for Journals by Title or ISSN
for Articles by Keywords
help
Similar Journals
Journal Cover
Aircraft Engineering and Aerospace Technology
Number of Followers: 234  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 0002-2667 - ISSN (Online) 1748-8842
This journal is no longer being updated because:
    RSS feed has been removed by the publisher
  • Reliability analysis for a hypersonic aircraft’s wing spar
    • Pages: 549 - 557
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 549-557, April 2019.
      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
       
  • Comparative analysis of lunar capture braking method based on particle
           swarm optimization
    • Pages: 558 - 566
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 558-566, April 2019.
      Purpose The purpose of this paper is to verify the feasibility of lunar capture braking through three methods based on particle swarm optimization (PSO) and compare the advantages and disadvantages of the three strategies by analyzing the results of the simulation. Design/methodology/approach The paper proposes three methods to verify capture braking based on PSO. The constraints of the method are the final lunar orbit eccentricity and the height of the final orbit around the Moon. At the same time, fuel consumption is used as a performance indicator. Then, the PSO algorithm is used to optimize the track of the capture process and simulate the entire capture braking process. Findings The three proposed braking strategies under the framework of PSO algorithm are very effective for solving the problem of lunar capture braking. The simulation results show that the orbit in the opposite direction of the trajectory has the most serious attenuation at perilune, and it should consume the least amount of fuel in theoretical analysis. The methods based on the fixed thrust direction braking and thrust uniform rotation braking can better ensure the final perilune control accuracy and fuel consumption. As for practice, the fixed thrust direction braking method is better realized among the three strategies. Research limitations/implications The process of lunar capture is a complicated process, involving effective coordination between multiple subsystems. In this article, the main focus is on the correctness of the algorithm, and a simplified dynamic model is adopted. At the same time, because the capture time is short, the lunar curvature can be omitted. Furthermore, to better compare the pros and cons of different braking modes, some influence factors and perturbative forces are not considered, such as the Earth’s flatness, light pressure and system noise and errors. Practical implications This paper presents three braking strategies that can satisfy all the constraints well and optimize the fuel consumption to make the lunar capture more effective. The results of comparative analysis demonstrate that the three strategies have their own superiority, and the fixed thrust direction braking is beneficial to engineering realization and has certain engineering practicability, which can also provide reference for lunar exploration orbit design. Originality/value The proposed capture braking strategies based on PSO enable effective capture of the lunar module. During the lunar exploration, the capture braking phase determines whether the mission will be successful or not, and it is essential to control fuel consumption on the premise of accuracy. The three methods in this paper can be used to provide a study reference for the optimization of lunar capture braking.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-04-05T08:24:36Z
      DOI: 10.1108/AEAT-09-2018-0250
       
  • Prediction of the development cost of general aviation aircraft
    • Pages: 567 - 574
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 567-574, April 2019.
      Purpose The purpose of this paper is to develop a flexible design-oriented development cost method for general aviation aircraft based on small sample and poor information. Design/methodology/approach To predict the development cost of general aviation aircraft accurately, the methodology is based on the collected cost data and actual technical, and then the cost prediction relationships derived from an exhaustive statistical and filtered from regression analysis are incorporated. A series of regression equations with high regression coefficient are yielded after the cost driving factors of the development cost are fixed. Next, several sets of equations with high regression coefficient are selected for final integration. It is a flexible method that can be used efficiently to predict the cost of general aviation aircraft. Findings The development of general aviation aircraft has relatively a late start and no cost prediction model has been suitable for small sample, the proposed method is expected and is rather desirable. Practical implications By comparing the approach with the ordinary regression model and back propagation (BP) neural network, the scheme in this work is more efficient and convenient. Originality/value The results obtained in this paper show that the proposed method not only has a certain degree of versatility, but also can provide a preliminary prediction of the development cost of general aviation aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-19T11:40:47Z
      DOI: 10.1108/AEAT-09-2018-0248
       
  • Flow characteristics of two-dimensional synthetic jets under diaphragm
           resonance excitation
    • Pages: 575 - 581
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 575-581, April 2019.
      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
       
  • Spin flight mode identification with OEEMD algorithm
    • Pages: 582 - 600
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 582-600, April 2019.
      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
       
  • Deflection analysis of the airship structure based on the tapered
           inflatable beam
    • Pages: 601 - 606
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 601-606, April 2019.
      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
       
  • Definition and representation of stiffened shell structures in the context
           of an integrated development process
    • Pages: 607 - 619
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 607-619, April 2019.
      Purpose A novel development process aims at finding solutions for lightweight stiffened shell structures and their efficient production. To respect the strong interdependency of structural design and production planning, particularly observed for composite structures, it is of high interest to start considering production effects in early development phases. This integrated approach requires an integrated representation of structure and production. The purpose of this study is to investigate the scope of relevant data and to find a structure for its representation. Design/methodology/approach The development task is analyzed and a system of so-called solution dimensions is presented, which covers all important aspects of stiffened shell structures and their production. An integrated product data model is developed to cover all of the solution dimensions. Findings The product data model consists of five coherent partial models. It is explained how these models are defined and how they are connected to each other. An academic example of an aircraft fuselage panel is used to demonstrate the definition process. It is shown how even complex structural concepts are defined systematically. Practical implications It is explained how this integrated product data model is used in a software project for the development of aircraft fuselage structures. Originality/value The presented approach for the definition and representation of stiffened shell structures enables the developer, e.g. of aircraft fuselage, to respect the crucial criterion of manufacturability from early development phases on. Further, new design approaches, e.g. as inspired by topology optimization, can be considered.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-04T10:58:59Z
      DOI: 10.1108/AEAT-07-2018-0205
       
  • RBF-based mesh morphing approach to perform icing simulations in the
           aviation sector
    • Pages: 620 - 633
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 620-633, April 2019.
      Purpose Numerical simulation of icing has become a standard. Once the iced shape is known, however, the analyst needs to update the computational fluid dynamics (CFD) grid. This paper aims to propose a method to update the numerical mesh with ice profiles. Design/methodology/approach The present paper concerns a novel and fast radial basis functions (RBF) mesh morphing technique to efficiently and accurately perform ice accretion simulations on industrial models in the aviation sector. This method can be linked to CFD analyses to dynamically reproduce the ice growth. Findings To verify the consistency of the proposed approach, one of the most challenging ice profile selected in the LEWICE manual was replicated and simulated through CFD. To showcase the effectiveness of this technique, predefined ice profiles were automatically applied on two-dimensional (2D) and three-dimensional (3D) cases using both commercial and open-source CFD solvers. Practical implications If ice accreted shapes are available, the meshless characteristic of the proposed approach enables its coupling with the CFD solvers currently supported by the RBF4AERO platform including OpenFOAM, SU2 and ANSYS Fluent. The advantages provided by the use of RBF are the high performance and reliability, due to the fast application of mesh smoothing and the accuracy in controlling surface mesh nodes. Originality/value As far as authors’ knowledge is concerned, this is the first time in scientific literature that RBF are proposed to handle icing simulations. Due to the meshless characteristic of the RBF mesh morphing, the proposed approach is cross solver and can be used for both 2D and 3D geometries.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-26T10:13:45Z
      DOI: 10.1108/AEAT-07-2018-0178
       
