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  Subjects -> AERONAUTICS AND SPACE FLIGHT (Total: 120 journals)
Showing 1 - 30 of 30 Journals sorted by number of followers
AIAA Journal     Hybrid Journal   (Followers: 1189)
SpaceNews     Free   (Followers: 826)
Journal of Spacecraft and Rockets     Hybrid Journal   (Followers: 771)
Journal of Propulsion and Power     Hybrid Journal   (Followers: 610)
Acta Astronautica     Hybrid Journal   (Followers: 494)
Advances in Space Research     Full-text available via subscription   (Followers: 457)
Aviation Week     Full-text available via subscription   (Followers: 437)
Aerospace Science and Technology     Hybrid Journal   (Followers: 427)
IEEE Transactions on Aerospace and Electronic Systems     Hybrid Journal   (Followers: 383)
Journal of Aircraft     Hybrid Journal   (Followers: 335)
Control Systems     Hybrid Journal   (Followers: 313)
IEEE Aerospace and Electronic Systems Magazine     Full-text available via subscription   (Followers: 279)
Journal of Navigation     Hybrid Journal   (Followers: 277)
Aircraft Engineering and Aerospace Technology     Hybrid Journal   (Followers: 261)
Gyroscopy and Navigation     Hybrid Journal   (Followers: 258)
Journal of Guidance, Control, and Dynamics     Hybrid Journal   (Followers: 204)
Space Science International     Open Access   (Followers: 199)
Space Science Reviews     Hybrid Journal   (Followers: 97)
International Journal of Aerospace Engineering     Open Access   (Followers: 82)
Progress in Aerospace Sciences     Full-text available via subscription   (Followers: 80)
Advances in Aerospace Engineering     Open Access   (Followers: 69)
Journal of Aerospace Engineering     Full-text available via subscription   (Followers: 69)
Propulsion and Power Research     Open Access   (Followers: 68)
Aerospace     Open Access   (Followers: 60)
Space Safety Magazine     Free   (Followers: 51)
Space Research Today     Full-text available via subscription   (Followers: 48)
Proceedings of the Institution of Mechanical Engineers Part G: Journal of Aerospace Engineering     Hybrid Journal   (Followers: 46)
International Journal of Aeroacoustics     Hybrid Journal   (Followers: 40)
IEEE Transactions on Circuits and Systems I: Regular Papers     Hybrid Journal   (Followers: 39)
International Journal of Aerodynamics     Hybrid Journal   (Followers: 37)
Canadian Aeronautics and Space Journal     Full-text available via subscription   (Followers: 34)
Journal of Aerospace Information Systems     Hybrid Journal   (Followers: 34)
International Journal of Aerospace Sciences     Open Access   (Followers: 32)
Journal of Aeronautics & Aerospace Engineering     Open Access   (Followers: 31)
Space Policy     Hybrid Journal   (Followers: 30)
CEAS Aeronautical Journal     Hybrid Journal   (Followers: 29)
Journal of Space Weather and Space Climate     Open Access   (Followers: 27)
Aviation Psychology and Applied Human Factors     Hybrid Journal   (Followers: 27)
Russian Aeronautics (Iz VUZ)     Hybrid Journal   (Followers: 24)
Egyptian Journal of Remote Sensing and Space Science     Open Access   (Followers: 24)
Artificial Satellites     Open Access   (Followers: 23)
International Journal of Aerospace Psychology     Hybrid Journal   (Followers: 23)
Journal of Aerospace Information Systems     Hybrid Journal   (Followers: 22)
Annual of Navigation     Open Access   (Followers: 22)
Chinese Journal of Aeronautics     Open Access   (Followers: 21)
Nonlinear Dynamics     Hybrid Journal   (Followers: 20)
Aerospace Medicine and Human Performance     Full-text available via subscription   (Followers: 19)
Journal of Aerospace Engineering & Technology     Full-text available via subscription   (Followers: 18)
Journal of Aerodynamics     Open Access   (Followers: 18)
Aerospace Scientific Journal     Open Access   (Followers: 18)
Journal of Wind Engineering and Industrial Aerodynamics     Hybrid Journal   (Followers: 17)
Aviation     Open Access   (Followers: 17)
International Journal of Space Structures     Full-text available via subscription   (Followers: 17)
Research & Reviews : Journal of Space Science & Technology     Full-text available via subscription   (Followers: 17)
Proceedings of the Human Factors and Ergonomics Society Annual Meeting     Hybrid Journal   (Followers: 16)
Fatigue of Aircraft Structures     Open Access   (Followers: 15)
International Journal of Satellite Communications Policy and Management     Hybrid Journal   (Followers: 13)
Frontiers in Astronomy and Space Sciences     Open Access   (Followers: 12)
Elsevier Astrodynamics Series     Full-text available via subscription   (Followers: 12)
Aeronautical Journal, The     Hybrid Journal   (Followers: 12)
International Journal of Crashworthiness     Hybrid Journal   (Followers: 12)
Journal of Airline and Airport Management     Open Access   (Followers: 12)
International Journal of Micro Air Vehicles     Full-text available via subscription   (Followers: 11)
Journal of Aviation Technology and Engineering     Open Access   (Followers: 11)
International Journal of Space Science and Engineering     Hybrid Journal   (Followers: 11)
Air Force Magazine     Full-text available via subscription   (Followers: 11)
COSPAR Colloquia Series     Full-text available via subscription   (Followers: 11)
International Journal of Space Technology Management and Innovation     Full-text available via subscription   (Followers: 10)
Aviation in Focus - Journal of Aeronautical Sciences     Open Access   (Followers: 10)
Journal of Aircraft and Spacecraft Technology     Open Access   (Followers: 9)
Population Space and Place     Hybrid Journal   (Followers: 9)
Journal of Aeronautical Materials     Open Access   (Followers: 9)
International Journal of Aviation Management     Hybrid Journal   (Followers: 9)
Transportmetrica A : Transport Science     Hybrid Journal   (Followers: 9)
Journal of the Astronautical Sciences     Hybrid Journal   (Followers: 9)
Advances in Aerospace Science and Technology     Open Access   (Followers: 8)
Air Medical Journal     Hybrid Journal   (Followers: 8)
Journal of Space Safety Engineering     Hybrid Journal   (Followers: 8)
International Journal of Aviation Technology, Engineering and Management     Full-text available via subscription   (Followers: 7)
Journal of the American Helicopter Society     Full-text available via subscription   (Followers: 7)
Journal of Aerospace Technology and Management     Open Access   (Followers: 7)
International Journal of Applied Geospatial Research     Hybrid Journal   (Followers: 7)
New Space     Hybrid Journal   (Followers: 6)
Aerospace Systems     Hybrid Journal   (Followers: 6)
International Journal of Turbo and Jet-Engines     Hybrid Journal   (Followers: 6)
RocketSTEM     Free   (Followers: 6)
Civil Aviation High Technologies     Open Access   (Followers: 5)
International Journal of Aviation, Aeronautics, and Aerospace     Open Access   (Followers: 5)
REACH - Reviews in Human Space Exploration     Full-text available via subscription   (Followers: 5)
International Journal of Sustainable Aviation     Hybrid Journal   (Followers: 5)
Aviation Advances & Maintenance     Open Access   (Followers: 5)
Cosmic Research     Hybrid Journal   (Followers: 5)
Unmanned Systems     Hybrid Journal   (Followers: 5)
Space and Polity     Hybrid Journal   (Followers: 4)
Life Sciences in Space Research     Hybrid Journal   (Followers: 4)
Aerotecnica Missili & Spazio : Journal of Aerospace Science, Technologies & Systems     Hybrid Journal   (Followers: 4)
Astrodynamics     Hybrid Journal   (Followers: 4)
Investigación Pecuaria     Open Access   (Followers: 3)
Aerospace technic and technology     Open Access   (Followers: 3)
Journal of Spatial Science     Hybrid Journal   (Followers: 3)
ASTRA Proceedings     Open Access   (Followers: 3)
Journal of KONBiN     Open Access   (Followers: 3)
Problemy Mechatroniki. Uzbrojenie, lotnictwo, inżynieria bezpieczeństwa / Problems of Mechatronics. Armament, Aviation, Safety Engineering     Open Access   (Followers: 3)
npj Microgravity     Open Access   (Followers: 3)
Journal of Astrobiology & Outreach     Open Access   (Followers: 3)
Journal of Aviation/Aerospace Education & Research     Open Access   (Followers: 2)
IEEE Journal on Miniaturization for Air and Space Systems     Hybrid Journal   (Followers: 2)
Microgravity Science and Technology     Hybrid Journal   (Followers: 2)
Вісник Національного Авіаційного Університету     Open Access   (Followers: 2)
International Journal of Aeronautical and Space Sciences     Hybrid Journal   (Followers: 2)
Ciencia y Poder Aéreo     Open Access   (Followers: 2)
MAD - Magazine of Aviation Development     Open Access   (Followers: 2)
Journal of the Australasian Society of Aerospace Medicine     Open Access   (Followers: 1)
Open Aerospace Engineering Journal     Open Access   (Followers: 1)
Advances in Astronautics Science and Technology     Hybrid Journal   (Followers: 1)
Journal of Engineering and Technological Sciences     Open Access   (Followers: 1)
Technical Soaring     Full-text available via subscription   (Followers: 1)
Spatial Information Research     Hybrid Journal   (Followers: 1)
Mekanika : Jurnal Teknik Mesin i     Open Access   (Followers: 1)
Perspectives of Earth and Space Scientists i     Open Access  

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Journal Cover
Proceedings of the Institution of Mechanical Engineers Part G: Journal of Aerospace Engineering
Journal Prestige (SJR): 0.422
Citation Impact (citeScore): 1
Number of Followers: 46  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 0954-4100 - ISSN (Online) 2041-3025
Published by Sage Publications Homepage  [1130 journals]
  • A new bladed assembly simulator and an improved two-parameter plot method
           for blade tip-timing numerical simulations
    • Authors: Masood Nikpour, Shapour Moradi, Iman Soodmand
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The blade tip-timing measurement technique is presently the most promising technique for monitoring the blades of axial turbines and aircraft engines in operating conditions. Due to the high cost of experimental simulations of blade tip-timing–based condition monitoring methods, a numerical simulator for the vibrational behavior of bladed assemblies can be helpful for researchers interested in this field. So far, in most of the numerical simulators, the centrifugal effect of rotational speed on the natural frequencies is neglected. In this study, a new bladed assembly considering the centrifugal effect of the rotational speed for blade tip-timing numerical simulations is proposed. Moreover, an improvement in the engine order estimation algorithm in a two-parameter plot method is accomplished. In the assembly, blades are assumed to be cantilevered Euler–Bernoulli beams coupled together using linear springs. The finite element method is used to extract mass and stiffness matrices from differential equations of the system. By using the two-parameter plot method, the engine order of the excitation is detected. To examine the performance of the algorithm, Monte–Carlo simulation is implemented. The new simulator fulfills both cyclic symmetry and increase in the natural frequencies with increase in rotational speed. Engine order estimation with the new formulation in the two-parameter plot method is accurate. Hence, the new simulator and formulation for two-parameter plot method are reliable for numerical simulations.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-13T11:02:58Z
      DOI: 10.1177/0954410021993082
       
  • Burning rate augmentation in solid fuel ramjets by swirling inlet air
           stream
    • Authors: Amir Mahdi Tahsini
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The influence of the inlet swirling flow on the regression rate of the fuel in the combustion chamber of the solid fuel ramjet is investigated in this study using numerical simulations. The finite-volume solver of the compressible turbulent reacting flow is developed to study the flow field where the burning rate is computed using the conjugate heat transfer method for the solid fuel. The correlation is found for the maximum regression rate versus an imposed inlet swirl when the linear distribution of the circumferential velocity is applied at the inlet stream. Although the regression rate augmentation is considerable due to the swirling flow field in the combustor, it is shown that the swirl is effective if is applied near the shear layer of the backstep flow in the combustor. The modified swirler with short blades is suggested to be used in solid fuel ramjets to increase the regression rate of the fuel and improve the performance, but with lower pressure loss.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-13T08:36:08Z
      DOI: 10.1177/09544100211001490
       
  • Numerical analysis and experimental verification of the induced waveform
           characteristics for aeroengine gas path debris electrostatic sensor
    • Authors: Jiachen Guo, Hongfu Zuo, Heng Jiang, Zhirong Zhong, Quan Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article introduces the principle of aeroengine gas path electrostatic monitoring and establishes a mathematical model of aeroengine gas path debris electrostatic sensor. In this study, we simulate particle’s movement based on the established model and perform numerical analysis of the induced charge pulse waveform. The simulation results show the quantitative relationship among particle’s charge amount, velocity, and pulse waveform’s features, and obtain the qualitative relationship between particle’s spatial position and pulse waveform’s features. A test rig is designed to verify the correctness of the mathematical model. A measurement mode based on dual-channel sensors has been proposed, and corresponding signal processing methods are used to calculate the velocity of the particle and reconstruct the charge pulse waveform from the measured voltage signal. The conclusions of this study not only avoid the shortcomings of traditional signal processing methods that directly use the measured voltage signal but also have important significance for improving the electrostatic monitoring capability.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-13T04:36:50Z
      DOI: 10.1177/0954410020986255
       
  • Large-scale flexible spacecraft mass properties determination using
           torque-generating control
    • Authors: Rui Liu, Jun Zhou, Minghe Chi, Gongjun Li, Jun Zhang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      It is necessary to determine the mass properties of an on-orbit large-scale flexible spacecraft in order to achieve high-precision attitude control. The mass properties of the spacecraft include the inertia, the center of mass, and the mass. Research in this area had been mostly focused on rigid spacecrafts. The coupling between the rigid body of the spacecraft and the large flexible appendages on it, however, could have significant effect on the mass properties determination. This article proposes a momentum conservation–based determination method that takes into consideration the rigid–flexible coupling factors of large-scale flexible spacecrafts.First, a control method is designed to ensure that the spacecraft stays motionless after the motivation incurred for mass properties determination. Second, an inertia matrix determination method is developed using a Kalman filter, in which the coupling factors are added in the amendment procedure. Third, the Kalman filter with input is applied in the determination of the center of mass: on the one hand, this method can use one the accelerometer placed at the flexible appendages, if the deformation can be measured; on the other side, the method can use three accelerometers placed at three orthogonal points of the rigid part of the spacecraft, when the deformation cannot be measured. Finally, the mass can be gained by estimating the center of mass twice. Simulations were carried out on a large-scale rigid–flexible coupling spacecraft. The results demonstrated that the determination errors in all cases are less than 10%, which meet the engineering requirements.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T12:52:29Z
      DOI: 10.1177/0954410020988466
       
  • Investigation on the influence of the structure parameters of blade tip
           recess on the performance of axial flow compressor
    • Authors: Song Yan, WuLi Chu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      As one of the important components of an aero engine, the compressor plays an important role in improving the performance of the aero engine. The blade tip recess (BTR) has great potential and advantages in improving the performance of the compressor. It is very important to clarify the influence of the structure parameters of the BTR on the performance of the compressor. In this study, the two-dimensional results of the BTR were analyzed by using the method of variance analysis, and the two-dimensional calculation results of the BTR were used to guide the design of the BTR of axial flow compressor rotor. In the NASA Rotor 35, the influence rules of the structure parameters of BTR on the recess effect that was basically the same as the two-dimensional conditions. The optimization of the rotor BTR structure parameters may be achieved by the two-dimensional calculation. The flow field analysis showed the BTR can retard the growth rate of the blockage area of the leading edge of blade tip by weakening the tip clearance leakage flow intensity that delayed the occurrence of blade tip blockage and improved the aerodynamic stability of the rotor.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T12:23:50Z
      DOI: 10.1177/0954410021998101
       
  • Investigation on lift characteristics of a double-delta wing pitching in
           various reduced frequencies
    • Authors: Tianxiang Hu, Yue Zhao, Peiqing Liu, Qiulin Qu, Hao Guo, Rinie AD Akkermans
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The unsteady lift characteristics of a double-delta wing were studied using both experimental and numerical approaches, which were also compared with a single-delta wing with the same main wing sweep angle. It was found that by increasing the reduced frequency of pitching, the hysteresis effect of lift was magnified. Moreover, in the high reduced frequency case k = 0.48, the difference between the lift coefficients of single- and double-delta wings became rather subtle. The wing surface pressure distribution results indicated the flow phenomenon of dramatic lift losses was due to the effect of lower surface suction during the wing being pitched downstroke. It was observed that, as the reduced frequency became sufficiently high, the virtual camber effect induced by pitching could dominate the flow field, which would mitigate the impact of wing geometry on the lift characteristics.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T11:56:46Z
      DOI: 10.1177/0954410021992201
       
  • Entropy and fractal perspectives of a flapping wing subjected to gust
    • Authors: Manabendra M De, Jaideep S Mathur, Sankaranarayanan Vengadesan
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Studies on entomopter’s performance under the influence of gust have received impetus in the past decade. There exists a dire need to ascertain the threshold of the frontal gusty conditions which would destabilize these anthropogenic flyers. This would help to devise methods to mitigate the detrimental effects of gust. In light of this aspect, the present study aims at analyzing the onsets of instability in a flapping wing system subjected to temporal gust by employing recurrence period density entropy (RPDE) and detrended fluctuation analysis (DFA). Simulation of the flapping wing along inclined stroke is carried out for a Reynolds number of 150. This Reynolds number lies in the typical operating regime of fruit flies and entomopters like the Pico aerial vehicle. Numerical simulations are carried out to solve the laminar, unsteady, and incompressible Navier–Stokes equations. The dynamic meshing technique is employed to model flapping kinematics. Nine gusts with a combination of frequency and velocity ratios of 0.1, 0.5, and 1.0 are considered. Instantaneous horizontal and vertical forces are estimated. Time series of these forces are analyzed using RPDE and DFA paradigms. These analyses indicate that gust frequency of an order of magnitude higher than flapping frequency and gust amplitude of the order of magnitude as the wing’s root mean square velocity induces a possible onset of instability.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T11:39:49Z
      DOI: 10.1177/0954410021993297
       
  • Improved uncertain method for safety analysis of aircraft landing gear
    • Authors: Xintian Liu, Shuanglong Geng, Xueguang Yu, Jiachi Tong, Yansong Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      There are various uncertain factors in most practical engineering applications, such as input loads, structural sizes, manufacturing tolerance, and initial and boundary conditions. The interval method and grey number theory are common methods to deal with uncertainty. In this article, the interval truncation method and grey number theory are improved. And a mixed method is proposed to represent the confidence interval of output result based on the improved interval truncation method and improved grey number theory. The proposed methods’ feasibility is verified by a stepped bar; the methods are applied to the analysis of aircraft landing gear safety uncertainty.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T11:26:12Z
      DOI: 10.1177/09544100211000582
       
  • Aeroacoustic noise calculations of non-compact bodies with permeable
           boundaries
    • Authors: Fang Wang, Qiuhong Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A hybrid computational aeroacoustic method with permeable boundary is developed to evaluate non-compact noise induced by low Mach number flow over arbitrarily shaped bodies. Based on Lighthill’s equation and the boundary element method, the unified integral equations are established in which the integral boundary surrounding the objects can be selected arbitrarily. Validation studies are developed for noise induced by two-dimensional NACA0012 airfoil and three-dimensional circular cylinder. For NACA0012 airfoil, the directivity patterns of calculated noise with different permeable boundaries agree well with Howe’s analytical solution for trailing edge model. The acoustic noise generated by circular cylinder has a good agreement with Revell’s experimental data and FW-H equation. It demonstrates that the noise predicted by different permeable boundary is as accurate as that calculated by the body surface.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T10:36:54Z
      DOI: 10.1177/0954410021996993
       
  • Early fault detection of gas turbine hot components under different
           ambient and operating conditions
    • Authors: Jiao Liu, Jinfu Liu, Daren Yu, Zhongqi Wang, Weizhong Yan, Michael Pecht
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Failure of hot components in gas turbines often causes catastrophic results. Early fault detection can prevent serious incidents and improve the availability. A novel early fault detection method of hot components is proposed in this article. Exhaust gas temperature is usually used as the indicator to detect the fault in the hot components, which is measured by several exhaust thermocouples with uniform distribution at the turbine exhaust section. The healthy hot components cause uniform exhaust gas temperature (EGT) profile, whereas the hot component faults could cause the uneven EGT profile. However, the temperature differences between different thermocouple readings are also affected by different ambient and operating conditions, and it sometimes has a greater influence on EGT than the faults. In this article, an accurate EGT model is presented to eliminate the influence of different ambient and operating conditions on EGT. Especially, the EGT profile swirl under different ambient and operating conditions is also included by considering the information of the thermocouples’ spatial correlations and the EGT profile swirl angle. Based on the developed EGT model, the detection performance of early fault detection of hot components in gas turbine is improved. The accuracy and effectiveness of the developed early fault detection method are evaluated by the real-world gas turbine data.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T09:57:51Z
      DOI: 10.1177/0954410020986890
       
  • Fault-tolerant control allocation of a flexible satellite with an
           infinite-dimensional model
    • Authors: Leila Ashayeri, Ali Doustmohammadi, Farhad Fani Saberi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Fault-tolerant control allocation (FTCA) strategy is proposed for attitude stabilization of a flexible satellite with actuator redundancy. The control scheme is based on the infinite-dimensional model of a flexible satellite with no discretization, so the spillover instability is eliminated. This is one of the important benefits of the proposed control scheme over the previous FTCA schemes that have been used for the flexible satellite. The proposed scheme contains two modules. The first module provides a virtual control law to meet stabilization and vibration control objectives in the presence of uncertainties and external disturbances. There is no need to implement in-domain actuators on panels to stabilize their vibration. In this module, the virtual control is designed using adaptive integral sliding mode approach where the sliding surface includes angular velocities, internal reaction torques, and nominal control for healthy system. The second module, based on fault/failure information and using a control allocation scheme, provides redistribution of the virtual control law among the available actuators. Due to simultaneous actuator faults and control constraints, there is an error between the actual virtual control and the designed control that affects the overall system stability. To eliminate this error, gain of the virtual control signal is adjusted by an adaptive updating law. The closed-loop system stability is guaranteed for small changes in a neighborhood of the sliding surface with simultaneous vibration damping. A numerical example illustrates the effectiveness of the proposed control strategy.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T07:01:53Z
      DOI: 10.1177/0954410021991273
       
  • Aerodynamics of a half-rotating wing in hovering flight: An integrated
           study
    • Authors: Qian Li, Jiwei Yuan, Huan Shen, Jiaguo Deng, Timothy R Jakobi, Sridhar Ravi, Xiaoyi Wang, Aihong Ji
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This study introduces a new quasi-flapping wing driving mechanism based on a half-rotating mechanism which is capable of pure rotational flapping rather than the more traditional oscillatory flapping method. Lift models for half-rotating wing (HRW) aircraft in hovering flight are proposed based on the kinematics of a HRW prototype and the flow characteristics near the surface of its wing. Alongside further analytical expressions for lift based on kinematic extractions, computational models and a novel lift validation mechanism are used to reinforce the aerodynamic characteristics of the HRW in hovering flight. The aerodynamics of the HRW are experimentally assessed for different wing layouts and wing materials. Results indicate that the flow field generated by the motion of the wing arranged symmetrically on both sides of the body interfere with each other, causing the average lift coefficient of the paired-wing HRW to be less than that of the single-wing HRW. The average lift coefficient of the flexible wing is larger than that of the rigid wing. In addition, the average lift of the flexible wing increases with increasing flexural compliance within a particular range. Lift forces in different flight conditions are calculated using derived formulas alongside representative computational models, through which the derivation of lift variation for the HRW in hovering flight is validated. The theoretical lift curves show reasonable agreement with numerical simulation results in terms of the time course over one stroke cycle. The mechanisms of the HRW for generation and shedding of vortices in hovering flight are further revealed in computed flow field characteristics results. The velocity vectors of the flow field between the HRW and the symmetrically rotating wing indicate that the HRW with asymmetric rotation can generate lift force effectively. The velocity difference between the wing and the fluid is the key factor influencing the structure of generated vortices. In detailed three-dimensional (3D) vortex flows, our computational fluid dynamics study shows that a horseshoe-shaped vortex is first generated in the early downstroke. The horseshoe-shaped vortex subsequently grows into a doughnut-shaped vortex ring, with a jet stream appearing in its core which forms the downwash. The doughnut-shaped vortex ring eventually elongates into a long arc-shaped wake vortex ring. A large increasing lift force is generated during the upstroke, most likely due to the stable distal attached vortices; and in accordance with this, downwash becomes evident in the vortex ring during the downstroke.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T06:58:33Z
      DOI: 10.1177/0954410021996545
       
