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Journal Cover Aircraft Engineering and Aerospace Technology
  [SJR: 0.391]   [H-I: 18]   [174 followers]  Follow
   Hybrid Journal Hybrid journal (It can contain Open Access articles)
   ISSN (Print) 0002-2667 - ISSN (Online) 1748-8842
   Published by Emerald Homepage  [335 journals]
  • 12th READ Conference & 6th SCAD Symposium
    • Authors: Tomasz Goetzendorf-Grabowski
      First page: 509
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.

      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:02Z
      DOI: 10.1108/AEAT-03-2017-0095
  • Design of novel aerial jet target
    • Authors: Zdobyslaw Jan Goraj, Marek Malinowski, Andrzej Frydrychewicz
      First page: 511
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose Aiming to present and discuss the requirements for flying targets which sometimes are contradictory to each other, and to perform a trade-off analysis before the design activity is started. To demonstrate conceptual and preliminary design processes using a practical example of PW-61 configuration. To show how results of experimental flight tests using a scaled flying target will be described and analysed before manufacturing the full scale flying target. Design/methodology/approach Important part of the paper consists of the selection of tailplane configuration of the flying target UAV in order to protect some expensive on-board systems against serious damages, and to obtain a sufficient dynamic stability, independently of the amount of the petrol in fuel tank. Inverted V-tail, U-tail and H-tail configurations were considered and compared both, theoretically and in flight experiments. Findings Flight dynamics models and associated computational procedures were useful both, in a preliminary design phase and during the final assessment of the configuration after flight tests. Selection of the tailplane configuration for the flying target UAV is very important in order to protect some expensive on-board systems against serious damages, and to obtain a sufficient dynamic stability, independent of the amount of the petrol in fuel tank. Practical implications Flying targets should be speedy, maneuverable, cheap, easy in deployment and multi-recoverable (if not destroyed by live ammunition), must have relatively low take-off weight and an endurance of at least 1 hour. This paper can be useful for proper selection of requirements and preliminary design parameters to make the design process more economically effective. Originality/value This paper presents very efficient methods of assessing the design parameters of flying targets, especially in an early stage of the design process. Stability computations are performed based on equations of motion and are supplemented by flight tests using the scaled flying models. It can be considered as an original, not typical but very practical approach, because it delivers lots of data in the early design stages at relatively low cost.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:20:56Z
      DOI: 10.1108/AEAT-10-2016-0174
  • Sizing implications of a regional aircraft for inner-city operations
    • Authors: Philipp Heinemann, Michael Schmidt, Felix Will, Sascha Kaiser, Christoph Jeßberger, Mirko Hornung
      First page: 520
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose The paper aims to assess the potential of aircraft operation from city centres to achieve shortened travel times and the involved aircraft design process. Design/methodology/approach The paper describes the methodical approach and iterative procedure of the design process. An assessment of potential technologies is conducted to provide the required enhancements to fulfil the constraints following an inner-city operation. Operational procedures were analysed to reduce the noise propagation through flight path optimization. Furthermore a ground based assisted take-off system was conceived to lower required take-off field length and to prevent engine sizing just for the take-off case. Cabin design optimization for a fast turnaround has been conducted to ensure a wide utilization spectrum. The results prove the feasibility of an aircraft developed for inner city operation. Findings A detailed concept for a 60 passenger single aisle aircraft is proposed for an Entry-Into-Service year 2040 with a design range of 1,500 nautical miles for a load factor of 90 %. Although the design for STOL and low noise operation had to be traded partly with cruise efficiency, a noteworthy reduction in fuel burn per passenger and nautical mile could be achieved against current aircraft. Practical implications The findings will contribute to the evaluation of the feasibility and impact of the Flightpath 2050 goal of a four hour door-to-door by providing a feasible but ambitious example. Furthermore it highlights possible bottlenecks and problems faced when realizing this goal. Originality/value The paper draws its value from the consideration of the overall sizing effects at aircraft level and from a holistic view on an inner-city airport/aircraft concept design for a four hour door-to-door goal.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:07Z
      DOI: 10.1108/AEAT-11-2016-0196
  • Morphing wing with skin discontinuity - kinematic concept
    • Authors: Andrzej Tarnowski
      First page: 535
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose This paper describes concept of morphing tailless aircraft with discontinuous skin, and its preliminary kinematic solution. Project assumptions, next steps and expected results are briefly presented. Design/methodology/approach Multidisciplinary numerical optimization will be used to determine control allocation for wing segments rotation. Wing demonstrator will be fabricated and tested in Wind Tunnel. Results will be used in construction of flying model and design of its control system. Flight data of morphing demonstrator and reference aircraft will result in comparative analysis of both technologies. Findings Proposed design combines advantages of wing morphing without complications of wing’s structure elastic deformation. Better performance, stability and maneuverability is expected due to wing’s construction which is entirely composed of unconnected wing segments. Independent control of each segment allows for free modeling of spanwise lift force distribution. Originality/value Nonlinear multipoint distribution of wing twist as the only mechanism for control and flight performance optimization has never been studied or constructed. Planned wind tunnel investigation of such complex aerodynamic structure has not been previously published and will be an original contribution to the development of aviation and in particular the aerodynamics of wing with discontinuous skin.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:08Z
      DOI: 10.1108/AEAT-11-2016-0208
  • A comprehensive review of vertical tail design
    • Authors: Fabrizio Nicolosi, Danilo Ciliberti, Pierluigi Della Vecchia, Salvatore Corcione, Vincenzo Cusati
      First page: 547
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose This work deals with a comprehensive review of design methods for aircraft directional stability and vertical tail sizing. The focus on aircraft directional stability is due to the significant discrepancies that classical semi-empirical methods, as USAF DATCOM and ESDU, provide for some configurations, since they are based on NACA wind tunnel tests about models not representative of an actual transport airplane. Design/methodology/approach The authors performed viscous numerical simulations to calculate the aerodynamic interference among aircraft parts on hundreds configurations of a generic regional turboprop aircraft, providing useful results that have been collected in a new vertical tail preliminary design method, named VeDSC. Findings The reviewed methods have been applied on a regional turboprop aircraft. The VeDSC method shows the closest agreement with numerical results. A wind tunnel test campaign involving more than 180 configurations has validated the numerical approach. Practical implications The investigation has covered both the linear and the non-linear range of the aerodynamic coefficients, including the mutual aerodynamic interference between the fuselage and the vertical stabilizer. Also, a preliminary investigation about rudder effectiveness, related to aircraft directional control, is presented. Originality/value In the final part of the paper, critical issues in vertical tail design are reviewed, highlighting the significance of a good estimation of aircraft directional stability and control derivatives.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:05Z
      DOI: 10.1108/AEAT-11-2016-0213
  • Adaptive design of experiments for efficient and accurate estimation of
           aerodynamic loads
    • Authors: Andrea Da Ronch, Marco Panzeri, M. Anas Abd Bari, Roberto d'Ippolito, Matteo Franciolini
      First page: 558
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose The purpose of this paper is to document an efficient and robust approach to generate aerodynamic tables using computational fluid dynamics. This is demonstrated in the context of a concept transport aircraft model. Design/methodology/approach Two design of experiments algorithms in combination with surrogate modelling are investigated. An adaptive algorithm is compared to an industry-standard algorithm used as benchmark. Numerical experiments are obtained solving the Reynolds-averaged Navier-Stokes equations on a large computational grid. Findings This study demonstrates that a surrogate model built upon an adaptive design of experiments strategy achieves a higher prediction capability than that built upon a traditional strategy. This is quantified in terms of the sum of the squared error between the surrogate model predictions and the computational fluid dynamics results. The error metric is reduced by about one order of magnitude compared to the traditional approach. Practical implications This work lays the ground to obtain more realistic aerodynamic predictions earlier in the aircraft design process at manageable costs, improving the design solution and reducing risks. This may be equally applied in the analysis of other complex and non-linear engineering phenomena. Originality/value This work explores the potential benefits of an adaptive design of experiment algorithm within a prototype working environment, whereby the maximum number of experiments is limited and a large parameter space is investigated.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:01Z
