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  Subjects -> AERONAUTICS AND SPACE FLIGHT (Total: 120 journals)
Showing 1 - 30 of 30 Journals sorted alphabetically
Acta Astronautica     Hybrid Journal   (Followers: 484)
Advances in Aerospace Engineering     Open Access   (Followers: 66)
Advances in Aerospace Science and Technology     Open Access   (Followers: 7)
Advances in Astronautics Science and Technology     Hybrid Journal  
Advances in Space Research     Full-text available via subscription   (Followers: 454)
Aeronautical Journal, The     Hybrid Journal   (Followers: 10)
Aerospace     Open Access   (Followers: 57)
Aerospace Medicine and Human Performance     Full-text available via subscription   (Followers: 14)
Aerospace Science and Technology     Hybrid Journal   (Followers: 422)
Aerospace Scientific Journal     Open Access   (Followers: 15)
Aerospace Systems     Hybrid Journal   (Followers: 3)
Aerospace technic and technology     Open Access   (Followers: 2)
Aerotecnica Missili & Spazio : Journal of Aerospace Science, Technologies & Systems     Hybrid Journal  
AIAA Journal     Hybrid Journal   (Followers: 1177)
Air Force Magazine     Full-text available via subscription   (Followers: 11)
Air Medical Journal     Hybrid Journal   (Followers: 8)
Annual of Navigation     Open Access   (Followers: 22)
Artificial Satellites     Open Access   (Followers: 23)
ASTRA Proceedings     Open Access   (Followers: 2)
Astrodynamics     Hybrid Journal   (Followers: 1)
Aviation     Open Access   (Followers: 15)
Aviation Advances & Maintenance     Open Access   (Followers: 3)
Aviation in Focus - Journal of Aeronautical Sciences     Open Access   (Followers: 10)
Aviation Psychology and Applied Human Factors     Hybrid Journal   (Followers: 26)
Aviation Week     Full-text available via subscription   (Followers: 438)
Canadian Aeronautics and Space Journal     Full-text available via subscription   (Followers: 33)
CEAS Aeronautical Journal     Hybrid Journal   (Followers: 29)
Chinese Journal of Aeronautics     Open Access   (Followers: 20)
Ciencia y Poder Aéreo     Open Access   (Followers: 2)
Civil Aviation High Technologies     Open Access   (Followers: 5)
Control Systems     Hybrid Journal   (Followers: 317)
Cosmic Research     Hybrid Journal   (Followers: 4)
COSPAR Colloquia Series     Full-text available via subscription   (Followers: 11)
Egyptian Journal of Remote Sensing and Space Science     Open Access   (Followers: 24)
Elsevier Astrodynamics Series     Full-text available via subscription   (Followers: 12)
Fatigue of Aircraft Structures     Open Access   (Followers: 15)
Frontiers in Astronomy and Space Sciences     Open Access   (Followers: 12)
Gyroscopy and Navigation     Hybrid Journal   (Followers: 255)
IEEE Aerospace and Electronic Systems Magazine     Full-text available via subscription   (Followers: 276)
IEEE Journal on Miniaturization for Air and Space Systems     Hybrid Journal   (Followers: 2)
IEEE Transactions on Aerospace and Electronic Systems     Hybrid Journal   (Followers: 383)
IEEE Transactions on Circuits and Systems I: Regular Papers     Hybrid Journal   (Followers: 39)
International Journal of Aeroacoustics     Hybrid Journal   (Followers: 39)
International Journal of Aerodynamics     Hybrid Journal   (Followers: 36)
International Journal of Aeronautical and Space Sciences     Hybrid Journal   (Followers: 2)
International Journal of Aerospace Engineering     Open Access   (Followers: 80)
International Journal of Aerospace Psychology     Hybrid Journal   (Followers: 23)
International Journal of Aerospace Sciences     Open Access   (Followers: 30)
International Journal of Applied Geospatial Research     Hybrid Journal   (Followers: 7)
International Journal of Aviation Management     Hybrid Journal   (Followers: 8)
International Journal of Aviation Technology, Engineering and Management     Full-text available via subscription   (Followers: 7)
International Journal of Aviation, Aeronautics, and Aerospace     Open Access   (Followers: 4)
International Journal of Crashworthiness     Hybrid Journal   (Followers: 12)
International Journal of Micro Air Vehicles     Full-text available via subscription   (Followers: 11)
International Journal of Satellite Communications Policy and Management     Hybrid Journal   (Followers: 13)
International Journal of Space Science and Engineering     Hybrid Journal   (Followers: 11)
International Journal of Space Structures     Full-text available via subscription   (Followers: 17)
International Journal of Space Technology Management and Innovation     Full-text available via subscription   (Followers: 10)