  • Improvement of structural characteristics of composite thin-walled beams
           using variable stiffness concept via curvilinear fiber placement
    • Authors: Touraj Farsadi, Mirac Onur Bozkurt, Demirkan Coker, Altan Kayran
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This study presents the use of variable stiffness concept via curvilinear fiber placement to achieve improved structural characteristics in composite thin-walled beams (TWBs). The TWB used in the study is constructed in circumferentially asymmetric stiffness (CAS) configuration. The variation of fiber angles along the span and the width of the TWB is included by defining two fiber path functions. A parametric study is performed to investigate the effects of different fiber paths on the structural performance metrics including frequency spacing, unit twist, and critical buckling load. For this purpose, a semi-analytical solution method is developed to conduct free vibration, deformation, and buckling analyses of the TWB with curvilinear fibers. The semi-analytical method is validated with several finite element (FE) analyses performed using ABAQUS. Elastic stress analyses of TWBs with selected fiber paths subjected to simplified distributed loading are also conducted using the FE method, and a ply failure criterion is applied to evaluate the strength of these TWBs. Overall results show that curvilinear fiber placement varied along the span leads to greater structural performance for a composite TWB than the straight fiber configuration.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T06:52:53Z
      DOI: 10.1177/0954410020988240
       
  • Oxygen concentration distribution in a pulse detonation engine with
           nozzle–ejector combinational structures
    • Authors: Zhiwu Wang, Lisi Wei, Weifeng Qin, Zijian Liang, Kun Zhang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Pulse detonation engines (PDEs) with three different types of nozzle–straight ejector combinational structures at three different ejector positions were simulated by the unsteady 2-D axisymmetric method to understand the influence of nozzle–ejector combinational structures on the performance of PDEs. Three types of nozzles included the straight nozzle, convergent nozzle, and convergent–divergent (CD) nozzle. Three ejector positions were considered according to the ratio of the distance between the nozzle outlet and the ejector inlet to the diameter of PDEs (Δx/d). Propane was used as the fuel and air as the oxidizer. The simulation results indicated that for the PDE with the straight nozzle, it took the shortest time for high-temperature burnt gas to exhaust from the detonation tube. For the PDE with the CD nozzle, the time at which the ejector was filled with external air was the fastest. Within the time range of t = 0–10 ms, the ejected air was less than the original air in the ejector among all the nine combinational structures. The maximum ejected air was obtained with the convergent nozzle, followed by the CD nozzle, and the minimum with the straight nozzle. For certain nozzles, the maximum air was ejected at the ejector position of Δx/d = +1, followed by the ejector position of Δx/d = 0, and the minimum at the ejector position of Δx/d = −1. For the convergent nozzle–ejector combinational structure, the air ejection speed was the fastest. Oxygen concentration distribution in the PDE with the CD nozzle was more uniform along the axial direction than the other nozzles.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T06:46:11Z
      DOI: 10.1177/0954410021991284
       
  • Modeling of fuel transport in pressure-fed systems with flow passage
           opening devices
    • Authors: Seong Ho Im
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This study presents a numerical model of a pressure-fed system with flow passage opening devices (FPODs) designed for an air vehicle with a high degree of maneuverability. The FPOD is a mechanical device that connects two separate fuel reservoirs and functions as a valve allowing liquid fuel to flow while minimizing the movement of pressurizing gas from upstream fuel tanks into downstream fuel tanks. A reduced-order model for the fuel motion in an annular fuel tank was developed to configure the depth and inclination angle of the free fuel surface on the cross-sectional plane of an annular fuel tank under accelerating conditions during flight. Furthermore, a newly proposed model that reflects the dynamic characteristics of the FPOD is used to determine the fluid type that is transported through the device. A simulation example shows that the full numerical model captures changes of the fuel transport condition over time in a complete pressure-fed system of annular fuel tanks with FPODs subject to acceleration.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T06:33:10Z
      DOI: 10.1177/0954410021994996
       
  • Near-surface particle image velocimetry measurements over a yawed slender
           delta wing
    • Authors: Ilyas Karasu, Sergen Tumse, Mehmat O. Tasci, Besir Sahin, Huseyin Akilli
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this study, extensive instantaneous velocity measurements were conducted within a flow area by stereo particle image velocimetry (SPIV) to investigate the influence of the yaw angle, β, on the vortical flow structure formed on a slender delta wing. This sideslip angle, β, in the yaw plane was varied from 4° up to 20° with an interval of 4° at two critical angles of attack, α = 25° and 35°, respectively. In order to reveal the influence of the yaw angle, β over the flow structure of the delta wing, time-averaged flow statistics, and instantaneous flow data obtained by the SPIV technique in the plan-view plane close to the suction surface of the delta wing were presented. It was observed that even a low yaw angle, for instance β = 8°, becomes to be effective on the flow characteristics of the delta wing, and this effect was augmented with increasing β. The influence of β is quite high on the vortical flow structure at α= 35° compared to the angle of attack of α = 25°. The flow structure that is symmetrical with respect to the centerline of the wing in the case of no yaw has disrupted with the existence β. Furthermore, the extent of the asymmetry enlarges with increasing β. The leading-edge vortex (LEV) on the windward side broken earlier and dominated the flow on the wing surface. It is concluded that this asymmetric flow structure can deteriorate the aerodynamic performance and cause other adverse effects such as unsteady loading.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T06:10:30Z
      DOI: 10.1177/0954410021999556
       
  • Laminar separation bubble dynamics and its effects on thin airfoil
           performance during pitching-up motion
    • Authors: Zhaolin Chen, Tianhang Xiao, Yan Wang, Ning Qin
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article reports an investigation into dynamic characteristics of the laminar separation bubbles (LSBs) associated with aerodynamic loads unsteadiness of a cambered thin airfoil in pitching-up motions at low Reynolds number flows. Unsteady Reynolds-averaged Navier–Stokes (URANS) simulations were conducted for a 4%c cambered thin airfoil at Reynolds number of 30,000 and 60,000. The airfoil pitches up from 0° to 25°angles of attack at dimensionless pitch rate [math] of 0.0398 and 0.0199. The [math] SST [math] turbulence transition model was used to account for the effect of transition on LSBs’ development. The LSBs are shown to evolve in their shape and size during the pitching motion. The influence of the LSBs on the airfoil upper surface during pitching motion continues to a higher incidence in comparison with that under static conditions before developing into a fully detached flow. Vortex merging is observed in the rear part of the LSBs in the turbulent portion for a Reynolds number of 30,000. At Reynolds number 60,000, the changing of the LSB length during pitching-up motion is similar to that of steady cases, except a delayed transition is observed as incidence increases. The results show further insight into the dynamic characteristics of the LSBs and their relation to the aerodynamic performance of the airfoil.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T05:49:51Z
      DOI: 10.1177/0954410021999529
       
  • A viscous blade body force model for computational fluid dynamics–based
           throughflow analysis of axial compressors
    • Authors: Jian Li, Jinfang Teng, Mingmin Zhu, Xiaoqing Qiang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In recent years, the computational fluid dynamics (CFD) techniques have attracted enormous interest in the throughflow calculations, and one of the major concerns in the CFD-based throughflow method is the modeling of blade forces. In this article, a viscous blade force model in the CFD-based throughflow program was proposed to account for the loss generation. The throughflow code is based on the axisymmetric Navier–Stokes equations. The inviscid blade force is determined by calculating a pressure difference between the pressure and suction surfaces, and the viscous blade force is related to the local kinetic energy through a skin friction coefficient. The viscous blade force model was validated by a linear controlled diffusion airfoil cascade, and the results showed that it can correctly introduce the loss into the CFD-based throughflow model. Then, the code was applied to calculate the transonic NASA rotor 67, and the calculated results were in good agreement with the measured results, which showed that the calculated shock losses reduce the dependence of the throughflow calculation on the empirical correlation. Last, the 3.5-stage compressor P&W3S1 at 85%, 100%, and 105% of the design speed was performed to demonstrate the reliability of the viscous blade force model in a multistage environment. The results showed that the CFD-based throughflow method can easily predict the spanwise mixing due to the inclusion of the turbulence model, and predicted results were in acceptable agreement with the experimental results. In a word, the proposed viscous blade force model and CFD-based throughflow model can achieve the throughflow analysis with an acceptable level of accuracy and a little time-consuming.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T04:32:42Z
      DOI: 10.1177/0954410021999557
       
  • Pressure measurements in detonation engines
    • Authors: Vladislav S Ivanov, Sergei M Frolov, Sergei S Sergeev, Yurii M Mironov, Andrei E Novikov, Iliya I Schultz
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The influence of waveguide tubes on the signals received by remote pressure sensors measuring pressure histories in pulsed detonation engines (PDEs) and rotating detonation engines (RDEs) is studied computationally. Two types of pressure sensors are considered: low-frequency static pressure sensors and high-frequency sensors of pressure pulsations. Three approaches to solving the problem are used: based on Euler (inviscid flow), Navier–Stokes (laminar flow), and unsteady Reynolds-averaged Navier–Stokes (turbulent flow) equations. The approaches based on the inviscid and laminar flow models are shown to provide the best predictive capability. The laminar flow model is applied to analyzing the readings of pressure sensors installed remotely in the waveguide tubes attached to the hydrogen-fueled PDE, RDE, and detonation ramjet (DR). It is shown that the measurements of static pressure and pressure pulsations by remote pressure sensors do not correspond to the time-averaged mean and local instantaneous values of pressure in the combustors. The pressure time histories can be recovered based on the measurements and computational fluid dynamics calculations of the operation process. The latter is demonstrated by analyzing the results of test fires with a dual-duct DR model.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T04:19:29Z
      DOI: 10.1177/0954410021993078
       
  • Flow behavior of skewed vortex generators on a backward-facing ramp
    • Authors: Atcha-uea Cheawchan, Yiming Wen, Zhen Wei Teo, Bing Feng Ng, Tze How New
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The impact of skewness angle on the effectiveness of vortex generators (VGs) and the behavior of streamwise vortices on flow separation behind a backward-facing ramp (BFR) with a sharp transition were experimentally investigated using surface oil flow visualizations, planar and stereoscopic particle image velocimetry measurements. Counter/corotating streamwise vortices were generated by a set of boundary layer-type rectangular VG located upstream of the BFR that comprised a flat- and 30-inclined sections with different skewness angles of 10°, 20°, and 30°. Local Reynolds number based on the VG location was Rex ≈ 3 × 106. Results show that the reattachment length was reduced by ∼45% when the VG was located five times its height ahead of the transition. Additionally, the behavior of the vortex core generated by the left vane displayed strong dependence on the skewness angle, whereby its vorticity magnitude and vortex instability increase with the skewness angle. Circulation magnitude and vortex radius of the left vortex core are also observed to be physically larger and less stable. In contrast, the vortex core produced by the right vane displays opposite behavior as the skewness angle increases. Lastly, the vortex core location is observed to fluctuate more in the vertical direction than horizontal direction.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T03:26:50Z
      DOI: 10.1177/0954410021996181
       
  • L1 neural network adaptive fault-tolerant controller for unmanned aerial
           vehicle attitude control system
    • Authors: Yan Zhou, Huiying Liu, Huijuan Guo, Jing Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this article, a L1 neural network adaptive fault-tolerant controller is exploited for an unmanned aerial vehicle attitude control system in presence of nonlinear uncertainties, such as system uncertainties, external disturbances, and actuator faults. A nonlinear dynamic inversion controller with sliding mode control law is designed as the outer-loop controller to track the attitude angles quickly and accurately which reduces dependence on model accuracy. A L1 neural network adaptive controller of the inner loop is introduced to compensate the nonlinear uncertainties and have a good attitude tracking. The radial basis function neural network technique is introduced to approximate a lumped nonlinear uncertainty and guarantee the stability and transient performance of the closed-loop system, instead of converting it to a half-time linear system by the parametric linearization method. Simulation results demonstrate the effectiveness of the proposed controller.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T03:05:22Z
      DOI: 10.1177/0954410020984098
       
  • Visualization of separation and reattachment in an incident shock-induced
           interaction
    • Authors: Yun Jiao, Chengpeng Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An experimental study is conducted on the qualitative visualization of the flow field in separation and reattachment flows induced by an incident shock interaction by several techniques including shear-sensitive liquid crystal coating (SSLCC), oil flow, schlieren, and numerical simulation. The incident shock wave is generated by a wedge in a Mach 2.7 duct flow, where the strength of the interaction is varied from weak to moderate by changing the angle of attack α of the wedge from 8° and 10° to 12°. The stagnation pressure upstream was set to approximately 607.9 kPa. The SSLCC technique was used to visualize the surface flow characteristics and analyze the surface shear stress fields induced by the initial incident shock wave over the bottom wall and sidewall experimentally which resolution is 3500 × 200 pixels, and the numerical simulation was also performed as the supplement for a clearer understanding to the flow field. As a result, surface shear stress over the bottom wall was visualized qualitatively by SSLCC images, and flow features such as separation/reattachment and the variations of position/size of separation bubble with wedge angle were successfully distinguished. Furthermore, analysis of shear stress trend over the bottom wall by a hue value curve indicated that the relative magnitude of shear stress increased significantly downstream of the separation bubble compared with that upstream. The variation trend of shear stress was consistent with the numerical simulation results, and the error of separation position was less than 2 mm. Finally, the three-dimensional schematic of incident shock-induced interaction has been achieved by qualitative summary by multiple techniques, including SSLCC, oil flow, schlieren, and numerical simulation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T02:41:42Z
      DOI: 10.1177/0954410020983495
       
  • Fast cooperative angular trajectory planning for multiple on-orbit service
           spacecraft based on the Bezier shape-based method
    • Authors: Zichen Fan, Mingying Huo, Ji Qi, Song Xu, Kang Sun, Naiming Qi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this article, a fast cooperative obstacle avoidance angular trajectory planning method is presented for multiple on-orbit service (OOS) spacecraft, which can obtain the spatial angular trajectories of multiple spacecraft in a considerably short time. For multiple OOS spacecraft equipped with the manipulator, under the constraints of the maximum control moment, multiple obstacles, dynamics, and geometry, the Bezier shape-based (SB) method is used to rapidly generate the cooperative obstacle avoidance attitude angular trajectories of all OOS spacecraft at the same time. The superiority of the proposed method is proved by comparing the results of Bezier SB method and the finite Fourier series (FFS) SB method under the same initial conditions. To further verify the effectiveness of the proposed method, the results of the Bezier method are substituted into the Gauss pseudospectral method (GPM) as the initial values to further optimize the angular trajectories in the simulation. Simulation results show that compared with the FFS SB method, the Bezier SB method can use shorter computation time to get the results of spacecraft attitude movement faster; compared with GPM, Bezier SB method only uses about 2% of computation time to get the results of movement time difference of only 2%. The method proposed in this article can be applied to complex space OOS missions, which requires multiple OOS spacecraft to rapidly realize the space obstacle avoidance attitude maneuver to complete cooperative operation missions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T02:13:41Z
      DOI: 10.1177/0954410021996992
       
  • Spacecraft attitude determination using electrostatically suspended
           gyroscope with rotating rings
    • Authors: Habib Ghanbarpourasl
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article presents a novel method for satellite attitude determination, using two electrostatically suspended gyroscopes (ESGs). In this method, two superconducting rings (like gimbals) rotated around each suspended spherical rotor, which causes a variable flux in the rings. A voltage will be induced in the rings according to Faraday’s law of induction. Satellite attitude is determined by integrating of the induced voltages and strap-down rate gyros using the extended Kalman filter. This gyroscope works precisely when the rotor is maintained at the center of the gyroscope cavity, between three pairs of the circular electrodes, and then a backstepping control algorithm is designed for this purpose. The performance of attitude determination using such sensor and designed control algorithm is evaluated by six-degree-of-freedom simulation of a satellite in the MATLAB software. The simulations show the proposed algorithm works in an excellent manner and has lower than a few hundred arc second errors. The accurately aligning of rotors in the known direction to space is essential and is another problem that we do not address here.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-11T01:54:18Z
      DOI: 10.1177/0954410021996152
       
  • Life Quality Index–based cost–benefit analysis of equipping life
           preservers aboard airplanes
    • Authors: Rui-liang Yang, Li-bin Yang, Li-jing Wang, Sha Li, Dong-han Geng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The life preservers aboard airplanes play a critical role in ensuring occupant safety in water-related accidents. However, several airlines off-load life preservers to save fuel and costs. In this study, a cost–benefit analysis was performed considering the Life Quality Index to examine the necessity of life preservers aboard aircraft. It was noted that the placement of life preservers aboard airplanes was reasonable and beneficial in the recent 15 years. Although life preservers are primarily required for extended overwater (EOW) operations, the distance from the shoreline for most of the water-related accidents was considerably smaller than that of an EOW operation, and most water-related accidents occurred close to an airport. In other words, air passengers were at risk of water-related accidents, regardless of whether the flight was classified as an EOW flight. Thus, life preservers must be made available for all the occupants on all passenger flights, regardless of the flight path.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T12:53:12Z
      DOI: 10.1177/0954410021993293
       
  • On modeling and dynamics of a multiple launch rocket system
    • Authors: Bo Li, Xiaoting Rui, Guoping Wang, Jianshu Zhang, Qinbo Zhou
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Dynamics analysis is currently a key technique to fully understand the dynamic characteristics of sophisticated mechanical systems because it is a prerequisite for dynamic design and control studies. In this study, a dynamics analysis problem for a multiple launch rocket system (MLRS) is developed. We particularly focus on the deductions of equations governing the motion of the MLRS without rockets by using a transfer matrix method for multibody systems and the motion of rockets via the Newton–Euler method. By combining the two equations, the differential equations of the MLRS are obtained. The complete process of the rockets’ ignition, movement in the barrels, airborne flight, and landing is numerically simulated via the Monte Carlo stochastic method. An experiment is implemented to validate the proposed model and the corresponding numerical results.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T11:46:00Z
      DOI: 10.1177/0954410020982243
       
  • An attitude updating algorithm using Picard iteration and higher degree
           polynomial
    • Authors: Xiaole Guo, Xixiang Liu, Miaomiao Zhao, Jie Yan, Wenqiang Yang, Xu Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      To accurately track body attitude under high dynamic environments, a new attitude updating algorithm for the strapdown inertial navigation system is proposed after further applying higher degree polynomial to the quaternion Picard iteration (QPI) algorithm. With QPI, calculation error introduced by Picard approximation can be eliminated, but the angular rate fitting error introduced by substituting polynomial for angular rate of body will still affect the accuracy of the attitude updating algorithms which are designed based on polynomial model. Hence, a five- rather three-degree polynomial constructing method using four samples of gyro outputs with coning motion constrain is designed and tested. Simulation results indicate the proposed method owns more accuracy than QPI, optimal coning algorithm, and Fourth4Rot under both low and high dynamic environments.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T05:35:27Z
      DOI: 10.1177/0954410020986631
       
  • Numerical study on the airfoil self-noise of three owl-based wings with
           the trailing-edge serrations
    • Authors: Dian Li, Xiaomin Liu, Lei Wang, Fujia Hu, Guang Xi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Previous publications have summarized that three special morphological structures of owl wing could reduce aerodynamic noise under low Reynolds number flows effectively. However, the coupling noise-reduction mechanism of bionic airfoil with trailing-edge serrations is poorly understood. Furthermore, while the bionic airfoil extracted from natural owl wing shows remarkable noise-reduction characteristics, the shape of the owl-based airfoils reconstructed by different researchers has some differences, which leads to diversity in the potential noise-reduction mechanisms. In this article, three kinds of owl-based airfoils with trailing-edge serrations are investigated to reveal the potential noise-reduction mechanisms, and a clean airfoil based on barn owl is utilized as a reference to make a comparison. The instantaneous flow field and sound field around the three-dimensional serrated airfoils are simulated by using incompressible large eddy simulation coupled with the FW-H equation. The results of unsteady flow field show that the flow field of Owl B exhibits stronger and wider-scale turbulent velocity fluctuation than that of other airfoils, which may be the potential reason for the greater noise generation of Owl B. The scale and magnitude of alternating mean convective velocity distribution dominates the noise-reduction effect of trailing-edge serrations. The noise-reduction characteristic of Owl C outperforms that of Barn owl, which suggests that the trailing-edge serrations can suppress vortex shedding noise of flow field effectively. The trailing-edge serrations mainly suppress the low-frequency noise of the airfoil. The trailing-edge serration can suppress turbulent noise by weakening pressure fluctuation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T05:25:48Z
      DOI: 10.1177/0954410020988229
       
  • Global flow visualization of transonic cavity flow with various yaw angles
    • Authors: Chih-Yung Huang, Chen-Yen Yeh, Yun-Fang Lin, Kung-Ming Chung
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This study experimentally investigated transonic cavity flows with different length-to-depth (L/h) ratios and yaw angles. Two rectangular models with L/h = 6.14 and 21.5 were examined with yaw angles of 10°, 30°, and 45° under a flow of Mach 0.83. The flow was visualized using pressure-sensitive paint (PSP) to obtain the detailed pressure distribution inside the cavity models. The acquired PSP data were compared with experimental data measured using Kulite transducers, and these data showed favorable agreement. Gradual pressure increases inside the cavity model with L/h = 6.14 were observed from the PSP measurements as open cavity flow. The flow impingement at the bottom of the cavity and the significant pressure rise inside the cavity model with L/h = 21.5 were observed as closed cavity flow. The present study quantitatively visualized the evolution of the pressure distribution from symmetric to asymmetric for different yaw angles using porous PSP sensors.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T05:11:26Z
      DOI: 10.1177/0954410020984086
       
  • Aerodynamic design optimization of a hypersonic rocket sled deflector
           using the free-form deformation technique
    • Authors: Tianjiao Dang, Bingfei Li, Dike Hu, Yachuan Sun, Zhen Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An aerodynamic design optimization of a hypersonic rocket sled deflector is presented using the free-form deformation (FFD) technique. The objective is to optimize the aerodynamic shape of the hypersonic rocket sled deflector to increase its negative lift and enhance the motion stability of the rocket sled. The FFD technique is selected as the aerodynamic shape parameterization method, and the continuous adjoint method based on the gradient method is used to search the optimization in the geometric shape parameter space; the computational fluid dynamics method for a hypersonic rocket sled is employed. An automatic design optimization method for the deflector is carried out based on the aerodynamic requirements of the rocket sled. The optimization results show that the optimized deflector meets the design requirement of increasing the negative lift under the constraint of drag. By improving the pressure distribution on the surface of the deflector, the negative lift is increased by 7.39%, which confirms the effectiveness of the proposed method.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T05:01:27Z
      DOI: 10.1177/0954410021994984
       
  • Showerhead film cooling injection orientation design on the turbine vane
           leading edge considering representative lean burn combustor outflow
    • Authors: Zhuang Wu, Hui-ren Zhu, Cun-liang Liu, Lin Li, Xu-yang Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The heat transfer performance of showerhead film cooling on the vane leading edge was numerically investigated considering representative lean burn combustor swirling outflow. Three cases with different inflow conditions (uniform inflow, positive swirling inflow, and negative swirling inflow) and three cases with different film injection angles (45°, 90°, and 135°) were studied. As the first study to explore the showerhead film design principle under swirling inflow, a newly designed asymmetrical counter-inclined (45° and 135°) film cooling was also proposed. To examine the design principles, the cooling effectiveness, heat transfer augmentation, and heat flux reduction of the newly designed asymmetrical case were evaluated compared with the traditional symmetrical case. The results show that the swirling inflow introduces obvious radial pressure gradient on the vane. The radial pressure gradient is the key influence factor to deflect the coolant migration, decrease the cooling effect, and degrade the homogeneity. The film with opposite orientation to the radial pressure gradient can weaken the deflect effect. The radial pressure gradient direction differs in different regions, making it impossible for the film with congruent injection orientation to simultaneously resist the pressure gradient on the entire vane. For the new design, the boundary line of the counter-inclined holes is consistent with the twisted stagnation line to guarantee that the injection orientation of all the film holes is opposite to the radial pressure gradient. As expected, the new design can effectively weaken the deflection effect and show uniform film distribution. The higher coolant mass ratio provides more obvious enhancement effect. At coolant mass ratio 3.71% and 4.56%, the overall area-averaged heat flux reduction (Δq) is increased by 0.311 and 0.576, and the overall area-averaged relative standard deviation is reduced by 12.17 and 11.66 compared with the traditional design. The results have confirmed the adaptability of the film design principle under swirling inflow.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T04:49:41Z
      DOI: 10.1177/0954410021996566
       