      DOI: 10.1108/AEAT-10-2016-0173
  • Practical problems of numerical optimization in aerospace sciences
    • Authors: Jacek Mieloszyk
      First page: 570
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose Application of numerical optimization to the aircraft design procedures applied in the airspace industry. Design/methodology/approach It is harder than ever to achieve competitive construction. This is why numerical optimization is becoming standard tool during the design process. Although, optimization procedures are becoming more mature, yet in the industry practice, fairly simple examples of optimization are present. The more complicated is the task to solve, the harder it is to implement automated optimization procedures. The article presents practical examples of optimization in aerospace sciences. The methodology is discussed in the article in great detail. Findings Encountered problems related to the numerical optimization are presented. Different approaches to the solutions of the problems are shown, which have impact on the time of optimization computations and quality of the obtained optimum. Achieved results are discussed in detail with relation to the used settings. Practical implications Investigated different aspects of handling optimization problems, improving quality of the obtained optimum, or speeding-up optimization by parallel computations can be directly applied in the industry optimization practice. Lessons learned from multidisciplinary optimization can bring industry products to higher level of performance and quality, i.e. more advanced, competitive, and efficient aircraft design procedures, which could be applied in the industry practice. This can lead to the new approach of aircraft design process. Originality/value Introduction of numerical optimization methods in aircraft design process. Showing how to solve numerical optimization problems related to advanced cases of conceptual and preliminary aircraft design.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:20:59Z
      DOI: 10.1108/AEAT-11-2016-0201
  • The electric-powered motorglider AOS-71 - the study of development
    • Authors: Jędrzej Marjanowski, Jan Tomasiewicz, Wojciech Frączek
      First page: 579
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose The purpose of this paper is to present the process of design and prototyping of a two-seat electric powered self-launching motorglider AOS-71 closely connected with the teaching process conducted by Warsaw University of Technology (WUT) academic staff within a unique educational ULS – Ultra Light Sailplanes program. Design/methodology/approach selected design methods and tools used during the development of the motorglider have been described. The CAD/CAM modules of the Siemens NX software were used to work on the structural design, tools and technical documentation. The core of the ULS educational program is to educate aerospace engineering students by providing an opportunity for them to participate in each phase of the aircraft life cycle - from conceptual drawings through structural design and prototyping to manufacturing, testing and maintenance. Findings The main innovations of the AOS-71 design are retractable ecological electric propulsion, spacious cockpit where seats are located side by side and the all-composite airframe in 90% made of advanced carbon epoxy composites. Practical implications The demonstrator of an electric motorglider that can be used as a multifunctional flying laboratory for flight research and student education. Originality/value The AOS-71 project and its continuation are a valuable example of involving aerospace students in each phase of the aircraft life cycle. It also contributes to the research in the field of utilizing innovative electrical propulsion system in aircraft design.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:04Z
      DOI: 10.1108/AEAT-11-2016-0218
  • Agricultural aircraft wing slat tolerance for bird strike
    • Authors: Adam Deskiewicz, Rafal Perz
      First page: 590
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose The aim of this study is to assess and describe possible consequences of a bird strike on a Polish designed PZL-106 Kruk agricultural aircraft. Due to its susceptibility to such events, a wing slat has been chosen for analysis. Design/methodology/approach Smooth Particle Hydrodynamics (SPH) formulation has been used for generation of the bird finite element model. The simulations were performed by the LS-Dyna explicit finite element analysis software. Several test cases have been analysed with differing parameters such as impact velocity, initial velocity vector direction, place of impact and bird mass. Findings Results of this study reveal that the structure remains safe after an impact at the velocity of 25 m/s. The influence of bird mass on slat damage is clearly observable when the impact velocity rises to 60 m/s. Another important finding was that in each case where the part did not withstand the applied load, it was the lug where first failure occurred. Some of the analysed cases indicated the possibility a consequent wing box damage. Practical implications This finding provides the manufacturer an important insight into the behaviour of the slat and suggests that more detailed analysis of the current lug design might improve the safety of the structure. Originality/value Even though similar analyses have been performed, they tended to focus on large transport aircraft components. This investigation will enhance our understanding of structural response of small, low-speed aircraft to a bird impact, which is a realistic scenario for the chosen case of an agricultural plane.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:00Z
      DOI: 10.1108/AEAT-11-2016-0220
  • Java framework for parametric aircraft design – ground performance
    • Authors: Vittorio Trifari, Manuela Ruocco, Vincenzo Cusati, Fabrizio Nicolosi, Agostino De Marco
      First page: 599
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose This paper introduces the take-off and landing performance analysis modules of the software library named JPAD (Java toolchain of Programs for Aircraft Design), dedicated to the aircraft preliminary design. An overview of JPAD is also presented. Design/methodology/approach The calculation of the take-off and landing distances have been implemented using a simulation-based approach. This expects to solve an appropriate set of ordinary differential equations (ODE), which describes the aircraft equations of motion during all the take-off and landing phases. Tests upon two aircraft models (ATR72 and B747-100B) have been performed in order to compare the obtained output with the performance data retrieved from the related flight manuals. Findings The tool developed has proven to be very reliable and versatile as it performs the calculation of the required performance with almost no computational effort and with a good accuracy, providing a less than the 5% difference with respect to the statistical trend and a difference from the flight manual or public brochure data around 10%. Originality/value The use of a simulation-based approach in order to have a more accurate estimation of the ground performance with respect to classic semi-empirical equations. Although performing the simulation of the aircraft motion, the approach shown is very time-saving and can be easily implemented in an optimization cycle.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:09Z
      DOI: 10.1108/AEAT-11-2016-0209
  • Robust design and optimization of UAV empennage
    • Authors: Witold Artur Klimczyk, Zdobyslaw Jan Goraj
      First page: 609
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose Aircraft design optimization often narrowed to analysis of cruise conditions does not take into account other flight phases (manoeuvres). These, especially in UAV sector, can be significant part of the whole flight. Empennage is a part of the aircraft, with crucial function for manoeuvres. It is important to consider robustness for highest performance. Design/methodology/approach Methodology for robust wing design is presented. Surrogate modelling using kriging is used to reduce the optimization cost for high-fidelity aerodynamic calculations. Analysis of varying flight conditions- angle of attack, is made to assess robustness of design for particular mission. Two cases are compared: global optimization of 11 parameters and optimization divided into 2 consecutive sub-optimizations. Findings Surrogate modelling proves its usefulness for cutting computational time. Optimum design found by splitting problem into sub-optimizations finds better design at lower computational cost. Practical implications It is demonstrated, how surrogate modelling can be used for analysis of robustness, and why it is important to consider it. Intuitive split of wing design into airfoil and planform sub-optimizations brings promising savings in the optimization cost. Originality/value Methodology presented in this paper can be used in various optimization problems, especially those involving expensive computations and requiring top quality design.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:20:57Z
      DOI: 10.1108/AEAT-11-2016-0221
  • Cross-flow effects regarding laminar flow control within conceptual
           aircraft design
    • Authors: Florian Schueltke, Eike Stumpf
      First page: 620
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 4, July 2017.
