain in pay load. Social implications The thrust developed by the double parabolic nozzle is seen to be on the higher side than that of conventional nozzles with better fuel economy. Originality/value The overall performance of the double parabolic nozzle is better than conical and bell nozzles for the same throat diameter and length.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T01:02:05Z
      DOI: 10.1108/AEAT-12-2017-0275
       
  • Oscillation mode flight data analysis based on FFT
    • Pages: 157 - 162
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 157-162, January 2019.
      Purpose The purpose of this paper is to propose an identification method of acquiring aircraft mode characteristics based on fast Fourier transform and half-power bandwidth method, aiming at the common oscillation met in flight test. Design/methodology/approach The feasibility of this method is demonstrated through derivation; the robustness analysis is conducted through three examples, and finally the method was applied on a set of sideslip angle record from flight test. Findings The derivation and numerical analysis both show that the presented method can have high accuracy and good robustness under coupled mode and noise condition. Practical implications The method proposed is of robustness, and it is concise and easy to apply on flight data record. Originality/value This paper demonstrates the feasibility of half power bandwidth to be applied on oscillation mode characteristics identification from flight data record, which is different from other method applied.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T01:05:22Z
      DOI: 10.1108/AEAT-04-2018-0139
       
  • Combustion characteristics of a two-stroke spark ignition UAV engine
           fuelled with gasoline and kerosene (RP-3)
    • Pages: 163 - 170
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 163-170, January 2019.
      Purpose The purpose of this paper is to compare the combustion characteristics, including the combustion pressure, heat release rate (HRR), coefficient of variation (COV) of indicated mean effective pressure (IMEP), flame development period and combustion duration, of aviation kerosene fuel, namely, rocket propellant 3 (RP-3), and gasoline on a two-stoke spark ignition engine. Design/methodology/approach This paper is an experimental investigation using a bench test to reflect the combustion performance of two-stroke spark ignition unmanned aerial vehicle (UAV) engine on gasoline and RP-3 fuel. Findings Under low load conditions, the combustion performance and HRR of burning RP-3 fuel were shown to be worse than those of gasoline. Under high load conditions, the average IMEP and the COV of IMEP of burning RP-3 fuel were close to those of gasoline. The difference in the flame development period between gasoline and RP-3 fuel was similar. Practical implications Gasoline fuel has a low flash point, high-saturated vapour pressure and relatively high volatility and is a potential hazard near a naked flame at room temperature, which can create significant security risks for its storage, transport and use. Adopting a low volatility single RP-3 fuel of covering all vehicles and equipment to minimize the number of different devices with the use of a various fuels and improve the application safeties. Originality/value Most two-stroke spark ignition UAV engines continue to combust gasoline. A kerosene-based fuel operation can be applied to achieve a single-fuel policy.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-19T08:10:07Z
      DOI: 10.1108/AEAT-03-2018-0112
       
  • Quantum-entanglement pigeon-inspired optimization for unmanned aerial
           vehicle path planning
    • Pages: 171 - 181
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 171-181, January 2019.
      Purpose The purpose of this paper is to propose a new algorithm for independent navigation of unmanned aerial vehicle path planning with fast and stable performance, which is based on pigeon-inspired optimization (PIO) and quantum entanglement (QE) theory. Design/methodology/approach A biomimetic swarm intelligent optimization of PIO is inspired by the natural behavior of homing pigeons. In this paper, the model of QEPIO is devised according to the merging optimization of basic PIO algorithm and dynamics of QE in a two-qubit XXZ Heisenberg System. Findings Comparative experimental results with genetic algorithm, particle swarm optimization and traditional PIO algorithm are given to show the convergence velocity and robustness of our proposed QEPIO algorithm. Practical implications The QEPIO algorithm hold broad adoption prospects because of no reliance on INS, both on military affairs and market place. Originality/value This research is adopted to solve path planning problems with a new aspect of quantum effect applied in parameters designing for the model with the respective of unmanned aerial vehicle path planning.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-15T09:49:29Z
      DOI: 10.1108/AEAT-03-2018-0107
       
  • A new energy-based model to predict spray droplet diameter in comparison
           with momentum-based models
    • Pages: 182 - 189
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 182-189, January 2019.
      Purpose To estimate mean droplet diameter (MDD) of a spray, three different numerical models were used in this paper. One of them is investigation of the surface instability of the liquid sheet producing from an injector. Design/methodology/approach First, the linear instability (LI) analysis introduced by Ibrahim (2006) is implemented. Second, the improved (ILI) analysis already introduced by the present authors is used. ILI analysis is different from the prior analysis, so that the instability of hollow-cone liquid sheet with different cone angles is investigated rather than a cylindrical liquid sheet. It means that besides the tangential and axial movements, radial movements of the liquid sheet and gas streams have been considered in the governing equations. Beside LI theory as a momentum-based approach, a new model as a theoretical energy-based (TEB) model based on the energy conservation law is proposed in this paper. Findings Based on the energy-based approach, atomization occurs because of kinetic energy loss. The resulting formulation reveals that the MDD is inversely proportional to the atomization efficiency and liquid Weber number. Research limitations/implications The results of these three models are compared with the available experimental data. Prediction obtained by the proposed TEB model is in reasonable agreement with the result of experiment. Practical implications The results of these three models are compared with the available experimental data. Prediction of the proposed energy-based theoretical model is in very good agreement with experimental data. Originality/value Comparison between the results of new model, experimental data, other previous methods show that it can be used as a new simple and fast model to achieve good estimation of spray MDD.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T01:09:47Z
      DOI: 10.1108/AEAT-06-2017-0155
       
  • Safety management systems: an opportunity and a challenge for military
           aviation organisations
    • Pages: 190 - 196
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 190-196, January 2019.
      Purpose Most military aviation organisations today have not evolved their safety management approach towards harmonising with civil aviation. Safety culture is the base for any civil aviation organisation, enabling employees to communicate effectively and be fully aware and extrovert on safety. Just culture and reporting culture both are related to safety culture. Both are parts of the awareness process, enhancing safety promotion. These distinct elements and the safety management systems (SMS) can serve well the military aviation. This paper aims to present and discuss the SMS philosophy, structure and elements as a solution for military aviation organisations. Design/methodology/approach The feature of civil aviation SMSs are presented and discussed, with reference to the applicable frameworks and regulations governing the SMS operation. A discussion on the challenges faced within the military aviation organisations, with a brief examination of a European Union military aviation organisation, is presented. Findings The European Military Airworthiness Requirements, which are based on the European Aviation Safety Agency set of rules, can act the basis for establishing military aviation SMSs. A civil-based approach, blended, as necessary, with military culture is workable, as this is the case for many defence forces that have adopted such aviation safety systems. Originality/value This viewpoint paper discusses the opportunities and challenges associated with the adoption of SMS by military aviation organisations. This is the first time that this issue is openly discussed and presented to the wider aviation community, outside military aviation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-26T09:53:19Z
      DOI: 10.1108/AEAT-05-2018-0146
       
  • Aircraft runway excursion features: a multiple correspondence analysis
    • Pages: 197 - 203
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 1, Page 197-203, January 2019.
      Purpose The risk of aircraft runway excursion, dependent on multiple factors, is related to operating conditions. The purpose of this paper is to identify the correspondence between features belonging to different aspects that occur in runway excursion events, distinguishing between take-off and landing phases. Design/methodology/approach To define the correspondence between the characteristic features of runway excursions, this study has applied multiple correspondence analysis (MCA). MCA is used to represent and model data sets as “clouds” of points in a multi-dimensional Euclidean space. There are five variables used in MCA: geographical region, potential cause, aircraft class, flight nature and aircraft damages. For the purpose of this research, the database contains only runway excursion accidents that took place between 2006 and 2016 among all categories of aircraft in all world regions. The events contained in the database were analyzed by separating those that occurred during take-off and those that occurred during landing. Findings With this method, this study identified a few particularly interesting variable combinations. Generally, the consequence of an aircraft runway excursion is substantial aircraft damage. Also, the most common cause of runway excursion during take-off is aircraft system faults, while during landing, it is weather conditions. Furthermore, the destruction of an aircraft is a result of a runway excursion due to bad weather conditions, both during take-off and landing. Practical implications The results of this study can be used by a broad range of civil aviation organizations for runway risk assessment and to select the most effective safety countermeasures for runway excursions. Originality/value The authors believe this study is original, especially for the statistical analysis method used.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-12T03:19:58Z
      DOI: 10.1108/AEAT-11-2017-0244
       
  • A novel improvement of Kriging surrogate model
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to introduce a method based on the optimizer of the particle swarm optimization (PSO) algorithm to improve the efficiency of a Kriging surrogate model. Design/methodology/approach PSO was first used to identify the best group of trend functions and to optimize the correlation parameter thereafter. Findings The Kriging surrogate model was used to resolve the fuselage optimization of an unmanned helicopter. Practical implications The optimization results indicated that an appropriate PSO scheme can improve the efficiency of the Kriging surrogate model. Originality/value Both the STANDARD PSO and the original PSO algorithms were chosen to show the effect of PSO on a Kriging surrogate model.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-19T09:10:52Z
      DOI: 10.1108/AEAT-06-2018-0157
       