  • Experimental and numerical investigation of a supersonic inlet with double
           S-bend diffuser
    • Authors: Jinsheng Zhang, Huacheng Yuan, Yunfei Wang, Guoping Huang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Design of a supersonic inlet with double S-bend diffuser was developed. Numerical simulation and wind tunnel experiment were carried out to investigate the aerodynamic performance and variable geometric rules of the inlet. The result indicates that the variable geometry scheme adopted solves the contradiction between starting performance at low Mach number and aerodynamic performance at high Mach number. The inlet works normally and stably over a wide speed range. At design point, the total pressure recovery coefficient reaches 0.47. In addition, two different kinds of inlets with double S-bend diffuser and single S-bend diffuser were studied. Compared with the double S-bend diffuser, the total pressure recovery coefficient of the single S-bend diffuser is higher at low Mach number (Ma0 < 3) and lower at high Mach number (Ma0> 3). With the increase of backpressure, shock train mainly moves upstream along the low-energy flow region in the diffuser. For the double S-bend diffuser, shock train will first move along the lower corner and then along the upper corner. For the single S-bend diffuser, it will only move along the upper corner. The strong secondary flow of the double S-bend is the main reason for the above phenomenon.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T04:14:30Z
      DOI: 10.1177/0954410020983981
       
  • Effect of hub clearance of cantilever stator on aerodynamic performance
           and flow field of a transonic axial-flow compressor
    • Authors: Botao Zhang, Bo Liu, Xiaochen Mao, Hejian Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      To investigate the effect of hub clearance of cantilever stator on the aerodynamic performance and the flow field of the transonic axial-flow compressor, the performance of single-stage compressors with the shrouded stator and cantilever stator was studied numerically. It is found that the hub corner separation on the stator blade suction surface (SS) was modified by introducing the hub leakage flow. The separation vortex on the SS of the stator blade root at about 10% axial chord length caused by the interaction of the shock wave and boundary layer was also controlled. Compared with the tip clearance size of the rotor blade, the stator hub clearance size (HCS) has a much less effect on the overall aerodynamic performance of the compressor, and there is no obvious effect on the flow field in the upstream blade row. With the increase of HCS, the leakage loss and the blockage degree in the flow field near the stator hub are increased and further make the adiabatic efficiency and the total pressure ratio of the compressor gradually decrease. Meanwhile, the stall margin of the compressor was changed slightly, but the response of the stall margin to the change of the HCS is nonlinear and insensitive. The stator hub leakage flow (HLF) can not only change the flow field near the hub but also redistribute the flow law within the range of the entire blade span. It will contribute to further understand the mechanism of the HLF and provide supports for the design of the cantilever stator of transonic compressors.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T04:00:49Z
      DOI: 10.1177/0954410021993700
       
  • An algorithm of pretrained fuzzy actor–critic learning applying in
           fixed-time space differential game
    • Authors: Xiao Wang, Peng Shi, Howard Schwartz, Yushan Zhao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Solving space differential game in an unknown environment remains a challenging problem. This article proposes a pretrained fuzzy actor–critic learning algorithm for dealing with the space pursuit-evasion game in fixed time. It is supposed that the research objects are two agents including one pursuer and one evader in space. A virtual environment, which is defined as the known part of the real environment, is utilized for deriving optimal strategies of the pursuer and the evader, respectively. Through employing the fuzzy inference system, a pretrained process, which is based on the genetic algorithm, is designed to obtain the initial consequent set of the pursuer and the evader. Besides, an actor–critic framework is applied to finely learn the suitable consequent set of the pursuer and evader in the real environment. Numerical experimental results validate the effectiveness of the proposed algorithms on improving the ability of the agents to adapt to the real environment.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:56:25Z
      DOI: 10.1177/0954410021992439
       
  • A path planning model of a tiltrotor for approaching an aircraft carrier
           during landing
    • Authors: Yu Wu, Haixu Li, Xichao Su
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A path planning model concerning a tiltrotor approaching an aircraft carrier is established in this study. In the model, the characteristic of the tiltrotor, the landing task, and the environment of the carrier are taken into account. First, the motion equations and the maneuverability of the tiltrotor in each flight mode are presented, and the constraints of control variables and flight envelope are given. The returning flight of the tiltrotor is divided into three phases corresponding to the three flight modes of the tiltrotor, and the constraints in each phase and the goal are set. Considering the flight safety of the tiltrotor, the environment of the carrier is described as flyable space and no-fly zones, and the no-fly zones are set taking the influences of turbulence and wind field induced by the moving aircraft carrier into account. The path planning issue is formulated into an optimization problem under the constraints of control variables and state variables. According to the characteristic of the established model, a pigeon inspired optimization (PIO)-based path planning algorithm is developed integrating the “step-by-step” and “one effort” path search strategies. Simulation results demonstrate that the tiltrotor can reach the target point with a reasonable landing path. Comparison among different algorithms is also conducted to verify that the PIO algorithm is capable of solving this online path planning problem.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:49:24Z
      DOI: 10.1177/0954410021996491
       
  • Effectiveness of multiple jets for a finned large slenderness ratio
           missile in supersonic crossflows
    • Authors: Longfei Li, Jiangfeng Wang, Ding Wang, Tianpeng Yang, Jiawei Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The reaction control system with multiple lateral jets shows great advantages in agility and maneuverability for supersonic air vehicles. Interactions among sonic jet plumes, X-shape fins, and supersonic crossflow at Mach 4.5 and Reynolds number 3.8 × 107 are numerically studied considering different number of jets for a large slenderness ratio missile with 7 jet exits. Three-dimensional Reynolds-averaged Navier–Stokes equations closed by Spalart–Allmaras turbulence model for the structured grid are validated and solved. The overall force and moment amplification factors of configurations with and without fins are analyzed and compared. Moreover, the force and moment amplification factors on fins and ratio of force and moment on fins are proposed and discussed to measure the jet effectiveness contributed from fins. The number of jet plumes is under consideration for all cases. Results show that the increment of effectiveness decreases as the number of jets increases for the finned configuration. Fins can significantly improve the jet effectiveness with more than 70% force and 50% moment increment, which shows great advantages to the jet effectiveness as well as the overall aerodynamic performance.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:39:05Z
      DOI: 10.1177/0954410020986898
       
  • Design of finite-time guidance law based on observer and head-pursuit
           theory
    • Authors: Chenqi Zhu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In order to improve the guiding accuracy in intercepting the hypersonic vehicle, this article presents a finite-time guidance law based on the observer and head-pursuit theory. First, based on a two-dimensional model between the interceptor and target, this study applies the fast power reaching law to head-pursuit guidance law so that it can alleviate the chattering phenomenon and ensure the convergence speed. Second, target maneuvers are considered as system disturbances, and the head-pursuit guidance law based on an observer is proposed. Furthermore, this method is extended to a three-dimensional case. Finally, comparative simulation results further verify the superiority of the guidance laws designed in this article.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:34:03Z
      DOI: 10.1177/0954410020984562
       
  • Evaluation of pilot and quadcopter performance from open-loop
           mission-oriented flight testing
    • Authors: Muhammad Junayed Hasan Zahed, Travis Fields
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Unmanned aircraft systems (UAS) have experienced tremendous growth through both commercial (i.e., toys and videography) and defense avenues. The rapid expansion, particularly in the consumer market, has outpaced regulatory bodies. Certification to commercially operate such vehicles currently requires the successful completion of a knowledge examination, without the need to physically operate a vehicle. The focus of the work presented herein is on quantifying the pilot and multi-rotor performance in an attempt to provide quantitative metrics that can be used to establish training and certification for pilots and aircraft. Test pilots were categorized based on their experience level, and the quadrotor unmanned aircraft was categorized based on the flight control mode. Cross-track command (CTC) and path error (PE) were calculated as potential time-domain metrics to quantify pilot and quadcopter performance. Individual binary logistic regression models were developed to identify the pilot experience level (PEL) and UAV control level (UCL) from the decision tree outcomes. A verification test case was included to evaluate the established regression models. Results show that the models can evaluate pilot and quadcopter performance individually, which can be used to develop the pilot training curriculum and/or evaluate pilot effectiveness in specific flight scenarios.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:27:41Z
      DOI: 10.1177/0954410020985987
       
  • Robust tracking for hypersonic vehicles subjected to mismatched
           uncertainties via fixed-time sliding mode control
    • Authors: Jianguo Guo, Shengjiang Yang, Zongyi Guo
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article investigates the robust tracking issue for the longitudinal dynamics of hypersonic vehicles subjected to mismatched uncertainties, and a novel sliding mode control approach is proposed to achieve the fixed-time convergence of tracking errors and satisfactory robustness against mismatched uncertainties. Establishing the control-oriented hypersonic vehicle model as velocity and altitude subsystems with mismatched uncertainties, the article introduces the nonlinear finite-time disturbance observer technique to estimate the uncertainties precisely. With the estimated uncertainties from the observer, the fixed-time sliding mode control is presented to track the velocity and altitude references. Consequently, the effect of the mismatched disturbances can be eliminated and the tracking performance can be improved. The stability of the closed-loop system is also analyzed. Numerical simulation results demonstrate the validity and superiority of the proposed control.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:22:37Z
      DOI: 10.1177/0954410021990239
       
  • Trajectory shaping guidance law design using constraint-combining
           multiplier
    • Authors: Min-Guk Seo, Chang-Hun Lee, Tae-Hun Kim
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A new design method for trajectory shaping guidance laws with the impact angle constraint is proposed in this study. The basic idea is that the multiplier introduced to combine the equations for the terminal constraints is used to shape a flight trajectory as desired. To this end, the general form of impact angle control guidance (IACG) is first derived as a function of an arbitrary constraint-combining multiplier using the optimal control. We reveal that the constraint-combining multiplier satisfying the kinematics can be expressed as a function of state variables. From this result, the constraint-combining multiplier to achieve a desired trajectory can be obtained. Accordingly, when the desired trajectory is designed to satisfy the terminal constraints, the proposed method directly can provide a closed form of IACG laws that can achieve the desired trajectory. The potential significance of the proposed result is that various trajectory shaping IACG laws that can cope with various guidance goals can be readily determined compared to existing approaches. In this study, several examples are shown to validate the proposed method. The results also indicate that previous IACG laws belong to the subset of the proposed result. Finally, the characteristics of the proposed guidance laws are analyzed through numerical simulations.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-03-10T03:16:37Z
      DOI: 10.1177/0954410020986371
       
  • Numerical investigation of wing–wing interaction and its effect on the
           aerodynamic force of a hovering dragonfly
    • Authors: Prafulla Kumar Swain, Siva Prasad Dora, Suryanarayana Murthy Battula, Ashok K Barik
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The present research focuses on the timing of wing–wing interaction that benefits the aerodynamic force of a dragonfly in hovering flight at Reynolds number 1350. A 3-D numerical simulation method, called the system coupling, was utilised by implementing a two-way coupling between the transient structural and flow analysis. We further explore the aerodynamic forces produced at different phase angles on the forewing and hindwing during the hovering flight condition of a dragonfly. A pair of dragonfly wings is simulated to obtain the force generated during flapping at a 60° inclination stroke plane angle with respect to the horizontal. The hovering flight is simulated by varying the phase angle and the inter-distance between the two wings. We observe a significant enhancement in the lift (16%) of the hindwing when it flaps in-phase with the forewing and closer to the forewing, maintaining an inter-wing distance of 1.2 cm (where centimetre is the mean chord length). However, for the same condition, the lift of the hindwing reduces by 9% when the wings are out of phase/counterstroke flapping. These benefits and drawbacks are dependent on the timing of the interactions between the forewing and hindwing. The time of interaction of wake capture, wing–wing interaction, dipole structure and development of root vortex are examined by 2-D vorticity of the flow field and isosurface of the 3-D model dragonfly. From the isosurface, we found that the root vortex elicited at the root of the hindwing in counter-flapping creates an obstacle for the shedding of wake vortices, which results in reduction of vertical lift during the upstroke of flapping. Hence, at the supination stage, a dragonfly uses a high rotation angle for the hovering flight mode. It is observed that the system coupling method was found to be more efficient and exhibited better performance. The present numerical methodology shows a very close match to the previously reported results.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-02-26T04:41:33Z
      DOI: 10.1177/0954410020982109
       
  • The inlet flow structure of a conceptual open-nose supersonic drone
    • Authors: Eiman B Saheby, Xing Shen, Anthony P Hays, Zhang Jun
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This study describes the aerodynamic efficiency of a forebody–inlet configuration and computational investigation of a drone system, capable of sustainable supersonic cruising at Mach 1.60. Because the whole drone configuration is formed around the induction system and the design is highly interrelated to the flow structure of forebody and inlet efficiency, analysis of this section and understanding its flow pattern is necessary before any progress in design phases. The compression surface is designed analytically using oblique shock patterns, which results in a low drag forebody. To study the concept, two inlet–forebody geometries are considered for Computational Fluid Dynamic simulation using ANSYS Fluent code. The supersonic and subsonic performance, effects of angle of attack, sideslip, and duct geometries on the propulsive efficiency of the concept are studied by solving the three-dimensional Navier–Stokes equations in structured cell domains. Comparing the results with the available data from other sources indicates that the aerodynamic efficiency of the concept is acceptable at supersonic and transonic regimes.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-02-26T03:16:29Z
      DOI: 10.1177/0954410020983043
       
  • Effects of Al, Zn, and rare earth elements on flammability of magnesium
           alloys subjected to sonic burner–generated flame by Federal Aviation
           Administration standards
    • Authors: Guangjian Wang, Zhiwei Zhao, Song Zhang, Lili Zheng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Mg alloys are promising structural materials in aerospace industry due to high strength to weight ratio. However, most Mg alloys are limited in aircraft cabins due to their susceptibility to ignition and burning. To improve fire resistance, adding alloying elements is a strategy. Thus, the goal of this study is to explore the effects of alloying elements Al, Zn, and rare earths on Mg-alloy flammability by experiment, using the system and procedures in compliance with the Federal Aviation Administration (FAA) standards for Mg-alloy flammability test. Six commercial Mg alloys with different alloying elements (AZ91E, ZK61A, ZE63A, EZ33A, WE43B, and EV31A) were tested. Results indicate that Mg alloys with Al or Zn elements were of short ignition time and high weight loss. With rare earths, Mg-alloy flammability was suppressed obviously. It appears that this suppression effect with rare earth addition was attributed to the formation of protective oxide film on the surface of molten alloy. Further, a heat transfer model was established to analyze the temperature evolution of the test specimen subjected to the sonic burner–generated flame by FAA standards, and ignition temperatures of all testing Mg alloys were predicted based on the experimental ignition time. The predicted results confirm that with rare earths addition, ignition was delayed after melting by the protective oxide film formed on the surface of the molten alloy.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2021-02-26T02:58:30Z
      DOI: 10.1177/0954410020987758
       
  • A flight mechanics-based justification of the unique range of Strouhal
           numbers for avian cruising flight
    • Authors: Diganta Bhattacharjee, Kamesh Subbarao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, analytical expressions for cycle-averaged aerodynamic forces generated by flapping wings are derived using a force model and flapping kinematics suitable for the forward flight of avian creatures. A strip theory-based formulation is proposed and the analytical expressions are found as functions of the amplitude of twist profile, mean twist angle, the flow separation point on the upper surfaces of the wings, and Strouhal number. Numerical results are obtained for a rectangular planform as well as for a representative avian wing planform. Utilizing these results, it is shown that there exists a narrow Strouhal number range where cycle-averaged net thrust, lift, and lift to drag ratio are optimal for a given flow pattern over the upper surfaces of the wings. This narrow Strouhal number range, found to be between 0.1 and 0.3, corresponds to the cruising range for a large number of avian creatures, as documented in current literature. An explanation, based on force constraints and local optimization in aerodynamic force generation, is provided for the unique range of Strouhal numbers utilized in avian cruising flight. The results and the approach outlined in the paper can be utilized to design efficient bio-inspired flapping vehicles.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-23T07:39:04Z
      DOI: 10.1177/0954410020976597
       
  • Neural network adaptive backstepping fault tolerant control for unmanned
           airships with multi-vectored thrusters
    • Authors: Shiqian Liu, James F Whidborne
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper presents the fault tolerant control (FTC) of an unmanned airship with multiple vectored thrusters in the presence of model parameter uncertainties and unknown wind disturbances. A fault tolerant control based on constrained adaptive backstepping (CAB) approach, combined with a radial basis function neural network (RBFNN) approximation, is proposed for the airship with thruster faults. A wind observer is designed to estimate the bounded wind disturbances. An adaptive fault estimator is proposed to estimate the unknown actuator faults. A weighted pseudo inverse based control allocation is incorporated to reconstruct and optimize the practical control inputs of the failed airship under constraints of actuator saturation. Rigorous stability analysis shows that trajectory tracking errors of the airship position and attitude converge to the desired set through Lyapunov theory. Numerical simulations demonstrate the fault tolerant trajectory tracking capability of the proposed NN-CAB controller under the actuator faults, even in the presence of aerodynamic coefficient uncertainties, and unknown wind disturbances.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-22T08:29:54Z
      DOI: 10.1177/0954410020976611
       
  • Investigation of infrared dim and small target detection algorithm based
           on the visual saliency feature
    • Authors: Shaoyi Li, Xiaotian Wang, Xi Yang, Kai Zhang, Saisai Niu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Infrared dim and small target detection has an important role in the infrared thermal imaging seeker, infrared search and tracking system, space-based infrared system and other applications. Inspired by human visual system (HVS), based on the fusion of significant features of targets, the present study proposes an infrared dim and small target detection algorithm for complex backgrounds. Firstly, in order to calculate the target saliency map, the proposed algorithm initially uses the difference of Gaussian (DoG) and the contourlet filters for the preprocessing and fusion, respectively. Then the multi-scale improved local contrast measure (ILCM) method is applied to obtain the interested target area, effectively suppress the background clutter and improve the target signal-to-clutter ratio. Secondly, the optical flow method is used to estimate motion regions in the saliency map, which matches with the interested target region to achieve the initial target screening. Finally, in order to reduce the false alarm rate, forward pipeline filtering and optical flow estimation information are applied to achieve the multi-frame target recognition and achieve continuous detection of dim and small targets in image sequences. Experimental results show that compared with the conventional Tophat (TOP-HAT) and ILCM algorithms, the proposed algorithm can achieve stable, continuous and adaptive target detection for multiple backgrounds. The area under curve (AUC) and the harmonic average measure F1 are used to measure the overall performance and optimal performance of the target detection effect. For sky, sea and ground backgrounds, the test results of proposed algorithm for most sequences are 1. It is concluded that the proposed algorithm significantly improves the detection effect.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-22T08:00:44Z
      DOI: 10.1177/0954410020980955
       
  • Experimental study on soft PSD material of dual pulse solid rocket motor
    • Authors: Chunguang Wang, Weiping Tian, Liwu Wang, Guiyang Xu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In order to study the failure reason of the soft PSD in the dual pulse solid rocket motor (SRM), the deformation process of the intermediate section of the second pulse combustion chamber was simplified to the two-dimensional plane strain state, and the calculation method of the circumferential strain of the soft PSD was obtained. The influencing factors of the circumferential strain of the soft PSD were studied. The main factors affecting the circumferential strain of the soft PSD are the gap between the soft PSD and the propellant grain, and the circumferential strain on the inner surface of the propellant grain. The calculation method could be used to initially estimate the circumferential strain of the soft PSD, and then predict the rationality and feasibility of the design scheme. The apparent morphology and area change rate of EPDM materials of PSD under different strains were studied by DIC tensile test. The variation of the porosity of the EPDM material with the increase of strain was obtained by micro-CT. By comparing the SEM results of the fracture and the slit of the tensile test piece, the failure mode of the EPDM material of PSDs was determined, and the failure mechanism of the PSD structure was obtained. The conclusions obtained in this paper can provide a useful reference for the design of the PSD in dual pulse SRM.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-21T03:17:18Z
      DOI: 10.1177/0954410020980958
       
  • Moving-horizon-estimator-integrated adaptive hierarchical sliding mode
           control for flexible hypersonic vehicles considering aeroservoelastic
           effect
    • Authors: Erkang Chen, Wuxing Jing, Changsheng Gao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In order to solve the attitude control problem of flexible hypersonic vehicles with consideration of aeroservoelastic effect, uncertainty and external disturbance, a novel moving-horizon-estimator-integrated adaptive hierarchical sliding mode control scheme is presented in this paper. First, the measurement model considering flexibility is established and the influence of aeroservoelastic effect on system stability is analyzed. Then moving horizon estimator is developed to reconstruct full state information from sensor measurements, while sliding mode disturbance observer and gain adaptation law is proposed to enhance the robustness and attenuate the chattering. Via combining moving horizon estimator, sliding mode disturbance observer, gain adaptation law and baseline hierarchical sliding mode controller, the moving-horizon-estimator-integrated adaptive hierarchical sliding mode control scheme that is able to achieve the control objective of both precise attitude control and active flexible vibration suppression is developed. Finally, Lyapunov theory is used to prove the stability of the proposed control scheme, and the numerical simulations are carried out, which further verify the effectiveness of the proposed control scheme against aeroservoelastic effect, uncertainty and external disturbance.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-18T06:20:31Z
      DOI: 10.1177/0954410020977553
       
  • Comprehensive design of an oleo-pneumatic nose landing gear strut
    • Authors: Muhammad Ayaz Ahmad, Syed Irtiza Ali Shah, Taimur Ali Shams, Ali Javed, Syed Tauqeer ul Islam Rizvi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A comprehensive design cycle of a nose landing gear strut having an oleo-pneumatic shock absorber for a lightweight aircraft is proposed. Design and analysis of a retractable nose landing gear according to Airworthiness Standards FAA FAR Part 23 have been carried out. This research is focused on mathematical modeling of an oleo-pneumatic strut with an analytical solution of design variables at static and dynamic loading conditions. The variation of spring and damping characteristics of an oleo-pneumatic shock absorber with the stroke length is also presented. Feasibility of equivalent mechanical spring and damper along with comparison of pneumatic as air spring and oleo as hydraulic damper is studied. Numerical integration technique was used to solve the dynamic model of an oleo-pneumatic strut with forcing function of an impact force during touch down scenario. Energy conservation principle was used to determine height required for drop tests. Parametric study of anteversion angle within the constraints of ground clearance and volume in the fuselage determined an optimized angle of nose landing gear strut. Based on the maximum pressure and impact force encountered during landing, the hydraulic cylinder and piston design was finalized. In order to validate the proposed design cycle for preliminary phase, a structural integrity of cylinder and piston assembly was carried out using finite element analysis. Deformation, maximum stresses and factor of safety validated the proposed design cycle of a nose landing gear strut specific to a general aviation aircraft having all up mass of 1600 Kg.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-14T07:05:28Z
      DOI: 10.1177/0954410020979378
       