      Purpose Laminarization of commercial aircraft surfaces is the most promising technology to reduce fuel consumption and ecological impact. Since laminar flow highly depends on cross-flow effects, there is the question in which way simple estimations and simplifications for application in conceptual aircraft design can be used to capture these cross-flow influences. This investigation aims to show the accuracy of 2D methods for estimating laminar flow regions on 3D wing objects. Design/methodology/approach Several methods, relating 3D and 2D flow conditions, are analyzed with regard to capture cross-flow influences. The 3D pressure distributions depending on utilized transformation method are compared to Reynolds-averaged Navier-Stokes (RANS) solutions. With the most precise transformation method, the laminar flow area on a conventional wing of a short range aircraft is determined and compared to the laminar area obtained with the RANS pressure distributions as input. Further, HLFC component sizing is carried out to obtain the net benefit in fuel reduction of simplified method compared to RANS method for a conventional short range aircraft. Findings In this particular case, the solutions calculated with the simplified methods show high deviations from those obtained with RANS. Originality/value This investigation underlines the need of proper methods for fast and accurate estimation of cross-flow effects to be able to assess the full potential of laminar flow control within conceptual aircraft design.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-05-30T11:21:04Z
      DOI: 10.1108/AEAT-11-2016-0210
  • Conjugate rotor-stator interaction procedure for film-cooled turbine
    • First page: 365
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The authors present a procedure for the parallel, steady and unsteady conjugate, Navier-Stokes/heat-conduction rotor-stator interaction analysis of multi-blade-row, film-cooled, turbine airfoil sections. A new grid generation procedure for multiple blade-row configurations including walls, thermal barrier coatings, plenums, and cooling tubes is discussed. Design/methodology/approach Steady, multi-blade-row interaction effects on the flow and wall thermal fields are predicted using a Reynolds’s-averaged Navier-Stokes simulation in conjuction with an inter-blade-row mixing plane. Unsteady, aero-thermal interaction solutions are determined using time-accurate sliding grids between the stator and rotor with an unsteady Reynolds’s-averaged Navier-Stokes model. Non-reflecting boundary condition treatments are utilized in both steady and unsteady approaches at all inlet, exit, and inter-blade-row boundaries. Parallelization techniques are also discussed. Findings The procedures developed in this research are compared against experimental data from the Air Force Research Laboratory’s Turbine Research Facility. Practical implications The software presented in this paper is useful as both a design and analysis tool for fluid system and turbomachinery engineers. Originality/value This research presents a novel approach for the simultaneous solution of fluid flow and heat transfer in film-cooled rotating turbine sections. The software developed in this research is validated against experimental results for 2D flow and the methods discussed are extendable to 3D.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:53Z
      DOI: 10.1108/AEAT-10-2014-0159
  • Optimization of turboprop ESFC and NOx emissions for UAV sizing
    • First page: 375
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose In this study, a genuine code was developed for multi-objective optimization of selected parameters of a turboprop UAV for minimum LTO NOx emissions and minimum ESFC at loiter (aerial reconnaissance phase of flight) by using a genetic algorithm. Design/methodology/approach The genuine code developed in this study, first makes computations on preliminary sizing of a UAV and its turboprop engine by analytical method for a given mission profile. Then, in order to minimize NOx emissions or ESFC or both of them, single and multi-objective optimization were done for the selected engine design parameters. Findings In single objective optimization, NOx emissions was reduced by 49% from baseline in given boundaries or constraints of compressor pressure ratio and compressor polytropic efficiency in the first case. In second case, ESFC was improved by 25% from baseline. In multi-objective optimization case, where previous two objectives were considered together, NOx emissions and ESFC decreased by 26.6% and 9.5% from baseline respectively. Practical implications Variation and trend in the NOx emission index and ESFC were investigated with respect to two engine design parameters, namely compressor pressure ratio and compressor polytropic efficiency. Engine designers may take into account the findings of this study to reach a viable solution for the bargain between NOx emission and ESFC. Originality/value UAVs have different flight mission profiles or characteristics compared to manned aircraft. Therefore they are designed in a different philosophy. As number of UAV flights increase in time, fuel burn and LTO NOx emissions worth investigating due to operating costs and environmental reasons. The study includes both sizing and multi-objective optimization of an unmanned air vehicle and its turboprop engine in coupled form; compared to manned aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:54Z
      DOI: 10.1108/AEAT-12-2015-0248
  • Mathematical model of one flexible transport category aircraft
    • First page: 384
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to present a mathematical model of one very flexible transport category airplane, whose structural dynamics was modelled with the strain based formulation. This model can be used to the analysis of couplings between the flight dynamics and structural dynamics. Design/methodology/approach The model was developed with the use of Hamiltonian mechanics and strain based formulation. Nonlinear flight dynamics, nonlinear structural dynamics and inertial couplings are considered. Findings The mathematical model allows the analysis of effects of high structural deformations on airplane flight dynamics. Research limitations/implications The mathematical model has more than sixty degrees of freedom. The computational burden is too high, if compared to the traditional rigid body flight dynamics simulations. Practical implications The mathematical model presented in this work allows a detailed analysis of the couplings between flight dynamics and structural dynamics in very flexible airplanes. The better comprehension of these couplings will contribute to the development of flexible airplanes.. Originality/value This work presents the application of NFNS_s (Nonlinear Flight Dynamics – Nonlinear Structural Dynamics - strain based formulation) methodology to model the flight dynamics of one very flexible transport category airplane. This paper addresses also the way as the analysis of results obtained in nonlinear simulations can be made. Comparisons of the NFNS_s and NFLS (Nonlinear Flight Dynamics – Linear Structural Dynamics) methodologies are presented in this work.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:16:00Z
      DOI: 10.1108/AEAT-12-2013-0230
  • Multiple fault-based FDI and reconfiguration for aircraft engine sensors
    • First page: 397
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose Condition monitoring and health management of an aircraft engine is of importance due to engine’s critical position in aircraft. Missions require uninterrupted and safer conditions during the flight or taxi operations. Hence, the deviations, abnormal situations or failures have to be under control. This paper proposes a cascade connected approach for an aircraft engine fault tolerant control. Design/methodology/approach The cascade connected structure includes a full-order Unknown Input Observer for fault detection and eliminating the unknown disturbance effect on system, a Generalized Observer Scheme for fault isolation and a Boolean Logic Mechanism for decision-making in reconfiguration process respectively. This combination is simulated on a linear turbojet engine model in case of unknown input disturbance and under various sensor failure scenarios. Findings The simulation results show that the suggested FDIR (Fault Detection Isolation Reconfiguration) approach works effectively for multiple sensor failures with various amplitudes. Originality/value Differ from other studies, proposed model is sensitive to unknown input disturbance and failures that have unknown amplitudes. One another notable feature of suggested FDIR approach is adaptability of structure against multiple sensor failures. Here, it is assumed that only a single fault is to be detected and isolated at a time. The simulation results show that the proposed structure can be suggested for linear models especially for physical redundancy-based real-time applications easily, quickly and effectively.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:52Z
      DOI: 10.1108/AEAT-04-2015-0100
  • A novel relative navigation method based on thrust on-line identification
           for tight formation keeping
    • First page: 406