  • On-line orbit planning and guidance for advanced upper stage
    • Pages: 634 - 647
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 634-647, April 2019.
      Purpose This paper aims to investigate the problem of on-line orbit planning and guidance for an advanced upper stage. Design/methodology/approach The double impulse optimal transfer orbit is planned by the Lambert algorithm and the improved particle swarm optimization (IPSO) method, which can reduce the total velocity increment of the transfer orbit. More specially, a simplified formula is developed to obtain the working time of the main engine for two phases of flight based on the theorem of impulse. Subsequently, the true anomalies of the start position and the end position for both two phases are planned by the Newton iterative algorithm and the Kepler equation. Finally, the first phase of flight is guided by a novel iterative guidance (NIG) law based on the true anomaly update with respect to the geometrical relationship. Also, a completely analytical powered explicit guidance (APEG) law is presented to realize orbital injection for the second phase of flight. Findings Simulations including Monte Carlo and three typical orbit transfer missions are carried out to demonstrate the efficiency of the proposed scheme. Originality/value A novel on-line orbit planning algorithm is developed based on the Lambert problem, IPSO optimization method and Newton iterative algorithm. The NIG and APEG are presented to realize the designed transfer orbit for the first and second phases of flight. Both two guidance laws achieve higher orbit injection accuracies than traditional guidance laws.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-21T10:28:52Z
      DOI: 10.1108/AEAT-08-2018-0225
       
  • Oil spill remote monitoring by using remotely piloted aircraft
    • Pages: 648 - 653
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 648-653, April 2019.
      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
       
  • Strategic approach to managing human factors risk in aircraft maintenance
           organization: risk mapping
    • Pages: 654 - 668
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 654-668, April 2019.
      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
       
  • Rapid trajectory optimization for hypersonic entry using convex
           optimization and pseudospectral method
    • Pages: 669 - 679
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 669-679, April 2019.
      Purpose This paper aims to develop a novel trajectory optimization algorithm which is capable of producing high accuracy optimal solution with superior computational efficiency for the hypersonic entry problem. Design/methodology/approach A two-stage trajectory optimization framework is constructed by combining a convex-optimization-based algorithm and the pseudospectral-nonlinear programming (NLP) method. With a warm-start strategy, the initial-guess-sensitive issue of the general NLP method is significantly alleviated, and an accurate optimal solution can be obtained rapidly. Specifically, a successive convexification algorithm is developed, and it serves as an initial trajectory generator in the first stage. This algorithm is initial-guess-insensitive and efficient. However, approximation error would be brought by the convexification procedure as the hypersonic entry problem is highly nonlinear. Then, the classic pseudospectral-NLP solver is adopted in the second stage to obtain an accurate solution. Provided with high-quality initial guesses, the NLP solver would converge efficiently. Findings Numerical experiments show that the overall computation time of the two-stage algorithm is much less than that of the single pseudospectral-NLP algorithm; meanwhile, the solution accuracy is satisfactory. Practical implications Due to its high computational efficiency and solution accuracy, the algorithm developed in this paper provides an option for rapid trajectory designing, and it has the potential to evolve into an online algorithm. Originality/value The paper provides a novel strategy for rapid hypersonic entry trajectory optimization applications.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-05-01T01:51:22Z
      DOI: 10.1108/AEAT-06-2018-0159
       
  • Disturbances rejection based on sliding mode control
    • Pages: 680 - 699
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 680-699, April 2019.
      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
       
  • Influence of Gurney flaps on aerodynamic characteristics of a
           canard-configuration aircraft
    • Pages: 700 - 707
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 700-707, April 2019.
      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
       
  • Fatigue stress analysis of the DV-2 engine turbine disk
    • Pages: 708 - 716
      Abstract: Aircraft Engineering and Aerospace Technology, Volume 91, Issue 4, Page 708-716, April 2019.
      Purpose The purpose of this study is to improve life prediction of certain components. Fatigue of the high-stressed structural elements is an essential parameter that affects the lifetime of such components. In particular, aviation engines are devices whose failure due to fatigue failure of one of the important components can lead to fatal consequences. Design/methodology/approach In this study, two analyses in the turbine disk of the jet engine during the simulated operating load were performed: The first one was the analysis of the heat-induced stresses using the finite element method. The goal of the second analysis was to determine the residual fatigue strength of a loaded disk by the software tool using the Palmgren - Miner Linear Damage Theory. Findings The results showed a high degree of similarity with the real tests performed on the aircraft engine and revealed the weak points in the design of the jet engine. Research limitations/implications It should be mentioned that without appropriate experiments, results of this analysis could not be verified. Practical implications These results are helpful in the re-designing of the jet engines to increase their technical feasibility. Originality/value Such analysis has been realized in the DV-2 jet engine research and development program for the first time in the history of jet engine manufacturing process in Slovakia and countries of Eastern Europe region.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-14T08:18:31Z
      DOI: 10.1108/AEAT-03-2018-0096
       
  • Exhaust toxicity evaluation in a gas turbine engine fueled by aviation
           fuel containing synthesized hydrocarbons
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to examine the toxicological impacts of exhaust generated during the combustion process of aviation fuel containing synthesized hydrocarbons. Design/methodology/approach Tests on aircraft turbine engines in full scale are complex and expensive. Therefore, a miniature turbojet engine was used in this paper as a source of exhaust gases. Toxicity was tested using innovative BAT–CELL Bio–Ambient Cell method, which consists of determination of real toxic impact of the exhaust gases on the human lung A549 and mouse L929 cells. The research was of a comparative nature. The engine was powered by a conventional jet fuel and a blend of conventional jet fuel with synthesized hydrocarbons. Findings The results show that the BAT–CELL method allows determination of the real exhaust toxicity during the combustion process in a turbine engine. The addition of a synthetic component to conventional jet fuel affected the reduction of toxicity of exhaust gases. It was confirmed for both tested cell lines. Originality/value In the literature related to the area of aviation, numerous publications in the field of testing the emission of exhaust gaseous components, particulates or volatile organic compounds can be found. However, there is a lack of research related to the evaluation of the real exhaust toxicity. In addition, it appears that the data given in aviation sector, mainly related to the emission levels of gaseous exhaust components (CO, Nox and HC) and particulate matters, might be insufficient. To fully describe the engine exhaust emissions, they should be supplemented with additional tests, i.e. in terms of toxicity.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-26T07:32:23Z
      DOI: 10.1108/AEAT-11-2018-0277
       
  • Sensing for aerospace combustor health monitoring
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper proposes new methods of fault detection for fuel systems in order to improve system availability. Novel fault systems are required for environmentally friendly lean burn combustion, but can carry high risk failure modes particularly through their control valves. The purpose of the developed technology is the rapid detection of these failure modes, such as valve sticking or impending sticking, and therefore to reduce this risk. However, sensing valve state is challenging due to hot environmental temperatures, which results in a low reliability for conventional position sensing. Design/methodology/approach Starting with the business needs elicited from stakeholders, a quality functional deployment process is performed to derive sensing system requirements. The process acknowledges the difference between test-bed and in-service aerospace needs through weightings on requirements and maps these customer requirements to systems performance metrics. The design of the system must therefore optimise the sensor suite, on- and off-board signal processing and acquisition strategy. Findings Against this systems engineering process, two sensing strategies are outlined which illustrate the span of solutions, from conventional gas path sensing with advanced signal processing to novel non-invasive sensing concepts. While conventional sensing may be feasible within a test cell, the constraints of aerospace in-service operation may necessitate more novel alternatives. Acoustic emission (detecting very high frequency surface vibration waves) sensing technology is evaluated to provide a non-invasive, remote and high temperature tolerant solution. Through this comparison, the considerations for the end-to-end system design are highlighted to be critical to sensor deployment success in-service. Practical implications The paper provides insight into different means of addressing the important problem of monitoring faults in combustor systems in gas turbines. By casting of the complex design problem within a systems engineering framework, the outline of a toolset for solution evaluation is provided. Originality/value The paper provides three areas of significant contributions: a diversity of methods to diagnosing fuel system malfunctions by measuring changes fuel flow distributions, through novel means, and the combustor exit temperature profiles (cause and effect); the use of analytical methods to support the selection (types and quantities) and placement of sensors to ensure adequate state awareness while minimising their impact on the engine system cost and weight; and an end-to-end data processing approach to provide optimised information for the engine maintainers allowing informed decision-making.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-26T07:27:43Z
      DOI: 10.1108/AEAT-11-2018-0283
       