  • CFD-CSD method for coupled rotor-fuselage vibration analysis with free
           wake-panel coupled model
    • Authors: Siwen Wang, Jinglong Han, Haiwei Yun, Xiaomao Chen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An efficient comprehensive vibration analysis method for a helicopter rotor–fuselage coupling system is presented. This loose computational fluid dynamics (CFD)/computational structural dynamics (CSD) coupling approach with a free wake–panel coupled model is used for system vibration response analysis. The CSD model of the helicopter consists of a fuselage model using a refined three dimensional (3 D) finite element model (FEM) and a rotor model consisting of nonlinear moderate deflection beam elements with 15 degrees of freedom. The unsteady Euler CFD solver is used for the flow field analysis of the entire vehicle. The induced inflow of the quasi-steady aerodynamic force is calculated with the free wake–panel coupled model, which is used to simulate rotor–fuselage aerodynamic interference. Using a full-scale helicopter as an example, the vibration responses of the typical fuselage position in hovering and level flights are analysed. When compared with the literature results and flight test data, the predictions of the proposed method are closer to the test data than those of the traditional method in hovering and low forward ratio flights, and the difference between the two methods is minimal in high forward ratio flight.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-11T11:25:44Z
      DOI: 10.1177/0954410020976512
       
  • Numerical study of the flow over the modified simple frigate shape
    • Authors: Tong Li, Yibin Wang, Ning Zhao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The simple frigate shape (SFS) as defined by The Technical Co-operative Program (TTCP), is a simplified model of the frigate, which helps to investigate the basic flow fields of a frigate. In this paper, the flow fields of the different modified SFS models, consisting of a bluff body superstructure and the deck, were numerically studied. A parametric study was conducted by varying both the superstructure length L and width B to investigate the recirculation zone behind the hangar. The size and the position of the recirculation zones were compared between different models. The numerical simulation results show that the size and the location of the recirculation zone are significantly affected by the superstructure length and width. The results obtained by Reynolds-averaged Navier-Stokes method were also compared well with both the time averaged Improved Delayed Detached-Eddy Simulation results and the experimental data. In addition, by varying the model size and inflow velocity, various flow fields were numerically studied, which indicated that the changing of Reynolds number has tiny effect on the variation of the dimensionless size of the recirculation zone. The results in this study have certain reference value for the design of the frigate superstructure.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-09T10:25:23Z
      DOI: 10.1177/0954410020977752
       
  • Time-marching solution of transonic flows at axial turbomachinery meanline
    • Authors: Simone Rosa Taddei
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A new blade force model is coupled to quasi-one dimensional Euler equations for a variable geometry flowpath. After analytical inclusion of the blade force, the flow equations take a strictly one-dimensional form with specific expressions of the convective flux and blade load source terms. Regardless of the flow turning, that is simply achieved by the load source term as an explicit function of the blade camber, the new form describes a perfect analogy between the average flow inside a blade passage and strictly one-dimensional flows, especially concerning wave propagation. This property allows capture of passage choking and shocks. Other types of shock more important for turbomachinery analysis, like leading edge strong shocks in compressors and trailing edge weak shocks in choked turbines, are modelled by properly matching the new set of equations inside blade regions with the standard quasi-one dimensional equations outside. Upon specification of viscous losses and subsonic deviations fitted from experimental results, the model predicts the choke mass flow of a transonic compressor stage (NASA stage 37) at a 0.1% to 0.4% accuracy both in the absence and in the presence of the leading edge shock. This result supports the effectiveness of the leading edge shock model. The accuracy on choke mass flow would decrease to around 1% if empirical input was specified from open-literature experimental correlations. The model captures the typical trend of exit angle with total pressure ratio for a choked turbine (NASA Lewis two-stage). This result involves satisfactory prediction of not only choke mass flow, but also trailing edge shock loss and supersonic deviation. The complete turbine operational map in terms of shaft torque and pressure ratio is also re-obtained with noticeable accuracy except in strong off-design conditions, where experimental correlations likely fail.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-09T10:20:31Z
      DOI: 10.1177/0954410020977350
       
  • Evaluation on the performance fluctuation after water ingestion for a
           turbofan engine compressor during flight descent
    • Authors: Lu Yang, Qun Zheng, Aqiang Lin
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Turbofan engine compressor is most severely threatened by the entry of liquid water during flight descent. This study aims to deeply understand the fluctuations of compressor performance parameters caused by water ingestion through frequency spectrum analysis. The water content and droplet diameter distribution are determined based on the real heavy rain environment. Results reveal that most of the droplets actually entering the core compressor have a particle size of less than 100 μm. In addition, the formation and motion of water film plays a critical role in affecting the fluctuation characteristics. Water ingestion deteriorates the compression performance and aggravates the unsteady fluctuations of the fan. However, the performance of the core compressor is less affected by water ingestion, but their fluctuations are still exacerbated. For some important parameters, such as inlet mass flow rate, total pressure ratio, total temperature ratio, compression work and efficiency, their main frequency of fluctuation are switched from the original blade passing frequency to the rotor passing frequency, and their amplitudes are correspondingly amplified to varying degrees. These phenomena can be observed in both the fluctuations of the fan and core compressor. Moreover, the operating point of them will be in the long-period and large-amplitude fluctuations, which leads them experiences the non-optimal state for a long time and threatens their operating stability.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-09T09:54:38Z
      DOI: 10.1177/0954410020977982
       
  • Experimental and numerical analysis of fluid-solid-thermal coupling on
           electric fuel pump
    • Authors: Renfeng Wei, Zhifeng Ye
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper designs an axial partition fuel cooling shell to solve the problem of temperature rise in the motor of the electric fuel pump (EFP). And describes a simplified method in conjunction with the computational fluid dynamics(CFD) to analyze heat generations and fuel cooling effects in integrated EFPs. Furthermore, CFD is used to numerically simulate the coupling effects among the fluid-solid-thermal based on multiple physical field. With varying different working conditions of the pump, cooling characteristics of the fuel cooling shell are obtained through CFD results. Finally, an experimental system for the EFP is established to verify reliability of the simplified method and the effectiveness of the fuel cooling scheme. Results show that fuel cooling shell plays an essential role in heat dissipating, with a maximum reduction of up to approximately 42 K in temperature. Temperature error between simulations and experiments is less than 4%, which indicates reliabilities of the simplified model and fuel cooling shell.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-09T09:24:11Z
      DOI: 10.1177/0954410020973925
       
  • A novel system identification algorithm for quad tilt-rotor based on
           neural network with foraging strategy
    • Authors: Zhigang Wang, Zhichao Lyu, Dengyan Duan, Jianbo Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Quad tilt-rotor(QTR) UAV is a nonlinear time-varying system in full flight mode. It is difficult and inaccurate to model the nonlinear time-varying system, which cannot fully reflect the problem of controlling input and system response output in the full flight mode. In order to solve the above problems, a novel neural network model was adopt to identify the nonlinear time-varying system of quad tilt-rotor in full flight mode. An adaptive learning rate algorithm based on foraging strategy is proposed based on the global error BP neural network. Corresponding to the nonlinear time-varying system, BP neural network is set as the time-invariant system structure with constant network structure and continuously changing weights at multiple times, and the nonlinear input-output relationship under the time-varying system is jointly described by fitting the network at all times. The extended Kalman filtering algorithm is used to track the network connection weights by modifying the network weights at the current moment with the input and output data at the next moment. The final identification result shows that the smaller mean square error of both only transition process and full flight mode, shows that using this optimization algorithm can well describe the input and output characteristics of the nonlinear time-varying systems. When the same network structure is adopted, no matter for transition mode or full mode, the BP optimization algorithm based on foraging strategy is better than the global BP algorithm for system identification of the full mode quad tilt-rotor. Therefore, when the BP neural network based on foraging strategy is adopted, the same network structure can be adopted to systematically identify the full mode of quad tilt-rotor by changing the weight.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-08T10:48:53Z
      DOI: 10.1177/0954410020976598
       
  • Numerical transient responses of cut-out borne composite panel and
           experimental validity
    • Authors: Hukum Chand Dewangan, Subrata Kumar Panda
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The effect of cut-out parameters (shapes: square and circular; position: concentric/eccentric) on the dynamic deflection values of the curved/flat layered composite panel are verified experimentally with the higher-order finite element solutions first-time. The solutions are obtained using the linear finite element model in the framework of cubic-order displacement filed functions. The necessity of higher-order kinematic model is verified by comparing the experimental transient data by conducting the different test to show the accuracy of the finite element solution. Moreover, the theoretical finite element solutions are obtained using the own experimental elastic property data for the comparison (numerical and experimental) purpose. Finally, the critical behaviour of the proposed numerical model for the dynamic analysis of damaged composite structure is examined by solving different types of example by varying the design constraint parameter including the cut-out factors (shape, size, location and eccentricity). The inclusiveness of each parameter on the time-dependent deflections is expressed in details from the various example including the comparison.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-08T10:38:59Z
      DOI: 10.1177/0954410020977344
       
  • An experimental study of rotor-stator wake unsteadiness in a multistage
           axial compressor
    • Authors: Jun Li, Jun Hu, Chenkai Zhang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The flow in a multistage axial compressor is highly unsteady, three-dimensional and turbulent. The interaction between compressor blade rows results in rotor/stator wake unsteadiness, which is not typically considered in the computational fluid dynamics (CFD) models. To gain depth and insight into the inner flow mechanism in multistage compressors, specifically the wake variability driven by the rotor/stator and stator/stator interactions, a compound total-pressure pneumatic probe with both high and low response-frequency were designed and manufactured. Unsteady rotor and stator wake measurements between blade rows for the third stage were carried out with this probe installing on a 3-DOF displacement mechanism, to deepen the knowledge of unsteady interactions in the embedded stages of a four-stage low-speed axial compressor. By performing frequency spectrum analysis and ensemble-average methods, higher spectral magnitude of the blade passing frequency (fBPF) and higher root mean square values of total pressure (PtRMS) at both sides of the stator wake region caused by the shedding of upstream boundary layer are revealed. In addition, the high-order harmonics are strengthened by the stator/stator interactions, especially near the blade tip. The individual contributions of rotor geometry variations/interactions of the upstream rotor wakes and the effects of downstream stator potential modulation to the wake variations can be understood.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-08T08:11:43Z
      DOI: 10.1177/0954410020976483
       
  • Far-field drag decomposition using hybrid formulas and vorticity based
           area sensors
    • Authors: L Qiao, XL He, Y Sun, JQ Bai, L Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Numerical simulation of flow-field has become an indispensable tool for aerodynamic design. Usually, wall surface integration is a tool used to calculate values of pressure drag and skin friction drag, but the aerodynamic mechanism of drag production is still confusing. In present work, in order to decompose the total drag into viscous drag, wave drag, induced drag, and spurious drag, a far-field drag decomposition (FDD) method is developed. This method depends on axial velocity defect and area sensor functions. The present work proposes three hybrid formulas for velocity defect to tackle the negative square root issue by analyzing the existing axial velocity defect formulas. For dealing with the issue of detection failure for near-wall cells, a novel vorticity based viscous area sensor function is proposed. The new area sensor function is also independent of the turbulence model, which ensures easy application to general simulation methods for flow-field. Three tests cases are there to validate the proposed FDD method. The three dimensional transonic CRM test case shows that the present improvement is crucial for accurate drag decomposition. Excellent agreement between total decomposed drags and results from the near-field method or experimental data further confirms the correctness.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-08T02:18:18Z
      DOI: 10.1177/0954410020973904
       
  • Thermal and structural response of aerospike mounted on blunt-nose body
    • Authors: Zhang ZhunHyok, Won CholJin, Ri CholUk, Kim CholJin, Kim RyongSop
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The inclusion of aerospike on blunt nose body of hypersonic vehicle has been considered to be the simplest and most efficient technique for a concurrent reduction of both aeroheating and wave drag due to hypersonic speed. However, the thermal and mechanical behavior of aerospike structure under the coupling effect of aerodynamic force and aeroheating remains unclear. In this study, the thermal and structural response of aerospike mounted on the blunt nose body of hypersonic vehicle was numerically simulated by applying 3 D fluid-thermal-structural coupling method based on loosely-coupled strategy. In the simulation, the angle-of-attack and the spike’s length and diameter are differently set as α = 0°–10°, L/D = 1–2 and d/D = 0.05–0.15, respectively. Through the parametric study, the following results were obtained. Firstly, the increase of vehicle’s angle-of-attack and spike’s length unfavorably affect the thermal and structural response of aerospike. Secondly, the increase of spike’s diameter can improve its structural response characteristic. Finally, the aerospike with the angle-of-attack of 0° and the length and diameter of L/D = 1 and d/D = 0.15, respectively, is preferred in consideration of the effect of flight angle-of-attack and spike’s geometrical structure on the thermal and structural response of spike and the drag reduction of vehicle. The numerical calculation results provide a technical support for the safe design of aerospike.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-07T07:55:31Z
      DOI: 10.1177/0954410020976488
       
  • An improved dynamic load-strength interference model for the reliability
           analysis of aero-engine rotor blade system
    • Authors: Bingfeng Zhao, Liyang Xie, Yu Zhang, Jungang Ren, Xin Bai, Bo Qin
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      As the power source of an aircraft, aero-engine tends to meet many rigorous requirements for high thrust-weight ratio and reliability with the continuous improvement of aero-engine performance. In this paper, based on the order statistics and stochastic process theory, an improved dynamic load-strength interference (LSI) model was proposed for the reliability analysis of aero-engine rotor blade system, with strength degradation and catastrophic failure involved. In presented model, the “unconventional active” characteristic of rotor blade system, changeable functioning relationships and system-component configurations, was fully considered, which is necessary for both theoretical analysis and engineering application. In addition, to reduce the computation cost, a simplified form of the improved LSI model was also built for convenience of engineering application. To verify the effectiveness of the improved model, reliability of turbojet 7 engine rotor blade system was calculated by the improved LSI model based on the results of static finite element analysis. Compared with the traditional LSI model, the result showed that there were significant differences between the calculation results of the two models, in which the improved model was more appropriate to the practical condition.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-07T07:50:33Z
      DOI: 10.1177/0954410020972898
       
  • Preliminary parameters design for a long endurance unmanned helicopter
           with low rotor-disc loading
    • Authors: Hong Zhao, Jian-Bo Li, Yuan Wang, Zhi-Gang Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper investigates the design of a long-endurance unmanned helicopter (LEUH) with low rotor disc loading (LRDL) and low rotor speed (LRS). Due to the flaws in flying qualities caused by the LRDL and the LRS, this paper establishes a flying quality evaluation model in which handling qualities (FQs) and flight control (FC) are introduced into the distributed multi-objective collaborative optimization (DMOCO) of the helicopters. The comprehensive design optimization on preliminary parameters of the LEUH in wind shear is also carried out. Numerical simulation results show that the LRDL and the LRS technologies are successfully applied to LEUH, with the FQs and the flight performance considered. Compared with A160 LEUH, the payload load ratio is significantly improved.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-01T06:21:46Z
      DOI: 10.1177/0954410020973899
       
  • Multi-physics simulation of an insect with flapping wings
    • Authors: Kabir Bakhshaei, Hoomaan MoradiMaryamnegari, Sadjad SalavatiDezfouli, Abdol Majid Khoshnood, Mani Fathali
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, the unsteady aerodynamic of an insect’s forward flight has been carried out with a novel approach. In order to fully utilize the available powerful solvers, an innovative intermediary MATLAB code has been written for the high-fidelity time-resolved multi-physics problems involving fluid flow and multi-body simulations. For simulating the insect’s flight, the FLUENT solver has been utilized to determine aerodynamic forces and moments of the wings and main body while ADAMS software has been employed to calculate translational and angular velocities. Overset grid technology accompanied with dynamic mesh method have been implemented for the movement of the insect. The code is responsible for the synchronization of the solvers at the end of each time step as well as the integration of the solutions. Three different simulations are done for two different insects’ geometries. For the first and second simulations, a simplified geometry of an insect is selected, due to the ease of manufacturing and testing. At first, all rotational and translational degrees of freedom are considered to be free. The motion path history shows the instability due to an inappropriate location of the center of gravity. Hence, in the second case, it is assumed that the insect’s main body is limited to the vertical motion. In the final simulation, a complicated model of a bee with exact geometry and wings kinematics extracts from the experimental data with the free translational degrees of freedom. According to the results, combining multiple software in which they can interact with each other at each time step, is the most accurate way for doing precise multi-physics simulations.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-01T06:20:31Z
      DOI: 10.1177/0954410020972581
       
  • Enhancement of air entrainment in ejector-diffuser using plate guidance at
           slots to reduce infrared emission
    • Authors: L Singh, SN Singh, SS Sinha
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Ejector-diffuser reduces infrared emissions and are installed in combat aircraft to counter the threat of heat-seeking missile. The specific role of an ejector-diffuser is to reduce the heat emissions without substantially affecting the engine performance. The present study investigates a new design of ejector-diffuser wherein straight-plates and hybrid-straight-plates are installed at each slot for improving the ejector-diffuser performance. The evaluation criteria of an ejector-diffuser is specified in terms of air entrainment through the slots, thermal characteristics, and recovery of pressure. This work is carried out in two stages. In the first part, the orientation of the plate at the slot is investigated by varying the angle between the slot and diffuser axis over the range [math]. The overall mass entrainment increases from 2.88 to 4.04 with the increase in plate angle. Further, the thermal characteristics also improves with increase in plate angle, but the pressure recovery decreases from 0.701 to 0.155. In the second part, the straight-plate at the slots are partially/fully replaced by hybrid-plate. Two configurations are proposed by first introducing a hybrid-plate at the first slot and straight-plate at the other slots, and subsequently by introducing hybrid-plate at all the slots. It is found that the pressure recovery in both the cases shows a significant improvement compared to the straight-plate case, the value being close to 0.75 for both the cases. However, the cumulative mass entrained by the first configuration of the hybrid-plate is better than the second configuration and is similar to the straight-plate guidance of 28°. Thus, the current study proposes an IRSS device having the hybrid-plate at the first slot and the straight-plate guidance at the remaining slots which reduces infrared emissions with minimum loading on the engine.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-01T06:17:28Z
      DOI: 10.1177/0954410020971938
       
  • A numerical study on the single pulsed energy addition based unsteady wave
           drag reduction at varied hypersonic flow regimes
    • Authors: Dathi SNV Rajasekhar Rao, Bibin John
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this study, unsteady wave drag reduction in hypersonic flowfield using pulsed energy addition is numerically investigated. A single energy pulse is considered to analyze the time-averaged drag reduction/pulse. The blast wave creation, translation and its interaction with shock layer are studied. As the wave drag depends only on the inviscid aspects of the flowfield, Euler part of a well-established compressible flow Navier-Stokes solver USHAS (Unstructured Solver for Hypersonic Aerothermodynamics) is employed for the present study. To explore the feasibility of pulsed energy addition in reducing the wave drag at different flight conditions, flight Mach numbers of 5.75, 6.9 and 8.0 are chosen for the study. An [math] apex angle blunt cone model is considered to be placed in such hypersonic streams, and steady-state drag and unsteady drag reductions are computed. The simulation results indicate that drag of the blunt-body can be reduced below the steady-state drag for a significant period of energy bubble-shock layer interaction, and the corresponding propulsive energy savings can be up to 9%. For energy pulse of magnitude 100mJ deposited to a spherical region of 2 mm radius, located 50 mm upstream of the blunt-body offered a maximum percentage of wave drag reduction in the case of Mach 8.0 flowfield. Two different flow features are found to be responsible for the drag reduction, one is the low-density core of the blast wave and the second one is the baroclinic vortex created due to the plasma energy bubble-shock layer interaction. For the same freestream stagnation conditions, these two flow features are noted to be very predominant in the case of high Mach number flow in comparison to Mach 5.75 and 6.9 cases. However, the ratio of energy saved to the energy consumed is noted as a maximum for the lower Mach number case.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-27T12:10:25Z
      DOI: 10.1177/0954410020973134
       
  • An experimental investigation on the use of a rectangular strut in a
           scramjet thruster for thrust vector control
    • Authors: DR Biju Ben Rose, BTN Sridhar
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An experimental investigation was carried out to find the effect of a strut spanning across the width of a single expansion ramp of a laboratory model scramjet thruster as a thrust vector control device. Cold flow tests were conducted with operating total pressures ranging from 4 bar to 10 bar and the thruster exhaust was to ambient atmosphere. The mass flow rate varied from 0.105 kg/s to 0.263 kg/s. Experiments were conducted by varying the strut height at different operational total pressures to find if any thrust vector control could be achieved to supplement the maneuverability of hypersonic vehicle with aerodynamic control. The laboratory model consisted of an isolator, a divergent combustor followed by a single expansion ramp. Except the side walls, the thruster was fabricated with stainless steel. A high quality acrylic sheet was used for internal flow visualization by a schilieren system. The wall pressure was recorded at different locations from the combustor inlet to ramp trailing edge. Shock pattern was studied from the schilieren images and it was observed that an increase in strut height caused a downward deflection of the exhaust. From the wall pressure distribution, two dimensional side force coefficient and pitching moment coefficient were calculated and the effect of strut height variation on the above coefficients was plotted. Results from experiments indicated that the presence of the strut yielded noticeable changes in side force and pitching moment. The increase in strut height provided exhaust stream directional changes which may be useful in maneuvering the vehicles employing scramjet propulsion system.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-24T02:53:08Z
      DOI: 10.1177/0954410020973128
       
  • Fixed-time three-dimensional guidance law with input constraint and
           actuator faults
    • Authors: Peng Li, Qi Liu, Chen-Yu He, Xiao-Qing Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper investigates the three-dimensional guidance with the impact angle constraint, actuator faults and input constraint. Firstly, an adaptive three-dimensional guidance law with impact angle constraint is designed by using the terminal sliding mode control and nonhomogeneous disturbance observer. Then, in order to solve the problem of the input saturation and actuator faults, an adaptive anti-saturation fault-tolerant three-dimensional law is proposed by using the hyperbolic tangent function based on the passive fault-tolerant control. Finally, the effectiveness of the designed guidance laws is verified by using the Lyapunov function and simulation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-24T02:51:28Z
      DOI: 10.1177/0954410020971973
       
  • A numerical study on the blow-off limit of premixed hydrogen/air flames in
           a cylindrical micro-combustor
    • Authors: Saeed Naeemi, Seyed Abdolmehdi Hashemi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In the current work, a numerical study on combustion of premixed H2–air in a micro-cylindrical combustor was carried out and the critical velocity of inlet flow that causes the blow-off was obtained. Furthermore, the effects the equivalence ratio, wall thickness, geometry of combustor and thermal properties of walls on the critical blow-off velocity were studied. The numerical results showed that, increasing the equivalence ratio results in higher critical blow-off velocity. A micro combustor with thicker wall had better flame stability. As the combustor dimeter is decreased the blow-off occur in lower inlet flow velocity. Higher thermal conductivity of walls increases the critical blow-off velocity. In addition, with varying heat convection coefficient (h) and emissivity coefficient [math] of the walls from 1 to 60 W/m2.K and 0.2 to 0.8 respectively, the critical blow-off velocity is reduced and shows the importance of wall thermal properties in the design and operation of micro-combustors.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-20T07:17:42Z
      DOI: 10.1177/0954410020971446
       
  • Effects of skin heat conduction on aircraft icing process
    • Authors: Xiaobin Shen, Yu Zeng, Guiping Lin, Zuodong Mu, Dongsheng Wen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      During the aircraft icing process caused by super-cooled droplet impingement, the surface temperature and heat flux distributions of the skin would vary due to the solid substrate heat conduction. An unsteady thermodynamic model of the phase transition was established with a time-implicit solution algorithm, in which the solid heat conduction and the water freezing were analyzed simultaneously. The icing process on a rectangular skin segment was numerically simulated, and the variations of skin temperature distribution, thicknesses of ice layer and water film were obtained. Results show that the presented model could predict the icing process more accurately, and is not sensitive to the selection of time step. The latent heat released by water freezing affects the skin temperature, which in turn changes the icing characteristics. The skin temperature distribution would be affected notably by the boundary condition of the inner skin surface, the lateral heat conduction and thermal property of the skin. It was found that the ice accretion rate of the case that the inner surface boundary is in natural convection at ambient temperature is much smaller than that with constant ambient temperature there; due to the skin lateral heat conduction, the outer skin surface temperature increases first and then decreases with uneven distribution, leading to an unsteady ice accretion rate and uneven ice thickness distribution; a smaller heat conductivity would lead to a more uneven temperature distribution and a lower ice accretion rate in most regions, but the maximum ice thickness could be larger than that of higher heat conductivity skin. Therefore, in order to predict the aircraft icing phenomenon more accurately, it is necessary to consider the solid heat conduction and the boundary conditions of the skin substrate, instead of applying a simple boundary condition of adiabatic or a fixed temperature for the outer skin surface.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-11T12:01:01Z
      DOI: 10.1177/0954410020972577
       