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to present a novel high-precision relative navigation method for tight formation keeping based on thrust on-line identification. Design/methodology/approach Considering thrust acceleration cannot be measured directly, an on-line identification method of thrust acceleration is explored via the estimated acceleration of major space perturbation and the inter-satellite relative states obtained from spaceborne acceleration sensors; then an effective identification model is designed to reconstruct thrust acceleration. Based on the identified thrust acceleration, relative orbit dynamics for tight formation keeping is established. Further, utilizing GPS measurement information, a modified Extended Kalman Filter (EKF) is suggested to obtain the inter-satellite relative position and relative velocity. Findings Compared with the normal EKF and the adaptive robust EKF, the proposed modified EKF has better estimation accuracy in radial and along-track directions due to accurate compensation of thrust acceleration. Meanwhile, high precision relative navigation results depend on high precision acceleration sensors. Finally, simulation studies on a chief-deputy formation flying control system are performed to verify the effectiveness and superiority of the proposed relative navigation algorithm. Originality/value This paper proposes a novel on-line identification method for thrust acceleration, and shows that thrust identification-based modified EKF is more efficient in relative navigation for tight formation keeping.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:16:00Z
      DOI: 10.1108/AEAT-10-2015-0224
  • Nonlinear integrated guidance and control based on adaptive backstepping
    • First page: 415
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to design adaptive nonlinear controller for a nonlinear system of integrated guidance and control. Design/methodology/approach A nonlinear integrated guidance and control approach is applied to a homing, tail controlled air vehicle. Adaptive backstepping controller technique is used to deal with the problem and Lyapanov theory is used in the stability analysis of nonlinear system. A nonlinear model of normal force coefficient is obtained from an existing nonlinear model of lift coefficient which was validated by open loop response. The simulation was performed in the pitch plane to prove the benefits of the proposed scheme however it can be readily extended to all three axes Findings Monte-Carlo simulations indicate that employing nonlinear adaptive backstepping formulation meaningfully improves performance of the system while it ensures stability of nonlinear system. Practical implications The proposed method could be used to obtain better performance of hit to kill accuracy without expense of control effort. Originality/value Nonlinear adaptive backstepping controller for nonlinear aerodynamic air vehicle is designed and guaranteed to be stable which is a novel based approach to the integrated guidance and control. This method makes noticeable performance improvement and it can be used with hit to kill accuracy.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:16:01Z
      DOI: 10.1108/AEAT-12-2014-0209
  • Aerodynamic parameter identification of hypersonic vehicle via
           Pigeon-inspired Optimization
    • First page: 425
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to propose a new approach for aerodynamic parameter identification of hypersonic vehicles, which is based on Pigeon-inspired Optimization (PIO) algorithm, with the objective of overcoming the disadvantages of traditional methods based on gradient such as New Raphson method, especially in noisy environment. Design/methodology/approach The model of hypersonic vehicles and PIO algorithm is established for aerodynamic parameter identification. Using the idea, identification problem will be converted into the optimization problem. Findings A new swarm optimization method, PIO algorithm is applied in this identification process. Experimental results demonstrated the robustness and effectiveness of the proposed method: it can guarantee accurate identification results in noisy environment without fussy calculation of sensitivity. Practical implications The new method developed in this paper can be easily applied to solve complex optimization problems when some traditional method is failed, and can afford the accurate hypersonic parameter for control rate design of hypersonic vehicles. Originality/value In this paper, we converted this identification problem into the optimization problem using the new swarm optimization method-PIO. This new approach is proved to be reasonable through simulation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:55Z
      DOI: 10.1108/AEAT-01-2015-0007
  • Hypersonic dynamic stability of wave riders
    • First page: 434
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to obtain close form expressions for the dynamic stability of conical wave riders with flat surfaces which could be equipped with lifting surfaces on its plain flat surface. Numerical simulation would require very large meshes to resolve flows at subscale level and the experimental evaluations would be equally difficult requiring expensive measurement facilities with challenging procedures to secure such vehicles in confined test sections to obtain satisfactory wind on and wind off oscillations.. Design/methodology/approach The design method uses appropriate pressure fields using small disturbance theory, which in turn is perturbed using the unsteady shock expansion theory to recover suitable expressions for the dynamic stability behaviour. Findings It was observed that the dynamic stability of the standard half-cone type wave riders with flat upper surfaces deteriorates with the axis position measured from the pointed apex reaching a minimum at around x/co = 0.666. The half-cone wave rider with flat upper surfaces is dynamically less stable than a pure cone. Research limitations/implications The method is typically less accurate when the similarity parameter M∞θ ≤ 1 or if the angle of attack is not small. With renewed interest in hypersonic future hypersonically would be designed as fast lifting bodies whose shapes would be close to the configurations of hypersonic wave riders especially if they are designed to operate at upper atmosphere altitudes. Practical implications With renewed interest in hypersonic future hypersonically would be designed as fast lifting bodies whose shapes would be close to the configurations of hypersonic wave riders especially if they are designed to operate at upper atmosphere altitudes. Originality/value The analytic approach outlined in this paper for evaluation of dynamic and static stability derivatives is original drawing from the strengths of the small disturbance theory and shock expansion techniques.. The method is particularly important as there are no reported theoretical, numerical or experimental results in literature.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:54Z
      DOI: 10.1108/AEAT-09-2015-0218
  • Continuous adjoint aerodynamic optimization design for multi-stage gas
           turbine with cooling air
    • First page: 444
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose Conduct the optimization of the multi-stage gas turbine with the effect of the cooling air injection based on the ajoint method. Design/methodology/approach Continuous adjoint method is combined with the S2 surface code. Findings The optimization of the stagger angles, stacking lines and the passage can improve the attack angles and restrain the development of the boundary reducing the secondary flow loss caused by the cooling air injection. Practical implications The aerodynamic performance of the gas turbine can be improved via the optimization of blade and passage based on the adjoint method. Originality/value The results of the first study on the adjoint method applied to the S2 surface through flow calculation including the cooling air effect are presented.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:58Z
      DOI: 10.1108/AEAT-05-2014-0058
  • Multiple-model adaptive estimator for spacecraft attitude sensor
    • First page: 457
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose This paper presents a multiple-model adaptive estimator (MMAE) to calibrate the star sensor low frequency error (LFE). The star sensor LFE, which is caused primarily by the periodic thermal distortion, has a great impact on spacecraft attitude determination accuracy. Design/methodology/approach The unfavorable effect of the LFE can be partly eliminated by using the calibration algorithm based on the augmented Kalman filter (AKF). However, the AKF may be worse than the traditional Kalman filter (KF) in the absence of the LFE. In order to cope with this problem, the MMAE is applied first time for combining the AKF and the KF in the spacecraft attitude determination system, such that satisfactory performance can be achieved in different operating scenarios. Findings The convergence of the presented MMAE is demonstrated through a formal derivation. A novel method is proposed to tune the MMAE design parameter, such that the convergence rate of the estimator is increased. It is shown via numerical studies that the presented algorithm outperforms the AKF and the KF. Practical implications The calibration algorithm is applicable for spacecraft attitude determination. Originality/value An effective star sensor LFE calibration algorithm based on the MMAE is developed. In addition, a novel method is proposed to increase convergence rate of the estimator.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:16:02Z
      DOI: 10.1108/AEAT-02-2015-0029
  • Fault tolerant control against actuator faults based on enhanced PID
           controller for a quadrotor
    • First page: 468