  • The effects of flow separation on a lambda wing aerodynamics
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose In this paper, experimental and numerical results of a lambda wing have been compared. The purpose of this paper is to study the behaviour of lambda wings using a CFD tool and to consider different numerical models to obtain the most accurate results. As far as the consideration of numerical methods is concerned, the main focus is on the evaluation of computational methods for an accurate prediction of contingent leading edge vortices’ path and the flow separation occurring because of the burst of these vortices on the wing. Design/methodology/approach Experimental tests are performed in a closed-circuit wind tunnel at the Reynolds number of 6 × 105 and angles of attack (AOA) ranging from 0 to 10 degrees. Investigated turbulence models in this study are Reynolds Averaged Navior–Stokes (RANS) models in a steady state. To compare the accuracy of the turbulence models with respect to experimental results, sensitivity study of these models has been plotted in bar charts. Findings The results illustrate that the leading edge vortex on this lambda wing is unstable and disappears soon. The effect of this disappearance is obvious by an increase in local drag coefficient in the junction of inner and outer wings. Streamlines on the upper surface of the wing show that at AOA higher than 8 degrees, the absence of an intense leading edge vortex leads to a local flow separation on the outer wing and a reverse in the flow. Research limitations/implications Results obtained from the behaviour study of transition (TSS) turbulence model are more compatible with experimental findings. This model predicts the drag coefficient of the wing with the highest accuracy. Of all considered turbulence models, the Spalart model was not able to accurately predict the non-linearity of drag and pitching moment coefficients. Except for the TSS turbulence model, all other models are unable to predict the aerodynamic coefficients corresponding to AOA higher than 10 degrees. Practical implications The presented results in this paper include lift, drag and pitching moment coefficients in various AOA and also the distribution of aerodynamic coefficients along the span. Originality/value The presented results include lift, drag and pitching moment coefficients in various AOA and also aerodynamic coefficients distribution along the span.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-26T07:24:23Z
      DOI: 10.1108/AEAT-12-2017-0271
       
  • Flow separation control of NACA-2412 airfoil with bio-inspired nose
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to achieve an optimum flow separation control over the airfoil using passive flow control method by introducing bio-inspired nose near the leading edge of the NACA 2412 airfoil. Design/methodology/approach Two distinguished methods have been implemented on the leading edge of the airfoil: forward facing step, which induces multiple accelerations at low angle of attack, and cavity/backward facing step, which creates recirculating region (axial vortices) at high angle of attack. Findings The porpoise airfoil (optimum bio-inspired nose airfoil) delays the flow separation and improves the aerodynamic efficiency by increasing the lift and decreasing the parasitic drag. The maximum increase in aerodynamic efficiency is 22.4 per cent, with an average increase of 8.6 per cent at all angles of attack. Research limitations/implications The computational analysis has been done for NACA 2412 airfoil at low subsonic speed. Practical implications This design improves the aerodynamic performance and increases structural strength of the aircraft wing compared to other conventional high-lift devices and flow-control devices. Originality/value Different bio-inspired nose designs which are inspired by the cetacean species have been analysed for NACA 2412 airfoil, and optimum nose design (porpoise airfoil) has been found.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:44:22Z
      DOI: 10.1108/AEAT-06-2018-0175
       
  • An integrated solution for space shuttle launching ramp
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this study is to select the most appropriate city in Turkey for space shuttle launching ramp. Design/methodology/approach In the proposed approach, an integrated methodology is used. The SWARA method is used in the first phase of the solution for determining criteria’s importance weights. Based on the criteria weights obtained by the SWARA method, the WASPAS method is used for selecting the best alternative. Findings Mugla is selected for the most suitable city for the first space shuttle launching pad according to determined criteria and proposed model. Research limitations/implications Although there are 81 cities in Turkey, 4 alternatives were selected for evaluation. It is possible to eliminate this limitation by the future studies with the implementation of proposed model to entire of Turkey. Practical implications This proposed model can be used by the countries which want to have a new or first space shuttle launching ramp in the world. Originality/value Although some climatic conditions are pointed out on the location of the space shuttle launching ramp, in literature, there is not a comprehensive and detailed evaluation example as much as the model proposed in this study. Therefore, this study is the first in terms of the proposed model and applied techniques in the sectoral sense. In addition, the study is also a guide for solving the original model which is revealed in the selection of the most suitable alternative city for space shuttle launching ramp by different multi-criteria decision-making methods.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:39:54Z
      DOI: 10.1108/AEAT-10-2018-0257
       
  • Digital holography interferometry for measuring the mass diffusion
           coefficients of N2 in RP-3 and RP-5 jet fuels
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this study is to measure the mass diffusion coefficient of nitrogen in jet fuel using digital holography interferometry for cost-effective designing and modeling of the aircraft tank inerting system. Design/methodology/approach The mass diffusion coefficients of N2 in RP-3 and RP-5 jet fuels were measured by digital holography interferometry at temperatures ranging from 278.15 to 343.15 K. The Arrhenius equation is used to adequately describe the relationship between mass diffusion coefficients and temperature. The viscosities of RP-3 and RP-5 jet fuels were also measured to examine the accuracy of the Stokes–Einstein model in calculating mass diffusion coefficients. Findings As temperature increases from 278.15 to 343.15 K, the mass diffusion coefficients increase 4.23-fold for N2 in RP-3 jet fuel and 5.13-fold for N2 in RP-5 jet fuel. The value of Dµ/T is not constant as the Stokes–Einstein equation expressed, but is a weak linear function of temperature. Practical implications A more accurate diffusion model is proposed by fitting the measured Dµ/T with the temperature and calculating the mass diffusion coefficients of N2 in RP-3 and RP-5 jet fuels within 10 per cent relative deviation. Originality/value A measurement system for mass diffusion coefficients of N2 in RP-3 and RP-5 jet fuels was constructed based on the digital holography interferometry. The mass diffusion coefficient can be expressed by a uniform polynomial function of temperature and viscosity.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:35:41Z
      DOI: 10.1108/AEAT-05-2018-0152
       
  • Numerical analysis of hub cavity leakage
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to predict the effect of the hub cavity leakage on the overall performance with numerical simulations in the fan/booster of a high bypass ratio turbofan engine. Design/methodology/approach Simulations are conducted for leakage at the fan, the outlet of guide vane and the three-stage booster, as well as hub leakage (contain cavities and sealing). The results obtained are compared to the corresponding simulations without hub leakage. Findings The rotor/stator interaction locations are evaluated to discover a better location. The results show that the seal tooth structure produces secondary flow and turbulence in the root of blade suction surface, which increases the aerodynamic loss. The sealing clearance should be controlled to shrink the turbulent region and decrease the leakage. Practical implications This work can provide a theoretical guidance and technical support for the compressor design, which avoid many repeated manufactures and reduce waste of resources. Originality/value This work improves the understanding of the impact mechanism of hub cavity leakage on the performance when the clearance size of seal is variable.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:21:18Z
      DOI: 10.1108/AEAT-06-2018-0154
       