International Journal of Sustainable Aviation     Hybrid Journal   (Followers: 5)
International Journal of Turbo and Jet-Engines     Hybrid Journal   (Followers: 6)
Investigación Pecuaria     Open Access   (Followers: 3)
Journal of Aerodynamics     Open Access   (Followers: 17)
Journal of Aeronautical Materials     Open Access   (Followers: 9)
Journal of Aeronautics & Aerospace Engineering     Open Access   (Followers: 28)
Journal of Aerospace Engineering     Full-text available via subscription   (Followers: 68)
Journal of Aerospace Engineering & Technology     Full-text available via subscription   (Followers: 16)
Journal of Aerospace Information Systems     Hybrid Journal   (Followers: 20)
Journal of Aerospace Information Systems     Hybrid Journal   (Followers: 32)
Journal of Aerospace Technology and Management     Open Access   (Followers: 7)
Journal of Aircraft     Hybrid Journal   (Followers: 337)
Journal of Aircraft and Spacecraft Technology     Open Access   (Followers: 8)
Journal of Airline and Airport Management     Open Access   (Followers: 12)
Journal of Astrobiology & Outreach     Open Access   (Followers: 3)
Journal of Aviation Technology and Engineering     Open Access   (Followers: 11)
Journal of Aviation/Aerospace Education & Research     Open Access   (Followers: 2)
Journal of Engineering and Technological Sciences     Open Access   (Followers: 1)
Journal of Guidance, Control, and Dynamics     Hybrid Journal   (Followers: 204)
Journal of KONBiN     Open Access   (Followers: 3)
Journal of Navigation     Hybrid Journal   (Followers: 279)
Journal of Propulsion and Power     Hybrid Journal   (Followers: 609)
Journal of Space Safety Engineering     Hybrid Journal   (Followers: 7)
Journal of Space Weather and Space Climate     Open Access   (Followers: 27)
Journal of Spacecraft and Rockets     Hybrid Journal   (Followers: 770)
Journal of Spatial Science     Hybrid Journal   (Followers: 3)
Journal of the American Helicopter Society     Full-text available via subscription   (Followers: 7)
Journal of the Astronautical Sciences     Hybrid Journal   (Followers: 8)
Journal of the Australasian Society of Aerospace Medicine     Open Access   (Followers: 1)
Journal of Wind Engineering and Industrial Aerodynamics     Hybrid Journal   (Followers: 16)
Life Sciences in Space Research     Hybrid Journal   (Followers: 3)
MAD - Magazine of Aviation Development     Open Access   (Followers: 2)
Mekanika : Jurnal Teknik Mesin i     Open Access   (Followers: 1)
Microgravity Science and Technology     Hybrid Journal   (Followers: 2)
New Space     Hybrid Journal   (Followers: 6)
Nonlinear Dynamics     Hybrid Journal   (Followers: 19)
npj Microgravity     Open Access   (Followers: 3)
Open Aerospace Engineering Journal     Open Access   (Followers: 1)
Population Space and Place     Hybrid Journal   (Followers: 9)
Problemy Mechatroniki. Uzbrojenie, lotnictwo, inżynieria bezpieczeństwa / Problems of Mechatronics. Armament, Aviation, Safety Engineering     Open Access   (Followers: 3)
Proceedings of the Human Factors and Ergonomics Society Annual Meeting     Hybrid Journal   (Followers: 16)
Proceedings of the Institution of Mechanical Engineers Part G: Journal of Aerospace Engineering     Hybrid Journal   (Followers: 45)
Progress in Aerospace Sciences     Full-text available via subscription   (Followers: 79)
Propulsion and Power Research     Open Access   (Followers: 67)
REACH - Reviews in Human Space Exploration     Full-text available via subscription   (Followers: 5)
Research & Reviews : Journal of Space Science & Technology     Full-text available via subscription   (Followers: 17)
RocketSTEM     Free   (Followers: 6)
Russian Aeronautics (Iz VUZ)     Hybrid Journal   (Followers: 24)
Science and Education : Scientific Publication of BMSTU     Open Access   (Followers: 1)
Space and Polity     Hybrid Journal   (Followers: 4)
Space Policy     Hybrid Journal   (Followers: 30)
Space Research Today     Full-text available via subscription   (Followers: 48)
Space Safety Magazine     Free   (Followers: 51)
Space Science International     Open Access   (Followers: 193)
Space Science Reviews     Hybrid Journal   (Followers: 97)
SpaceNews     Free   (Followers: 824)
Spatial Information Research     Hybrid Journal   (Followers: 1)
Technical Soaring     Full-text available via subscription   (Followers: 1)
Transport and Aerospace Engineering     Open Access   (Followers: 1)
Transportmetrica A : Transport Science     Hybrid Journal   (Followers: 8)
Unmanned Systems     Hybrid Journal   (Followers: 5)
Вісник Національного Авіаційного Університету     Open Access   (Followers: 2)