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to present fault tolerant control of a quadrotor based on the enhanced PID structure in the presence of one or more actuator faults. Design/methodology/approach Mathematical model of the quadrotor is derived by parameter identification of the system for the simulation of the UAV dynamics and flight control in MATLAB/Simulink. An improved PID structure is used to provide the stability of the nonlinear quadcopter system both for attitude and path control of the system. Not only for the healthy system but also for the faulty conditions simulation results of the whole system including motor dynamics are presented in the study. Findings In this study, actuator faults are considered to show that a robust controller design handles the loss of effectiveness in motors up to some extent. For the loss of control effectiveness of 20% in first and third motors, psi state follows the reference with steady state error, it does not go unstable. Motor 1 and motor 3 respond to given motor fault quickly. When it comes to one actuator fault, steady state errors remain in some states but the system does not become unstable. Originality/value In this paper, an enhanced PID controller is proposed to keep the quadrotor stable in case of actuator faults. Proposed method demonstrates the effectiveness of the control system against motor faults.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:57Z
      DOI: 10.1108/AEAT-04-2015-0096
  • Thermoelastic vibration and maneuver control of smart satellites
    • First page: 477
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to analyze and control the thermally induced vibration of orbiting smart satellite panels, which have been modeled as functionally graded material (FGM) beams. Design/methodology/approach It is assumed that the satellite moves in a circular orbit and has pitch angle rotation maneuver. Rapid temperature changes at day-night transitions in orbit, generate time dependent bending moments that induce vibrations in the appendages. So the heat radiation effects on the appendages should be considered. The thermally induced vibrations of the appendages and the nonlinear heat transfer equation are coupled and should be solved simultaneously. So, the governing equations of the motion are nonlinear and very complicated ones. A robust passivity-based controller is proposed to control the satellite maneuver and appendages vibrations, using piezoelectric sensors/actuators. Findings After the simulation, the effects of the heat radiation, piezoelectric actuators, and piezoelectric locations on the response of the system are studied. The results of dynamic response and thermal analysis show that the radiation thermal effects are coupled with structure dynamic. These effects induce the vibration. Also, the effectiveness and the capability of the controller are analyzed. The results of the simulation show that the robust passivity-based control can ensure that the satellite rotates in the desired trajectory and vibrations of the appendages are damped. It demonstrates that the proposed control scheme is feasible and effective. Originality/value The paper is the basis of deriving the governing equations, thermal analysis and a robust control system design of a smart satellite with FGM panels.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:56Z
      DOI: 10.1108/AEAT-11-2015-0241
  • Aerodynamic robust optimization of flying wing aircraft based on interval
    • First page: 491
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to propose a robust optimization strategy to deal with the aerodynamic optimization issue, which does not need a large sum of information on the uncertainty of input parameters. Design/methodology/approach Interval numbers were adopted to describe the uncertain input, which only requires bounds, and does not necessarily need probability distributions. Based on the method, model outputs were also regarded as intervals. In order to identify a better solution, an order relation was used to rank interval numbers. Findings Based on intervals analysis method, the uncertain optimization problem was transformed into nested optimization. The outer optimization was used to optimise the design vector, and inner optimization was used to compute the interval of model outputs. A flying wing aircraft was used as a basis for uncertainty optimization through the suggested optimization strategy, and optimization results demonstrated the validity of the method. Originality/value In aircraft conceptual design, the uncertain information of design parameters are often insufficient. Interval number programming method used for uncertainty analysis is effective for aerodynamic robust optimization for aircraft conceptual design.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:57Z
      DOI: 10.1108/AEAT-09-2016-0145
  • Analysis on the attitude dynamics of a PhoneSat during deployment
    • First page: 498
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 3, May 2017.
      Purpose The purpose of this paper is to establish the dynamics model of a Z-folded PhoneSat considering hinge friction and to investigate the influence of disturbances, such as friction, stiffness asymmetry, deployment asynchronicity and initial disturbance angular velocity on the attitude of PhoneSat during and after deployment. Design/methodology/approach For the Z-folded PhoneSat, the dynamics model considering hinge friction is established and the dynamics simulation is carried out. The effects of friction, stiffness asymmetry, deployment asynchronicity and initial disturbance angular velocity on the attitude motion of the PhoneSat are studied and the attitude motion regularities of the PhoneSat considering the disturbance factors mentioned above are discussed. Findings Friction has a main contribution to reducing the oscillation of attitude motion and damping out the residual oscillation, ultimately decreasing the deployment time. An increasing length of deployment time is required with the increasing stiffness asymmetry and time difference of asynchronous deployment, which also have slight disturbances on the attitude angle and angular velocity of PhoneSat after the deployment. The initial disturbance angular velocity in the direction of deployment would be proportionally weakened after the deployment, while initial disturbance angular velocity in other direction induces angular velocities of other axes, which dramatically enhances the complexity of attitude control. Originality/value The paper is a useful reference for engineering design of small satellites attitude control system.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-28T12:15:59Z
      DOI: 10.1108/AEAT-01-2016-0013
  • Model predictive control of an unmanned aerial vehicle
    • Pages: 193 - 202
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 193-202, March 2017.
      Purpose The purpose of this paper is to design a controller for the unmanned aerial vehicle (UAV). Design/methodology/approach In this study, the constrained multivariable multiple-input and multiple-output (MIMO) model predictive controller (MPC) has been designed to control all outputs by manipulating inputs. The aim of the autopilot of UAV is to keep the UAV around trim condition and to track airspeed commands. Findings The purpose of using this control method is to decrease the control effort under the certain constraints and deal with interactions between each output and input while tracking airspeed commands. Originality/value By using constraint, multivariable (four inputs and seven outputs) MPC unlike the relevant literature in this field, the UAV tracked airspeed commands with minimum control effort dealing with interactions between each input and output under disturbances such as wind.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:02Z
      DOI: 10.1108/AEAT-03-2015-0074
  • Simulation-based steady-state aero-thermal model for small-scale turboprop
    • Pages: 203 - 210
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 203-210, March 2017.
      Purpose An aircraft engine control system consists of a large scale of control parameters and variables because of the complex structure of aero-engine. Monitoring and adjusting control variables and parameters such as detecting, isolating and reconfiguring the system faults/failures depend on the controller design. Developing a robust controller is based on an accurate mathematical model. Design/methodology/approach In this study, a small-scale turboprop engine is modeled. Simulation is carried out on MATLAB/Simulink for design and off-design operating conditions. Both steady-state and transient conditions (from idle to maximum thrust levels) are tested. The performance parameters of compressor and turbine components are predicted via trained Neuro-Fuzzy model (ANFIS) based on component maps. Temperature, rotational speed, mass flow, pressure and other parameters are generated by using thermodynamic formulas and conservation laws. Considering these calculated values, error calculations are made and compared with the cycle data of the engine at the related simulation conditions. Findings Simulation results show that the designed engine model’s simulation values have acceptable accuracy for both design and off-design conditions from idle to maximum power operating envelope considering cycle data. The designed engine model can be adapted to other types of gas turbine engines. Originality/value Different from other literature studies, in this work, a small-scale turboprop engine is modeled. Furthermore, for performance prediction of compressor and turbine components, ANFIS structure is applied.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:41:54Z
      DOI: 10.1108/AEAT-02-2015-0062
  • Artificial neural networks to predict aerodynamic coefficients of
           transport airplanes
    • Pages: 211 - 230
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 211-230, March 2017.