  • 4D-trajectory time windows: definition and uncertainty management
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The use of the 4D trajectory operational concept in the future air traffic management (ATM) system will require the aircraft to meet very accurately an arrival time over a designated checkpoint. To do this, time intervals known as time windows (TW) are defined. The purpose of this paper is to develop a methodology to characterise these TWs and to manage the uncertainty associated with the evolution of 4D trajectories. Design/methodology/approach 4D trajectories are modelled using a point mass model and EUROCONTROL’s BADA methodology. The authors stochastically evaluate the variability of the parameters that influence 4D trajectories using Monte Carlo simulation. This enables the authors to delimit TWs for several checkpoints. Finally, the authors set out a causal model, based on a Bayesian network approach, to evaluate the impact of variations in fundamental parameters at the chosen checkpoints. Findings The initial results show that the proposed TW model limits the deviation in time to less than 27 s at the checkpoints of an en-route segment (300 NM). Practical implications The objective of new trajectory-based operations is to efficiently and strategically manage the expected increase in air traffic volumes and to apply tactical interventions as a last resort only. We need new tools to support 4D trajectory management functions such as strategic and collaborative planning. The authors propose a novel approach for to ensure aircraft punctuality. Originality/value The main contribution of the paper is the development of a model to deal with uncertainty and to increase predictability in 4D trajectories, which are key elements of the future airspace operational environment.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-19T08:56:30Z
      DOI: 10.1108/AEAT-01-2018-0031
       
  • Interstellar medium magnetic field in bow shock modelling
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to demonstrate the impact of interstellar (IS) magnetic field on stellar shocks existence, shape and size in the stellar wind (SW) vs interstellar medium (ISM) numerical models. Design/methodology/approach Comparison of hydrodynamics (HD) and magnetohydrodynamic (MHD) models results with or without ISM magnetic field, its intensity and ISM parameters. Findings ISM magnetic field facilitates formation and stabilises bow shocks around all astrophysical objects. ISM magnetic field may also be one of the reasons for a bow shock existence around the Sun. Practical implications ISM magnetic field should be implemented in MHD and future kinetic numerical models of the SW interaction with ISM plasma. Originality/value This paper presents the results of HD and MHD models of bow shocks and the importance of ISM magnetic field implementation, according to astronomical bow shock observations. The study also presents a review of the most important papers showing the numerical results of bow shock formation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-18T01:38:45Z
      DOI: 10.1108/AEAT-01-2018-0004
       
  • Assessment of a small UAV speed polar graph by conducting flight tests
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present an approach to a polar graph measurement by a flight testing technique and to propose a baseline research method for future tests of UAV polar graphs. The method presented can be used to demonstrate a conceptual and preliminary design process using a scaled, unmanned configuration. This shows how results of experimental flight tests using a scaled flying airframe may be described and analysed before manufacturing the full scale aircraft. Design/methodology/approach During the research, the flight tests were conducted for two aerodynamic configurations of a small UAV. This allowed the investigation of the influence of winglets and classic vertical stabilizers on the platform stability, performance and therefore polar graphs of a small unmanned aircraft. Findings A methodology of flight tests for the assessment of a small UAV’s polar graph has been proposed, performed and assessed. Two aerodynamic configurations were tested, and it was found that directional stability had a large influence on the UAV’s performance. A correlation between the speed and inclination of the altitude graph was found – i.e. the higher the flight speed, the steeper the altitude graph (higher descent speed, steeper flight path angle). This could be considered as a basic verification that the recorded data have a physical sense. Practical implications The polar graph and therefore glide ratio of the aircraft is a major factor for determining its performance and power required for flight. Using the right flight test procedure can speed-up the process of measuring glide ratio, making it easier, faster, robust, more effective and accurate in future research of novel, especially unorthodox configurations. This paper also can be useful for the proper selection of requirements and preliminary design parameters for making the design process more economically effective. Originality/value This paper presents a very efficient method of assessing the design parameters of UAVs, especially the polar graph, in an early stage of the design process. Aircraft designers and producers have been widely performing flight testing for years. However, these procedures and practical customs are usually not wide spread and very often are treated as the company’s “know how”. Results presented in this paper are original, relatively easily be repeated and checked. They may be used either by professionals, highly motivated individuals and representatives of small companies or also by ambitious amateurs.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-18T01:38:38Z
      DOI: 10.1108/AEAT-03-2018-0099
       
  • Innovative time-based separation procedures for civil RPAS integration
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to suggest feasible solutions to overcome the problem of unmanned aerial vehicles integration within the existing airspace. Design/methodology/approach It envisages innovative time-based separation procedures that will enhance the integration in the future air traffic management (ATM) system of next generation of large remotely piloted aircraft system (RPAS). 4D navigation and dynamic mobile area concepts, both proposed in the framework of Single European Sky ATM Research program, are brought together to hypothesize innovative time-based separation procedures aiming at promoting integration of RPAS in the future ATM system. Findings Benefits of proposed procedures, mainly evaluated in terms of volume reduction of segregated airspace, are quantitatively analyzed on the basis of realistic operational scenarios focusing on monitoring activities in both nominal and emergency conditions. Eventually, the major limits of time-based separation for RPAS are investigated. Practical implications The implementation of the envisaged procedures will be a key enabler in RPAS integration in future ATM integration. Originality/value In the current ATM scenario, separation of RPAS from air traffic is ensured by segregating a large amount of airspace areas with fixed dimensions, dramatically limiting the activities of these vehicles.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-12-18T01:38:36Z
      DOI: 10.1108/AEAT-08-2018-0235
       
  • Effect of scrubbing efficiency on fuel scrubbing inerting for aircraft
           fuel tanks
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to improve the previous fuel scrubbing model and find out the relationship between bubble diameter and scrubbing efficiency (ƞ). Design/methodology/approach A fuel tank scrubbing test bench was established to verify the accuracy of this model. Ullage and dissolved oxygen concentration were measured, and images of bubble size and distribution were collected and analyzed using image analysis software. Findings The bubble diameter has a great influence on ullage and dissolved oxygen concentration during the fuel scrubbing process. The scrubbing efficiency (ƞ) has an exponential relationship with bubble diameter and decreases rapidly as the bubble diameter increases. Practical implications The variation of the ullage and dissolved oxygen concentration predicted by this model is more accurate than that of the previous model. In addition, the study of bubble size can provide a guidance for the design of fuel scrubber. Originality/value This study not only improves the previous fuel scrubbing model but also develops a method to calculate scrubbing efficiency (ƞ) based on bubble diameter. In addition, a series of tests and analyses were conducted, including numerical calculation, experiment and image analysis.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-19T12:25:27Z
      DOI: 10.1108/AEAT-10-2017-0215
       
  • Sensitivity of automatic control of emergency manoeuvre to parametric
           uncertainty of the airplane model
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The main purpose of this work was elaboration and verification of a method of assessing the sensitivity of automatic control laws to parametric uncertainty of an airplane’s mathematical model. The linear quadratic regulator (LQR) methodology was used as an example design procedure for the automatic control of an emergency manoeuvre. Such a manoeuvre is assumed to be pre-designed for the selected airplane. Design/methodology/approach The presented method of investigating the control systems’ sensitivity comprises two main phases. The first one consists in computation of the largest variations of gain factors, defined as differences between their nominal values (defined for the assumed model) and the values obtained for the assumed range of parametric uncertainty. The second phase focuses on investigating the impact of the variations of these factors on the behaviour of automatic control in the manoeuvre considered. Findings The results obtained allow for a robustness assessment of automatic control based on an LQR design. Similar procedures can be used to assess in automatic control arrived at through varying design methods (including methods other than LQR) used to control various manoeuvres in a wide range of flight conditions. Practical implications It is expected that the presented methodology will contribute to improvement of automatic flight control quality. Moreover, such methods should reduce the costs of the mathematical nonlinear model of an airplane through determining the necessary accuracy of the model identification process, needed for assuring the assumed control quality. Originality/value The presented method allows for the investigation of the impact of the parametric uncertainty of the airplane’s model on the variations of the gain-factors of an automatic flight control system. This also allows for the observation of the effects of such variations on the course of the selected manoeuvre or phase of flight. This might be a useful tool for the design of crucial elements of an automatic flight control system.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-19T12:25:10Z
      DOI: 10.1108/AEAT-01-2018-0017
       
  • Leakage flow analysis in the gas turbine shroud gap
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of the study is to measure the mass flow in the flow through the labyrinth seal of the gas turbine and compare it to the results of numerical simulation. Moreover the capability of two turbulence models to reflect the phenomenon will be assessed. The studied case will later be used as a reference case for the new, original design of flow control method to limit the leakage flow through the labyrinth seal. Design/methodology/approach Experimental measurements were conducted, measuring the mass flow and the pressure in the model of the labyrinth seal. It was compared to the results of numerical simulation performed in ANSYS/Fluent commercial code for the same geometry. Findings The precise machining of parts was identified as crucial for obtaining correct results in the experiment. The model characteristics were documented, allowing for its future use as the reference case for testing the new labyrinth seal geometry. Experimentally validated numerical model of the flow in the labyrinth seal was developed. Research limitations/implications The research studies the basic case, future research on the case with a new labyrinth seal geometry is planned. Research is conducted on simplified case without rotation and the impact of the turbine main channel. Practical implications Importance of machining accuracy up to 0.01 mm was found to be important for measuring leakage in small gaps and decision making on the optimal configuration selection. Originality/value The research is an important step in the development of original modification of the labyrinth seal, resulting in leakage reduction, by serving as a reference case.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-16T09:45:41Z
      DOI: 10.1108/AEAT-01-2018-0038
       