  • Virtual guidance-based finite-time path-following control of underactuated
           autonomous airship with error constraints and uncertainties
    • Authors: Yan Wei, Pingfang Zhou, Yueying Wang, Dengping Duan, Zheng Chen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper addresses the finite-time three-dimensional path-following control problem for underactuated autonomous airship with error constraints and uncertainties. First, a five degrees-of-freedom path-following error model in the Serret-Frenet coordinate frame is established. By applying the finite-time stability theory, a virtual guidance-based finite-time adaptive neural backstepping path-following control approach is proposed. Barrier Lyapunov functions (BLFs) are introduced to deal with attitude error constraints. Neural networks (NNs) are presented to compensate for the uncertainties. To prevent the “explosion of complexity” in the design of the backstepping method, a finite-time convergent differentiator (FTCD) is introduced to estimate the time derivatives of virtual control signals. Stability analysis showed that all closed-loop signals are uniformly ultimately bounded, the constrained requirements on the airship attitude errors are never violated, and the path-following errors converge to a small neighborhood of the origin in a finite time. At last, simulation studies are provided to demonstrate the effectiveness of the proposed control approach.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-06T06:36:20Z
      DOI: 10.1177/0954410020969319
       
  • Attitude trajectory tracking of quadrotor UAV using super-twisting
           observer-based adaptive controller
    • Authors: Ai-Jun Chen, Ming-Jian Sun, Zhen-Hua Wang, Nai-Zhang Feng, Yi Shen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The successful implementation of high-level decision algorithm on quadrotor depends on the accurate trajectory tracking performance. In this paper attitude estimation and trajectory tracking control problem of quadrotor unmanned aerial vehicle (UAV) with endogenous and exogenous disturbance are considered, where the lumped disturbance characteristic does not have a probabilistic illustration but instead the dynamics are known to have a bound. The problem is handled by developing disturbance estimator and control strategy. In order to estimate lumped disturbance precisely, a globally finite time stable extended state observer is proposed based on super-twisting algorithm. Stability analysis and observer’s parameters selection rule are discussed by using Lyapunov’s stability theory. The proposed observer strategy achieves accurate observing performance of disturbance without increasing observer’s order, and chattering effect is also reduced by applying super-twisting algorithm. Furthermore, a super-twisting sliding mode control law is proposed to guarantee the asymptotic convergence of the drone’s orientation with respect to the reference. Finally, a numerical study based on simulations is presented to analyze the performance of proposed approach.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-05T07:19:26Z
      DOI: 10.1177/0954410020966476
       
  • Leading edge redesign of dual-peak type variable inlet guide vane and its
           effect on aerodynamic performance
    • Authors: Hengtao Shi, Lucheng Ji
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Recently, a new type airfoil for variable inlet guide vane (VIGV), featuring “dual-peak” surface velocity pattern at high incidence, is proposed and shows wide low-loss operation range. To further improve its performance, this paper researches the influence of leading edge (LE) thickness and shape on the loss level and surface velocity features of the “dual-peak” type airfoil. Firstly, a polynomial-based continuous-curvature leading edge design method was briefly introduced and used in the LE redesign of sample airfoils. Then, steady simulations based on Reynolds-Averaged Navier-Stokes method (RANS), carried out by commercial software CFX after grid independent study, were used to determine the aerodynamic performance, surface velocity distribution and boundary-layer behaviors of all research airfoils. Simulation results indicate that there exists an optimized range of LE relative thickness that can achieve lower airfoil loss level at high incidence condition. For Case 1 ([math]) and Case 2 ([math]), the optimized LE relative thickness range is [math] and [math]. The LE shape optimization can further reduce the maximum incidence condition loss coefficient with proportion up to 18% for airfoils with optimal LE thickness. Analysis of flow mechanism indicates that the optimized LE thickness and shape can reduce the suction spike height and subsequent adverse pressure gradient, therefore, decrease the LE separation scale and result in a lower loss coefficient. As an application, a dual peak VIGV with circular LE, presented in previous paper as the optimized VIGV, is redesigned in the LE portion according to the research findings and achieved 0.6 percent improvement in passage-averaged total pressure recovery coefficient [math] at extreme high stagger angle point and the low-loss operation range extends with about 5°, which confirms the effectiveness of the research findings in three-dimensional environment.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-11-05T07:18:07Z
      DOI: 10.1177/0954410020966168
       
  • Entrance length effects on the flow features of rectangular liquid jets
    • Authors: MH Aliyoldashi, M Tadjfar, A Jaberi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An experimental study was carried out to investigate the effects of entrance length on the main characteristics of rectangular liquid jets discharged into the stagnant atmosphere. Six rectangular nozzles, all with the same aspect ratio of 3 but with different entrance length ratios ranging from 3.3 to 60 were constructed. The physics of the fluid flows was visualized by the aid of backlight shadowgraph technique and high speed photography. Flow visualizations revealed that in the mid-range of Weber numbers, the perturbations induced over the liquid surface remarkably depended on the entrance length ratio. Moreover, the characteristics of the axis-switching instability of rectangular liquid jets were measured. It was found that axis-switching wavelength was independent of the entrance length, while the amplitude of axis-switching was directly influenced. For entrance length ratios smaller than 10, the amplitude was increased with increase of entrance length, whereas for entrance length ratios higher than 10, this trend was reversed. Measurements of breakup length also showed that the transition of flow regimes was not perceptibly affected by the entrance length.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-28T04:03:07Z
      DOI: 10.1177/0954410020968445
       
  • Characteristics of open cavity flow with floor inclinations at
           M = 2.0
    • Authors: VS Saranyamol, Priyank Kumar, Sudip Das
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Experimental studies on open cavity flows at supersonic speed of M = 2.0 were carried out. Oil flow visualization tests were made to understand the steady features of the surface flow field. Unsteady pressure measurements were done at five locations inside the cavity and pressure spectrum of these measurements were obtained. Cavity floor was made inclined to influence the flow directing towards the cavity leading edge with both, a favourable and adverse slope, by giving a positive and negative inclination angles to the floor, respectively. It is observed that the negative inclinations to the cavity floor behaves in a similar way to the base cavity, but a positive inclination helps to reduce the fluctuating pressures by 80% and reduce OASPL to the order of 14 dB and more.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-28T04:02:47Z
      DOI: 10.1177/0954410020966493
       
  • Free vibration analysis of rotating thin-walled cylindrical shells with
           hard coating based on Rayleigh-Ritz method
    • Authors: Dongxu Du, Wei Sun, Xianfei Yan, Kunpeng Xu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper focuses on the free vibration analyses of the rotating hard-coating cylindrical shells with various boundary conditions and the effects of hard coating on the vibration behaviors of the rotating shells. To accurately predict more natural characteristics, the characteristic orthogonal polynomials are taken as the admissible displacement functions. Considering the influences of Coriolis force, centrifugal force and initial hoop tension caused by rotation, the equations of motion of the shells are established by the use of the Rayleigh-Ritz method. Based on the state vector method, an efficient method is developed to solve the equations. By comparing with the results of both the finite element analysis and published literatures, the high accuracy and good convergence of the proposed model are verified. In addition, the effects of the boundary conditions, parameters of hard coating, rotating speed and number of circumferential waves on the vibration behaviors of the hard-coating shells are evaluated. This study may provide a reference for the application of hard-coating damping treatment to the vibration suppression of rotating thin-walled structures.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T09:00:31Z
      DOI: 10.1177/0954410020967243
       
  • Analysis of influencing parameters in ion thruster plume simulation
    • Authors: Baiyi Zhang, Guobiao Cai, Hongru Zheng, Bijiao He, Huiyan Weng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Effects from ion thruster plume have long raised concerns. However, little work has systematically analyzed the influencing parameters in the ion thruster plume simulation. This paper analyzes LIPS 200 ion thruster plume simulations about the influencing parameters. The numerical simulations are carried out by a hybrid particle-in-cell (PIC) method and the direct simulation Monte Carlo (DSMC) method. The PIC method is employed for the plasma dynamics, and the DSMC method is used for collisions. Simulation results were compared in detail to obtain the variation of the results with the parameters. Besides, experimental data were compared to simulation results to optimize the parameters. Finally, with these researches, the optimal parameters are obtained.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:58:31Z
      DOI: 10.1177/0954410020967220
       
  • On the reductions of aerofoil-turbulence interaction noise through
           multi-wavelength leading edge serrations
    • Authors: S Narayanan, Sushil Kumar Singh
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper provides an experimental study into the use of multi-wavelength sinusoidal leading edge (LE) serrations for enhancing the aerofoil-broadband noise reductions. The noise reduction performances of multi-wavelength serration profiles introduced on a flat plate are compared against those generated by single-wavelength profiles when applied separately. The multi-wavelength leading edge serration is made in such a way that its maximum amplitude is kept same as that of each single-wavelength ones to be compared. The present study reveals that the dual-wavelength serrations provide higher noise reductions over a narrow band of frequencies as compared to single and triple wavelength ones. Further, it reveals that the noise reduction characteristics of dual-wavelength serrated airfoils are similar to the flat plates. It shows that the baseline plate generate higher noise radiations for all emission angles as compared to leading edge serrated plates, but the common feature among them is the downstream directivity. For the range of frequencies 0.9 to 5 kHz, the highest directivity is seen at an emission angle of 55° for the baseline, while it occurs at 75° for the serrated plates. The dual wavelength serrations generate lowest acoustic radiations as compared to single and triple ones for all the emission angles. Also, it is noticed that the radiation levels of the dual serrations decrease with increase in amplitude of the serration, which shows that the longer dual serrations generate lowest acoustic radiations. Thus, the present study illustrates that the dual wavelength leading edge serrations act as the best passively modified serration profiles for achieving the highest noise reductions over a wide range of frequencies as compared to single and triple wavelength ones.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:55:50Z
      DOI: 10.1177/0954410020965747
       
  • Fast fixed-time convergent smooth adaptive guidance law with terminal
           angle constraint for interception of maneuvering targets
    • Authors: Peng Zhang, Xiaoyu Zhang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper introduces a fast fixed-time guidance law with terminal angle constraint for interception of maneuvering targets, which is based on the structure of singularity-free fast terminal sliding mode and the fixed-time stability theory. Different from the finite-time stability, the fixed-time stability can predefine the maximum stabilization time of system states which is independent on the initial value of system states. Under the proposed guidance law, the guidance system can achieve stabilization within settling time which decides by the parameters of controller. In addition, an adaptive law is proposed which alleviate the chattering of sliding mode and smooths the guidance law. Meanwhile, the proof of the sliding mode manifold and system states fixed-time convergence is given by Lyapunov stability theory. Finally, numerical simulations demonstrate the performance of the proposed guidance law is satisfying.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:51:07Z
      DOI: 10.1177/0954410020965094
       
  • Numerical study on aerodynamic performance of waverider with a new
           bluntness method
    • Authors: Zhipeng Qu, Houdi Xiao, Mingyun Lv, Guangli Li, Cui Kai
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      AbastrctThe waverider is deemed the most promising configuration for hypersonic vehicle with its high lift-to-drag ratio at design conditions. However, considering the serious aero-heating protection, the sharp leading edge must be blunted. The existing traditional bluntness methods including the following two types: “reducing material method” and “adding material method”. Compared to the initial waverider, the volume will be smaller or larger using the traditional methods. With the fixed blunted radius, the volume and aerodynamic performance is determined. In this paper, a new bluntness method which is named “mixing material method” is developed. In this new method, a new parameter is introduced based on the traditional two bluntness methods. Under fixed blunted radius, the volume and aerodynamic performance can be changed within a wide range by adjusting the parameter. When the parameter is 0 and 1, the novel blunted method degenerated into the “reducing material method” and “adding material method” respectively. The influence of new parameter on the aerodynamic characteristics and volume are studied by numerical simulation. Results show that the volume, lift and lift-to-drag ratio increases with the increase of the parameter under the fixed blunt radius, but simultaneously, the drag will also increase. Therefore, considering the different requirements of the air-breathing hypersonic aircrafts for the balance of thrust and drag, lift and weight, a suitable bluntness parameter can be selected to achieve a balance. This research can provide reference for hypersonic waverider vehicle design.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:30:46Z
      DOI: 10.1177/0954410020968419
       
  • Structure and aerodynamic characteristics of a coaxial quad-wing flapper
    • Authors: Huan Shen, Qian Li, Kun Hu, Zhuoqun Feng, Aihong Ji
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      As a special type of micro ornithopter, the coaxial quad-wing flapper (CQWF) enables greater flight speed and higher stability than the paired-wing flapper. These characteristics are closely related to the unique pneumatic mechanism of the CQWF. Therefore, the aerodynamic generation mechanism of the CQWFs has been actively researched in recent years. This study verifies the reliability of flow-field simulations in a CQWF prototype with an aerodynamically optimized driving mechanism. For the selected motion parameters and shape dimensions of the flapping-wing aircraft, the vorticity fields at different elevation angles are observed in flow-field simulations. The elevation angle strongly affects the lift. Moreover, the wing movement based on the Clap–Fling mechanism significantly affected the acquisition of the lift, which explains the higher stability of the CQWF than that of the paired-wing flapper and provides a theoretical basis for the optimization of the flapping prototype. When tested on a wind-tunnel platform, the prototype yields slightly higher lift compared with those obtained in the simulation study. In addition to confirming the phenomenon revealed in flow visualization, it also showed that the unsteady mechanism of the two-pair wing is more powerful than calculated.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:29:46Z
      DOI: 10.1177/0954410020967546
       
  • Shock train control by boundary layer suction in a scramjet isolator
    • Authors: Vignesh Ram Petha Sethuraman, Tae Ho Kim, Heuy Dong Kim
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The isolator plays a critical role in the scramjet engine situated between the inlet and the combustion chamber. The flow field is more complex with shock–shock interaction and shock boundary layer interaction result in a series of compression waves reffered to as “shock train”. The presence of such flow inside the isolator can degrade the performance of the scramjet engine. The present study focus on the characteristic of the shock train flow field in an isolator and its control by partial removal of the boundary layer. The results examine the variation of the inlet to outlet pressure ratio along with different suction flow ratio. Numerical results indicate that boundary layer suction will cause the slight downstream movement of shock train location and the length of the shock train is reduced. Also when the suction flow gets choked, the transformation of shock train into a single curved normal shock is observed. The effect of varying the upstream boundary layer plays a major role in the suction flow ratio. Furthermore, a significant improvement in the total pressure loss and static pressure rise is obtained by boundary layer suction. The location of the shock train has a greater impact on the performance of the isolator.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:28:45Z
      DOI: 10.1177/0954410020967537
       
  • Compressible large eddy simulation of the unsteady evolution process in a
           LPT Cascade with incoming wakes
    • Authors: Yunfei Wang, Huanlong Chen, Huaping Liu, Yanping Song, Fu Chen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An in-house large eddy simulation (LES) code based on three-dimensional compressible N-S equations is used to research the impact of incoming wakes on unsteady evolution characteristic in a low-pressure turbine (LPT) cascade. The Mach number is 0.4 and Reynolds number is 0.6 × 105 (based on the axial chord and outlet velocity). The reduced frequency of incoming wakes is Fred = 0 (without wakes), 0.37 and 0.74. A detailed analysis of Reynolds stresses and turbulent kinetic energy inside the boundary layer has been carried out. Particular consideration is devoted to the transport process of incoming wakes and the intermittent property of the unsteady boundary layer. With the increase of reduced frequency, the inhibiting effect of wakes on boundary layer separation gradually enhances. The separation at the rear part of the suction side is weakened and the separation point moves downstream. However, incoming wakes lead to an increase in dissipation and aerodynamic losses in the main flow area. Excessive reduced frequency (Fred = 0.74) causes the main flow area to become one of the main source areas of loss. An optimal reduced frequency exists to minimize the aerodynamic loss of the linear cascade.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-26T08:28:06Z
      DOI: 10.1177/0954410020967535
       
  • Effect of air jet with injection pressure on the performance of mixed
           compression air intake
    • Authors: NK Gahlot, NK Singh
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A computational study on the performance of mixed compression supersonic air intake has been carried out with and without air jet at various operating conditions. Commercially available software ANSYS was used and K-ω SST turbulence model was selected to capture the turbulent flow inside the air intake. All the simulations were simulated at a design Mach number of 2.2. Two Air jet of 1 mm diameter each and perpendicular to the local ramp surface have been placed in longitudinal direction at 0.47 times the total length of the air intake. Effect of variation of injection pressure on the flow field of air intake has been studied. Injection pressure has been varied with respect to the free stream pressure. Four different cases of injection pressure have been investigated. Three different positions (1.far away before the air jet, 2. immediately after the air jet and 3. far away behind the air jet) of normal shock were simulated to study the effect of air jet by varying the back pressure of the supersonic air intake. Significant reduction in the flow separation due the normal shock wave was noticed for all the cases of injection pressure, which further helps in improving the performance of the supersonic air intake. Important performance parameters such as flow distortion, mass flow ratio and total pressure recovery were calculated to measure the efficacy of supersonic air intake with air jet at various operating conditions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-23T05:50:20Z
      DOI: 10.1177/0954410020966507
       
  • Characteristics of helicopter engine exhaust through scaled experiments
           using stereoscopic particle image velocimetry
    • Authors: Zhen Wei Teo, Wai Hou Wong, Zhi Wen Lee, Tze How New, Bing Feng Ng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Helicopter engines are often mounted atop the fuselage to keep the aircraft footprint small and optimal for operations. As a result, hot gases produced by the engines may inadvertently impinge upon the tail boom or dissipate inefficiently that compromises on operation safety. In this study, a scaled fuselage model with a hot air blower was used to simulate hot exhaust gases. The velocity field immediately outside the exhaust port was measured through stereoscopic particle image velocimetry to capture the trajectory and flow behaviour of the gases. Two cases were considered: freestream to exhaust velocity ratios of 0 (no freestream velocity) and 0.46 (co-flowing free stream), respectively. The formation of a counter-rotating vortex pair was detected for both cases but were opposite in the rotational sense. For the case without freestream, the plume formed into a small “kidney” shape, before expanding and dissipating downstream. For the case with freestream, the plume formed into a slenderer and more elongated “reversed-C” shape as compared to the case without freestream. It also retained its shape further downstream and maintained its relative position. These observations on the trajectory and shape of plume provide basis to understanding the nature and interaction of the plume with its surroundings.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-23T05:43:20Z
      DOI: 10.1177/0954410020966471
       
  • Aircraft flat-spin recovery using sliding-mode based attitude and altitude
           control
    • Authors: Salahudden, Vijay S Dwivedi, Prasiddha N Dwivedi, Dipak K Giri, Ajoy K Ghosh
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In the present paper, a control command to recover steady-straight-level flight from flat-oscillatory-stable-left-spin is developed using a sliding-mode based attitude and altitude control. Direct spin recovery, using a spin solution by bifurcation results, to low angle-of-attack is achieved in finite-time without any separation in dynamics. The exponential convergence of errors is discussed by invoking Barbalat’s Lemma theorem. Thereafter settling time is obtained thereby making the system a finite-time stable to reach the sliding surface. The novelty of this work lies in the proposed control strategy, wherein expressions for all four primary control inputs are obtained in a closed-loop form without any approximation and altitude margin required for flat-spin recovery is investigated based on a heuristic approach for a fixed controller gains. Additionally, results of this research indicate the proposed controller first stops the spin by controlling the attitude of the aircraft thereby rotation stops about the body axis and then reaches the commanded altitude to attain the horizontal flight.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-23T05:37:40Z
      DOI: 10.1177/0954410020964674
       
  • Calibration of the CFD code based on testing of a standard AGARD-B model
           for determination of aerodynamic characteristics
    • Authors: Čedomir Kostić, Aleksandar Bengin, Boško Rašuo, Dijana Damljanović
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The goal of this work is to build a unique numerical method to obtain the basic aerodynamic characteristics of the aircraft and to enable a wide application of the method in the analysis of some aerodynamic characteristics of the aircraft, without use of empirical methods. The Computational fluid dynamics (CFD) simulation method was being calibrated based on test results of the standard AGARD-B (Advisory Group for Aerospace Research and Development) test model, which were obtained in the T-38 trisonic wind tunnel facility of the Military Technical Institute (VTI) in Belgrade, Serbia.The paper presents the CFD simulation through a description of the conditions of flow, geometry of the computer domain, grid density and mesh strategy, boundary conditions, initial strategy and turbulence model. The CFD simulation was carried out for flow cases with similarity parameters M = 0.6, M = 0.85 and M = 1.6 and Re = from 7.7(x106) to 9.9(x106) . The results of calculations were compared with the appropriate experimental ones and presented in the form of comparative diagrams for the drag, lift and pitching moment coefficients. The results of investigation presented in divergence diagrams show very good agreement between numerical and experimental ones. Simulated flows are illustrated by the distribution of pressure and velocity components on the surface of the tested model and the computational domain. This CFD simulation will be applied to other similar aerodynamic designs for a wide range angles of attack and Mach numbers and can be a strong point for the development of different aerodynamic designs.The ultimate aim of the work is to use the previous calibrated CFD simulation method as the basis for future determination of the aerodynamic characteristics of aircraft in non-stationary flight modes, caused by motion of the aircraft and/or by changing the free-velocity vector.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-23T05:33:32Z
      DOI: 10.1177/0954410020966859
       
  • Effects of propeller flow on the longitudinal and lateral dynamics and
           model couplings of a fixed-wing micro air vehicle
    • Authors: K Harikumar, Jinraj V Pushpangathan, Suresh Sundaram
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper analyzes the effects of propeller flow on the linear coupled longitudinal and lateral dynamics of a 150 mm wingspan fixed wing micro air vehicle (MAV). The effects propeller flow on the lift, drag, pitching moment and side force is obtained through wind tunnel tests. The aerodynamic forces and moments are modeled as a function of angle of attack, sideslip angle, control surface deflection and propeller rotation per minute. The nonlinear six degrees of freedom model is linearized about straight and constant altitude flight conditions for different trim airspeed to obtain linear coupled longitudinal and lateral state space model. The eigenvalues and eigenvectors of linear coupled longitudinal and lateral state space model are compared with and without propeller flow effects. The variation in the natural frequencies and damping ratios of short period mode, phugoid mode and Dutch roll mode are analyzed for various trim airspeed. An increase in the natural frequency is observed for phugoid mode and Dutch roll mode with propeller effects. The stability of the spiral mode is enhanced by the propeller flow and also the response of the roll subsidence mode is faster with propeller effects. Detailed analysis of eigenvalues and eigenvectors shows the importance of incorporating propeller flow in analyzing the dynamics of the MAV.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-22T09:24:02Z
      DOI: 10.1177/0954410020966155
       
  • Investigation of rotating detonation fueled by pre-combustion cracked
           kerosene under different channel widths
    • Authors: Xingkui Yang, Yun Wu, Yepan Zhong, Feilong Song, Shida Xu, Di Jin, Xin Chen, Shunli Wang, Jianping Zhou
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this study, the effects of channel widths on the characteristics of the rotating detonation wave (RDW) were investigated. Pre-combustion cracked kerosene and 50% oxygen-enriched air were taken as the propellant. Keeping the outer diameter (D = 150mm) constant, the channel widths (W) of the combustor range from 15 mm to 50 mm in the experiments. The results indicate that the time for the formation of a stable RDW is longer under the wider channel, while the velocity of the RDW increases significantly with a wider channel. Increasing the ER has a positive effect on the wave velocity and the flow rate has little effect on wave velocity. The wave pressure increases under the higher ER and flow rate. Under the same flow rate and ER, the RDW pressure tends to reach the maximum value when the channel width is 25 mm, and the pressure range is 2 bar to 6 bar. Five kinds of the RDW modes were observed in the experiments, namely the failure “pop-out”, single-wave mode, two-counter rotating waves mode, and two-co rotating waves mode. The two-counter rotating waves mode seems to be an intermediate mode of single-wave mode and two-co rotating waves mode in the conducted experiments, and the multi-wave mode is more likely to occur under the narrower channel and the higher oxygen content.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-22T08:21:07Z
      DOI: 10.1177/0954410020965773
       