      Purpose Multidisciplinary design frameworks elaborated for aeronautical applications require considerable computational power that grows enormously with the utilization of higher fidelity tools to model aeronautical disciplines like aerodynamics, loads, flight dynamics, performance, structural analysis and others. Surrogate models are a good alternative to address properly and elegantly this issue. With regard to this issue, the purpose of this paper is the design and application of an artificial neural network to predict aerodynamic coefficients of transport airplanes. The neural network must be fed with calculations from computational fluid dynamic codes. The artificial neural network system that was then developed can predict lift and drag coefficients for wing-fuselage configurations with high accuracy. The input parameters for the neural network are the wing planform, airfoil geometry and flight condition. An aerodynamic database consisting of approximately 100,000 cases calculated with a full-potential code with computation of viscous effects was used for the neural network training, which is carried out with the back-propagation algorithm, the scaled gradient algorithm and the Nguyen–Wridow weight initialization. Networks with different numbers of neurons were evaluated to minimize the regression error. The neural network featuring the lowest regression error is able to reduce the computation time of the aerodynamic coefficients 4,000 times when compared with the computing time required by the full potential code. Regarding the drag coefficient, the average error of the neural network is of five drag counts only. The computation of the gradients of the neural network outputs in a scalable manner is possible by an adaptation of back-propagation algorithm. This enabled its use in an adjoint method, elaborated by the authors and used for an airplane optimization task. The results from that optimization were compared with similar tasks performed by calling the full potential code in another optimization application. The resulting geometry obtained with the aerodynamic coefficient predicted by the neural network is practically the same of that designed directly by the call of the full potential code. Design/methodology/approach The aerodynamic database required for the neural network training was generated with a full-potential multiblock-structured code. The training process used the back-propagation algorithm, the scaled-conjugate gradient algorithm and the Nguyen–Wridow weight initialization. Networks with different numbers of neurons were evaluated to minimize the regression error. Findings A suitable and efficient methodology to model aerodynamic coefficients based on artificial neural networks was obtained. This work also suggests appropriate sizes of artificial neural networks for this specific application. We demonstrated that these metamodels for airplane optimization tasks can be used without loss of fidelity and with great accuracy, as their local minima might be relatively close to the minima of the original design space defined by the call of computational fluid dynamics codes. Research limitations/implications The present work demonstrated the ability of a metamodel with artificial neural networks to capture the physics of transonic and subsonic flow over a wing-fuselage combination. The formulation that was used was the full potential equation. However, the present methodology can be extended to model more complex formulations such as the Euler and Navier–Stokes ones. Practical implications Optimum networks reduced the computation time for aerodynamic coefficient calculations by 4,000 times when compared with the full-potential code. The average absolute errors obtained were of 0.004 and 0.0005 for lift and drag coefficient prediction, respectively. Airplane configurations can be evaluated more quickly. Social implications If multidisciplinary optimization tasks for airplane design become more efficient, this means that more efficient airplanes (for instance less polluting airplanes) can be designed. This leads to a more sustainable aviation. Originality/value This research started in 2005 with a master thesis. It was steadily improved with more efficient artificial neural networks able to handle more complex airplane geometries. There is a single work using similar techniques found in a conference paper published in 2007. However, that paper focused on the application, i.e. providing very few details of the methodology to model aerodynamic coefficients.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:07Z
      DOI: 10.1108/AEAT-05-2014-0069
  • First principle analysis of Coandă Micro Air Vehicle aerodynamic
           forces for preliminary sizing
    • Pages: 231 - 245
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 231-245, March 2017.
      Purpose The purpose of this paper is to reformulate the governing equations incorporating major variables and parameters for the design a Micro Air Vehicle (MAV), to meet the desired mission and design requirements. Design/methodology/approach Mathematical models for various spherical and cylindrical Coandă MAV configurations were rederived from first principles, and the performance measures were defined. To verify the theoretical prediction to a certain extent, a computational fluid dynamic (CFD) simulation for a Coandă MAV generic models was performed. Findings The major variables and parameters of Coandă MAV have been formulated into practical guidelines, which relate the lift (or thrust) produced for certain input variables, particularly the Coandă MAV jet momentum coefficient. The influences of the geometrical parameters are elaborated. Research limitations/implications The present analysis on Coandă jet-configured MAV is focused on the lift generation due to the Coandă jet effect through a meticulous analysis. The effects of viscosity, the Coandă jet thickness, the radius of curvature of the surface and the stability of Coandă jet are not considered and will be the subject of the following work. Practical implications The results obtained can be used for sizing in the preliminary design of Coandă MAVs. Originality/value Physical and mathematical models were developed which can describe the physical phenomena of the flow field near the Coandă MAV surfaces influenced by Coandă jet sheets and for obtaining a relationship between relevant variables and parameters to the lift of practical interest.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:42Z
      DOI: 10.1108/AEAT-03-2015-0080
  • Satellite formation keeping via chaotic artificial bee colony
    • Pages: 246 - 256
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 246-256, March 2017.
      Purpose Keeping satellite position within close tolerances is key for the utilization of satellite formations for space missions. The presence of perturbation forces makes control inevitable if such mission objective is to be realised. Various approaches have been used to obtain feedback controller parameters for satellites in a formation; this paper aims to approach the problem of estimating the optimal feedback parameter for a leader–follower pair of satellites in a small eccentric orbit using nature-based search algorithms. Design/methodology/approach The chaotic artificial bee colony algorithm is a variant of the basic artificial bee colony algorithm. The algorithm mimics the behaviour of bees in their search for food sources. This paper uses the algorithm in optimizing feedback controller parameters for a satellite formation control problem. The problem is formulated to optimize the controller parameters while minimizing a fuel- and state-dependent cost function. The dynamical model of the satellite is based on Gauss variational equations with J2 perturbation. Detailed implementation of the procedure is provided, and experimental results of using the algorithm are also presented to show feasibility of the method. Findings The experimental results indicate the feasibility of this approach, clearly showing the effective control of the transients that arise because of J2 perturbation. Originality/value This paper applied a swarm intelligence approach to the problem of estimating optimal feedback control parameter for a pair of satellites in a formation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:01Z
      DOI: 10.1108/AEAT-02-2014-0019
  • Performance optimization of aircraft in frontline squadrons
    • Pages: 257 - 261
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 257-261, March 2017.
      Purpose The purpose of this paper is to indicate the increase in operational efficiency of the aircraft in frontline squadrons by implementation of effective condition-based predictive maintenance (CBPM) philosophy using data link data transfer in real-time operational environment. Design/methodology/approach Real-time data transfer from the aircraft to the maintenance hub using data link is used as the key feature behind achieving an increase in the operational efficiency. Findings Considerable amount of increase in the operational efficiency and decrease in down time could be achieved on utilization of real-time aircraft parameters in the decision-making process of CBPM. Practical implications Incorporation of this methodology in frontline operating environment would result in achieving maintenance optimization, thereby reducing the downtime of the aircraft. Social implications Helps in achieving performance optimization with comparatively reduced downtime. Originality/value Up to 20 per cent of reduction in downtime could be achieved if real-time data transfer using data link is used parameters in the decision-making process of CBPM.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:11Z
      DOI: 10.1108/AEAT-07-2015-0178
  • Aerodynamic wing shape optimization based on the computational design
           framework CEASIOM
    • Pages: 262 - 273
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 262-273, March 2017.