  • Preliminary stability analysis methods for PrandtlPlane aircraft in
           subsonic conditions
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The present paper aims to assess the reliability and the limitations of analysing flight stability of a box-wing aircraft configuration known as PrandtlPlane by means of methods conceived for conventional aircraft and well known in the literature. Design/methodology/approach Results obtained by applying vortex lattice methods to PrandtlPlane configuration, validated previously with wind tunnel tests, are compared to the output of a “Roskam-like” method, here defined to model the PrandtlPlane features. Findings The comparisons have shown that the “Roskam-like” model gives accurate predictions for both the longitudinal stability margin and dihedral effect, whereas the directional stability is always overestimated. Research limitations/implications The method here proposed and related achievements are valid only for subsonic conditions. The poor reliability related to lateral-directional derivatives estimations may be improved implementing different models known from the literature. Practical implications The possibility of applying a faster method as the “Roskam-like” one here presented has two main implications: it allows to implement faster analyses in the conceptual and preliminary design of PrandtlPlane, providing also a tool for the definition of the design space in case of optimization approaches and it allows to implement a scaling procedure, to study families of PrandtlPlanes or different aircraft categories. Social implications This paper is part of the activities carried out during the PARSIFAL project, which aims to demonstrate that the introduction of PrandtlPlane as air transport mean can fuel consumption and noise impact, providing a sustainable answer to the growing air passenger demand envisaged for the next decades. Originality/value The originality of this paper lies in the attempt of adopting analysis method conceived for conventional airplanes for the analysis of a novel configuration. The value of the work is represented by the knowledge concerning experimental results and design methods on the PrandtlPlane configuration, here made available to define a new analysis tool.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-16T09:22:00Z
      DOI: 10.1108/AEAT-12-2017-0284
       
  • Propulsion model for (hybrid) unmanned aircraft systems (UAS)
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper deals with the estimation of the necessary masses of propulsion components for multirotor UAS. Originally, within the design process of multirotors, this is an iterative problem, as the propulsion masses contribute to the total takeoff mass. Hence, they influence themselves and cannot be directly calculated. The paper aims to estimate the needed propulsion masses with respect to the requested thrust because of payload, airframe weight and drag forces and with respect to the requested flight time. Design/methodology/approach Analogue to the well-established design synthesis of airplanes, statistical data of existing electrical motors, propellers and rechargeable batteries are evaluated and analyzed. Applying Rankine and Froude’s momentum theory and a generic model for electro motor efficiency factors on the statistical performance data provides correlations between requested performance and, therefore, needed propulsion masses. These correlations are evaluated and analyzed in the scope of buoyant-vertical-thrusted hybrid UAS. Findings This paper provides a generic mathematical propulsion model. For given payloads, airframe structure weights and a requested flight time, appropriate motor, propeller and battery masses can be modelled that will provide appropriate thrust to lift payload, airframe and the propulsion unit itself over a requested flight time. Research limitations/implications The model takes into account a number of motors of four and is valid for the category of nano and small UAS. Practical implications The presented propulsion model enables a full numerical design process for vertical thrusted UAS. Hence, it is the precondition for design optimization and more efficient UAS. Originality/value The propulsion model is unique and it is valid for pure multirotor as well as for hybrid UAS too.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-15T09:56:27Z
      DOI: 10.1108/AEAT-01-2018-0033
       
  • Conflict-resolution algorithms for RPAS in non-segregated airspace
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to focus on the development of conflict-resolution algorithms between Remotely Piloted Aircraft System (RPAS) and conventional aircraft. The goal of the conflict-resolution algorithm is to estimate the minimum protection distance (MPD) which is required to avoid a potential conflict. Design/methodology/approach The conflict-resolution algorithms calculate the last location at which an RPAS must start climbing to avoid a separation minima infringement. The RPAS maneuvers to prevent the conventional aircraft based on the kinematic equations. The approach selects two parameters to model the conflict-geometry: the path-intersection angle and the Rate of Climb (ROCD). Findings Results confirmed that the aircraft pair flying in opposition was the worst scenario because the MPD reached its maximum value. The best value of the MPD is about 12 Nautical Miles to ensure a safe resolution of a potential conflict. Besides, variations of the ROCD concluded that the relation between the ROCD and the MPD is not proportional. Research limitations/implications The primary limitation is that the conflict-resolution algorithms are designed in a theoretical framework without bearing in mind other factors such as communications, navigation capacity, wind and pilot errors among others. Further work should introduce these concepts to determine how the MPD varies and affects air traffic safety. Moreover, the relation between an ROCD requirement and the MPD will have an impact on regulations. Practical implications The non-linear relation between the MPD and the ROCD could be the pillar to define a standardized MPD in the future for RPAS systematic integration. To accomplish this standard, RPAS could have to fulfil a requirement of minimum ROCD until a specified flight level. Originality/value This paper is the first approach to quantify the Minimum Protection Distance between RPAS and conventional aircraft, and it can serve the aeronautical community to define new navigation requirements for RPAS.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-15T09:52:47Z
      DOI: 10.1108/AEAT-01-2018-0024
       
  • New hail impact simulation models on composite laminated wing leading edge
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The risk of hail-impact occurrence that can decrease local strength property must be taken into account in the design of primary airframe structures in aviation, energy and space industries. Because of the high-speed of hail impact in operation, it can affect the load carrying capacity. Testing all impact scenarios onto real structure is expensive and impractical. The purpose of this paper is to present a cost-effective hybrid testing regime including experimental tests and FEM-based simulations for airframe parts that are locally exposed to the impacting hail in flight. Design/methodology/approach Tested samples (specimens) are flat panels of laminated and sandwich carbon/epoxy composites that are used in designing lightweight new airframes. The presented numerical simulations provide a cost effective and convenient tool for investigating the hail impact scenarios in the design process. The smoothed particle hydrodynamics (SPH) technique was selected for the simulation of projectiles. The most commonly used shape of projectiles in hail impact tests is the ice ball with a defined diameter. The proposed simulation technique was verified and validated in tests on flat composite panels (specimens). Findings Integration of the numerical analyses with high-speed impact tests of hail onto flat laminated and sandwich composite shells has been presented, and a developed simulation model for impact results assessment was obtained. Originality/value The tested coupons (specimens) are flat panels as representative of structural design deployed in real aircraft structures. These numerical simulations provide a cost effective and convenient tool for hail impact scenarios in the design process.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-06T11:34:02Z
      DOI: 10.1108/AEAT-02-2018-0089
       
  • Wall distance effect on heat transfer at high flow velocity
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present the results of experimental and numerical research on heat transfer distribution under the impinging jets at various distances from the wall and high jet velocity. This work is a part of the INNOLOT Program financed by National Centre for Research and Development. Design/methodology/approach The air jets flow out from the common pipe and impinge on a surface which is cooled by them, and in this way, all together create a model of external cooling system of low-pressure gas turbine casing. Measurements were carried out for the arrangement of 26 in-line jets with orifice diameter of 0.9 mm. Heat transfer distribution was investigated for various Reynolds and Mach numbers. The cooled wall, made of transparent PMMA, was covered with a heater foil on which a layer of self-adhesive liquid crystal foil was placed. The jet-to-wall distance was set to h = from 4.5 to 6 d. Findings The influence of various Reynolds and Mach numbers on cooled flat plate and jet-to-wall distance in terms of heat transfer effectiveness is presented. Experimental results used for the computational fluid dynamics (CFD) model development, validation and comparison with numerical results are presented. Practical implications Impinging air jets is a commonly used technique to cool advanced turbines elements, as it produces large convection enhancing the local heat transfer, which is a critical issue in the development of aircraft engines. Originality/value The achieved results present experimental investigations carried out to study the heat transfer distribution between the orthogonally impinging jets from long round pipe and flat plate. Reynolds number based on the jet orifice exit conditions was varied between 2,500 and 4,000; meanwhile, for such Re, the flow velocity in jets was particularly very high, changing from M = 0.56 to M = 0.77. Such flow conditions combination, i.e. the low Reynolds number and very high flow velocity cannot be found in the existing literature.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-02T12:48:28Z
      DOI: 10.1108/AEAT-01-2018-0022
       
  • Fast identification of transonic buffet envelope using computational fluid
           dynamics
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a numerical method based on computational fluid dynamics that allows investigating the buffet envelope of reference equivalent wings at the equivalent cost of several two-dimensional, unsteady, turbulent flow analyses. The method bridges the gap between semi-empirical relations, generally dominant in the early phases of aircraft design, and three-dimensional turbulent flow analyses, characterised by high costs in analysis setups and prohibitive computing times. Design/methodology/approach Accuracy in the predictions and efficiency in the solution are two key aspects. Accuracy is maintained by solving a specialised form of the Reynolds-averaged Navier–Stokes equations valid for infinite-swept wing flows. Efficiency of the solution is reached by a novel implementation of the flow solver, as well as by combining solutions of different fidelity spatially. Findings Discovering the buffet envelope of a set of reference equivalent wings is accompanied with an estimate of the uncertainties in the numerical predictions. Just over 2,000 processor hours are needed if it is admissible to deal with an uncertainty of ±1.0° in the angle of attack at which buffet onset/offset occurs. Halving the uncertainty requires significantly more computing resources, close to a factor 200 compared with the larger uncertainty case. Practical implications To permit the use of the proposed method as a practical design tool in the conceptual/preliminary aircraft design phases, the method offers the designer with the ability to gauge the sensitivity of buffet on primary design variables, such as wing sweep angle and chord to thickness ratio. Originality/value The infinite-swept wing, unsteady Reynolds-averaged Navier–Stokes equations have been successfully applied, for the first time, to identify buffeting conditions. This demonstrates the adequateness of the proposed method in the conceptual/preliminary aircraft design phases.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-11-02T12:21:43Z
      DOI: 10.1108/AEAT-01-2018-0057
       