  • Cooperative guidance law for active aircraft defense with intercept angle
           constraint
    • Authors: Min He, Xiaofang Wang, Hai Lin, Nianyuan Xiao, Zonglin Du
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, the three-body engagement scenario is considered where a target aircraft fires a defender missile to intercept the attacker missile. A cooperative guidance law with the intercept angle constraint for both of the defender missile and target has been presented. With the assumption that the attacker missile uses augmented proportional navigation guidance law, the nonlinear relative motion model of target-attacker-defender engagement is built. Considering the requirement of miss distance and satisfying the intercept angle constraint, the function index is established. The cooperative guidance law is derived based on optimal control theory. Moreover, given initial launch condition, the feasible intercept angle region of defender is analyzed, considering the limited maneuverability of defender and target and the intercept time constraint which means the attacker must be intercepted by the defender prior to hitting the target. Similarly, the feasible launch region of defender is obtained with the given designated intercept angle, variable overload of defender and target, and intercept time constraint. The simulation results further demonstrate that within the feasible region of designated intercept angle and launching condition, the defender can intercept the attacker with designated intercept angle successfully despite of the limited maneuverability. Compared with conventional uncooperative situation, the target-defender cooperation could significantly reduce the maneuverability requirements for the defender.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-22T08:06:38Z
      DOI: 10.1177/0954410020965095
       
  • Grid transformation and dynamic scattering for tail rotor radar cross
           section analysis
    • Authors: Zeyang Zhou, Jun Huang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      With the promotion and enhancement of stealth technology of helicopter rotor components, the research on the dynamic radar cross section (RCS) of helicopter rotor is becoming more and more important and imminent. In order to facilitate the calculation and analysis of the electromagnetic scattering characteristics during rotor rotation, a dynamic scattering calculation (DSC) method based on quasi-static principle (QSP) and grid coordinate transformation is presented. After analyzing the advantages and disadvantages of QSP, the dynamic principle is used to describe the rotation process of the rotor. Combined with the grid coordinate transformation method, the RCS of the rotor is accurately calculated by physical optics (PO) and physical theory of diffraction (PTD). Then the influence of azimuth, elevator angle and observation distance on rotor dynamic RCS is analyzed. The results show RCS of the tail rotor is indeed dynamic and periodic and its main influencing factors include azimuth and elevation angle. The proposed DSC method is efficient and effective for studying the dynamic RCS of tail rotor.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-21T07:32:40Z
      DOI: 10.1177/0954410020966175
       
  • Cooling effectiveness of matrix, pin fin array and hybrid structure: A
           comparative study
    • Authors: Lianfeng Yang, Yigang Luan, Shi Bu, Haiou Sun, Franco Magagnato
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In modern gas turbines, the trailing edge of turbine blades must be cooled by compact heat transfer structures. The basic problems in the design of cooling ducts include enhancing heat transfer, reducing pressure loss and obtaining uniform temperature distribution. The purpose is to improve energy efficiency and guarantee the engine lifespan. In this work, both experiment and numerical simulation are employed to study pressure drop and heat transfer of various kinds of cooling configurations. Pin fin array, matrix and hybrid structures are investigated in a comparative study. Thermochromic liquid crystal technique is applied to obtain heat transfer distribution on the channel surface. The results show that matrix creates much stronger heat transfer than pin fin array with increased pressure loss penalty. Performances of matrix structures are quite different due to the configurations (dense or sparse). Hybrid structures are always worse than the baseline matrix in terms of average thermal performance, due to the higher pressure loss, however, heat transfer can be improved. The performance of hybrid structure depends on the arrangement and diameter of the pin fins. Pin fins in central area provide not only larger pressure loss but also stronger heat transfer than pin fins near the bend region. Cases with larger diameter result in the thermal performance degradation. Compared with sparse matrix, the hybrid structures can compensate for the lower heat transfer enhancement. As for the dense hybrid structures, the average heat transfer capacity can be improved with reasonable pin fin arrangement.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-21T06:13:01Z
      DOI: 10.1177/0954410020965401
       
  • Integrated guidance and control framework for the waypoint navigation of a
           miniature aircraft with highly coupled longitudinal and lateral dynamics
    • Authors: K Harikumar, Jinraj V Pushpangathan, Sidhant Dhall, M Seetharama Bhat
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A solution to the waypoint navigation problem for the fixed wing micro air vehicles (MAV) having a severe coupling between longitudinal and lateral dynamics, in the framework of integrated guidance and control (IGC) is addressed in this paper. IGC yields a single step solution to the waypoint navigation problem, unlike conventional multiple loop design. The pure proportional navigation (PPN) guidance law is integrated with the coupled MAV dynamics. A multivariable static output feedback (SOF) controller is designed for the linear state space model formulated in IGC framework. A waypoint navigation algorithm is proposed that handles the minimum turn radius constraint of the MAV and also evaluates the feasibility of reaching a waypoint. Non-linear simulations with and without wind disturbances are performed on a high fidelity 150 mm wingspan MAV model to demonstrate the proposed waypoint navigation algorithm.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-21T05:42:45Z
      DOI: 10.1177/0954410020964992
       
  • Advances in coupled axial turbine and nonaxisymmetric exhaust volute
           aerodynamics for turbomachinery
    • Authors: Jie Gao, Chunde Tao, Dongchen Huo, Guojie Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Marine, industrial, turboprop and turboshaft gas turbine engines use nonaxisymmetric exhaust volutes for flow diffusion and pressure recovery. These processes result in a three-dimensional complex turbulent flow in the exhaust volute. The flows in the axial turbine and nonaxisymmetric exhaust volute are closely coupled and inherently unsteady, and they have a great influence on the turbine and exhaust aerodynamic characteristics. Therefore, it is very necessary to carry out research on coupled axial turbine and nonaxisymmetric exhaust volute aerodynamics, so as to provide reference for the high-efficiency turbine-volute designs. This paper summarizes and analyzes the recent advances in the field of coupled axial turbine and nonaxisymmetric exhaust volute aerodynamics for turbomachinery. This review covers the following topics that are important for turbine and volute coupled designs: (1) flow and loss characteristics of nonaxisymmetric exhaust volutes, (2) flow interactions between axial turbine and nonaxisymmetric exhaust volute, (3) improvement of turbine and volute performance within spatial limitations and (4) research methods of coupled turbine and exhaust volute aerodynamics. The emphasis is placed on the turbine-volute interactions and performance improvement. We also present our own insights regarding the current research trends and the prospects for future developments.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-21T04:53:10Z
      DOI: 10.1177/0954410020966454
       
  • Numerical study of the pseudo-boiling phenomenon in the transcritical
           liquid oxygen/gaseous hydrogen flame
    • Authors: Hamed Zeinivand, Mohammad Farshchi
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The interactions and effects of turbulent mixing, pseudo-boiling phenomena, and chemical reaction heat release on the combustion of cryogenic liquid oxygen and gaseous hydrogen under supercritical pressure conditions are investigated using RANS simulations. Comparisons of the present numerical simulation results with available experimental data reveal a reasonably good prediction of a supercritical axial shear hydrogen-oxygen flame using the standard k-ε turbulence model and the eddy dissipation concept combustion model with a 23 reaction steps kinetics for H2-O2 reaction. The present simulation qualitatively reproduced oxygen injection and its reaction with the co-flowing hydrogen, which is characterized by rapid flame expansion, downstream flame propagation, and expansion induced flow recirculation. Several turbulence models were used for numerical simulations. It is shown that the selection of an appropriate turbulence model for transcritical reacting flows is crucial and far more important than for subcritical reacting flows. It is indicated that the pseudo-boiling phenomena is the main reason for the considerable differences between the turbulence models in a transcritical flame. Also, it is demonstrated that the liquid oxygen core disappears faster in a non-reacting flow than in a reacting flow. The shear layer in the non-reacting flow is much stronger than reacting case; providing a large transfer of energy from the outer layer to the inner layer. At the supercritical injection conditions, the difference between the turbulence models is much less than the transcritical injection conditions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-20T08:24:59Z
      DOI: 10.1177/0954410020964692
       
  • Influence of tip clearance and cavity depth on heat transfer in a cutback
           squealer tip
    • Authors: Weijie Wang, Shaopeng Lu, Hongmei Jiang, Qiusheng Deng, Jinfang Teng, and Wensheng Yu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Numerical simulations are conducted to present the aerothermal performance of a turbine blade tip with cutback squealer rim. Two different tip clearance heights (0.5%, 1.0% of the blade span) and three different cavity depths (2.0%, 3.0%, and 6.0% of the blade span) are investigated. The results show that a high heat transfer coefficient (HTC) strip on the cavity floor appears near the suction side. It extends with the increase of tip clearance height and moves towards the suction side with the increase of cavity depth. The cutback region near the trailing edge has a high HTC value due to the flush of over-tip leakage flow. High HTC region shrinks to the trailing edge with the increase of cavity depth since there is more accumulated flow in the cavity for larger cavity depth. For small tip clearance cases, high HTC distribution appears on the pressure side rim. However, high HTC distribution is observed on suction side rim for large tip clearance height. This is mainly caused by the flow separation and reattachment on the squealer rims.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-20T08:19:18Z
      DOI: 10.1177/0954410020964690
       
  • Analysis of rotor aerodynamic response during ramp collective pitch
           increase by CFD method
    • Authors: Kai Zhang, Qijun Zhao, Xiayang Zhang, Guohua Xu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In order to study the aerodynamic response characteristics of the helicopter rotor during ramp collective pitch increase, the moving-embedded grid technique is employed for numerical simulation. The governing equations are modeled via Navier-Stokes equations, as well as one equation S-A turbulence model. In order to improve the precision of unsteady simulation of the rotor flowfield, the three-order scheme known as Roe-WENO scheme is employed for the spatial discretization of convective fluxes, and the implicit LU-SGS scheme is adopted for the temporal discretization. The flowfield and aerodynamic characteristics of the SA349/2 Gazelle helicopter rotor are computed for verification, and thereafter, the present method is used to simulate the transient aerodynamic response of the rotor under different collective pitch increment rates. The unsteady flowfield and aerodynamic characteristics of the rotor under ramp collective pitch increase are obtained and compared with the experimental data. The results show that the numerical method not only can accurately predict the unsteady aerodynamic loads of the rotor in steady state, but also is capable of effectively simulating the transient aerodynamic response of the rotor, characterized by overshoot and delay phenomenon, during ramp collective pitch increase. Finally, the opposite ramp decrease in collective pitch and the influence of pre-twist on aerodynamic response are analyzed. The result shows that the transient aerodynamic response of the rotor under ramp collective pitch increase and decrease present a certain of symmetry. The change in pre-twist of blades only affects the thrust coefficient in steady state, while have little influence on the transient maneuvering process of collective pitch.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-20T07:13:39Z
      DOI: 10.1177/0954410020965404
       
  • Adaptive compliant controller for space robot stabilization in
           post-capture phase
    • Authors: Pengcheng Xia, Jianjun Luo, Mingming Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Safety and reliability are the primary prerequisites of space robotic manipulation. Due to the inaccurate inertial parameters of the tumbling target, tracking the desired trajectory directly will lead to the build up of large contact force and torque and damage the grasping point. The measurement noise in the contact wrenches will disturb the application of compliant stabilization strategy, and lead to mission failure. In order to coordinate the desired motion and contact, a compliant stabilization is required for realistic application. However, the measurement noise in the measured contact will disturb the application of compliant control scheme. According to these facts, herein, an adaptive compliant stabilization control scheme is proposed for a safe and reliable stabilization process. With the reference of the unsafe desired motion, a safe admittance motion is generated with an adaptive stiffness virtual spring. In consideration of the parameter selection and the presence of the contact wrenches measurement noise, a neural network-based coordinated adaptive impedance tracking controller is designed to track the safe motion and consume the transmitted energy from the tumbling target at the same time. With the benefit of the combination of the admittance motion and the coordinated adaptive impedance tracking controller, interactions at the grasping point can be controlled and the target can be stabilized under the influence of the measurement noise in the contact wrenches. Furthermore, safety and reliability of the proposed control scheme are validated via digital simulations.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-12T07:20:08Z
      DOI: 10.1177/0954410020964983
       
  • Towards a methodology for new technologies assessment in aircraft
           operating cost
    • Authors: Valeria Vercella, Marco Fioriti, Nicole Viola
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The need for a greener and competitive aircraft is leading to the use of new technologies. A thorough assessment of these technologies is mandatory from the initial phases of aircraft design to understand their feasibility and to select the most promising one both in terms of performances and in terms of costs. This paper proposes a methodology to assess the operating cost of innovative technologies for regional aircraft. In particular, two NASA studies have been adopted to determine the impact onto costs of MEA and AEA technologies and advanced ECS solutions for two innovative regional aircraft concepts developed during the European Clean Sky 2 research. The proposed methodology is able to assess the effect of on-board systems electrification level in terms of fuel and maintenance costs savings. The methodology, which allows to evaluate the effect of specific technological improvements onto costs, is applied exploiting the results provided by a reliable cost model and gives the opportunity to quantify operating cost savings for different regional aircraft. Applying the modified cost model to the reference aircraft under study, savings ranging from 1.6 to 3.1% of direct operating cost are estimated for MEA and AEA technologies. Greater savings are estimated for the individual cost items involved. More specifically, a reduction of fuel cost ranging from 6 to 14.5% is envisaged as a consequence of the lower SFC associated to innovative ECS technologies.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-10-12T07:06:06Z
      DOI: 10.1177/0954410020964675
       
  • Performance improvement of a high-speed aero-fuel centrifugal pump through
           active inlet injector
    • Authors: Jia Li, Xin Wang, Wancheng Wang, Yue Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper presents a high-speed aero-fuel centrifugal pump with an active inlet injector for an aero-engine aiming at regulating the internal flow field and improving overall hydraulic performance. Unlike most of the existing centrifugal pumps for aero-engines, an injector is designed and integrated with the pump to accomplish the active flow control. Firstly, by employing the energy equation in the pump, reasonable geometrical parameters of the injector are calculated. Then, a validation study is conducted with three known turbulence models, showing that simulations with the RNG κ-ε turbulence model can accurately predict the head and efficiency of the experimental pump. Finally, simulation results with the determined turbulence model are discussed. The results show that the static pressure is uniformly distributed inside the impeller, the volute and the injector. The flow field is significantly ameliorated by improving the pressure inside the suction pipe and controlling the flow direction via the injector. Furthermore, the head and efficiency of the designed pump with an active inlet injector are improved compared to the one without an injector.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-24T04:47:52Z
      DOI: 10.1177/0954410020960961
       
  • Sparse identification of nonlinear unsteady aerodynamics of the
           oscillating airfoil
    • Authors: Chong Sun, Tian Tian, Xiaocheng Zhu, Zhaohui Du
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Reduced-order models are widely used in aerospace engineering. A model for unsteady aerodynamics is desirable for designing the blades of wind turbines. Recently, sparse identification of nonlinear dynamics with control was introduced to identify the parameters of an input-output dynamical system. In this paper, two models for attached flows and one for separated flows are identified through this technique. For the unsteady lift of the attached flow, Model I is a linear model that presents the dynamic change of an unsteady lift to a static lift. Model II was built based on Model I in order to obtain a more general system with closed-loop control. It has a first-order inert element that delays the overall input of the static lift. The Model II results replicate the training data very well and give an accurate prediction of other oscillating cases with different oscillation amplitudes, reduced frequency or mean angle of attack. For the unsteady lift of the separated flow, Model III is identified as a nonlinear model, which also has a first-order inert element. This model captures the nonlinear aerodynamics of the separated flow and replicates the training cases well. In addition, the prediction of Model III has good agreement with the numerical results.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-23T09:07:29Z
      DOI: 10.1177/0954410020959873
       
  • DSMC simulation of rarefied gas flow over a 2D backward-facing step in the
           transitional flow regime: Effect of Mach number and wall temperature
    • Authors: Deepak Nabapure, Ram Chandra Murthy K
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Rarefied gas flow over a backward-facing step (BFS) is often encountered in separating flows prevalent in aerodynamic flows, engine flows, condensers, space vehicles, heat transfer systems, and microflows. Direct Simulation Monte Carlo (DSMC) is a powerful tool to investigate such flows. The purpose of this research is to assess the impact of Mach number and wall temperature on the flow and surface properties in the transitional flow regime. The Mach numbers considered are 5, 10, 25, 30, and the ratio of the temperature of the wall to that of freestream considered are 1, 2, 4, 8. The Reynolds number for the cases studied is 8.6, 17.2, 43, and 51.7, respectively. Typically the flow properties near the wall are found to increase with both Mach number and wall temperature owing to compressibility and viscous dissipation effects. The variation in flow properties is more sensitive to Mach number than the wall temperature. The surface properties are found to decrease with Mach number and increase with wall temperature. Moreover, in the wake of the step, the vortex’s recirculation length is reasonably independent of both free stream Mach number and wall temperature, whereas it decreases with Knudsen number.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-23T09:02:47Z
      DOI: 10.1177/0954410020959872
       
  • Fuel efficiency optimization of high-aspect-ratio aircraft via variable
           camber technology considering aeroelasticity
    • Authors: Liqiang Guo, Jun Tao, Cong Wang, Miao Zhang, Gang Sun
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this study, variable camber technology is applied to improve the fuel efficiency of high-aspect-ratio aircraft with aeroelasticity considered. The nonlinear static aeroelastic analyses are conducted for CFD/CSD (computational fluid dynamics/computational structural dynamics) numerical simulations. The RBF (radial basis function) method is adopted for the transmission of aerodynamic loads and structural displacements, the diffusion smoothing method is employed for grid deformation in each iteration of CFD/CSD coupling, and the FFD (free-form deformation) method is introduced for the parameterization of variable camber wing. Based on the aerodynamic characteristic curves under different cambers, the discrete variable camber control matrix for the high-aspect-ratio aircraft during the cruise phase is established. The Fibonacci method is employed to optimize the fuel efficiency by utilizing the control matrix. The results indicate that the drag during the cruise phase is reduced obviously and the fuel efficiency is improved evidently comparing to the original configuration.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-21T10:46:29Z
      DOI: 10.1177/0954410020959964
       
  • Experimental evaluation and numerical simulation of performance of the
           bypass dual throat nozzle
    • Authors: Mohammad Hadi Hamedi-Estakhrsar, Hossein Mahdavy-Moghaddam
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Bypass dual throat nozzle (BDTN) is a modern concept of fluidic thrust vector control. This method able to solve the problem of thrust loss without need the secondary mass flow from other part of engine. Internal nozzle performance and thrust vector angles have been measured in the BDTN experimentally and numerically. A new simple approach is proposed to detect the thrust deflection angle. Numerical simulation of 3-D turbulent air flow is carried out by using the RNG k-e turbulence model. The obtained results of thrust coefficient, discharge coefficient and thrust deflection angle have been validated by comparing with measured experimental data. The results show that for nozzle pressure ratio of 1–4 the tested nozzle able to deflect the thrust vector of 26.5°-19°. By increasing NPR from 2 up to 4, the thrust coefficient values will change in the range of 0.85-0.93. Also the effect of different positions of the bypass channel on the BDTN performance parameters has been investigated numerically. The predicted results show that the BDTN configuration with bypass duct on the first nozzle throat has the highest value of thrust deflection angle over the range of NPRs.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-21T10:44:09Z
      DOI: 10.1177/0954410020959886
       
  • Attitude control of nanosatellite with single thruster using relative
           displacements of movable unit
    • Authors: Anton V Doroshin, Alexander V Eremenko
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The attitude dynamics of a nanosatellite (NS) with one movable unit changing its angular position relative to a main body of the nanosatellite is considered. This relative movability of the unit can be implemented with the help of flexible rods of variable length connecting the unit with the main body. Change of the relative position of the movable unit shifts the center of mass of the entire mechanical system. The NS has a single jet engine rigidly mounted into the NS main body. The shift of the mass center creates an arm of the jet-engine thrust and a corresponding control torque. Schemes to control the attitude dynamics of the satellite using the movability of its unit are developed, using both the torque from the engine and inertia change.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-21T05:40:10Z
      DOI: 10.1177/0954410020959868
       
  • The influence of yaw on the unsteady surface pressures over a two-wheeled
           landing-gear model
    • Authors: WR Graham, A Gatto
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Landing-gear noise is an increasing issue for transport aircraft. A key determinant of the phenomenon is the surface pressure field. Previous studies have described this field when the oncoming flow is perfectly aligned with the gear. In practice, there may be a cross-flow component; here its influence is investigated experimentally for a generic, two-wheel, landing-gear model. It is found that yaw angles as small as 5° cause significant changes in both overall flow topology and unsteady surface pressures. Most notably, on the outboard face of the leeward wheel, large-scale separation replaces predominantly attached flow behind a leading-edge separation bubble. The effect on unsteady surface pressures includes marked shifts in the content at frequencies in the audible range, implying that yaw is an important parameter for landing-gear noise, and that purely unyawed studies may not be fully representative of the problem.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-21T01:37:06Z
      DOI: 10.1177/0954410020959881
       
  • Compressible flow characteristics in bent duct with constant flow section
    • Authors: Xiao-lin Sun, Shan Ma, Zhan-xue Wang, Jing-wei Shi, Li Zhou
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The components with large curvature features are widely applied in aero-engines. Complex flow features are induced due to large curvature under high-subsonic even transonic incoming flow condition. In this study, the formation mechanisms of local acceleration in bent duct are investigated. To this end, the cold fluid test of a 60° bent duct with constant flow section was conducted. The surface static pressures and the schlieren flow visualizations were obtained. Then the three-dimensional numerical simulations based on the experimental model were computed using computational fluid dynamics software. The simulations were conducted using five different turbulence models to compare with the experimental data. The validation study shows that the shear stress transfer (SST) κ-ω turbulence model is suitably used for the simulations. Results show that three different flow situations were shown for the bent duct at diverse nozzle pressure ratios (NPRs). One situation was shown by the case at NPR = 1.5, in which the whole flow field is subsonic, and just two jet edges are shown by the schlieren images. One situation was shown by the case at NPR = 1.8, in which a local supersonic region is induced near the lower wall at the hind side of the bent section, and a small shock wave is observed. The other one situation was shown by the cases at NPR = 2.0, 2.5 and 3.0, in which the air flow in the whole passage reaches supersonic speeds and an oblique shock wave is shown for each case.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-14T04:46:02Z
      DOI: 10.1177/0954410020958592
       
  • Hit-to-kill accurate minimum time continuous second-order sliding mode
           guidance for worst-case target maneuvers
    • Authors: Jinraj V Pushpangathan, Harikumar Kandath, Ajithkumar Balakrishnan
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The recent research is focused on the development of an advanced interceptor missile that has hit-to-kill accuracy against ballistic targets performing evasive maneuvers. In this paper, a guidance law that achieves hit-to-kill accuracy against ballistic target executing worst-case maneuvers is developed using second-order sliding mode control and optimal control. The guidance law thus developed is continuous and has minimum time convergence for worst-case target maneuvers. The performance of the continuous guidance law with minimum time convergence is evaluated through numerical simulations against ballistic targets executing step maneuvers with changing polarity and weaving maneuvers.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-04T07:47:06Z
      DOI: 10.1177/0954410020954977
       
  • Linear amplification factor transport equation for stationary crossflow
           instabilities in supersonic boundary layers
    • Authors: Jiakuan Xu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Based on the database from linear stability theory (LST) analysis, a local amplification factor transport equation for stationary crossflow (CF) waves in low-speed boundary layers was developed in 2019. In this paper, the authors try to extend this transport equation to compressible boundary layers based on local flow variables. The similarity equations for compressible boundary layers are introduced to build the function relations between non-local variables and local flow parameters. Then, compressibility corrections are taken into account to modify the source term of the transport equation. Through verifications of different sweep angles, Reynolds numbers, angles of attack, Mach numbers, and different cross-section geometric shapes, the rationality and correctness of the new transport equation established in this paper are illustrated.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-09-02T07:23:27Z
      DOI: 10.1177/0954410020954999
       