      Purpose A collaborative design environment is needed for multidisciplinary design optimization (MDO) process, based on all the modules those for different design/analysis disciplines, and a systematic coupling should be made to carry out aerodynamic shape optimization (ASO), which is an important part of MDO. Design/methodology/approach Computerized environment for aircraft synthesis and integrated optimization methods (CEASIOM)-ASO is developed based on loosely coupling all the existing modules of CEASIOM by MATLAB scripts. The optimization problem is broken down into small sub-problems, which is called “sequential design approach”, allowing the engineer in the loop. Findings CEASIOM-ASO shows excellent design abilities on the test case of designing a blended wing body flying in transonic speed, with around 45 per cent drag reduction and all the constraints fulfilled. Practical implications
      Authors built a complete and systematic technique for aerodynamic wing shape optimization based on the existing computational design framework CEASIOM, from geometry parametrization, meshing to optimization. Originality/value CEASIOM-ASO provides an optimization technique with loosely coupled modules in CEASIOM design framework, allowing engineer in the loop to follow the “sequential approach” of the design, which is less “myopic” than sticking to gradient-based optimization for the whole process. Meanwhile, it is easily to be parallelized.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:20Z
      DOI: 10.1108/AEAT-04-2015-0098
  • Effects of material property variation in composite panels
    • Pages: 274 - 279
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 274-279, March 2017.
      Purpose The purpose of this paper is to determine which of the ten material properties of the Hashin progressive damage model significantly affect the maximum load-carrying ability of center-notched carbon fiber panels under in-plane tension and out-of-plane bending. Design/methodology/approach The approach used is to calculate the maximum load using a finite element model for a range of material property values as specified by a fraction factorial design. The finite element model used has been experimentally validated in prior work. Findings Results showed that for the laminates considered, at most three and as few as one of the ten Hashin material properties significantly affected the magnitude of the maximum load. Practical implications While the results of this paper only specifically apply to the laminates included in the study, the results suggest that, in general, only a small number of the Hashin material properties affect laminate load-carrying ability. Originality/value Knowing which properties are significant is of value in selecting materials to optimize performance and also in determining which properties need to be known to a high accuracy.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:49Z
      DOI: 10.1108/AEAT-10-2014-0177
  • Sigma overbound for aircraft landing in presence of day-to-day multipath
    • Pages: 280 - 289
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 280-289, March 2017.
      Purpose The purpose of this work is to obtain an overbounded broadcast sigma from actual (non-Gaussian) correction error distribution under the stringent navigation integrity requirements for aircraft precision approach and landing. Design/methodology/approach Approach is statistically to overbound satellite pseudorange correction error distribution with the use of numerical solution of Fisher-Z transformation. Inflation factors for overbounding broadcast sigma are extracted from Fisher-Z transformation based on measured correlation and counted independent identically distributed (iid) sample sizes of true empirical data. Findings New overbounded broadcast sigma values for eight long-pass satellites were obtained based on measured actual empirical data and ensured integrity risk at 10−8 probability level. Proposed methodology successfully overbounds ground reflection multipath-type systematic and temporal errors sources. Originality/value This paper introduced a new method of accounting for ground reflection multipath for local area augmentation system/ground-based augmentation system navigation integrity. The method is also applicable to statistically overbound any other serially correlated temporal variation in measured data if both correlation values and finite iid sample sizes are known.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:33Z
      DOI: 10.1108/AEAT-03-2015-0081
  • LQR/LQG attitude stabilization of an agile microsatellite with CMG
    • Pages: 290 - 296
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 290-296, March 2017.
      Purpose The purpose of the paper is to design a new attitude stabilization system for a microsatellite based on single gimbal control moment gyro (SGCMG) in which the gimbal rates are selected as controller parameters. Design/methodology/approach In the stability mode, linear quadratic regulator (LQR) and linear quadratic Gaussian (LQG) control strategies are presented with the gimbal rates as a controller parameters. Instead of developing a control torque to solve the attitude problem, the attitude controller is developed in terms of the control moment gyroscope gimbal angular velocities. Attitude control torques are generated by means of a four SGCMG pyramid cluster. Findings Numerical simulation results are provided to show the efficiency of the proposed controllers. Simulation results show that this method could stabilize satellite from initial condition with large angles and with more accuracy in comparison with feedback quaternion and proportional-integral-derivative controllers. These results show the effect of filtering the noisy signal in the LQG controller. LQG in comparison to LQR is more realistic. Practical implications The LQR method is more appropriate for the systems that have project models reasonably exact and ideal sensors/actuators. LQG is more realistic, and it can be used when not all of the states are available or when the system presents noises. LQR/LQG controller can be used in the stabilization mode of satellite attitude control. Originality/value The originality of this paper is designing a new attitude stabilization system for an agile microsatellite using LQR and LQG controllers.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:17Z
      DOI: 10.1108/AEAT-07-2014-0102
  • Fast time-delay measurement for integrated pulsar pulse profiles
    • Pages: 297 - 303
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 297-303, March 2017.
      Purpose The authors proposed a new method of fast time delay measurement for integrated pulsar pulse profiles in X-ray pulsar-based navigation (XNAV). As a basic observation of exact orientation in XNAV, time of arrival (TOA) can be obtained by time delay measurement of integrated pulsar pulse profiles. Therefore, the main purpose of the paper is to establish a method with fast time delay measurement on the condition of limited spacecraft’s computing resources. Design/methodology/approach Given that the third-order cumulants can suppress the Gaussian noise and reduce calculation to achieve precise and fast positioning in XNAV, the proposed method sets the third-order auto-cumulants of standard pulse profile, the third-order cross-cumulants of the standard and the observed pulse profile as basic variables and uses the cross-correlation function of these two variables to estimate the time delay of integrated pulsar pulse profiles. Findings The proposed method is simple, fast and has high accuracy in time delay measurement for integrated pulsar pulse profiles. The result shows that compared to the bispectrum algorithm, the method improves the precision of the time delay measurement and reduced the computation time significantly as well. Practical implications To improve the performance of time delay estimation in XNAV systems, the authors proposed a novel method for XNAV to achieve precise and fast positioning. Originality/value Compared to the bispectrum algorithm, the proposed method can improve the speed and precision of the TOA’s calculation effectively by using the cross-correlation function of integrated pulsar pulse profile’s third-order cumulants instead of Fourier transform in bispectrum algorithm.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:23Z
      DOI: 10.1108/AEAT-02-2015-0030
  • Cycle optimization of the staged combustion rocket engines
    • Pages: 304 - 313
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 304-313, March 2017.
      Purpose The purpose of this paper is to introduce new and modified “staged combustion” cycles in the form of engineering algorithm as a possible propulsion contender for future aerospace vehicle to achieve the highest possible “total impulse” to “mass” of propulsion system. Design/methodology/approach In this regard, the mathematical cycle model is formed to calculate the engine’s parameters. In addition, flow conditions (pressure, temperature, flow rate, etc). in the chamber, nozzle and turbopump are assessed based on the results of turbo machinery power balance and initial data such as thrust, propellant mixture ratio and specifications. The developed code has been written in the modern, object-oriented C++ programming language. Findings The results of the developed code are compared with the Russian RD180 engine which demonstrates the superiority and capability of new “thermodynamic diagrams”. Research limitations/implications This algorithm is under constraint to control the critical variation of combustion pressure, turbine rpm, pump cavitation and turbine temperature. It is imperative to emphasize that this paper is limited to “oxidizer-rich staged combustion” engines with “single pre-burner”. Originality/value This study sheds light on using fuel booster turbopump and the second-stage fuel pump to moderate the effect of cavitation on pumps which reduces tank pressure and, as a consequence, decreases the propulsion system weight.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:44Z
      DOI: 10.1108/AEAT-12-2013-0229
  • A missile guidance law tolerant to unestimated evasive maneuvers
    • Pages: 314 - 319
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 314-319, March 2017.