  • Composite technology development based on helicopter rotor blades
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to document the approach, effort and cost of advance composite technology implementation suited for small and medium enterprises on the example of composite main rotor blade development for ILX-27 helicopter. Design/methodology/approach This work was carried out as part of a development project for main rotor blades used on the ILX 27 helicopter. The paper presents all stages of the design of the blade structure in parallel with composite technology development. The data were gathered and documented during project execution. The stages of R&D work in terms of labor intensity and important processes influencing quality and efficiency were assessed. Findings The paper provides key aspects for successful composite capability introduction. The incurred cost of equipment and staff training is evaluated. The paper also summarized the cost of composite parts manufactured with developed technology. Practical implications The paper provides detail example of composite capability development including basic technologies, processes, equipment and cost of the project. Presented details can be great guidelines for small and medium enterprises with the goal of composite technology introduction for aerostructures design and manufacturing. Originality/value This paper present clear, complete and verified process of composite capability development for aerostructures design and build suited for small and medium enterprises. It presents detail cost, calculated in Polish economy environment, of each phase and final cost of the product.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T12:58:23Z
      DOI: 10.1108/AEAT-12-2017-0260
       
  • Less-skilled pilot decision support
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to overview the systems and their elements developing for supporting the less-skilled pi-lots. Design/methodology/approach Several European (like EPATS, SAT-Rdmp, Pplane, Esposa, Clean Sky2) and national projects (NASA SATS, Hungarian SafeFly) develop the personal/small aircraft and personal/small aircraft transportation systems. The projects had analysed the safety aspects, too, and they underlined the aircraft will be controlled by so-called less-skilled pilots (owners, renters), having less experiences. The paper defines the cross-connected controls, introduces the methods of subjective analysis in pilot decision processes, improves the pilot workload model, defines the possible workload management and describes the developing pilot decision support system. Findings Analysing the personal/small aircraft safety aspects, a unique and important safety problem induced by less-skilled pilots has been identified. The considerable simplification of the air-craft control system, supporting the pilot subjective decisions and introducing the pilot work-load management, may eliminate this problem. Research limitations/implications Only the system elements have been used in concept validation tests. Practical implications The developing pilot supporting system in its general form has on - board and ground sub-systems, too, except a series of elements integrated into the pilot cockpit environment and control system. Several system elements (sensors, integrated controls, etc.) might be implement now, but the total system need further studies. The subjective decision process needs further development of the methodology and concept validation. Social implications The system may catalyse the society acceptance of the personal aircraft and their safer piloting, applicability. Originality/value The paper introduces an original supporting system for less-skilled pilots.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T12:56:40Z
      DOI: 10.1108/AEAT-12-2017-0269
       
  • Modeling of surface spectra with and without dust from Martian infrared
           data: new aspects
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to characterize the mineral composition of Martian surfaces based on Thermal Emission Spectrometer (TES; Mars Global Surveyor) as measured in the infrared thermal range. It presents modeling and interpreting of TES spectral data from selected Martian regions from which the atmospheric influences had been removed using radiative transfer algorithm and deconvolution algorithm. The spectra from the dark area of Cimmeria Terra and the bright Isidis Planitia were developed in Philip Christensen’s and Joshua Bandfield’s publications, where these spectra were subjected to spectral deconvolution to estimate the mineral composition of the Martian surface. The results of the analyses of these spectra were used for the modeling of dusty and non-dusty surface of Mars. As an additional source, the mineral compositions of Polish basalts and mafic rocks were used for these surfaces as well as for modeling Martian meteorites Shergottites, Nakhlites and Chassignites. Finally, the spectra for the modeling of the Hellas region were obtained from the Planetary Fourier Spectrometer (PFS) – (Mars Express) and the mineralogical compositions of basalts from the southern part of Poland were used for this purpose. The Hellas region was modeled also using simulated Martian soil samples Phyllosilicatic Mars Regolith Simulant and Sulfatic Mars Regolith Simulant, showing as a result that the composition of this selected area has a high content of sulfates. Linear spectral combination was chosen as the best modeling method. The modeling was performed using PFSLook software written in the Space Research Centre of the Polish Academy of Sciences. Additional measurements were made with an infrared spectrometer in thermal infrared spectroscopy, for comparison with the measurements of PFS and TES. The research uses a kind of modeling that successfully matches mineralogical composition to the measured spectrum from the surface of Mars, which is the main goal of the publication. This method is used for areas where sample collection is not yet possible. The areas have been chosen based on public availability of the data. Design/methodology/approach The infrared spectra of the Martian surface were modeled by applying the linear combination of the spectra of selected minerals, which then are normalized against the measured surface area with previously separated atmosphere. The minerals for modeling are selected based on the expected composition of the Martian rocks, such as basalt. The software used for this purpose was PFSLook, a program written in C++ at the Space Research Centre of the Polish Academy of Sciences, which is based on adding the spectra of minerals in the relevant percentage, resulting in a final spectrum containing 100 per cent of the minerals. Findings The results of this work confirmed that there is a relationship between the modeled, altered and unaltered, basaltic surface and the measured spectrum from Martian instruments. Spectral deconvolution makes it possible to interpret the measured spectra from areas that are potentially difficult to explore or to choose interesting areas to explore on site. The method is described for mid-infrared because of software availability, but it can be successfully applied to shortwave spectra in near-infrared (NIR) band for data from the currently functioning Martian spectroscopes. Originality/value This work is the only one attempting modeling the spectra of the surface of Mars with a separated atmosphere and to determine the mineralogical composition.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T01:09:22Z
      DOI: 10.1108/AEAT-01-2018-0051
       
  • The effect of wing-tip propulsors on Icaré 2 aeroelasticity
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The separation of energy conversion and propulsor is a promising aspect of hybrid-electric propulsion systems, allowing for increased installation efficiencies and setting the basis for distributed propulsion concepts. University of Stuttgart’s Institute of Aircraft Design has a long experience with electrically powered aircraft, starting with Icaré 2, a solar-powered glider flying, since 1996. Icaré 2 recently has been converted to a three-engine motor glider with two battery-powered wing-tip propellers, in addition to the solar-powered main electric motor. This adds propulsion redundancy and will allow analyzing yaw control concepts with differential thrust and the propeller-vortex interaction at the wing-tip. To ensure airworthiness for this design modification, new ground vibration tests (GVTs) and flutter calculations are required. The purpose of this paper is to lay out the atypical approach to test execution due to peculiarities of the Icaré 2 design such as an asymmetrical aileron control system, the long wing span with low frequencies of the first mode and elevated wing tips bending under gravity and thus affecting the accuracy of the wing torsion frequency measurements. Design/methodology/approach A flutter analysis based on GVT results is performed for the aircraft in basic configuration and with wing tip propulsors in pusher or tractor configuration. Apart from the measured resonant modes, the aircraft rigid body modes and the control surface mechanism modes are taken into consideration. The flutter calculations are made by a high-speed, low-cost software named JG2 based on the strip theory in aerodynamics and the V-g method of flutter problem solution. Findings With the chosen atypical approach to GVT the impact of the suspension on the test results was shown to be minimal. Flutter analysis has proven that the critical flutter speed of Icaré 2 is sufficiently high in all configurations. Practical implications The atypical approach to GVT and subsequent flutter analysis have shown that the effects of wing-tip propulsors on aeroelasticity of the high aspect ratio configuration do not negatively affect flutter characteristics. This analysis can serve as a basis for an application for a permit to fly. Originality/value The presented methodology is valuable for the flutter assessment of aircraft configurations with atypical aeroelastic characteristics.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-31T01:03:45Z
      DOI: 10.1108/AEAT-12-2017-0279
       
  • Developing the pilots’ load measuring system
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The main purpose of this study is to introduce the pilots’ load model and developed concept of load measuring system for operator load management. Design/methodology/approach In future aeronautical system, the role of operators (pilots and air traffic controllers [ATCOs]) will be in transition from active controlling to passive monitoring. Therefore, the operators’ load (task, information, work and mental) model was developed. There were developed measuring systems integrating into the pilot and ATCOs working environment eye tracking system outside measuring equipment. Operator load management was created by using the measurement. Findings In future system depending on time and automation level, the role of information and mental load will be increased. In flight simulator practice, developed load management method serves as a good tool for improving the quality of pilot training. According to the test results, the load monitoring and management system increase the safety of operators’ action in an emergency situation. Research limitations/implications The developed method were tested in two flight simulators (one developed for scientific investigation and other one applied for pilot training) and ATM management laboratory. Practical implications By deployment of the develop load monitoring and management system, the safety of aircraft flights and air transport management will be increased, especially in an emergency situation. Social implications People and society’s acceptance of future highly automated system will be increased. Originality value The analysis focuses on the following: developing operator’s load model as improved situation awareness model of Endsley, developing monitoring system integrated into operator’s working environment, creating load management system.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-26T09:24:16Z
      DOI: 10.1108/AEAT-01-2018-0080
       
  • Control and monitoring assistant for pilot
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Cooperation of a pilot with an automated aircraft control and monitoring systems is a problem which should be solved designing the whole system. The method of design, which creates an assistant of a pilot, is the purpose of this study. Design/methodology/approach An analysis of human factors shows demands for working environment. An integration method for various technological systems and algorithms is searched. Findings It is possible to make the whole system to become a pilot assistant, which has ability to exchange information with pilot by a dialogue. Structural flexibility is obtained in multi-agent system structure. Practical implications Proposed approach is a solution of how to integrate increasing amount of aircraft systems. It is expected that new form of cooperation fits to human features. Proposed methodology solves problem of simultaneous control by two controllers and cooperative making decisions. Social implications Dialogue between human and the system proposed in this solution will change perception of machines. Originality/value New abilities of machines and proposition of their realisation are presented. Presented solution of simultaneous control and decision-making during aircraft control is a novel approach to human–machine cooperation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-22T07:40:04Z
      DOI: 10.1108/AEAT-01-2018-0012
       