  • Redesign of morphing UAV's winglet using DS algorithm based ANFIS model
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present a novel approach based on the differential search (DS) algorithm integrated with the adaptive network-based fuzzy inference system (ANFIS) for unmanned aerial vehicle (UAV) winglet design. Design/methodology/approach The winglet design of UAV, which was produced at Faculty of Aeronautics and Astronautics in Erciyes University, was redesigned using artificial intelligence techniques. This approach proposed for winglet redesign is based on the integration of ANFIS into the DS algorithm. For this purpose, the cant angle (c), the twist angle (t) and taper ratio (λ) of winglet are selected as input parameters; the maximum value of lift/drag ratio (Emax) is selected as the output parameter for ANFIS. For the selected input and output parameters, the optimum ANFIS parameters are determined by the DS algorithm. Then the objective function based on optimum ANFIS structure is integrated with the DS algorithm. With this integration, the input parameters for the Emax value are obtained by the DS algorithm. That is, the DS algorithm is used to obtain both the optimization of the ANFIS structure and the necessary parameters for the winglet design. Thus, the UAV was reshaped and the maximum value of lift/drag ratio was calculated based on new design. Findings Considerable improvements on the max E are obtained through winglet redesign on morphing UAVs with artificial intelligence techniques. Research limitations/implications It takes a long time to obtain the optimum Emax value by the computational fluid dynamics method. Practical implications Using artificial intelligence techniques saves time and reduces cost in maximizing Emax value. The simulation results showed that satisfactory Emax values were obtained, and an optimum winglet design was achieved. Thus, the presented method based on ANFIS and DS algorithm is faster and simpler. Social implications The application of artificial intelligence methods could be used in designing more efficient aircrafts. Originality/value The study provides a new and efficient method that saves time and reduces cost in redesigning UAV winglets.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:15:28Z
      DOI: 10.1108/AEAT-09-2018-0255
       
  • Design-airworthiness integration method for general aviation aircraft
           during early development stage
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Major changes of an aircraft configuration are conducted during the early design stage. It is important to include the airworthiness regulations at this stage while there is extensive freedom for designing. The purpose of this paper is to introduce an efficient design framework that integrates airworthiness guidelines and documentation at the early design stage. Design/methodology/approach A new design and optimization process is proposed that logically includes the airworthiness regulations as design parameters and constraints by constructing a certification database. The design framework comprises requirements analysis, preliminary sizing, conceptual design synthesis and loads analysis. A design certification relation table (DCRT) describes the legal regulations in terms of parameters and values suitable for use in design optimization. Findings The developed framework has been validated and demonstrated for the design of a Federal Aviation Regulations (FAR) 23 four-seater small aircraft. The validation results show an acceptable level of accuracy to be applied during the early design stage. The total mass minimization problem has been successfully solved while satisfying all the design requirements and certification constraints specified in the DCRT. Moreover, successful compliance with FAR 23 subpart C is demonstrated. The proposed method is a useful tool for design optimization and compliance verifications during the early stages of aircraft development. Practical implications The new certification database proposed in this research makes it simpler for engineers to access a large amount of legal documentation related to airworthiness regulations and provides a link between the regulation text and actual design parameters and their bounds. Originality/value The proposed design optimization framework integrates the certification database that is built of several types of legal documents such as regulations, advisory circulars and standards. The Engineering Requirements and Guide summarizes all the documents and design requirements into a single document. The DCRT is created as a summary table that indicates the design parameters affected by a given regulation(s), the design stage at which the parameter can be evaluated and its value bounds. The introduction of the certification database into the design optimization framework significantly reduces the engineer’s load related for airworthiness regulations.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:10:16Z
      DOI: 10.1108/AEAT-05-2018-0143
       
  • Evaluation of load-carrying capabilities of friction stir welded, TIG
           welded and riveted joints of AA2014-T6 aluminium alloy
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Presently, the materials used in light combat aircraft structures are aluminium alloys and composites. These structures are joined together through riveted joints. The weight of these rivets for the entire aircraft is nearly one ton. In addition to weight, the riveted connection requires a lot of tools, equipments, fixtures and manpower, which makes it an expensive and time-consuming process. Moreover, Al alloy is also welded using tungsten inert gas (TIG) welding process by proper control of process parameters. This process has limitations such as porosity, alloy segregation and hot cracking. To overcome the above limitations, an alternative technology is required. One such technology is friction stir welding (FSW), which can be successfully applied for welding of aluminium alloy in LCA structures. Therefore, this paper aims to compare the load carrying capabilities of FSW joints with TIG welded and riveted joints. Design/methodology/approach FSW joints and TIG welded joints were fabricated using optimized process parameters, followed by riveted joints using standard shop floor practice in the butt and lap joint configurations. Findings The load-carrying capabilities of FSW joints are superior than those of other joints. FSW joints exhibited 75 per cent higher load-carrying capability compared to the riveted joints and TIG-welded joints. Practical implications From this investigation, it is inferred that the FSW joint is suitable for the replacement of riveted joints in LCA and TIG-welded joints. Originality/value Friction stir butt joints exhibited 75 per cent higher load-carrying capability than riveted butt joints. Friction stir welded lap joints showed 70 per cent higher load-carrying capability than the riveted lap joints. Friction stir butt joints yielded 41 per cent higher breaking load capabilities than the TIG-welded butt joints. Moreover, Friction stir lap weld joints have 57 per cent more load-carrying capabilities than the TIG-welded lap joints.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-20T09:00:36Z
      DOI: 10.1108/AEAT-11-2017-0233
       
  • Systematic evaluation of the helicopter rotor blades: design variables and
           interactions
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on figure of merit (FM) and lift-to-drag responses developed for hover and forward flight, respectively. Design/methodology/approach The approach is to couple a general flight dynamic simulation code, previously validated in the time domain, with design of experiment (DOE) required for the response surface development. DOE includes I-optimality criteria to preselect the data and improve data acquisition process. Desirability approach is also implemented for a better understanding of the optimum rotor blade planform in both hover and forward flight. Findings The resulting system provides a systematic manner to examine the rotor blade design variables and their interactions, thus reducing the time and cost of designing rotor blades. The obtained results show that the blade taper ratio of 0.3, the point of taper initiation of about 0.64 R within a SC1095R8 airfoil satisfy the maximum FM of 0.73 and the maximum lift-to-drag ratio of about 5.5 in hover and forward flight. Practical implications The work shows the practical possibility to implement the proposed optimization process that can be used for the advanced rotor blade design. Originality/value The work presents the rapid and reliable optimization process efficiently used for designing advanced rotor blades in hover and forward flight.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-13T02:05:14Z
      DOI: 10.1108/AEAT-06-2018-0163
       
  • 3D path planning, routing algorithms and routing protocols for unmanned
           air vehicles: a review
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a comprehensive review in major research areas of unmanned air vehicles (UAVs) navigation, i.e. three degree-of-freedom (3D) path planning, routing algorithm and routing protocols. The paper is further aimed to provide a meaningful comparison among these algorithms and methods and also intend to find the best ones for a particular application. Design/methodology/approach The major UAV navigation research areas are further classified into different categories based on methods and models. Each category is discussed in detail with updated research work done in that very domain. Performance evaluation criteria are defined separately for each category. Based on these criteria and research challenges, research questions are also proposed in this work and answered in discussion according to the presented literature review. Findings The research has found that conventional and node-based algorithms are a popular choice for path planning. Similarly, the graph-based methods are preferred for route planning and hybrid routing protocols are proved better in providing performance. The research has also found promising areas for future research directions, i.e. critical link method for UAV path planning and queuing theory as a routing algorithm for large UAV networks. Originality/value The proposed work is a first attempt to provide a comprehensive study on all research aspects of UAV navigation. In addition, a comparison of these methods, algorithms and techniques based on standard performance criteria is also presented the very first time.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-13T01:03:14Z
      DOI: 10.1108/AEAT-01-2019-0023
       
  • Laser velocimetry for turbofan inlet distortion applications
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to assess state-of-the-art techniques for quantifying flow distortion in the inlets of turbofan engines, particularly with respect to the prospects for future flight applications. Design/methodology/approach To adequately characterize the flow fields of complex aircraft inlet distortions, the author has incorporated laser velocimetry techniques, namely, stereoscopic particle image velocimetry (PIV) and Doppler velocimetry based on filtered Rayleigh scattering (FRS), into inlet distortion studies. Findings Overall, the results and experience indicate that the pathway for integration of FRS technologies into flight systems is clearer and more robust than that of PIV. Practical implications While always a concern, the topic of inlet distortion has grown in importance as contemporary airframe designers seek extremely compact and highly integrated inlets. This research offers a means for gaining new understanding of the in situ aerodynamic phenomena involved with complex inlet distortion. Originality/value This paper presents unique applications of turbofan inlet velocimetry methods while providing an original assessment of technological challenges involved with progressing advanced velocimetry techniques for flight measurements.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-06T12:55:43Z
      DOI: 10.1108/AEAT-11-2018-0285
       