  • Transonic flutter characteristics of an airfoil with morphing devices
    • Authors: Shun He, Shijun Guo, Wenhao Li
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An investigation into transonic flutter characteristic of an airfoil conceived with the morphing leading and trailing edges has been carried out. Computational fluid dynamics (CFD) is used to calculate the unsteady aerodynamic force in transonic flow. An aerodynamic reduced order model (ROM) based on autoregressive model with exogenous input (ARX) is used in the numerical simulation. The flutter solution is determined by eigenvalue analysis at specific Mach number. The approach is validated by comparing the transonic flutter characteristics of the Isogai wing with relevant literatures before applied to a morphing airfoil. The study reveals that by employing the morphing trailing edge, the shock wave forms and shifts to the trailing edge at a lower Mach number, and aerodynamic force stabilization happens earlier. Meanwhile, the minimum flutter speed increases and transonic dip occurs at a lower Mach number. It is also noted that leading edge morphing has negligible effect on the appearance of the shock wave and transonic flutter. The mechanism of improving the transonic flutter characteristics by morphing technology is discussed by correlating shock wave location on airfoil surface, unsteady aerodynamics with flutter solution.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-31T04:44:04Z
      DOI: 10.1177/0954410020953046
       
  • Fuzzy PD hybrid control of low frequency vibration of annular antenna
    • Authors: Xinghui Zhai, Yajun Luo, Yahong Zhang, Shilin Xie
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The annular antenna is a typical large flexible space truss structure featured by small damping and low modal frequencies. The external disturbances, e.g. impulse load resulting from satellite attitude adjusting, may induce the low frequency large amplitude vibrations of annular antenna for a long time and thus reduce working precision and cause even its damage. The active control of vibration of annular antenna under impulse excitation is investigated in the paper. The voice coil actuator instead of piezoelectric stack actuator is used in order to meet the demand of large output displacement. The governing equation of active vibration control system is established by use of finite element method. The proportional differential (PD) control and fuzzy control algorithms are firstly studied in the active control. The results show that the fuzzy control exhibits worse control performance than PD control due to weak control function near structural equilibrium position. To circumvent the drawback of fuzzy control, a fuzzy PD hybrid control strategy is proposed which can combine the merits of both control methods. The simulated and experimental results show that the fuzzy PD hybrid control can yield the best control effect under impulse excitation comparing with the PD control and ordinary fuzzy control. The work provides a promising control way for active control of low frequency and large displacement vibration of annular antenna in satellite engineering.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-28T06:01:11Z
      DOI: 10.1177/0954410020955005
       
  • A decentralized method for collision detection and avoidance applied to
           civil aircraft
    • Authors: Haotian Niu, Cunbao Ma, Pei Han
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      With the increasing density level of airspace, the flawed logic of resolution in air conflict has become a potential hazard to keep flight safety for civil aviation. A powerful decision-support system is needed to identify and resolve potential conflicts on planned trajectory in advance. Existing studies on this subject mainly focus on the centralized means, but seldom consider the decentralized approaches. In this paper, a decentralized method is proposed so that each aircraft can generate the collision-free Reference Business Trajectory (RBT) autonomously, and resolve potential conflicts while conforming to the unified rules. Firstly, a Synchronous Discrete-Time-Discrete-Space trajectory modeling is developed to divide the continuous planned trajectory into multiple trajectory segments according to motion state. Thus, the collision can be accurately located at one certain risky segment, and the corresponding collision time can be acquired precisely. Through a weight analysis of collision time, the critical trajectory segment is determined to implement the task of conflict resolution. Then, the Optimal Reciprocal Collision Avoidance (ORCA) algorithm is adopted and extended to determine the collision-free maneuver with the consideration of direction selectivity. At last, the Trajectory Change Points (TCPs) are achieved by the quadratic program for each aircraft. The proposed method can help aircraft generate collision-free RBT in decentralized way successfully. Several simulations are conducted to confirm the validity and efficiency of the proposed approach.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-28T05:58:31Z
      DOI: 10.1177/0954410020953045
       
  • Adaptive fault-tolerant control for hybrid attitude tracking control
           system with quantized control torque and measurement
    • Authors: Min Li, Yingchun Zhang, Yunhai Geng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, the problem of fault tolerant control for spacecraft attitude tracking control system in the presence of actuator faults/failures, quantized control torque and measurement, uncertain inertial matrix and external disturbances is taken into account. The dynamical uniform quantizers are developed to quantize the signals of control torque and measurement, which can reduce the data transmission rate. In combination with the CA and FTC technique, a robust adaptive fault tolerant control scheme is proposed to cope with the effects of quantization errors in control torque and measurement, the unknown actuator faults/failures, uncertain inertial matrix and external disturbances. The developed control strategy combined with quantized control torque and measurement can guarantee the stability of overall closed-loop system and achieve satisfactory attitude tracking performance. Finally, simulation results are presented to verify the effectiveness of the proposed methods.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-28T02:48:52Z
      DOI: 10.1177/0954410020953303
       
  • Periodic flow structures in a turbofan fan stage in windmilling
    • Authors: Nicolás García Rosa, Adrien Thacker, Guillaume Dufour
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In a fan stage under windmilling conditions, the stator operates under negative incidence, leading to flow separation, which may present an unsteady behaviour due to rotor/stator interactions. An experimental study of the unsteady flow through the fan stage of a bypass turbofan in windmilling is proposed, using hot-wire anemometry. Windmilling conditions are reproduced in a ground engine test bed by blowing a variable mass flow through a bypass turbofan in ambient conditions. Time-averaged profiles of flow coefficient are independent of the mass flow, demonstrating the similarity of velocity triangle. Turbulence intensity profiles reveal that the high levels of turbulence production due to local shear are also independent of the inlet flow. A spectral analysis confirms that the flow is dominated by the blade passing frequency, and that the separated regions downstream of the stator amplify the fluctuations locked to the BPF without adding any new frequency. Phase-locked averaging is used to capture the periodic wakes of the rotor blades at the rotor/stator interface. A spanwise behaviour typical of flows through windmilling fans is evidenced. Through the inner sections of the fan, rotor wakes are thin and weakly turbulent, and the turbulence level remains constant through the stage. The rotor wakes thicken and become more turbulent towards the fan tip, where flow separation occurs. Downstream of the stator, maximum levels of turbulence intensity are measured in the separated flow. Large periodical zones of low velocity and high turbulence intensity are observed in the outer parts of the separated stator wake, confirming the pulsating motion of the stator flow separation, locked at the blade passing frequency. Space-time diagrams show that the flow is chorochronic, and a 2 D non-linear harmonic simulation is able to capture the main interaction modes, however, the stator incidence distribution could be affected by 3 D effects.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-28T02:47:53Z
      DOI: 10.1177/0954410020948297
       
  • Numerical investigation of aerodynamic characteristics of free-spinning
           tail projectile with canards roll control
    • Authors: Jiawei Zhang, Juanmian Lei, Jianping Niu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      To reduce aerodynamic coupling between the canards and the tail fins of a canard-controlled projectile, the afterbody of the projectile is decoupled from the forebody by a bearing structure, namely, a free-spinning tail. A series of numerical simulations was conducted for different angles of attack using NASA’s canard-controlled projectile with a free-spinning tail. The results were then compared with the wind tunnel test data. The spin rate of the free-spinning tail shows that, with the canard roll control, the tail section will rotate at lower angles of attack and “lock-in” at higher ones, demonstrating nonlinearization between the rotating rate and the angle of attack. According to a flow structure analysis, the circular flow velocity induced by canards is responsible for the non-linear characteristics of the tail. Moreover, the change in position of the circular flow velocity results in a reverse of the rolling moment of the “+” fixed tail projectile at different angles of attack. Furthermore, a comparison of the aerodynamic characteristics of the fixed (“+” and “x”) and free-spinning tail configurations proves that when the tail is spinning, all the aerodynamic coefficients of the free-spinning tail projectile are between those of the “+” and “x” fixed tail projectiles. The longitudinal difference in aerodynamic characteristics is related to the rolling angle, whereas the lateral difference is related to both the rolling angle and rotation rate. When the tail section “locks-in,” different rolling angles lead to different characteristics in both the longitudinal and lateral directions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-27T10:22:58Z
      DOI: 10.1177/0954410020953316
       
  • Alpha-SIM: A quick 3D geometry model simplification approach to support
           aircraft EWIS routing
    • Authors: Zaoxu Zhu, G La Rocca, Yao Zheng, Jianjun Chen
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Routing design of aircraft Electrical Wiring Interconnection System (EWIS) is time-consuming and error-prone. A solution, which automatically routes the EWIS inside the aircraft Digital MockUp (DMU), has been proposed and presented in the previous publications. The DMU, however, includes over-detailed features, which hardly influence the routing results but significantly increase the geometry-involved computational time thus hampering any automated routing. These features cannot be easily and efficiently suppressed. Therefore, a quick 3 D geometry simplification method, named Alpha-SIM, is proposed to enable a quick simplification of the airframe components included in the DMU and improve the benefit of the aforementioned automatic EWIS routing approach. The method is inspired by Descriptive Geometry techniques and the 3 D modelling approach using 2 D sketches, and aims at removing very detailed and/or internal features while preserving the intuitive notional shape of the given CAD model. The intuitive notional shape is represented by a 3 D point cloud of the model outer boundary and their 2 D projections on user-defined planes. These 2 D projections are then processed such to generate a set of 2 D profiles, called Alpha-Shapes, which are used, eventually, to re-build the 3 D model of the DMU components in a simplified/de-featured manner. By controlling the density of the 3 D points and the Alpha value to generate the 2 D profiles from the point projections, various geometric approximation levels can be achieved. The results of the test cases demonstrate the efficiency and effectiveness of the proposed method on the geometry simplification for automatic EWIS routing.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-26T08:12:47Z
      DOI: 10.1177/0954410020952922
       
  • Geomagnetic signal de-noising method based on improved empirical mode
           decomposition and morphological filtering
    • Authors: Hongqi Zhai, Lihui Wang, Qingya Liu, Nan Qiao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      To solve the problem that geomagnetic signals are susceptible to random noise and instantaneous pulse interference in geomagnetic navigation, a geomagnetic signal de-noising method based on improved empirical mode decomposition (IEMD) and morphological filtering (MF) is proposed. The instantaneous pulse interference is eliminated by designing different structural elements according to the characteristics of the pulse signal. The signal after filtering the instantaneous pulse interference is decomposed by EMD, and the intrinsic mode functions (IMFs) obtained from the decomposition are determined as two modes (i.e. noise IMFs and mixed IMFs) by the cross-correlation coefficient criterion. The noise IMFs are removed directly, and a normalized least means square filter (NLMS) is designed to remove noise from mixed IMFs, which can adaptively adjust the filtering parameters according to the noise level of different IMF components. The noise-reduced mixed IMFs and residual are reconstructed to obtain the final geomagnetic signal. Experiment results illustrate that the proposed MF-IEMD method can effectively achieve noise reduction. Comparing with the traditional EMD and MF-EMD de-noising methods, the root mean square errors(RMSE) decreased by 49.27% and 24.79%, respectively.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-25T10:47:32Z
      DOI: 10.1177/0954410020951022
       
  • Investigation on the flow-control strategy for an aggressive turbine
           transition ducts
    • Authors: Jun Liu, Hongrui Liu, Guang Liu, Qiang Du, Pei Wang, Sheng Chang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      After being studied for years, aggressive intermediate turbine duct is being attempted to be applied in turbine design to further improve the engine-performance. With such design, the shaft could be shortened effectively. However, under the influence of the more distorted coming-flow and stronger pressure-gradient in a real engine, the flow field would be more complicated definitely. Besides that, the upstream-rotor tip-leakage flow is a key loss-source by inducing separation. Flow-control strategies are necessary in this situation. In this paper, the flow field in an aggressive duct has been analyzed to declare the source of separation primarily. Then wide-chord blade design concept has been adopted as a control strategy firstly to realize the purpose of improving the areo-performance. After being verified, numerical method has been used in this study. Under the same aero-condition, the prototype and the modified turbine are analyzed. With this novel flow-control strategy, separation has been improved, even diminished. However, the flow structures within the blade passage are altered correspondingly. An instrumental conclusion is that the pressure loss could be decreased successfully by designing the wide-chord blade specially.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-25T10:47:27Z
      DOI: 10.1177/0954410020950848
       
  • Cooling structure design of gas turbine blade by using multi-level highly
           efficient design platform
    • Authors: Zhiqi Zhao, Lei Luo, Shouzuo Li, Dandan Qiu, Songtao Wang, Zhongqi Wang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, the internal cooling structure of a second-stage rotor blade is designed by using a multi-level highly efficient design platform. The design process is divided into schematic design and detailed design in sequence. The calculations of pipe-network and heat conduction are presented to preliminary evaluate the cooling structures derived from the schematic design stage. The flow field and heat transfer characteristics of the revised cooling structures are analyzed in the detailed design by using the three-dimensional conjugated heat transfer calculation method. Topological structure, mass flow rate, pressure distribution, heat transfer coefficient and temperature distribution of the cooling channels are presented. It is found that the schematic design results based on one-dimensional to three-dimensional solution method are in good agreement with the detailed design results. Meanwile, the introduction of the schematic design is helpful to shorten the cooling design cycle and reduce the dependence of the design experience. In this work, a five-pass serpentine passage with single cooling air inlet in the cooling system may lead to low flow rate at the trailing edge, which is prone to cause hot gas back-flow and local high heat load. The cooling system with a right-angle channel and a three-pass serpentine channel helps to distribute the flow reasonably and reduce the thermal gradient on the blade surface. The optimal cooling structure meet the requirements well. Compared with the uncooled blade, the average temperature of the blade decrease over 530 K with limited cooling air.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-21T09:44:00Z
      DOI: 10.1177/0954410020951680
       
  • Investigation on tip clearance control for the high-pressure rotor of an
           uncooled vaneless counter-rotating turbine
    • Authors: Wei Zhao, Xiuming Sui, Kai Zhang, Zeming Wei, Qingjun Zhao
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In order to develop a tip clearance control system for an uncooled vaneless counter-rotating turbine, tip clearance variation of its high pressure rotor blade at off-design conditions is analyzed. Aero-thermal interaction simulation is performed to predict the temperature and deformation of the solid blade. At operating conditions with rotating speeds greater than 60% design value and expansion ratios greater than 85% design value, the blade tip clearance height at leading edge remains unchanged when the expansion ratio decreases, meanwhile that at trailing edge decreased obviously. However, the tip clearance height variations at the leading edge and trailing edge are almost the same in a conventional subsonic turbine at such conditions. The cause is that the flow in the high-pressure rotor is choked at these conditions. The choked flow results in that the fluid and solid blade temperatures upstream of the throat are not affected by the back pressure and only those downstream of the throat increases with the back pressure. Consequently, the blade height at leading edge keeps constant, and that at trailing edge varies because of thermal expansion. To avoid the rubbing of the blade and case, the blade height at trailing edge is diminished by 30%. As a result, the blade tip clearance height at low speed operating conditions increases in axial direction. Such a design leads to a stronger tip leakage flow. More flow losses might be generated. Therefore, a casing cooling method is proposed to control the blade tip clearance height at leading edge and trailing edge respectively. The deformations of the casing with different mass flow rate of cooling air at design and off-design conditions are calculated. It shows that the blade tip clearance heights at leading edge and at trailing edge of the rotor can be well controlled with appropriate amount of cooling air.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-19T05:32:31Z
      DOI: 10.1177/0954410020950509
       
  • Gradient-like minimization methods for aeroengines diagnosis and control
    • Authors: L Sánchez de León, J Rodrigo, JM Vega, JL Montañés
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Nowadays, there is an ever growing interest for gas turbine and aeroengines prognostics. The capability to assess not only the current state of an asset, but also to be able to predict its remaining useful life (RUL), and hence to perform condition-based maintenance (CBM) —if, and only when, it is needed— can represent a huge deal in the manufacturer profits. Against the plethora of data-driven methods that have arisen in the past few years, there is still some knowledge to be gained in terms of understanding the underlying phenomenology of engine degradation. In fact, it is certainly a non-trivial problem, to realize what has happened to the rotating components of an engine just by observing the pressure being measured by certain sensor rise, or some other temperature measured along the main gas-path decrease its value. In this regard, model-based approaches —and, in particular, gas path analysis (GPA)— can assist us in gaining such knowledge. In this paper, a non-linear GPA technique is revisited, introducing some novelties to the solver, and making use of current computational methods and resources, to establish a solid ‘foundation’ that will serve as the basis for further research.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-18T07:15:31Z
      DOI: 10.1177/0954410020946991
       
  • New atmospheric data model for constant altitude accelerated flight
           performance prediction calculations and flight trajectory optimization
           algorithms
    • Authors: Radu I Dancila, Ruxandra M Botez
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article presents a new method for storing and computing the atmospheric data used in time-critical flight trajectory performance prediction calculations, such as flight performance prediction calculations in flight management systems and/or flight trajectory optimization, of constant altitude cruise segments. The proposed model is constructed based on the forecast data provided by Meteorological Service Agencies, in a GRIB2 data file format, and the set of waypoints that define the lateral component of the evaluated flight profile(s). The atmospheric data model can be constructed/updated in the background or off-line, when new atmospheric prediction data are available, and subsequently used in the flight performance computations. The results obtained using the proposed model show that, on average, the atmospheric parameter values are computed six times faster than through 4D linear interpolations, while yielding identical results (value differences of the order of 10e−14). When used in flight trajectory performance calculations, the obtained results show that the proposed model conducts to significant computation time improvements. The proposed model can be extended to define the atmospheric data for a set of cruise levels (usually multiple of 1000 ft).
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-14T07:30:44Z
      DOI: 10.1177/0954410020945555
       
  • Experimental and computational investigation on comparison of micro-scale
           open rotor and shrouded rotor hovering in ground effect
    • Authors: Han Han, Changle Xiang, Bin Xu, Yong Yu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Various investigations on open rotor (OR) hovering in-ground effect (IGE) are carried out, but few papers report shrouded rotor (SR) hovering IGE. This paper compares aerodynamic performance and flowfield characteristics of OR and SR hovering IGE by both experimental measurements and computational fluid dynamic (CFD) simulations. Experimental results reveal that in IGE flight, the aerodynamic performance of SR is more sensitive than that of OR. And at constant power, SR offers more thrust than OR at the same ground distance. Ground has a great influence on thrust for OR below 2.2 rotor radius distance, while for SR it shows obvious effect below 1.5 rotor radius distance. It is also shown that normalized aerodynamic coefficients of OR and SR are independent on rotor speed. In addition, for OR the rotor thrust coefficient changes nearly linearly with the logarithmic distance from ground, while for SR it changes nonlinearly. Flowfield analysis by CFD shows that shroud changes the tip flow features and expands the slipstream area of SR. When ground distance gets small, back pressure below the rotor-disk plane increases, which is more obvious for SR than OR. Furthermore, shroud thrust of SR decreases because of tip leakage flow and flow separation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-13T07:12:59Z
      DOI: 10.1177/0954410020949292
       
  • The genetic algorithm-radial basis function neural network to quickly
           predict aerodynamic performance of compressors
    • Authors: Tianquan Tang, Bo Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this paper, compressor aerodynamic performance has been predicted based on throughflow theory, combined with a surrogate model, which is a combination of the Genetic Algorithm (GA) and generalized Radial Basis function (RBF) neural network. And the predicting results have been compared with those from the traditional models and spanwise mixing model, which still widely be used to predict the aerodynamic performance. We first predicted the deviation angle and total-pressure loss coefficient (TPLC) by the surrogate model, and then using these two intermediate variables connected the model with throughflow theory. The pressure ratio and efficiency, representing the compressors’ total performance parameters, are predicted and compared with experimental data. In order to increase the accuracy of prediction, a data augmentation method based on the piecewise cubic Hermite interpolation (PCHIP) algorithm is introduced to enlarge the training database. At the same time, considering the vast differences of deviation angle and loss in different working conditions as well as aerodynamic and geometric differences of rotor and stator, the database and the network should be split into six components based on the choke, the normal and the stall conditions as well as rotor and stator. Then, the performance curves of pressure ratio and efficiency can be determined by an iteration process. The predicting results are compared with experimental data, which shows that the surrogate model matches experiments much better than those from the traditional models and spanwise mixing model.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-13T07:12:29Z
      DOI: 10.1177/0954410020948977
       
  • Harnaś-3, new generation of aerobatic airplane, comprehensive
           structure strength analysis
    • Authors: Wojciech Grendysa, Marek Jonas
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The object of the static strength analysis presented in the following paper is an aerobatic airplane Harnaś-3. It is a new generation of an aerobatic airplane and its unusual arrangement makes it possible to make aerobatic maneuvers that are not possible to do by other airplanes. The untypical arrangement of the aerobatic plane Harnaś-3 causes that the strength analysis of its structure is particularly complex. A spatially developed structure requires a comprehensive approach, taking into account both the specific properties of composite materials and the need to analyze the strength ratio for various cases of external loads, appropriate for aviation regulations. The methodology presented in this article allowed to improve the structure of the Harnaś-3 aircraft to reach the weight of a complete structure of only 235 kg, which allows building an aircraft lighter than the competitors.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-06T03:01:12Z
      DOI: 10.1177/0954410020948642
       
  • Numerical study of secondary mass flow modulation in a Bypass Dual-Throat
           Nozzle
    • Authors: MH Hamedi-Estakhrsar, M Ferlauto, H Mahdavy-Moghaddam
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The fluidic thrust-vectoring modulation on a Bypass Dual-Throat Nozzle (BDTN) is studied numerically. The thrust vectoring modulation is obtained by varying the secondary mass flow, introducing different area contraction ratios of the bypass duct. The scope of present study is twofold: (i) to set up a model for the control of the secondary mass flow that is consistent with the resolution of the nozzle main flow and (ii) to derive a simplified representation of a valve system embedded in the bypass channel. The simulations of the turbulent airflow inside the BDTN and its efflux in the external ambient have been simulated by using RANS approach with RNG [math] turbulence modeling. The numerical results have been validated with experimental and numerical data available in the open literature. The nozzle performance and thrust vector angle are computed for different values of the bypass area contraction ratio. The effects of different secondary mass flow rates on the system resultant thrust ratio and discharge coefficient of the bypass dual-throat nozzle have been investigated. By using the proposed approach to the secondary mass flow modulation, the thrust pitch angle has been controlled up to 27°.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-06T03:00:10Z
      DOI: 10.1177/0954410020947920
       
  • Experimental investigation on the structures and induced drag of wingtip
           vortices for different wingtip configurations
    • Authors: Ze-Peng Cheng, Yang Xiang, Hong Liu
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      As an effective method to reduce induced drag and the risk of wake encounter, the winglet has been an essential device and developed into diverse configurations. However, the structures and induced drag, as well as wandering features of the wingtip vortices (WTVs) generated by these diverse winglet configurations are not well understood. Thus, the WTVs generated by four typical wingtip configurations, namely the rectangular wing with blended/raked/split winglet and without winglet (Model BL/RA/SP/NO for short), are investigated in this paper using particle image velocimetry technology. Comparing with an isolated primary wingtip vortex generated by Model NO, multiple vortices are twisted coherently after installing the winglets. In addition, the circulation evolution of WTVs demonstrates that the circulation for Model SP is the largest, while Model RA is the smallest. By tracking the instantaneous vortex center, the vortex wandering behavior is observed. The growth rate of wandering amplitude along with the streamwise location from the quickest to the slowest corresponds to Model SP, Model NO, Model BL, Model RA in sequence, implying that the WTVs generated by model SP exhibit the quickest mitigation. Considering that the induced drag scales as the lift to power 2, the induced drag and lift are estimated based on the wake integration method, and then the form factor λ, defined by [math], is calculated to evaluate the aerodynamic performance. Comparing with the result of Model NO, the form factor decreases by 7.99%, 4.80%, and 2.60% for Model RA, Model BL, Model SP, respectively. In sum, Model RA and BL have a smaller induced drag coefficient but decay slower; while Model SP has a larger induced drag coefficient but decays quicker. An important implication of these results is that reducing the strength of WTVs and increasing the growth rate of vortex wandering amplitude can be the mutual requirements for designing new winglets.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-05T09:13:17Z
      DOI: 10.1177/0954410020947911
       