  • Research and selection of MALE wing profile
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this research is a preliminary selection of wing section, which would be the best suited for PW-100 – a MALE class UAV of 600 kg weight. PW-100 will be used as a testing platform in different institutions such as research institutes, industry research centers or universities of technology (phase 1) to enable the in-flight testing of various on-board systems (mobile radars, thermovision sensors, chemical sensors, antennas, teledetection systems and others). Untypical layout of PW-100 resulted from the plans of further development of this configuration for a military application. Design/methodology/approach Important role in the research described in this paper plays the selection of main wing section to fulfil the preliminary requirements regarding maximum lift coefficient, minimum drag, aerodynamic efficiency etc. Two different wing sections (R1082 and SA19) were tested in wind tunnel, both with flaps deflected at the range of 0°-30°. Experimental measurements were performed in the low turbulence wind tunnel with closed test section of 45 cm × 35 cm. Numerical simulations of the flow around the wing sections were performed using MSES code. Boundary conditions were assumed basing on the typical mission of PW-100 for flight altitude around 9,000 m, speed of 110 km/h what results in Re = 956,000. Findings Lift coefficients obtained from both experimental and numerical methods for single slatted airfoil SA19 are much higher than that of get for Ronch R1082 airfoil. PW-100 aircraft with SA19 airfoils will be able to be trimmed and fly at any altitude up to 9,000 m and with an arbitrary weight up to 600 kg. Aerodynamic characteristics of SA19 remain smoother and more predictable than that of R1082 airfoil. The very promising properties of SA19 airfoil are well known to the authors since the beginning of last decade when PW team worked together with IAI team on CAPECON project and now it was fully confirmed by this research. Practical implications It was confirmed that selection of the proper wing section for the special mission performed by UAV is of the highest importance decision to be taken at the preliminary design phase. Because there is a limited access to the base of technical parameters in many different UAVs classes and the classical analysis of trends cannot be fully applied, the wing section analysis, either experimental or numerical, must be performed. The situation is much worse than in the case of manned aircrafts because most of the modern UAVs are made of carbon or glass fiber, and therefore, there is no chance for analysis of trends. Originality/value This paper presents a very efficient method of assessing the influence of wing section on aircraft performance adopted for MALE class UAV, especially in an early stage of preliminary design process. The assessment is built mainly on three requirements: Maximum 2D lift coefficient for take-off configuration with flap deflected on 20 degrees should be greater than 2.4. Endurance factor CL1.5/CD for loitering conditions (Ma = 0.5 and CD0 = 0.008) should be greater than 110. The relative wing section thickness should be greater/equal than 19 per cent (it is required for high volume fuel tank located in the wings).
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-19T12:56:14Z
      DOI: 10.1108/AEAT-02-2018-0092
       
  • Satellite angular motion classification for active on-orbit debris removal
           using robots
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a novel method for identification and classification of rotational motion for uncontrolled satellites. These processes are shown in context of close proximity orbital operations. In particular, it includes a manipulator arm mounted on chaser satellite and used to capture target satellites. In such situations, a precise extrapolation of the target’s docking port position is needed to determine the manipulator arm motion. The outcome of this analysis might be used in future debris removal or servicing space missions. Design/methodology/approach Nonlinear, and in some special cases, chaotic nature of satellite rotational motion was considered. Four parameters were defined: range of motion toward docking port, dominant frequencies, fractal dimension of the motion and its time dependencies. Findings The qualitative analysis was performed for presented cases of spacecraft rotational motion and for each case the respective parameters were calculated. The analysis shows that it is possible to detect the type of rotational motion. Originality/value A novel procedure allowing to estimate the type of satellite rotational motion based on fractal approach was proposed.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-19T08:56:50Z
      DOI: 10.1108/AEAT-01-2018-0049
       
  • Hybridization of training aircraft with real world flight profiles
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to analyze real-world flight data of a piston engine training aircraft collected from an internet-based radar service, along with wind data provided by a weather forecast model, and to use such data to design a hybrid electric power system. Design/methodology/approach The modeling strategy starts from the power demand imposed by a real-world wind-corrected flight profile, where speed and altitude are provided as functions of time, and goes through the calculation of the efficiency of the powertrain components when they meet such demand. Each component of the power system and, in particular, the engine and the propeller, is simulated as a black box with an efficiency depending on the actual working conditions. In the case of hybrid electric power system, the battery charging and discharging processes are simulated with the Shepherd model. Findings The variability of power demand and fuel consumption for a training aircraft is analyzed by applying the proposed methodology to the Piper PA-28-180 Cherokee, a very popular aircraft used for flight training, air taxi and personal use. The potentiality of hybridization is assessed by analyzing the usage of the engine over more than 90 flights. A tentative sizing of a hybrid electric power system is also proposed. It guarantees a fuel saving of about 5%. Originality/value The scientific contribution and the novelty of the investigation are related to the modeling methodology, which takes into account real-world flight conditions, and the application of hybridization to a training aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-19T08:02:54Z
      DOI: 10.1108/AEAT-01-2018-0036
       
  • Aerodynamics analysis of rotor’s impact on the aircraft in the
           tandem wing configuration
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present the results of aerodynamic calculation of the aircraft in tandem wing configuration called VTOL. A presented vehicle combines the capabilities of the classic aircraft and helicopters. The aircraft is equipped with two pairs of tilt-rotors mounted on the tips of the front and the rear wing. The main goal of the presented research was to find the aerodynamic impact of both pairs of tilt-rotors on aerodynamic coefficients of the aircraft. Moreover, the rotors impact on the static stability of the aircraft was investigated too. Design/methodology/approach The CFD analysis was made for the complete aircraft in the tandem wing configuration. The computation was performed for the model of aircraft which was equipped with the four sub-models of the front and rear rotors. They were modeled as the actuator discs. This method allows for computing the aerodynamic impact of rotating components on the aircraft body. All aerodynamic analysis was made by the MGAERO software. The numerical code of the software was based on the Euler flow model. The used numerical method allows for the quick computation of very complex model of aircraft with a satisfied accuracy. Findings The result obtained by computation includes the aerodynamic coefficients which described the impact of the tilt rotors on the aircraft aerodynamic. The influence of the angle of attack, sideslip angle and the change of rotor tilt angle was investigated. Evaluation of the influence was made by the stability margin analysis and the selected stability derivatives computation. Practical implications Presented results could be very useful in the computation of dynamic stability of unconventional aircraft. Moreover, results could be helpful during designing the aircraft in the tandem wing configuration. Originality/value This paper presents the aerodynamic analysis of the unconventional configuration of the aircraft which combines the tandem wing feature with the tilt-rotor advantages. The impact of disturbance generated by the front and rear rotors on the flow around the aircraft was investigated. Moreover, the impact of rotors configuration on the aircraft static stability was found too.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-19T07:56:14Z
      DOI: 10.1108/AEAT-01-2018-0065
       
  • The mathematical model of UAV vertical take-off and landing
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a mathematical model of the dynamics of the unmanned aerial vehicle (UAV) vertical take-off and landing (VTOL). It will be used to develop control laws to a multirotor that is inherently unstable. Also, the model will be used to design algorithms to estimate the attitude of an object. Design/methodology/approach The physical model of UAV assumes that it is a rigid body with six degrees of freedom acted by forces generated by the propellers, motors, aerodynamic forces, gravity and disturbance forces. The mathematical model was described by differential equations. However, drive system (propeller, BLDC motor and BLDC motor controller) was described by six transfer functions. These transfer functions were demarcated with Matlab/Simulink identification toolbox from data received from a specially designed laboratory stand. Moments of inertia of the platform have been analytically determined and compared with empirical results from the pendulum. The mathematical model was implemented in Matlab/Simulink. Findings The paper confirms the need of designing mathematical models. Moreover, mathematical models show that some parts of the object are better to be replaced by experimental results than by equations, which is proved by the data. The paper also shows advantages of using Matlab/Simulink. What is more the simulation of the model proves that multirotor is an unstable object. Research limitations/implications The test results show that drive units are strongly dependent on ambient conditions. An additional problem is the different response of the drive set to increasing and decreasing the control signal amplitude. Next tests will be done at different temperatures and air densities of the environment, also it is need to explore drag forces. Practical implications The mathematical model is a simplification of the physical model expressed by means of equations. The results of simulation like accelerations and angular rate are noise-free. However, available sensors always have their errors and noise. To design control loops and attitude estimation algorithms, there is a need for identification of sensors’ errors and noise. These parameters have to be measured. Originality/value The paper describes a solution of correct identification of drive unit, which is a main component of the UAV.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T07:25:09Z
      DOI: 10.1108/AEAT-01-2018-0041
       