  • Characteristics of a co-flowing jet with varying lip thickness and
           constant bypass ratio
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose Subsonic commercial aircraft operate with turbo-fan engines that operate with moderate bypass ratio (BR) co-flowing jets (CFJ). This study aims to analyse CFJ with constant BR 6.3 and varying primary nozzle lip thickness (LT) to find a critical LT in CFJ below which mixing enhances and beyond which mixing inhibits. Design/methodology/approach CFJ were characterized with a constant BR of 6.3 and varying lip thicknesses. A single free jet with a diameter equal to that of a primary nozzle of the co-flowing jet was also studied for comparison. Findings The results show that within a critical limit, the mixing enhanced with an increase in LT. This was signified by a reduction in potential core length (PCL). Beyond this limit, mixing inhibited leading to the elongation of PCL. This limit was controlled by parameters such as LT and magnitude of BR. Practical implications The BR value of CFJ in the present study was 6.3. This lies under the moderate BR value at which subsonic commercial turbofan operates. Hence, it becomes impervious to study its mixing behavior. Originality/value This is the first effort to find the critical value of LT for a constant BR for compressible co-flow jets. The CFJ with moderate BR and varying LT has not been studied in the past. The present study focuses on finding a critical LT below which mixing enhances and above which mixing inhibits.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-06T01:06:05Z
      DOI: 10.1108/AEAT-01-2019-0007
       
  • A stochastic detailed scheduling model for periodic maintenance of
           military rotorcraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to develop a stochastic detailed schedule for a preventive/scheduled/periodic maintenance program of a military aircraft, specifically a rotorcraft or helicopter. Design/methodology/approach The new model, entitled the military “periodic aviation maintenance stochastic schedule” (PAM-SS), develops a stochastic detailed schedule for a PUMA SA 330SM helicopter for the 50-h periodic inspection, using cyclic operation network (CYCLONE) and Monte Carlo simulation (MCS) techniques. The PAM-SS model identifies the different periodic inspection tasks of the maintenance schedule, allocates the resources required for each task, evaluates a stochastic duration of each inspection task, evaluates the probability of occurrence for each breakdown or repair, develops the CYCLONE model of the stochastic schedule and simulates the model using MCS. Findings The 50-h maintenance stochastic duration follows a normal probability distribution and has a mean value of 323 min and a standard deviation of 23.7 min. Also, the stochastic maintenance schedule lies between 299 and 306 min for a 99 per cent confidence level. Furthermore, except the pilot and the electrical team (approximately 90 per cent idle), all other teams are around 40 per cent idle. A sensitivity analysis is also performed and yielded that the PAM-SS model is not sensitive to the number of technicians in each team; however, it is highly sensitive to the probability of occurrence of the breakdowns/repairs. Practical implications The PAM-SS model is specifically developed for military rotorcrafts, to manage the different resources involved in the detailed planning and scheduling of the periodic/scheduled maintenance, mainly the 50-h inspection. It evaluates the resources utilization (idleness and queue), the stochastic maintenance duration and identifies backlogs and bottlenecks. Originality/value The PAM-SS tackles military aircraft planning and scheduling in a stochastic methodology, considering uncertainties in all inspection task durations and breakdown or repair durations. The PAM-SS, although developed for rotorcrafts can be further developed for any other type of military aircraft or any other scheduled maintenance program interval.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-06T01:00:24Z
      DOI: 10.1108/AEAT-09-2018-0254
       
  • Multidisciplinary wing design of a light long endurance UAV
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is finding the optimal geometric parameters and developing of a method for optimizing a light unmanned aerial vehicle (UAV) wing, maximizing, at the same time, its endurance with the assumed parameters of aircraft mission. Design/methodology/approach The research is based on the experience gained by the author’s contribution to the project of building medium-altitude, long-endurance class, light UAV called “Samonit”. The author was responsible for the structure design, wind tunnel tests and flight tests of the “Samonit” aircraft. Based on the experience, the author was able to develop an optimization process considering various disciplines involved in the whole aircraft design topics such as aerodynamics, flight mechanics, structural stiffness and weight, aircraft stability and maneuverability. The presented methodology has a multidisciplinary nature, as in the process of optimization both aerodynamic aspects and the influence of wing geometric parameters on the wing structure and weight and the aircraft payload were taken into account. The optimal wing configuration was obtained using the genetic algorithms. Findings As a result, a set of wing geometrical parameters has been obtained that allowed for achieving twice as long endurance as compared with the initial one. Practical implications Using the methodology presented in the paper, an aircraft designer can easily find the optimum wing configuration of a designed aircraft, satisfying the mission requirements in a best way. Originality/value An original procedure has been developed, based on the actual design, wind tunnel tests and numerical calculations of “Samonit” aircraft, enabling the determination of optimum wing configuration for a small unmanned aircraft.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-05T10:36:03Z
      DOI: 10.1108/AEAT-09-2018-0256
       
  • Challenges of turboprop engine installation on small aircraft
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present the challenges of turbine engine installation on small aircraft. The work was a part of the European Union project Efficient Systems and Propulsion for Small Aircraft, FP7 EU – Activity, 7.1.4. improving cost efficiency. Design/methodology/approach Few of the most interesting issues associated with replacing a piston engine with a turboprop engine were chosen: changes in engine bay cooling, air inlet, exhaust system, nacellès weight and parts reduction, flight tests and performance. The publication presents an approach to: design, assemble and test the small aircraft with a turboprop engine. Findings Replacement of piston engine by turbine was carried out. The full program of ground and flight test small aircraft has been successfully completed. Pros and cons of the new design are described in the paper. Practical implications Currently, aviation gasoline (AVGAS ) is increasingly being replaced by JET-A1 (kerosene-type fuels) or diesel fuel. The change concerns engine replacement and all the necessary additional components on the aircraft. This is consistent with the new directions of development of aviation: clean, green and eco design. Replacing the piston engine with a turbine allows improvement to performance and reduces operation cost. Originality/value The achieved results allow for identifying and highlighting new directions of aviation technology development. A significant added value is to draw attention to the necessity of preparing for future requirements and amendments in regulations for the new class of aircraft: general aviation SET(L) – single engine turboprop.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-06-05T10:30:19Z
      DOI: 10.1108/AEAT-09-2017-0198
       
  • Inertia parameter identification of anunknown captured space target
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present a method to obtain the inertia parameter of a captured unknown space target. Design/methodology/approach An inertia parameter identification method is proposed in the post-capture scenario in this paper. This method is to resolve parameter identification with two steps: coarse estimation and precise estimation. In the coarse estimation step, all the robot arms are fixed and inertia tensor of the combined system is first calculated by the angular momentum conservation equation of the system. Then, inertia parameters of the unknown target are estimated using the least square method. Second, in the precise estimation step, the robot arms are controlled to move and then inertia parameters are once again estimated by optimization method. In the process of optimization, the coarse estimation results are used as an initial value. Findings Numerical simulation results prove that the method presented in this paper is effective for identifying the inertia parameter of a captured unknown target. Practical implications The presented method can also be applied to identify the inertia parameter of space robot. Originality/value In the classic momentum-based identification method, the linear momentum and angular momentum of system, both considered to be conserved, are used to identify the parameter of system. If the elliptical orbit in space is considered, the conservation of linear momentum is wrong. In this paper, an identification based on the conservation of angular momentum and dynamics is presented. Compared with the classic momentum-based method, this method can get a more accurate identification result.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-05-28T09:10:38Z
      DOI: 10.1108/AEAT-04-2018-0128
       