  • Systematic reduced order model development of a pitching NACA0012 airfoil
    • Authors: Jaclynn Mohrfeld Halterman, Mesbah Uddin
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Simple reduced order models (ROMs) for the aerodynamic coefficients - lift, drag, and pitch moment - of a pitching NACA0012 airfoil are presented. The ROMs are designed for quick computation of the transient aerodynamic characteristics of the airfoil and are developed utilizing computational fluid dynamics (CFD) simulation results. The entire aerodynamic system is modeled as a single input, multi output system yielding three independent systems to be characterized. A systematic, two step process is employed to develop the ROMs for each aerodynamic system. First, a CFD simulation is conducted to determine the linearity of each system, and any nonlinear system is restructured as a nonlinear operator followed by a linear system to allow for the use of linear system identification techniques. A second CFD simulation is conducted to determine the frequency response of each linear system, and the coefficients of each ROM are extracted by fitting a second order model to each frequency response function. The ROMs are validated against an independent CFD simulation of a pitching airfoil and are shown to accurately model each aerodynamic coefficient.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-04T12:13:53Z
      DOI: 10.1177/0954410020946912
       
  • Effects of smart flap on aerodynamic performance of sinusoidal
           leading-edge wings at low Reynolds numbers
    • Authors: AA Mehraban, MH Djavareshkian, Y Sayegh, B Forouzi Feshalami, Y Azargoon, AH Zaree, M Hassanalian
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Sinusoidal leading-edge wings have shown a high performance after the stall region. In this study, the role of smart flaps in the aerodynamics of smooth and sinusoidal leading-edge wings at low Reynolds numbers of 29,000, 40,000 and 58,000 is investigated. Four wings with NACA 634-021 profile are firstly designed and then manufactured by a 3 D printer. Beam bending equation is used to determine the smart flap chord deflection. Next, wind tunnel tests are carried out to measure the lift and drag forces of proposed wings for a wide range of angles of attack, from zero to 36 degrees. Results show that using trailing-edge smart flap in sinusoidal leading-edge wing delays the stall point compared to the same wing without flap. However, a combination of smooth leading-edge wing and smart flap advances the stall. Furthermore, it is found that wings with smart flap generally have a higher lift to drag ratio due to their excellent performance in producing lift.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-08-04T12:07:17Z
      DOI: 10.1177/0954410020946903
       
  • Post-stall flight dynamics of commercial transport aircraft configuration:
           A nonlinear bifurcation analysis and validation
    • Authors: Fei Cen, Qing Li, Zhitao Liu, Lei Zhang, Yong Jiang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Loss-of-control has become the largest fatal accident category for worldwide commercial jet accidents, and any initiative aimed at preventing such events requires an understanding of the fundamental aircraft behavior, especially the flight dynamics at post-stall region at which loss-of-control usually occurred. A series of low-speed static and dynamic wind tunnel tests of the Common Research Model over a large angle of attack/sideslip envelope was conducted and a non-linear aerodynamic model was developed. The bifurcation analysis, complemented by time-history simulation was used to understand the post-stall flight dynamics and the numerical analysis results were preliminary validated by wind tunnel virtual flight test. Several representative post-stall behaviors for the transport aircraft have been identified, including departure, periodic oscillation, post-stall gyration and steep spiral, etc. Furthermore, the predicted periodic oscillation in pitch motion has been perfectly duplicated in wind tunnel virtual flight test. The approach used in this work shows a promising way to uncover the flight dynamics of transport aircraft at extreme and loss-of-control flight conditions, as well as to apply to nonlinear unsteady aerodynamics modeling and validation, flight accident investigation, advanced flight control law design or studying initiative for loss-of-control prevention or mitigation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-29T10:35:21Z
      DOI: 10.1177/0954410020944085
       
  • Multi-unmanned aerial vehicle multi acoustic source localization
    • Authors: Suresh Manickam, Sufal Chandra Swar, David W Casbeer, Satyanarayana Gupta Manyam
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper addresses a multisource localization problem with multiple unmanned aerial vehicles equipped with appropriate sensors coordinating with each other, wherein the sources are simultaneously emitting identical acoustic signals. Distributed coordinated localization algorithms based on multiple range and direction measurements are presented and performances are evaluated in different practically significant mission scenarios. Non-deterministic polynomial (NP) hardness to determine optimal number of unmanned aerial vehicles for a given mission scenario is discussed. Group coordination, tactical path, and goal replan strategies to enable efficient localization of single and multiple acoustic sources have been designed. The localization algorithm along with coordination strategy is verified in the presence of realistic error conditions through simulation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-29T10:28:25Z
      DOI: 10.1177/0954410020943086
       
  • Effect of flight/structural parameters and operating conditions on dynamic
           behavior of a squeeze-film damped rotor system during diving–climbing
           maneuver
    • Authors: Xi Chen, Xiaohua Gan, Guangming Ren
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      During aircraft maneuvering flights, engine's rotor-bearing systems are subjected to parametric excitations and additional inertial forces, which may cause severe vibration and abnormal operation. Based on Lagrange's principle combined with finite element modeling, the differential equations of motion for a squeeze film damped rotor-bearing system mounted on an aircraft in maneuvering flight are derived. Using Newmark–Hilber–Hughes–Taylor integration method, dynamic characteristics of the nonlinear rotor system under maneuvering flight are investigated. The factors are considered, involving mass unbalance, oil–film force, gravity, parametric excitations and additional inertial forces, and instantaneous static eccentricity of journal induced by maneuvering loads. The effects of forward velocity, radius of curvature, rotating speed, mass unbalance, oil–film clearance, and elastic support stiffness on transient responses of rotor system are discussed during diving–climbing maneuver.The results indicate that when the aircraft performs a diving–climbing maneuver in the vertical plane, the journal deviates from the center of oil–film outer ring, and the excursion direction of whirl orbit is determined by centrifugal acceleration and additional gyroscopic moment. The journal whirls asynchronously around the instantaneous static eccentricity and its magnitude is related to the maneuvering loads and the supporting stiffness. Increasing forward velocity or decreasing pitching radius, the rotor vibration will enter earlier into or withdraw later from the relatively large eccentricity. Rotating near critical speeds or excessive mass unbalances should be prevented during maneuvering flights. For large maneuver, the oil–film radial clearance needs to be enlarged properly to avoid hard contact between journal and outer ring. In addition, the stiffness of elastic support needs to be appropriately determined for damping performance. Overall, it provides a flexible approach with good expandability to predict dynamic characteristics of on-board squeeze-film damped rotor system during maneuvering flights in the design process.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-29T10:28:24Z
      DOI: 10.1177/0954410020942610
       
  • Study on integrated control for supersonic inlet and turbofan engine model
    • Authors: Haoying Chen, Haibo Zhang, Yao Du, Qiangang Zheng
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Considering the supersonic inlet model with normal shock position feedback, the integrated control method of inlet and turbofan engine is studied. The integrated model includes the supersonic inlet model and the component level model of engine. Combining the relationship between the normal shock position and the total pressure recovery coefficient, the supersonic inlet and engine model is constructed. On the basis of this model, the normal shock position closed-loop control simulation is carried out, which shows that the normal shock position matching point could be stabilized near the optimal value while restraining the inlet stream disturbance. Furthermore, based on the H∞ control algorithm, an inlet and engine integrated control is designed to control the installation thrust and turbine pressure ratio with fuel, nozzle throat area, and normal shock position as control variables. The simulation results show that the response time of the integrated control is faster than the independent control. The integrated control has stronger ability to restrain the atmospheric disturbance, which could ensure the stable and reliable operation of the propulsion system.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-27T12:24:33Z
      DOI: 10.1177/0954410020944068
       
  • A novel de-icing strategy combining electric-heating with plasma synthetic
           jet actuator
    • Authors: TianXiang Gao, Zhen Bing Luo, Yan Zhou, Sheng Ke Yang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Traditional electro-thermal de-icing strategy has the drawback of high energy consumption and its heat knife accounts for a considerable amount of the cost. Compared to electro-thermal de-icing, thermo-mechanical expulsion de-icing system eliminates the heat knife by combining electric-heating with electro-magnetic expulsion de-icing system. The consumption is significantly lower but the system has complex mechanical structures. In this article, a novel de-icing strategy combining electric-heating with plasma synthetic jet actuator is proposed for the first time. Its main idea is to replace heat knife with a simple mechanical device. During the de-icing process, the electric-heating is used to remove the adhesion force, then a single pulse of plasma synthetic jet actuator exerts a rapid force on ice and makes it fracture. Schlieren imaging shows plasma synthetic jet actuator can make free ice columns fracture into pieces and powder. The ice can even be completely penetrated by pressurized air when the discharge energy is relatively large. And compared with the non-deicing process, the intensities of waves and jets are significantly weakened. During the hybrid de-icing process, high-speed photography shows that plasma synthetic jet actuator can divide an ice layer 200 mm in diameter and 10 mm in thickness into multiple blocks completely in tens of milliseconds after electric-heating removes the adhesion force. Besides, energy consumption of plasma synthetic jet actuator in a de-icing cycle accounts for only 0.27% of the whole system. Compared with the free ice columns of the same size, the ice debonded by electric-heating fragmented into smaller blocks after activating plasma synthetic jet actuator. However, heating for too long does not bring more beneficial effects on the fracture of ice in this experiment. At last, a new “micro piston ice-breaker” which is waterproof is proposed to meet the needs of in-flight de-icing.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-24T09:58:40Z
      DOI: 10.1177/0954410020944728
       
  • Influence of axial skewed slots on the rotating instability of a low-speed
           axial compressor
    • Authors: Tao Li, Yadong Wu, Hua Ouyang
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Experimental and numerical analyses were performed on a low-speed axial compressor rotor to investigate the aerodynamic and acoustic effects of axial skewed slots casing treatment on the rotating instability. The experimental results showed that the stall margin could be improved by 8.0% and the frequency broadband hump owing to the rotating instability was suppressed effectively. In the noise spectra, the two dominant broadband humps on both sides of the blade-passing frequency also reduced in amplitude. Full-annulus unsteady computational fluid dynamics simulations were performed near the design condition. Time- and frequency-domain analyses as well as a proper orthogonal decomposition method were applied to obtain the oscillation, frequency, energy and flow characteristics of the rotating instability. Axial skewed slots casing treatment causes a distinct reduction in the amplitude of the pressure fluctuations and frequency spectra with a decrease in the energy of the rotating instability modes. The slots alleviated the tip flow blockage by the periodic injection and removal of the fluid from the passage, which enabled a high tip clearance flow downstream with little impingement on the neighbouring blade tip.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-22T04:13:39Z
      DOI: 10.1177/0954410020944331
       
  • Mechanism of affecting the performance and internal flow field of an axial
           flow subsonic compressor with self-recirculation casing treatment
    • Authors: Hao G Zhang, Fei Y Dong, Wei Wang, Wu L Chu, Song Yan
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This investigation aims to understand the mechanisms of affecting the axial flow compressor performance and internal flow field with the application of self-recirculation casing treatment. Besides, the potentiality of further enhancing the compressor performance and stability by optimizing the geometric structure of self-recirculation casing treatment is discussed in detail. The results show that self-recirculation casing treatment generates about 7.06, 7.89% stall margin improvements in the experiment and full-annulus unsteady calculation, respectively. Moreover, the compressor total pressure and isentropic efficiency are improved among most of operating points, and the experimental and calculated compressor peak efficiencies are increased by 0.7% and 0.6%, respectively. The comparisons between baseline shroud and self-recirculation casing treatment show that the flow conditions of the compressor rotor inlet upstream are improved well with self-recirculation casing treatment, and the degree of the pressure enhancement in the blade top passage for self-recirculation casing treatment is higher than that for baseline. Further, self-recirculation casing treatment can restrain the leading edge-spilled flows made by the blade tip clearance leakage flows and weaken the blade tip passage blockage. Hence, the flow loss near the rotor top passage is reduced after the application of self-recirculation casing treatment. The rotor performance and stability for self-recirculation casing treatment are greater than those for baseline. The flow-field analyses also indicate that the adverse effects caused by the clearance leakage flows of the blades tip rear are greater than those made by the clearance leakage flows of the blades leading edge. When one injecting part of self-recirculation casing treatment is aligned with the inlet of one blade tip passage, the flow-field quality in the passage is not the best among all the passages between two adjacent injecting parts of self-recirculation casing treatment. Further, the flow-field analyses also indicate that the effect of the relative position between the blade and self-recirculation casing treatment on the flows in the self-recirculation casing treatment may be ignored during the optimization of the recirculating loop configuration.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-21T08:48:23Z
      DOI: 10.1177/0954410020942675
       
  • Effect of Gurney flap on the aerodynamic behavior of an airfoil in
           mutational ground effect
    • Authors: Elham Roozitalab, Masoud Kharati-Koopaee
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this research, the effect of Gurney flap on the aerodynamic behavior of an airfoil in mutational ground effect is investigated. To perform this, lift and drag coefficients of NACA 4412 airfoil section in the presence and absence of Gurney flap in mutational ground effect are evaluated and compared. To provide a better illustration of the effect of Gurney flap on the aerodynamic behavior of the airfoil section in the mutational ground effect, results are obtained during the takeoff and landing processes. Validation of the used numerical model is also performed by comparison of the obtained results with those of other works and reasonable agreements were seen. Results show that inclusion of Gurney flap to the airfoil leads to higher variations of lift and drag coefficients during takeoff and landing process. During the takeoff process, the flapped airfoil results in a higher lift decrement and drag increment although an increase in distance of airfoil from the ground or angle of attack causes the lift decrement of the flapped airfoil to get close to that of clean airfoil. It is shown that during takeoff process, downwash is generated around the airfoil as the airfoil leaves the step and as the airfoil gets away from the step, the generated downwash decreases. During the landing process, at each distance of airfoil form the ground and angle of attack, the lift decrement and drag increment of the flapped airfoil are significant compared to that of the clean airfoil. Results exhibit that during landing process, upwash is generated around the airfoil as the airfoil reaches the step and further airfoil move on the step leads to the decrease in the upwash. The decrement in lift coefficient and also the increment in drag coefficient during landing process are more remarkable than those in takeoff process.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-20T10:18:51Z
      DOI: 10.1177/0954410020943754
       
  • Conceptual design and analysis of an affordable truss-braced wing regional
           jet aircraft
    • Authors: Saeed Hosseini, Mohammad Ali Vaziri-Zanjani, Hamid Reza Ovesy
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A regional, turbofan-powered, 72-passenger, transport aircraft with very high aspect ratio truss-braced wings is developed with an affordable methodology from an existing 52 passenger, conventional twin-turboprop aircraft. At first, the ration behind the selection of the truss-braced wing configuration is discussed. Next, the methodologies for the sizing, weight, aerodynamics, performance, and cost analysis are presented and validated against existing regional aircraft. The variant configurations and their design features are then discussed. Finally, sensitivity analysis is carried out to investigate the effects of the wing aspect ratio and engine bypass ratio on the aircraft weight, aerodynamics, and cost. It has been found that the penalties associated with the wing weight will prevent the acceptable realization of the high aspect ratio wing benefits, but when it is combined with the very high bypass ratio engines, a 17% reduction in the mission fuel weight is achieved. In contrast, the cost analysis has revealed that the application of higher aspect ratio wings in the truss-braced wing configuration may increase the development and maintenance costs. Consequently, with aspect ratios higher than 24, eventually, these costs may outperform the associated fuel cost reductions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-06T10:30:47Z
      DOI: 10.1177/0954410020923060
       
  • Extended validation of a ground-based three-axis spacecraft simulator
           model
    • Authors: H Sh Ousaloo, Gh Sharifi, B Akbarinia
      First page: 151
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The ground-based spacecraft dynamics simulator plays an important role in the implementation and validation of attitude control scenarios before a mission. The development of a comprehensive mathematical model of the platform is one of the indispensable and challenging steps during the control design process. A precise mathematical model should include mass properties, disturbances forces, mathematical models of actuators and uncertainties. This paper presents an approach for synthesizing a set of trajectories scenarios to estimate the platform inertia tensor, center of mass and aerodynamic drag coefficients. Reaction wheel drag torque is also estimated for having better performance. In order to verify the estimation techniques, a dynamics model of the satellite simulator using MATLAB software was developed, and the problem reduces to a parameter estimation problem to match the experimental results obtained from the simulator using a classical Lenevnberg-Marquardt optimization method. The process of parameter identification and mathematical model development has implemented on a three-axis spherical satellite simulator using air bearing, and several experiments are performed to validate the results. For validation of the simulator model, the model and experimental results must be carefully matched. The experimental results demonstrate that step-by-step implementation of this scenario leads to a detailed model of the platform which can be employed to design and develop control algorithms.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-14T02:41:28Z
      DOI: 10.1177/0954410020938968
       
  • Multi-scale optimisation of thin-walled structures by considering a
           global/local modelling approach
    • Authors: Michele I Izzi, Marco Montemurro, Anita Catapano, Daniele Fanteria, Jérôme Pailhès
      First page: 171
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this work, a design strategy for optimising thin-walled structures based on a global-local finite element (FE) modelling approach is presented. The preliminary design of thin-walled structures can be stated in the form of a constrained non-linear programming problem (CNLPP) involving requirements of different nature intervening at the different scales of the structure. The proposed multi-scale optimisation (MSO) strategy is characterised by two main features. Firstly, the CNLPP is formulated in the most general sense by including all design variables involved at each pertinent scale of the problem. Secondly, two scales (with the related design requirements) are considered: (a) the structure macroscopic scale, where low-fidelity FE models are used and (b) the structure mesoscopic scale (or component level), where more accurate FE models are involved. In particular, the mechanical responses of the structure are evaluated at both global and local scales, avoiding the use of approximated analytical methods. The MSO is here applied to the least-weight design of an aluminium fuselage barrel of a wide-body aircraft. Fully parametric global and local FE models are interfaced with an in-house metaheuristic algorithm. Refined local FE models are created only for critical regions of the structure, automatically detected during the global analysis, and linked to the global one, thanks to the implementation of a sub-modelling approach. The whole process is completely automated, and once set, it does not need any further user intervention.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-14T02:43:49Z
      DOI: 10.1177/0954410020939338
       
  • Embedded large eddy simulation of transitional flow over NACA0012 aerofoil
    • Authors: Yujing Lin, Jian Wang, Mark Savill
      First page: 189
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An accurate computation of near-field unsteady turbulent flow around aerofoil is of outstanding importance for aerofoil trailing edge noise source prediction, which is a representative of main contributor to airframe noise and fan noise in modern commercial aircraft. In this study, an embedded large eddy simulation (ELES) is fully implemented in a separation-induced transitional flow over NACA0012 aerofoil at a moderate Reynolds number. It aims to evaluate the performance of the ELES method in aerodynamics simulation for wall-bounded aerospace flow in terms of accuracy, computational cost and complexity of implementation. Some good practice is presented including the special treatments at RANS-LES interface to provide more realistic turbulence generation in LES inflow. A comprehensive validation of the ELES results is performed by comparing with the experimental data and the wall-resolved large eddy simulation results. It is concluded that the ELES method could provide sufficient accuracy in the transitional flow simulations around aerofoil. It is proved to be a promising alternative to the pure LES for industrial flow applications involving wall boundary layer due to its significant computational efficiency.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-06T12:16:28Z
      DOI: 10.1177/0954410020939797
       
  • Finite-time velocity-free prescribed performance control for Halo orbit
           autonomous rendezvous
    • Authors: Dandan Zheng, Jianjun Luo, Zeyang Yin, Zhaohui Dang
      First page: 205
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper studies the problem of autonomous rendezvous in the libration point orbit without relative velocity measurement information. The proposed rendezvous algorithm consists of the finite-time-convergent differentiator and the finite-time prescribed performance controller. Wherein, the differentiator is used to compute the unknown relative velocity between the target and the chaser spacecraft. The novel differentiator-based finite-time prescribed performance controller ensures that the rendezvous error converges to an arbitrarily small prescribed region in finite time in spite of the presence of additive bounded disturbances. Furthermore, the prescribed convergence rate can be also achieved simultaneously. The associated stability proof is constructive and accomplished by the development of a Lyapunov function candidate. Numerical simulations on a final rendezvous approach example are provided to demonstrate the effectiveness and robustness of the proposed control algorithm.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-14T02:39:48Z
      DOI: 10.1177/0954410020940892
       
  • Trimming analysis of flexible aircrafts based on computational fluid
           dynamics/computational structural dynamics coupling methodology
    • Authors: Hua Ruhao, Chen Hao, Yuan Xianxu, Tang Zhigong, Bi Lin
      First page: 219
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A numerical methodology based on the coupling of computational fluid dynamics (CFD) and computational structural dynamics is established to obtain the trimming characteristics of flexible aircrafts in this paper. Reynolds-averaged Navier–Stokes equations are solved through CFD technique. Based on the frame of unstructured mesh, techniques of dynamic chimera mesh and morphing mesh are adopted to treat the data transfer between different computational zones and structure deformation caused by aeroelasticity, respectively. When it is applied to a projectile model with large slenderness ratio constructed in this paper, convergence histories of various initial conditions demonstrate the efficiency and robustness of the algorithm. The influence of the structural rigidity and normal loads on the trimming condition of flexible projectiles is investigated, and the locations of the aerodynamic center with various rigidities present the explanation that elastic deformation can move the aerodynamic center forward and weaken the margin of the stability. Furthermore, the trimming condition of flexible projectiles with propulsion is researched, which indicates that thrust misalignment will increase the effect of elastic deformation on the trimming condition, and the stability margin will be further weakened because of thrust misalignment. The conclusion provided in this paper can provide guidance for the structural design, control system design, and stability analysis for modern aircrafts with small stability margin and low rigidity.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-15T08:37:01Z
      DOI: 10.1177/0954410020941943
       
  • Field-of-view-constrained impact angle control guidance with error
           convergence before interception considering speed changes
    • Authors: Jinrae Kim, Namhoon Cho, Youdan Kim
      First page: 238
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An impact angle control guidance law is proposed for stationary target interception considering missile's field-of-view limit and speed changes. The proposed impact angle control guidance law is structured as a biased proportional navigation with a time-varying bias. The proposed guidance law does not involve any switching logic for maintaining lock-on; hence, the guidance command is continuous during the entire engagement. Unlike the most existing studies, the proposed method guarantees that the impact angle error converges to zero before interception without the constant-speed assumption. To realize these desirable properties, the positive invariance of the bounded look angle interval and the change of independent variable are utilized. Numerical simulations are conducted to demonstrate the performance of the proposed guidance law.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-22T10:16:22Z
      DOI: 10.1177/0954410020942010
       
  • Quantized feedback particle filter for unmanned aerial vehicles tracking
           with quantized measurements
    • Authors: Yu Wang, Xiaogang Wang, Naigang Cui
      First page: 257
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Many existing state estimation approaches assume that the measurement noise of sensors is Gaussian. However, in unmanned aerial vehicles tracking applications with distributed passive radar array, the measurements suffer from quantization noise due to limited communication bandwidth. In this paper, a novel state estimation algorithm referred to as the quantized feedback particle filter is proposed to solve unmanned aerial vehicles tracking with quantized measurements, which is an improvement of the feedback particle filter (FPF) for the case of quantization noise. First, a bearing-only quantized measurement model is presented based on the midriser quantizer. The relationship between quantized measurements and original measurements is analyzed. By assuming that the quantization satisfies [math], Sheppard’s correction is used for calculating the variances of the measurement noise. Then, a set of controlled particles is used to approximate the posterior distribution. To cope with the quantization noise of passive radars, a new formula of the gain matrix is derived by modifying the measurement noise covariance. Finally, a typical two-passive radar unmanned aerial vehicles tracking scenario is performed by QFPF and compared with the three other algorithms. Simulation results verify the superiority of the proposed algorithm.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-21T08:48:26Z
      DOI: 10.1177/0954410020942682
       
 
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