  • Passive control of cavity acoustics via the use of surface waviness at
           subsonic flow
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Aircraft noise is dominant for residents near airports when planes fly at low altitudes such as during departure and landing. Flaps, wings, landing gear contribute significantly to the total sound emission. This paper aims to present a passive flow control (in the sense that there is no power input) to reduce the noise radiation induced by the flow over the cavity of the landing gear during take-off and landing. Design/methodology/approach The understanding of the noise source mechanism is normally caused by the unsteady interactions between the cavity surface and the turbulent flows as well as some studies that have shown tonal noise because of cavity resonances; this tonal noise is dependent on cavity geometry and incoming flow that lead us to use of a sinusoidal surface modification application upstream of a cavity as a passive acoustics control device in approach conditions. Findings It is demonstrated that the proposed surface waviness showed a potential reduction in cavity resonance and in the overall sound pressure level at the majority of the points investigated in the low Mach number. Furthermore, optimum sinusoidal amplitude and frequency were determined by the means of a two-dimensional computational fluid dynamics analysis for a cavity with a length to depth ratio of four. Research limitations/implications The noise control by surface waviness has not implemented in real flight test yet, as all the tests are conducted in the credible numerical simulation. Practical implications The application of passive control method on the cavity requires a global aerodynamic study of the air frame is a matter of ongoing debate between aerodynamicists and acousticians. The latter is aimed at the reduction of the noise, whereas the former fears a corruption of flow conditions. To balance aerodynamic performance and acoustics, the use of the surface waviness in cavity leading edge is the most optimal solution. Social implications The proposed leading-edge modification it has important theoretical basis and reference value for engineering application it can meet the demands of engineering practice. Particularly, to contribute to the reduce the aircraft noise adopted by the “European Visions 2020”. Originality/value The investigate cavity noise with and without surface waviness generation and propagation by using a hybrid approach, the computation of flow based on the large-eddy simulation method, is decoupled from the computation of sound, which can be performed during a post-processing based on Curle’s acoustic analogy as implemented in OpenFOAM.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-18T02:30:41Z
      DOI: 10.1108/AEAT-01-2018-0061
       
  • Structural model with controls of a very light airplane for numerical
           flutter calculations
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The finite element model developed for a new-designed aircraft was used to solve some problems of structural dynamics. The key purpose of the task was to estimate the critical flutter velocities of the light airplane by performing numerical analysis with application of MSC Software. Design/methodology/approach Flutter analyses processed by Nastran require application of some complex aeroelastic model integrating two separate components – structural model and aerodynamic model. These sub-models are necessary for determining stiffness, mass and aerodynamic matrices, which are involved in the flutter equation. The aircraft structural model with its non-structural masses was developed in Patran. To determine the aerodynamic coefficient matrix, some simplified aerodynamic body-panel geometries were developed. The flutter equation was solved with the PK method. Findings The verified aircraft model was used to determine its normal modes in the range of 0-30 Hz. Then, some critical velocities of flutter were calculated within the range of operational velocities. As there is no certainty that the computed modes are in accordance with the natural ones, some parametric calculations are recommended. Modal frequencies depend on structural parameters that are quite difficult to identify. Adopting their values from the reasonable range, it is possible to assign the range of possible frequencies. The frequencies of rudder or elevator modes are dependent on their mass moments of inertia and rigidity of controls. The critical speeds of tail flutter were calculated for various combinations of stiffness or mass values. Practical implications The task described here is a preliminary calculational study of normal modes and flutter vibrations. It is necessary to prove the new airplane is free from flutter to fulfil the requirement considered in the type certification process. Originality/value The described approach takes into account the uncertainty of results caused by the indeterminacy of selected constructional parameters.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T12:59:38Z
      DOI: 10.1108/AEAT-01-2018-0059
       
  • Source term model for rod vortex generator
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The simulations of grid-resolved rod vortex generators (RVGs) require high computational cost and time. Additionally, the computational mesh topology must be adjusted to rods geometries. The purpose of this study is to propose the new source term model for RVG. Design/methodology/approach The model was proposed by modification of Bender, Anderson, Yagle (BAY) model used to predict flows around different type of vortex generators (VGs) – vanes. Original BAY model was built on lifting line theory. The proposed model was implemented in ANSYS Fluent by means of the user-defined function technique. Additional momentum and energy sources are imposed to transport equations. Findings The computational results of source term model were validated against experimental data and numerical simulation results for grid-resolved rod. It was shown that modified BAY model can be successfully used for RVG in complex cases. An example of BAY model application for RVG on transonic V2C airfoil with strongly oscillating shock waves is presented. Aerodynamic performance predicted numerically by means of both approaches (grid resolved RVG and modeled) is in good agreement, what indicates application opportunity of the proposed model to complex cases. Practical implications Modified BAY model can be used to simulate the influence of RVGs in complex real cases. It allows for time/cost reduction if the location or distribution of RVG has to be optimized on a profile, wing or in the channel. Originality/value In the paper, the new modification of BAY model was proposed to simulate RVGs. The presented results are innovative because of original approach to model RVGs.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T12:58:37Z
      DOI: 10.1108/AEAT-01-2018-0072
       
  • Neutral component of the local interstellar medium
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to review the current knowledge about the neutral component of the local interstellar medium (LISM), which due to the resonant charge exchange, photoionization and electron impact ionization processes has a profound impact on the heliosphere structure. Design/methodology/approach This work is based on the heliospheric literature review. Findings The summary of four major effects of neutral hydrogen atoms penetrating solar wind (SW), i.e. the disappearance of the complicated flow structure; the emergence of “hydrogen wall” in front of the heliopause (HP); decreasing distance of termination shock (TS), HP and bow shock (BS) layer from the Sun; and recently discovered by the Interstellar Boundary Explorer mission, a region of enhanced energetic neutral atom (ENA) emission seen in all sky maps as a ribbon. Practical implications In the context of constantly developing space technologies in aerospace engineering and prospective deep space missions, there is a need of general reviews about the interstellar space surroundings of the Sun and gathering the knowledge to help in theoretical, numerical and experimental investigations such as the optimization of the scientific equipment and spacecraft structure to work in specific conditions. Originality/value The survey encapsulate basic and relevant processes playing an important role in the physics of the nearest surroundings of the Sun and the latest results of numerical and experimental investigations focused on the neutral LISM component and its influence on the heliosphere, which is strongly desired in future works. Until now, not many of such reviews have been done.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T12:57:38Z
      DOI: 10.1108/AEAT-01-2018-0021
       
  • Aircraft model for the automatic taxi directional control design
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a concept of an automatic directional control system of remotely piloted aerial system (RPAS) during the taxiing phase. In particular, it shows the initial stages of the control laws synthesis-mathematical model and simulation of taxiing aircraft. Several reasons have emerged in recent years that make the automation of taxiing an important design challenge including decreased safety, performance and pilot workload. Design/methodology/approach The adapted methodology follows the model-based design approach in which the control system and the aircraft are mathematically modelled to allow control laws synthesis. The computer simulations are carried out to analyse the model behaviour. Findings Chosen methodology and modelling technique, especially tire-ground contact model, resulted in a taxiing aircraft model that can be used for directional control law synthesis. Aerodynamic forces and moments were identified in the wind tunnel tests for the full range of the slip angle. Simulations allowed to compute the critical speeds for different taxiway conditions in a 90° turn. Practical implications The results can be used for the taxi directional control law synthesis and simulation of the control system. The computed critical speeds can be treated as a safety limits. Originality/value The taxi directional control system has not been introduced to the RPAS yet. Therefore, the model of taxiing aircraft including aerodynamic characteristics for the full range of the slip angle has a big value in the process of design and implementation of the future auto taxi systems. Moreover, computed speed safety limits can be used by designers and standards creators.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T12:51:16Z
      DOI: 10.1108/AEAT-01-2018-0025
       
  • Design and analysis of a feedback loop to regulate the basic parameters of
           the unmanned aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The motivation to perform research on feedback control system for unmanned aerial vehicles, a fact that each quadrocopter is unstable. Design/methodology/approach For this reason, it is necessary to design a control system which is capable of making unmanned aerial vehicle vertical take-off and landing (UAV VTOL) stable and controllable. For this purpose, it was decided to use a feedback control system with cascaded PID controller. The main reason for using it was that PID controllers are simple to implement and do not use much hardware resources. Moreover, cascaded control systems allow to control object response using more parameters than in a standard PID control. STM32 microcontrollers were used to make a real control system. The rapid prototyping using Embedded Coder Toolbox, FreeRTOS and STM32 CubeMX was conducted to design the algorithm of the feedback control system with cascaded PID controller for unmanned aerial vehicle vertical take-off and landings (UAV VTOLs). Findings During research, an algorithm of UAV VTOL control using the feedback control system with cascaded PID controller was designed. Tests were performed for the designed algorithm in the model simulation in Matlab/Simulink and in the real conditions. Originality/value It has been proved that an additional control loop must have a full PID controller. Moreover, a new library is presented for STM32 microcontrollers made using the Embedded Coder Toolbox just for the research. This library enabled to use rapid prototyping while developing the control algorithms.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-17T01:07:58Z
      DOI: 10.1108/AEAT-01-2018-0039
       
  • Cold spraying and laser cladding as an alternative to electroplating
           processes
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to propose cold spraying and laser cladding processes as alternatives to cadmium and chromium electroplating, respectively. There are many substances or chemicals within the coating technology that can be identified as substances of very high concern because of their carcinogenic or mutagenic nature. Cadmium and chromium undoubtedly belong to these items and are the basic constituents of electrolytic coating processes. Finding an alternative and adapting to the existing restrictions of the usage of such hazardous products stands for many to be or not to be in the market. Design/methodology/approach The research work was focused on down selecting the appropriate materials, producing the coating samples, testing their properties and optimizing process parameters by statistical method. On the one hand, the high-pressure cold spray system and spraying of the titanium coating on the landing gear component, and on the other hand, the high-energy laser cladding facility and the wear resistant cobalt-based coating deposited onto the shock absorber piston. Substrates of these two applications were made of the same material, 4330 – high-strength low-carbon steel. Findings Meeting the requirements of Registration, Evaluation, Authorization and Restriction of Chemicals implies undertaking research and implementation work to identify alternative processes. The work provides the technical characteristics of new coatings justifying application readiness of the researched processes. Originality/value Taguchi’s design of experiment method was combined with the measurements and analysis of specified coating properties for the optimization of the cold spray process parameters. There is also laser cladding process development presented as a fast rate technology generating coatings with the unique properties.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-15T02:01:57Z
      DOI: 10.1108/AEAT-01-2018-0071
       