      Purpose This note aims to introduce a terminal guidance law that is able to compensate for evasive target maneuvers without estimating their acceleration. Design/methodology/approach The new guidance law is derived in the framework of linear-quadratic optimal control to ensure interception with minimum energy even in the presence of a target maneuver. Findings An explicit closed-form expression for the missile acceleration command is provided, which turns out to be a non-trivial extension of proportional navigation guidance. Simulation results against evasive maneuvers of various intensities are provided to compare the new law to classical ones and thus show the benefits of the proposed approach. Originality/value The proposed guidance law was not reported so far in the literature and provides a simple way to deal with evasive maneuvers.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:46Z
      DOI: 10.1108/AEAT-04-2014-0044
  • A nonlinear control approach for a hypersonic vehicle
    • Pages: 320 - 329
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 320-329, March 2017.
      Purpose The purpose of the paper is to design a robust control system for a generic hypersonic vehicle which includes dynamic nonlinear, open loop unstable and parametric uncertainties. Design/methodology/approach For a complex longitudinal model of a generic hypersonic vehicle which includes dynamic nonlinear, open loop unstable and parametric uncertainties, a nonlinear dynamic inverse (NDI) approach combined with proportional differential (PD) control is used to design a strong robust control system to deal with the sensitivity to changes of atmosphere condition. In this way, a simple genetic algorithm is used to search a group of parameters of the control system to satisfy the specific performance indices. Then parametric uncertainties are considered to verify the robustness of the control system. Findings The PD hypersonic vehicle control system using NDI approach can satisfy the specific flight performance. And it has strong robustness under the parametric uncertainties. Originality/value The paper fulfills a complete process of the nonlinear control system design for a generic hypersonic vehicle. And, the simulation results show the efficiency and robustness of the control system.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:05Z
      DOI: 10.1108/AEAT-06-2013-0119
  • Improvement of aircraft crashworthy performance using inversion failure
           strut system
    • Pages: 330 - 337
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 330-337, March 2017.
      Purpose The purpose of this paper is to improve the crashworthiness of aircraft by using the strut system as an energy absorption device without redesigning other components. Design/methodology/approach The novel strut system consists of metal stepped thin-walled tubes and articulated connecting hinges. The strut is suffering axial load during impact process for rotating of hinges, and the metal stepped tube has an inversion failure behaviour. Findings The metal stepped tube has lower initial impact load and more stable failure behaviour. The geometrical factors have a great influence on the impact load and energy absorption efficiency. The best length ratio between upper and lower sections is about 2:1 and 1:1 for the metal stepped circular and square tubes, respectively. Practical implications The metal stepped tube with inversion mechanism is suitable for aircraft strut system to improve crashworthiness performance. Originality/value A new strut system is provided using metal inversion failure stepped tubes and articulated connecting hinges to improve crash worthiness of aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:47Z
      DOI: 10.1108/AEAT-09-2015-0205
  • Imperfect maintenance model for estimating aircraft fleet availability
    • Pages: 338 - 346
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 338-346, March 2017.
      Purpose The purpose of this paper is to define conditions under which improved availability of fleet of G-4 jet trainers is obtained, and optimization of intermediate-level maintenance through imperfect maintenance model application. This research has been conducted based on available knowledge, and experience gained by performing intermediate-level maintenance of Serbian Air Force aircrafts. Design/methodology/approach Analysis of the data collected from daily maintenance reports, and the analysis of maintenance technology and organization, was performed. Based on research results, a reliability study was performed. Implementation of imperfect maintenance with its models of maintenance policies (especially a quasi-renewal process and its treating of reliability and optimal maintenance) was proposed to define new maintenance parameters so that the greater level of availability could be achieved. Findings The proposed methodology can potentially be applied as a simple tool to estimate the present maintenance parameters and to quickly point out some deficiencies in the analyzed maintenance organization. Validation of this process was done by conducting a reliability case study of G-4 jet trainer fleet, and numerical computations of optimal maintenance policy. Research limitations/implications The methodology of the availability estimation when reliability parameters were not tracked by the maintenance organization, and optimization of intermediate-level maintenance, has so far been applied on G-4 jet trainers. Moreover, it can be potentially applied to other aircraft types. Originality/value Availability estimation and proposed optimization of intermediate maintenance is based on a survey of data for three years of aircraft fleet maintenance. It enables greater operational readiness (due to a military rationale) with possible cost reduction as a consequence but not as a goal.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:52Z
      DOI: 10.1108/AEAT-10-2015-0221
  • Application of harmonic balance method to aerodynamic characteristics
           prediction of spinning vehicle
    • Pages: 347 - 357
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 347-357, March 2017.
      Purpose The purpose of this paper is to examine the ability of the harmonic balance method for predicting the aerodynamic characteristics of rigid finned spinning vehicle. Design/methodology/approach The aerodynamic characteristics of a rigid four-finned spinning vehicle at Mach number 2.5 and angle of attack of 20 degrees are simulated using the harmonic balance method and the unsteady time-accurate approach based on the dual-time method. The numerical results are analyzed, and the computed aerodynamic coefficients of the harmonic balance method are compared with those of the dual-time method. The influence of the number of harmonics is presented. The computed Magnus force and moment coefficients are compared with the experimental data. The flow fields at different roll angles are presented. The computational efficiency of harmonic balance method is analyzed. Findings The results show that the aerodynamic coefficients of spinning vehicle could be predicted by the harmonic balance method with reasonable accuracy compared with the dual-time method. For the harmonic balance method, the accuracy of the computed leeward side flow is relatively poor compared with that of the computed windward side flow. Meanwhile, the computational efficiency is influenced by initial guess and the intensity of unsteady effect. Practical implications The harmonic balance method could be used for the aerodynamic prediction of spinning vehicle, which may improve the efficiency of vehicle design. Originality/value This paper presents the results of the harmonic balance method for simulating the aerodynamic characteristics of finned spinning vehicle. The accuracy and efficiency of the method are analyzed.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:50Z
      DOI: 10.1108/AEAT-09-2015-0202
  • Experimental facilities for modal testing
    • Pages: 358 - 363
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 89, Issue 2, Page 358-363, March 2017.
      Purpose The experimental modal analysis requires good knowledge of various engineering fields, such as mechanical vibrations, transducers used in vibration measurement, transducers and system calibration methods, data acquisition systems, digital signal processing and system identification. Test facilities constitute a key factor for improving the quality of the estimated modal model. This paper aims to describe the experimental facilities at the Institute of Aeronautics and Space (IAE) Modal Testing Laboratory in terms of associated instrumentation and data acquisition system, metrological aspects and computational resources. The discussion is completed with a practical application showing a ground vibration testing (GVT) of an unmanned aerial vehicle (UAV). Design/methodology/approach The experimental facilities were evaluated in a typical GVT, using three shakers in both vertical and horizontal excitations and 88 response measurement points. The global excitation method was used to excite all desired modes. The reliability of the experimental modal model was validated by an auto modal assurance criterion matrix for the measured modes of the structure. Findings The experimental facilities were successfully used for validating the dynamical characteristics of the UAV under testing. Originality/value The modal test facilities of the Modal Testing Laboratory at the IAE, the main research center of the Brazilian Air Force, are described in this paper.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2017-03-07T08:42:30Z
      DOI: 10.1108/AEAT-04-2015-0099
School of Mathematical and Computer Sciences
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