  • High redundancy electromechanical actuator for thrust vector control of a
           launch vehicle
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to design an electromechanical actuator which can inherently tolerate a stuck or loose failure without any need for fault detection isolation and reconfiguration. Design/methodology/approach Generalized design methodology for a thrust vector control application is adopted to reduce the design iterations during the initial stages of the design. An optimum ball screw pitch is selected to minimize the motor sizing and maximize the load acceleration. Findings A high redundancy electromechanical actuator for thrust vector control has lower self-inertia and higher reliability than a direct drive simplex configuration. This configuration is a feasible solution for thrust vector control application because it offers a more acceptable and graceful degradation than a complete failure. Research limitations/implications Future work will include testing on actual hardware to study the transient disturbances caused by a fault and their effect on launch vehicle dynamics. Practical implications High redundancy electromechanical actuator concept can be extended to similar applications such as solid motor nozzle in satellite launch vehicles and primary flight control system in aircraft. Social implications High redundancy actuators can be useful in safety critical applications involving human beings. It can also reduce the machine downtime in industrial process automation. Originality/value The jam tolerant electromechanical actuator proposed for the launch vehicle application has a unique configuration which does not require a complex fault detection isolation and reconfiguration logic in the controller. This enhances the system reliability and allows a simplex controller having a lower cost.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-05-28T09:05:37Z
      DOI: 10.1108/AEAT-06-2018-0165
       
  • Aeroelastic stability analysis of hingeless rotor blades in hover using
           fully intrinsic equations and dynamic wake model
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to develop a new approach for aeroelastic analysis of hingeless rotor blades. Design/methodology/approach The aeroelastic approach developed here is based on the geometrically exact fully intrinsic beam equations and three-dimensional unsteady aerodynamics. Findings The developed approach is accurate, fast and very useful in rotorcraft aeroelastic analysis. Originality/value This beam formulation has been never combined with three-dimensional aerodynamic model to be used for aeroelastic analysis of blades. In addition, it is possible to handle the composite blades, as well as blades with initial curvatures and twist with this proposed formulation.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-05-28T09:00:57Z
      DOI: 10.1108/AEAT-07-2018-0212
       
  • Conceptual design of an aircraft for Mars mission
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present the results of a conceptual design of Martian aircraft. This study focuses on the aerodynamic and longitudinal dynamic stability analysis. The main research questions are as follows: Does a tailless aircraft configuration can be used for Martian aircraft' How to the short period characteristic can be improved by side plates modification' Design/methodology/approach Because of a conceptual design stage of this Martian aircraft, aerodynamic characterises were computed by the Panukl package by using the potential flow model. The longitudinal dynamic stability was computed by MATLAB code, and the derivatives computed by the SDSA software were used as the input data. Different aircraft configurations have been studied, including different wing’s aerofoils and configurations of the side plate. Findings This paper presents results of aerodynamic characteristics computations and longitudinal dynamic stability analysis. This paper shows that tailless aircraft configuration has potential to be used as Martian aircraft. Moreover, the study of the impact of side plates’ configurations on the longitudinal dynamic stability is presented. This investigation reveals that the most effective method to improve the short period damping ratio is to change the height of the bottom plate. Practical implications The presented result might be useful in case of further design of the aircrafts for the Mars mission and designing the aircrafts in a tailless configuration. Social implications It is considered by the human expedition that Mars is the most probable planet to explore. This paper presents the conceptual study of aircraft which can be used to take the high-resolution pictures of the surface of Mars, which can be crucial to find the right place to establish a potential Martian base. Originality/value Most of aircrafts proposed for the Mars mission are designed in a configuration with a classic tail; this paper shows a preliminary calculation of the tailless Martian aircraft. Moreover, this paper shows the results of a dynamic stability analysis, where similar papers about aircrafts for the Mars mission do not show such outcomes, especially in the case of the tailless configuration. Moreover, this paper presents the results of the dynamic stability analysis of tailless aircraft with different configurations of the side plates.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-04-05T08:41:24Z
      DOI: 10.1108/AEAT-08-2018-0231
       
  • Fast high fidelity CFD/CSM fluid structure interaction using RBF mesh
           morphing and modal superposition method
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present a fast and effective approach to tackle complex fluid structure interaction problems that are relevant for the aeronautical design. Design/methodology/approach High fidelity computer-aided engineering models (computational fluid dynamics [CFD] and computational structural mechanics) are coupled by embedding modal shapes into the CFD solver using RBF mesh morphing. Findings The theoretical framework is first explained and its use is then demonstrated with a review of applications including both steady and unsteady cases. Different flow and structural solvers are considered to showcase the portability of the concept. Practical implications The method is flexible and can be used for the simulation of complex scenarios, including components vibrations induced by external devices, as in the case of flapping wings. Originality/value The computation mesh of the CFD model becomes parametric with respect to the modal shape and, so, capable to self-adapt to the loads exerted by the surrounding fluid both for steady and transient numerical studies.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-04-05T08:37:15Z
      DOI: 10.1108/AEAT-09-2018-0246
       
  • High-precision navigation and positioning of celestial exploration rover
           based on depth camera
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to verify the correctness and feasibility of simultaneous localization and mapping (SLAM) algorithm based on red-green-blue depth (RGB-D) camera in high precision navigation and localization of celestial exploration rover. Design/methodology/approach First, a positioning algorithm based on depth camera is proposed. Second, the realization method is described from the five aspects of feature detection method, feature point matching, point cloud mapping, motion estimation and high precision optimization. Feature detection: taking the precision, real-time and motion basics as the comprehensive consideration, the ORB (oriented FAST and rotated BRIEF) features extraction method is adopted; feature point matching: solves the similarity measure of the feature descriptor vector and how to remove the mismatch point; point cloud mapping: the two-dimensional information on RGB and the depth information on D corresponding; motion estimation: the iterative closest point algorithm is used to solve point set registration; and high precision optimization: optimized by using the graph optimization method. Findings The proposed high-precision SLAM algorithm is very effective for solving high precision navigation and positioning of celestial exploration rover. Research limitations/implications In this paper, the simulation validation is based on an open source data set for testing; the physical verification is based on the existing unmanned vehicle platform to simulate the celestial exploration rover. Practical implications This paper presents a RGB-D camera-based navigation algorithm, which can be obtained by simulation experiment and physical verification. The real-time and accuracy of the algorithm are well behaved and have strong applicability, which can support the tests and experiments on hardware platform and have a better environmental adaptability. Originality/value The proposed SLAM algorithm can deal with the high precision navigation and positioning of celestial exploration rover effectively. Taking into account the current wide application prospect of computer vision, the method in this paper can provide a study foundation for the deep space probe.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-04-05T08:30:21Z
      DOI: 10.1108/AEAT-09-2017-0200
       
  • A review of the analytical methods used for seaplanes’ performance
           prediction
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to investigate the different analytical methods used to predict the performance of seaplanes to define the weaknesses in each method and be able to extend the analytical approach to include the nonlinear terms (unsteadiness). Design/methodology/approach First, the elemental hydrodynamic characteristics of seaplanes are discussed. Second, five different analytical methods are reviewed. The advantages and disadvantages of each method are stated. After that, the heave and pitch equations of seaplane motion are illustrated. The procedure of obtaining the solution of the heave and pitch equations of seaplane motion is explained. Finally, the results obtained from the most common methods are compared. Findings The results show that the methods are based on different assumptions and considerations. As a result, no method is optimal for all types of seaplanes. Moreover, some of the analytical methods do not study the stability of the seaplane, which is a major issue in the design of seaplanes. In addition, all methods consider the motion as steady and linear. The objective is to extend the work to include the nonlinear effects. Originality/value This paper presents some of the analytical methods used in describing the performance of seaplanes and explains how can they be applied. Moreover, it summarises the advantages and disadvantages of each method.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-27T10:14:04Z
      DOI: 10.1108/AEAT-07-2018-0186
       