  • Operations reliability study of small aircraft powered by piston engines
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Europe has adopted Flight Path 2050 (FP 2050) challenge with an objective of 90 per cent of the travelers being able to reach door-to-door European destinations within 4 hours by 2050. The aim can be achieved by reliable, well-organized small aircraft transport (SAT). Analysis of the currently operating small aircraft operational reliability data will support the development of future aircraft designs as well as reliability and safety requirements necessary for commercial operations. Design/methodology/approach The paper provides results of a statistical analysis of small aircraft current operations based on the reported events contained in the Database named European Coordination Centre for Aviation Incident Reporting Systems database. It presents identified safety indicators and focuses particularly on those related to the aviation technology. Findings It has been found that certain airframe and powerplant systems have the biggest influence on flight safety. Practical implications Multidisciplinary analysis of the operational and aircraft components reliability data will help in a proper preparation of the SAT supporting facilities, a design process of new aircraft and improvements of the existing airframe and powerplant systems. Originality/value Presented results are valuable for further developments of the statistical tools facilitating new product introduction.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-15T01:57:36Z
      DOI: 10.1108/AEAT-01-2018-0005
       
  • Application of computed tomography to hole expansion measurements
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to investigate a possibility of determination of the rivet hole expansion with the use of computed tomography (CT). This method offers several benefits in comparison to the traditionally used destructive methods. Design/methodology/approach The measurements of rivet hole expansion were performed on three specimens with the use of CT. Then, the same specimens were measured with the use of the conventional destructive method. This allows to estimate accuracy of the proposed method and characterize its advantages and limitations. Findings Good correlation with the destructive method has been obtained. The proposed method enables more detailed analysis of a joint as arbitrarily oriented cross-section for analysed area can be easily generated and increase of measurements number is always possible and simple. The disadvantage of the method is lower accuracy of diameter determination than in the case of conventional methods. Research limitations/implications The measurements were performed only on one type of specimens. Probably, if a rivet and sheets were made of the same alloy, the measurements would be barely possible. The rivets were installed with squeezing ratio D/Do = 1.7 whose value is close to maximum as defined in riveting instructions (Kaniowski, 2015). This means that measured hole expansions were higher than in typical joint. The proposed method is appropriate for simple specimens (one rivet at a specimen width). Practical implications The investigation shows that rivet hole expansion can be measured with the use of CT. This method is useful especially when destruction of a specimen is not allowed or more detailed analysis is required (e.g. measurements on many depth levels). Originality/value The paper presents measurements of rivet hole expansion with the method which has not been used before for this application. Advantages and limitations of the proposed approach compared to conventional methods are discussed.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-11T09:53:16Z
      DOI: 10.1108/AEAT-01-2018-0040
       
  • Partially feasible solution space for integrated SATS operations
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to investigate the feasibility of solving an integrated flight scheduling, fleet assignment and crew pairing problem for an on-demand service using a small, up to 19-seater, aircraft. Design/methodology/approach Evolutionary algorithm is developed to solve the problem. Algorithm design assumes indirect solution representation that allows to evaluate partially feasible solutions only and speed up calculations. Tested algorithm implementation takes advantage of the graphic processing unit. Findings Performed tests confirm that the algorithm can successfully solve the defined integrated scheduling problem. Practical implications The presented algorithm allows to optimise on-demand transport service operation within minutes. Social implications Optimisation of operation cost contributes to better accessibility of transport. Originality/value The presented integrated formulation allows to avoid sub optimal solutions that are results of solving sequential sub problems. Indirect representation and evaluation strategy can be applied to speed up calculations in other problems as well.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-11T09:46:37Z
      DOI: 10.1108/AEAT-01-2018-0045
       
  • Aircraft model for automatic taxi directional control system design
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a concept of an automatic directional control system of remotely piloted aerial system (RPAS) during the taxiing phase. In particular, it shows the initial stages of the control laws synthesis – mathematical model and simulation of taxiing aircraft. Several reasons have emerged in recent years that make the automation of taxiing an important design challenge including decreased safety, performance and pilot workload. Design/methodology/approach The adapted methodology follows the model-based design approach in which the control system and the aircraft are mathematically modelled to allow control laws synthesis. The computer simulations are carried out to analyse the model behaviour. Findings Chosen methodology and modelling technique, especially tire-ground contact model, resulted in a taxing aircraft model that can be used for directional control law synthesis. Aerodynamic forces and moments were identified in the wind tunnel tests for the full range of the slip angle. Simulations allowed to compute the critical speeds for different taxiway conditions in a 90° turn. Practical implications The results can be used for the taxi directional control law synthesis and simulation of the control system. The computed critical speeds can be treated as safety limits. Originality/value The taxi directional control system has not been introduced to the RPAS yet. Therefore, the model of taxiing aircraft including aerodynamic characteristics for the full range of the slip angle has a big value in the process of design and implementation of the future auto taxi systems. Moreover, computed speed safety limits can be used by designers and standard creators.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-11T02:42:40Z
      DOI: 10.1108/AEAT-06-2018-0161
       
  • Flow-separation-control system operating in feedback closed loop
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this study is to develop the concept of self-adapting system which would be able to control a flow on the wing-high-lift system and protect the flow against strong separation. Design/methodology/approach The self-adapting system has been developed based on computational approach. The computational studies have been conducted using the URANS solver. The experimental investigations have been conducted to verify the computational results. Findings The developed solution is controlled by closed-loop-control (CLC) system. As flow actuators, the main-wing trailing-edge nozzles are proposed. Based on signals received from the pressure sensors located at the flap trailing edge, the CLC algorithm changes the amount of air blown from the nozzles. The results of computational simulations confirmed good effectiveness and reliability of the developed system. These results have been partially confirmed by experimental investigations. Research limitations/implications The presented research on an improvement of the effectiveness of high-lift systems of modern aircraft was conducted on the relatively lower level of the technology readiness. However, despite this limitation, the results of presented studies can provide a basis for developing innovative self-adaptive aerodynamic systems that potentially may be implemented in future aircrafts. Practical implications The studies on autonomous flow-separation control systems, operating in a closed feedback loop, are a great hope for significant advances in modern aeronautical engineering, also in the UAV area. The results of the presented studies can provide a basis for developing innovative self-adaptive aerodynamic systems at a higher level of technological readiness. Originality/value The presented approach is especially original and valuable in relation to the innovative concept of high-lift system supported by air-jets blown form the main-wing-trailing-edge nozzles; the effective and reliable flow sensors are the pressure sensors located at the flap trailing edge, and the effective and robust algorithm controlling the self-adapting aerodynamic system – original especially in respect to a strategy of deactivation of flow actuators.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-11T02:17:57Z
      DOI: 10.1108/AEAT-12-2017-0270
       
  • Comparison of flutter calculation methods based on ground vibration test
           result
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      PurposeA low-cost but credible method of low-subsonic flutter analysis based on ground vibration test (GVT) results is presented. The purpose of this paper is a comparison of two methods of immediate flutter problem solution: JG2 – low cost software based on the strip theory in aerodynamics (STA) and V-g method of the flutter problem solution and ZAERO I commercial software with doublet lattice method (DLM) aerodynamic model and G method of the flutter problem solution. In both cases, the same sets of measured normal modes are used. Design/methodology/approachBefore flutter computation, resonant modes are supplied by some non-measurable but existing modes and processed using the author’s own procedure. For flutter computation, the modes are normalized using the aircraft mass model. The measured mode orthogonalization is possible. The flutter calculation made by means of both methods are performed for the MP-02 Czajka UL aircraft and the Virus SW 121 aircraft of LSA category. FindingsIn most cases, both compared flutter computation results are similar, especially in the case of high aspect wing flutter. The Czajka T-tail flutter analysis using JG2 software is more conservative than the one made by ZAERO, especially in the case of rudder flutter. The differences can be reduced if the proposed rudder effectiveness coefficients are introduced. Practical implicationsThe low-cost methods are attractive for flutter analysis of UL and light aircraft. The paper presents the scope of the low-cost JG2 method and its limitations. Originality/valueIn comparison with other works, the measured generalized masses are not used. Additionally, the rudder effectiveness reduction was implemented into the STA. However, Niedbal (1997) introduced corrections of control surface hinge moments, but the present work contains results in comparison with the outcome obtained by means of the more credible software.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-09T09:53:38Z
      DOI: 10.1108/AEAT-03-2018-0102
       
  • Suppression of nonlinear aeroelastic vibrations by learned neural network
           controller
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of the paper is to analyze the active suppression of the aeroelastic vibrations of ailerons with strongly nonlinear characteristics by neural network/reinforcement learning (NN/RL) control method and comparing it with the classic robust methods of suppression. Design/methodology/approach The flexible wing and aileron with hysteresis nonlinearity is treated as a plant-controller system and NN/RL and robust controller are used to suppress the nonlinear aeroelastic vibrations of aileron. The simulation approach is used for analyzing the efficiency of both types of methods in suppressing of such vibrations. Findings The analysis shows that the NN/RL controller is able to suppress the nonlinear vibrations of aileron much better than linear robust method, although its efficiency depends essentially on the NN topology as well as on the RL strategy. Research limitations/implications Only numerical analysis was carried out; thus, the proposed solution is of theoretical value, and its application to the real suppression of aeroelastic vibrations requires further research. Practical implications The work shows the NN/RL method has a great potential in improving suppression of highly nonlinear aeroelastic vibrations, opposed to the classical robust methods that probably reach their limits in this area. Originality/value The work raises the questions of controllability of the highly nonlinear aeroelastic systems by means of classical robust and NN/RL methods of control.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-08T11:39:34Z
      DOI: 10.1108/AEAT-01-2018-0019
       