  • Aircraft flow angles calibration via observed-based wind estimation
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to propose a novel approach, in which the reference data for the flow angles calibration are obtained by using measurements coming from an inertial navigation system and an air data sensor. Design/methodology/approach This is obtained by using the Kalman filter theory for the evaluation of the reference angle-of-attack and angle-of-sideslip. Findings The designed Kalman filter has been implemented in Matlab/Simulink and validated using flight data coming from two very different aircraft, the Piaggio Aerospace P1HH medium altitude long endurance unmanned aerial system and the Alenia-Aermacchi M346 Master™ transonic trainer. This paper illustrates some results where the filter satisfactory behaviour is verified by comparing the filter outputs with the data coming from high-accuracy nose-boom vanes. Practical implications The methodology aims to lower the calibration costs of the air data systems of an advanced aircraft. Originality/value The calibration of air-data systems for the evaluation of the flow angles is based on the availability of high-accuracy reference measurements of angle-of-attack and angle-of-sideslip. Typically, these are obtained by auxiliary sensors directly providing the reference angles (e.g. nose-boom vanes). The proposed methodology evaluates the reference angle-of-attack and angle-of-sideslip by analytically reconstructing them using calibrated airspeed measurements and inertial data.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-22T08:48:01Z
      DOI: 10.1108/AEAT-06-2017-0145
       
  • Study on multi-loop control strategy of three-shaft gas turbine for
           electricity generation
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to develop a dynamic performance model of three-shaft gas turbine for electricity generation and to study a multi-loop control strategy of three-shaft gas turbine for electricity generation. Design/methodology/approach In this paper, the dynamic performance model of the three-shaft gas turbine is established and developed. A novel approach, variable partial differential coefficient deviation linearization method is used to simulate the dynamic performance of the three-shaft gas turbine. Single-loop control system, feed-forward feedback control system and cascade system are assessed to control the engine during transient operation. Findings A novel approach, variable partial differential coefficient deviation linearization method is used to simulate the dynamic performance of the three-shaft gas turbine. According to the results shown, the cascade control system is most satisfactory due to its fastest response and the best stability and robustness. Originality/value The method of variable partial linearization is adopted to make the dynamic simulation of the model achieve higher precision, better steady state and less computation time. This paper provides a theoretical study for the multi-loop control system of a marine three-shaft gas turbine.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-22T08:44:01Z
      DOI: 10.1108/AEAT-05-2018-0149
       
  • Piston-electric propulsion system as reliable alternative for classic and
           nonconventional piston engine configurations
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to define reliability requirements to be imposed on electric engines to assure similar or higher value of mean time between failures (MTBF) for mixed piston-electric propulsion configurations when compared to classic and unconventional piston engine configurations. Design/methodology/approach Reliability estimation was done using mathematical model of safety of light aircraft commercial operations. The model was developed on the basis of Federal Aviation Administration and National Transport Safety Board data. The analysis was conducted for numerous piston and electric configurations. It allowed comparison of selected solutions and definition of relation between electric engine MTBF and MTBF calculated for entire mixed piston-electric propulsion system. Findings It was found that, from reliability point of view, mixed piston-electric engine propulsion is attractive alternative for classic single- and twin-piston configuration. It would allow to at least doubling of MTBF for propulsion without increase of operational cost. Practical implications Rationale behind exploiting electric propulsion in aviation is provided. Relation between electric engine reliability and entire propulsion reliability was identified and defined. Minimum requirements concerning MTBF value for electric engine application in aviation was assessed. Conclusions from this study can be used for definition of requirements for new aircraft and by the regulatory authorities. Originality/value Originality consists in use of real accident statistics included in mathematical model of safety for assessment of MTBF for various classic and novel piston and piston-electric engine configurations of light aircraft. Output from the study can be exploited by the industry.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-19T09:41:20Z
      DOI: 10.1108/AEAT-01-2018-0037
       
  • Dynamics and control of a flexible rotating clamped-free beam by SDRE
           strategy
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is the dynamic analysis of the coupled rotation and vibration motion of a system containing a central rigid body to which is attached a flexible beam. Design/methodology/approach The methodology includes the Lagrange’s formulation by using the extended Hamilton’s Principle in conjunction with the assumed modes method to describe the system of equations by ordinary differential equations. The first unconstrained mode of vibration was considered as the solution for the transversal displacement. Such mode emerges as the eigenvalue problem solution associated to the dynamics of the system. The control strategy adopted is a nonlinear analogy of the linear quadratic regulator problem as the Riccati equation is solved at every integration step during the numerical solutions. This strategy is known as state-dependent Riccati equation. Findings By means of computational simulations, it was found the relation between controlled motion and inertia ratio. Research limitations/implications This work is limited to planar case and fixed hub. Practical implications Practical implications of this work realize the design of lighter yet dexterous structures. Originality/value The contribution of this paper is the position and vibration control of a flexible beam accounting for nonlinearity effects and the fact that the structure to where it is clamped has a comparable inertia.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-14T08:51:49Z
      DOI: 10.1108/AEAT-11-2017-0240
       
  • Enhancements in conceptual electric aircraft design
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to describe the enhancement of the numerical method for conceptual phase of electric aircraft design. Design/methodology/approach The algorithm provides a balance between lift force and weight of the aircraft, together with drag and thrust force equilibrium, while modifying design variables. Wing geometry adjustment, mass correction and performance estimation are performed in an iterative process. Findings Aircraft numerical model, which is most often very simplified, has a number of new improvements. This enables to make more accurate analyses and to show relationships between design parameters and aircraft performance. Practical implications The presented approach can improve design results. Originality/value The new methodology, which includes enhanced numerical models for conceptual design, has not been presented before.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-14T08:43:07Z
      DOI: 10.1108/AEAT-07-2018-0192
       
  • An automatic system for a helicopter autopilot performance evaluation
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to elaborate and develop an automatic system for automatic flight control system (AFCS) performance evaluation. Consequently, the developed AFCS algorithm is implemented and tested in a virtual environment on one of the mission task elements (MTEs) described in Aeronautical Design Standard 33 (ADS-33) performance specification. Design/methodology/approach Control algorithm is based on the Linear Quadratic Regulator (LQR) which is adopted to work as a controller in this case. Developed controller allows for automatic flight of the helicopter via desired three-dimensional trajectory by calculating iteratively deviations between desired and actual helicopter position and multiplying it by gains obtained from the LQR methodology. For the AFCS algorithm validation, the objective data analysis is done based on specified task accomplishment requirements, reference trajectory and actual flight parameters. Findings In the paper, a description of an automatic flight control algorithm for small helicopter and its evaluation methodology is presented. Necessary information about helicopter dynamic model is included. The test and algorithm analysis are performed on a slalom maneuver, on which the handling qualities are calculated. Practical implications Developed automatic flight control algorithm can be adapted and used in autopilot for a small helicopter. Methodology of evaluation of an AFCS performance can be used in different applications and cases. Originality/value In the paper, an automatic flight control algorithm for small helicopter and solution for the validation of developed AFCS algorithms are presented.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-14T08:32:46Z
      DOI: 10.1108/AEAT-07-2018-0190
       