  • UAV application for precision agriculture
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this study is to show the potentials of a cost-effective unmanned aerial vehicles (UAV) system for agriculture industry. The current population growth rate is so vast that farming industry must be highly efficient and optimized. As a response for high quality food demands, the new branch of the agriculture industry has been formed – the precision agriculture. It supports farming process with sensors, automation and innovative technologies. The UAV advantages over regular aviation are withering. Not only they can fly at lower altitude and are more precise but also offer same high quality and are much cheaper. Design/methodology/approach The main objective of this project was to implement an exemplary cost-effective UAV system with electronic camera stabilizer for gaining useful data for agriculture. The system was based on small, unmanned flying wing able to perform fully autonomous missions, a commercially available camera and an own-design camera stabilizer. The research plan was to integrate the platform and run numerous experimental flights over farms, fields and woods collecting aerial pictures. All the missions have been planned to serve for local farming and forest industries and cooperated with local business authorities. Findings In preliminary flight tests, the variety of geodetic, forest and agriculture data have been acquired, placed for post processing and applied for the farming processes. The results of the research were high quality orthophoto maps, 3D maps, digital surface models and images mosaics with normalized difference vegetation index. The end users were astonished with the high-quality results and claimed the high importance for their business. Originality/value The case study results proved that this kind of a small UAV system is exceptional to manage and optimize processes at innovative farms. So far only professional, high-cost UAV platforms or traditional airships have been applied for agriculture industry. This paper shows that even simple, commercially available equipment could be used for professional applications.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-08T11:35:33Z
      DOI: 10.1108/AEAT-01-2018-0056
       
  • Analysis and optimization of morphing wing aerodynamics
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present a method for analysis and optimization of morphing wing. Moreover, a numerical advantage of morphing airfoil wing, typically assessed in simplified two-dimensional analysis is found using higher fidelity methods. Design/methodology/approach Because of multi-point nature of morphing wing optimization, an approach for optimization by analysis is presented. Starting from naïve parametrization, multi-fidelity aerodynamic data are used to construct response surface model. From the model, many significant information are extracted related to parameters effect on objective; hence, design sensitivity and, ultimately, optimal solution can be found. Findings The method was tested on benchmark problem, with some easy-to-predict results. All of them were confirmed, along with additional information on morphing trailing edge wings. It was found that wing with morphing trailing edge has around 10 per cent lower drag for the same lift requirement when compared to conventional design. Practical implications It is demonstrated that providing a smooth surface on wing gives substantial improvement in multi-purpose aircrafts. Details on how this is achieved are described. The metodology and results presented in current paper can be used in further development of morphing wing. Originality/value Most of literature describing morphing airfoil design, optimization or calculations, performs only 2D analysis. Furthermore, the comparison is often based on low-fidelity aerodynamic models. This paper uses 3D, multi-fidelity aerodynamic models. The results confirm that this approach reveals information unavailable with simplified models.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-08T11:20:46Z
      DOI: 10.1108/AEAT-12-2017-0289
       
  • A numerical study into the longitudinal dynamic stability of the tailless
           aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this research is a study into a mathematical approach of a tailless aircraft dynamic stability analysis. This research is focused on investigation of influence of elevons (elevator) on stability derivatives and consequently on the aircraft longitudinal dynamic stability. The main research question is to determine whether this impact should be taken into account on the conceptual and preliminary stage of the analysis of the longitudinal dynamic stability. Design/methodology/approach Aerodynamic coefficients and longitudinal stability derivatives were computed by Panukl (panel methods). The analysis of the dynamic stability of the tailless aircraft was made by the Matlab code and SDSA package. Findings The main result of the research is a comparison of the dynamic stability of the tailless aircraft for different approaches, with and without the impact of elevator deflection on the trim drag and stability derivatives. Research limitations/implications This paper presents research that mostly should be considered on the preliminary stage of aircraft design and dynamic stability analysis. The impact of elevons deflection on the aircraft moment of inertia has been omitted. Practical implications The results of this research will be useful for the further design of small tailless unmanned aerial vehicles (UAVs). Originality/value This research reveals that in case of the analysis of small tailless UAVs, the impact of elevons deflection on stability derivatives is bigger than the impact of a Mach number. This impact should be taken into consideration, especially for a phugoid mode.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-02T01:26:19Z
      DOI: 10.1108/AEAT-01-2018-0032
       
  • Aerodynamic and flight dynamic interaction in spin
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to describe an integrated approach to spin analysis based on 6-DOF (degrees of freedom) fully nonlinear equations of motion and a three-dimensional multigrid Euler method used to specify a flow model. Another purpose of this study is to investigate military trainer performance during a developed phase of a deliberately executed spin, and to predict an aircraft tendency while entering a spin and its response to control surface deflections needed for recovery. Design/methodology/approach To assess spin properties, the calculations of aerodynamic characteristics were performed through an angle-of-attack range of −30 degrees to +50 degrees and a sideslip-angle range of −30 degrees to +30 degrees. Then, dynamic equations of motion of a rigid aircraft together with aerodynamic loads being premised on stability derivatives concept were numerically integrated. Finally, the examination of light turboprop dynamic behaviour in post-stalling conditions was carried out. Findings The computational method used to evaluate spin was positively verified by comparing it with the experimental outcome. Moreover, the Euler code-based approach to lay down aerodynamics could be considered as reliable to provide high angles-of-attack characteristics. Conclusions incorporate the results of a comparative analysis focusing especially on comprehensive assessment of output data quality in relation to flight tests. Originality/value The conducted calculations take into account aerodynamic and flight dynamic interaction of an aerobatic-category turboprop in spin conditions. A number of manoeuvres considering different aircraft configurations were simulated. The computational outcomes were subsequently compared to the results of in-flight tests and the collected data were thoroughly analysed to draw final conclusions.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-10-02T01:09:38Z
      DOI: 10.1108/AEAT-01-2018-0042
       
  • Novel coupled model for power loss prediction in a record-breaking
           electric aircraft motor
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to devise an analytical approach to calculate conductor winding losses, considering multiple contributing aspects simultaneously. These include the geometric configuration of coil windings, frequency of the electric current and the dependency on the coil temperature, derived studying a coupled fluid–solid model considering the cooling system characteristics. The obtained results allow identifying power loss trends according to such system variables as coolant inlet temperature or overall flow rate of the motor. Design/methodology/approach An easy-to-use coupled analytical approach is applied, which is suitable for rapid estimations of the impact of parameter variation on the resulting conductor winding power losses that facilitates decision-making in the design process of electric aircraft engines. Findings In the considered cooling parameters, the overall conductor winding power losses vary approximately between 6 kW and 7.2 kW. More than 95 per cent of this loss is because of direct current losses. These losses cause the variation in maximal coil temperature ranging between 115°C and 170°C. Practical implications The SP260D motor is set and was currently tested in Extra 330. It recently broke two world records. Social implications One of the current trends in aircraft engineering is electric aircraft. Advantages of electric aircraft include improved manoeuvrability because of greater torque from electric motors, increased safety because of decreased chance of mechanical failure, less risk of explosion or fire in the event of a collision and less noise. There will be environmental and cost benefits associated with the elimination of dependency on fossil fuels and resultant emissions. Originality/value The use of a novel fluid–solid interaction model for predicting conductor winding power loss of the SP260D electric aircraft motor has not been done earlier. A novel alternative derivation of the widely applied Dowell’s formula (Dowell, 1966) is presented for the estimation of proximity losses in square winding conductors.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-09-27T02:44:38Z
      DOI: 10.1108/AEAT-12-2017-0278
       
  • Multirotor UAV sensor fusion for precision landing
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose In relation to rapid development of possible applications of unmanned vehicles, new opportunities for their use are emerging. Among the most dynamic, we can distinguish package shipments, rescue and military applications, autonomous flights and unattended transportation. However, most of the UAV solutions have limitations related to their power supplies and the field of operation. Some of these restrictions can be overcome by implementing the cooperation between unmanned aerial vehicles (UAVs) and unmanned ground vehicles (UGVs). The purpose of this paper is to explore the problem of sensor fusion for autonomous landing of a UAV on the UGV by comparing the performance of precision landing algorithms using different sensor fusions to have precise and reliable information about the position and velocity. Design/methodology/approach The difficulties in this scenario, among others, are different coordination systems and necessity for sensor data from air and ground. The most suitable solution seems to be the use of widely available Global Navigational Satellite System (GNSS) receivers. Unfortunately, the position measurements obtained from cheap receivers are encumbered with errors when desiring precision. The different approaches are based on the usage of sensor fusion of Inertial Navigation System and image processing. However most of these systems are very vulnerable to lightning. Findings In this paper, methods based on an exchange of telemetry data and sensor fusion of GNSS, infrared markers detection and others are used. Different methods are compared. Originality/value The subject of sensor fusion and high-precision measurements in reference to the autonomous vehicle cooperation is very important because of the increasing popularity of these vehicles. The proposed solution is efficient to perform autonomous landing of UAV on the UGV.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2018-09-18T01:15:59Z
      DOI: 10.1108/AEAT-01-2018-0070
       
 
 
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