  • The approximation method in the problem on a flow of viscous fluid around
           a thin plate
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The paper aims to obtain an effective solution to the problem on a flow of viscous fluid around a thin plate using a new approximation method based on the exact Navier–Stokes equations. Also, correction factors are proposed to improve the obtained solution at high Reynolds numbers. Design/methodology/approach The paper has opted for a method that is based on an approximation scheme for certain perturbations concerning the velocity of the oncoming unperturbed flow behind a leading edge of the plate as a zero approximation step. The perturbations are assumed to be small, far from the plate when compared to the basic flow to justify the linearization. Numerical methods are used for the integral equations at each approximation step. Findings This paper provides the friction force coefficient compared with the classical Blasius solution and the ANSYS results. Also, some diagrams of the velocity distribution in the flow are presented. The first and second approximation steps provide a sufficiently high degree of accuracy. Research limitations/implications Because of the chosen research approach, the results may lack accuracy for low and average Reynolds numbers. Thus, researchers are encouraged to improve the proposed method further. Practical implications The paper includes implications for the development of an aircraft design or a wind turbine design considering a wing as a thin plate at the first approximation. Originality/value This paper provides a new approximation method based on the exact Navier–Stokes equations, in contrast to the known solutions.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-12T12:26:08Z
      DOI: 10.1108/AEAT-07-2018-0196
       
  • Magna-Lok rivet joint and the stiffness-equivalent FE model
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present a methodology for the determination of the stiffness when using simplified substitutive model of the joint. The usage of detailed finite element (FE) model of the joint in complex assemblies is not convenient; therefore, the substitutive model of the joint is used in FE models. Design/methodology/approach The detailed and simplified FE model of the joint is created in ABAQUS software and the analysis as well. The results of displacements are used for the determination of the stiffness of connecting element in simplified substitutive FE model. The approach is presented based on the general view on the different regions in the joint. Findings A simple FE modelling approach for the joint including the equivalent stiffness is presented. The particular solution is performed for Magna-Lok type of the rivet. The results show the same displacement for the detailed and simplified FE models. The analytical formula for stiffness determination in the load case with minimal secondary bending is introduced. Practical implications The approach for stiffness determination is straightforward and so no stiffness “tuning” is necessary in the simplified FE model. Originality/value The new approach for definition of simple FE model of the joint is introduced. It is not necessary to model a complex structure with detailed joints. The equivalent stiffness can be determined by presented procedure for every joint without limitation of the type.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-12T10:26:35Z
      DOI: 10.1108/AEAT-07-2018-0188
       
  • Establishment of the Swedish Aeronautical Research Center (SARC)
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to present the newly founded Swedish Aeronautical Research Center (SARC), based on the triple helix theory, to foster the seamless Swedish aerospace research interplay between academia, research organizations and industry. Design/methodology/approach The paper is a technical paper, mainly relating and explaining sources and concepts for research planning and organization. Used concepts are the triple helix approach (for socioeconomic effects), the role of academia and industry interplay for education and the technology readiness level (TRL) concept for strategic research planning. Focusing on the establishment of a graduate school, lessons learned from previous national research schools are also presented. Findings The paper gives an overview of and explains the interplay between politics, social welfare and industrial R&D needs, with the academic viewpoint of aeronautical research and education. Shortcomings in both the use of TRL for research program planning and the Swedish competence cluster system are identified and remedies suggested. The main findings are suggestions for future actions to be conducted by SARC in the fields of research and education. Practical implications The paper includes implications for the seamless interplay between academia, research organizations and industry. Originality/value So far, no publication about the newly founded SARC has been made yet. It is unique in the way that it makes substantial use of national technical documents so that this information becomes available for non-Swedish speakers. Additionally, the perhaps-unique system of industrial competence clusters is presented.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-12T10:26:05Z
      DOI: 10.1108/AEAT-07-2018-0201
       
  • FDM 3D printing method utility assessment in small RC aircraft design
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to investigate the possibility of manufacturing fused deposition modelling (FDM) 3D printed structures such as wings or fuselages for small remote control (RC) air craft and mini unmaned aerial vehicles (UAVs). Design/methodology/approach Material tests, design assumptions and calculations were verified by designing and manufacturing a small radio-controlled motor-glider using as many printed parts as possible and performing test flights. Findings It is possible to create an aircraft with good flight characteristics using FDM 3D printed parts. Current level of technology allows for reasonably fast manufacturing of 3D printed aircraft with good reliability and high success ratio of prints; however, only some of the materials are suitable for printing thin wall structures such as wings. Practical implications The paper proves that apart from currently popular small RC aircraft structural materials such as composites, wood and foam, there is also printed plastic. Moreover, 3D printing is highly competitive in some aspects such as first unit production time or production cost. Originality/value The presented manufacturing technique can be useful for quick and cost-effective creating scale prototypes of the aircraft for performing test flights.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-03-12T10:06:53Z
      DOI: 10.1108/AEAT-07-2018-0189
       
  • Visual and microscopic examination of the rocket engine combustion chamber
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose This paper aims to enhance the selection of the best material of the rocket engine combustion chamber. The chamber has been destroyed during dynamometer tests, and the goal of this inspection is to verify the nature of the damage in the context of checking the usefulness of this type of graphite for the combustion chamber construction. Design/methodology/approach This paper presents the results of visual and microscopic inspection of the rocket engine combustion chamber of Ø50 × 165 mm in dimension, which was made of R type graphite. Findings An analysis of the fracture surface shows that in the inspected combustion chamber voids and inclusions are present. EDS analysis of the fracture surface shows that in the inspected combustion chamber inclusions are present which have a relatively high amount of elements like: Ti, C, S, V, Si, O and a relatively small amount of Fe and Ni. Research limitations/implications Research limitations is concerned the failure analysis by a scanning electron microscope (SEM) Zeiss EVO 25 MA with EDS detector: Brüker X Flash Detector 5010 125 eV and Espirit 1.9.0.2176 EDS software. Practical implications Designing of the engine combustion chamber the researches can select the best of the rocket engine combustion chamber, made of R type graphite, with the minimum voids and inclusions to decrease the possibility of bursting of this chamber. Originality/value The most dangerous issues in the inspected combustion chamber during an outflow are hot gases as a result of high fuel combustion temperature, so it causes the nozzle heating and the engine stress increase of visible inclusions in cross-sections.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-28T03:30:49Z
      DOI: 10.1108/AEAT-06-2018-0166
       
  • Short range rocket-target: research, development and implementation
    • Abstract: Aircraft Engineering and Aerospace Technology, Ahead of Print.
      Purpose The purpose of this paper is to present the aerodynamic analysis and external ballistics modeling used in the development of a rocket-target for short range air defence missile systems. Design/methodology/approach A computational fluid dynamics (CFD) analysis of the airflow around the rocket-target was carried out to estimate the drag, which was needed to develop a mathematical model for external ballistics of the rocket-target. Field-experimental testing was conducted to compare the model results to the data obtained experimentally using various additional measurement techniques such as global positioning system (GPS) coordinates marking of the crash and launch sites, air defence surveillance radar tracking and installing equipment for telemetric data capturing and transmission. Findings Various ballistic parameters such as the velocity and trajectory of the rocket-target were obtained taking into account the CFD analysis results and internal ballistics data. The field-experimental testing showed a good agreement between the model results and the results obtained by the experimental techniques. Practical implications The presented computational models and the experimental techniques could be used in future developments of similar aircraft. Originality/value This paper presents a research approach for developing a rocket-target. The results of the research were used as a basis for developing a rocket-target for short range air defence rocket systems. The developed rocket-target was successfully implemented in practice.
      Citation: Aircraft Engineering and Aerospace Technology
      PubDate: 2019-02-25T12:52:18Z
      DOI: 10.1108/AEAT-07-2018-0177
       
  • 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
       
  • 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
       
  • 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
       
  • 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
       
  • 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
       
  • 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
       
 
 
JournalTOCs
School of Mathematical and Computer Sciences
Heriot-Watt University
Edinburgh, EH14 4AS, UK
Email: journaltocs@hw.ac.uk
Tel: +00 44 (0)131 4513762
Fax: +00 44 (0)131 4513327
 
Home (Search)
Subjects A-Z
Publishers A-Z
Customise
APIs
Your IP address: 18.207.252.123
 
About JournalTOCs
API
Help
News (blog, publications)
JournalTOCs on Twitter   JournalTOCs on Facebook

JournalTOCs © 2009-