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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: 45  
 
  Hybrid Journal Hybrid journal (It can contain Open Access articles)
ISSN (Print) 0954-4100 - ISSN (Online) 2041-3025
Published by Sage Publications Homepage  [1099 journals]
  • 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
       
  • Corrigendum to Optimality considerations for propulsive fuselage power
           savings
    • Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.

      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-12-21T11:08:30Z
      DOI: 10.1177/0954410020982417
       
  • 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
       
  • Field-of-view-constrained impact angle control guidance with error
           convergence before interception considering speed changes
    • Authors: Jinrae Kim, Namhoon Cho, Youdan Kim
      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
       
  • 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
       
  • Quantized feedback particle filter for unmanned aerial vehicles tracking
           with quantized measurements
    • Authors: Yu Wang, Xiaogang Wang, Naigang Cui
      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
       
  • 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
       
  • 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
      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
       
  • 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
      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
       
  • Extended validation of a ground-based three-axis spacecraft simulator
           model
    • Authors: H Sh Ousaloo, Gh Sharifi, B Akbarinia
      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
       
  • Finite-time velocity-free prescribed performance control for Halo orbit
           autonomous rendezvous
    • Authors: Dandan Zheng, Jianjun Luo, Zeyang Yin, Zhaohui Dang
      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
       
  • Conception of developing the dynamically similar downscaled medium-range
           passenger airplane model for in-flight testing
    • Authors: Aleksander Olejnik, Stanisław Kachel, Robert Rogólski, Jarosław Milczarczyk
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The article describes the proposition for developing the similar dynamically downscaled model of a medium range passenger aircraft and its application for researches in area of aerodynamics and flight dynamics. Computer simulations of aerodynamic flows are commonly used in advanced aircraft designing. Numerous data on the characteristics of an airplane can be obtained from tunnel tests of geometrically scaled models. To get complete characteristics from both stable and unstable flight conditions, dynamically scaled models are constructed and applied in static or dynamic tests. The dynamically similar scaled model is, in fact, a reduced model of the real airplane which has specific qualities similar to qualities of a real aircraft and these relations are strictly defined with specific similarity numbers (factors). The article presents methodology for determining scale factors in relation to geometric, aerodynamic, and structural properties (mass, stiffness) of the aircraft. The methodology will be presented on the example of Tu-154 aircraft that crashed in April 2010 with the President of the Republic of Poland and many military and government officials on board. Dynamically similar downscaled model was designed in Faculty of Mechatronics and Aerospace of the Military University of Technology (FMA MUT, Warsaw, Poland) and is still being developed in the framework of ongoing research project. This paper presents only general assumptions taken for the process and the current stage of successively developed model.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-10T11:21:23Z
      DOI: 10.1177/0954410020934301
       
  • Embedded large eddy simulation of transitional flow over NACA0012 aerofoil
    • Authors: Yujing Lin, Jian Wang, Mark Savill
      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
       
  • Application of ground vibration testing in small manned or unmanned
           aircraft prototyping
    • Authors: Aleksander Olejnik, Stanisław Kachel, Robert Rogólski, Michał Szcześniak
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The article describes the vibration measurement technology used in experimental investigation of light aircraft and some series of exemplary results obtained during the prototyping process. The aim of investigations presented herein was to determine the resonant frequencies and natural modes of an aircraft or its selected structural components. Ground vibration testing is an essential dynamic structural test necessary to carry out before the aircraft certification. This test should be performed on the aircraft example which is predicted to be tested in flight. The measuring system used for ground vibration testing in the Institute of Aviation Technology of the Military University of Technology consists of a multi-channel LMS SCADAS analyzer, a set of piezoelectric accelerometers, two vibration exciters equipped with impedance heads and a computer with the Test.Lab Software. The aim of the article was to present the methodology of performing ground vibration testing tests. Having applied the equipment to measure an airplane or its airframe component, key vibration characteristics corresponding to specific resonant points can be determined. Not only completed aircraft can be tested but also its isolated fragments (wings, stabilizers, tail units) or just empty airframe. Testing separately supported components allows examining their aeroelastic properties at early stage of prototyping. Ground vibration testing technology applied in various stages of the prototyping process was demonstrated in four peculiar research cases. The testing examples presented herein were the following: the isolated strut-braced wing of a light reconnaissance airplane, the light drone imitating an aerial target for some on-ground anti-aircraft artillery sets, the empty airframe of a very light jet and the miniature UAV. Some exemplary results obtained from testing these objects were presented. At the end, some observations and conclusions were noted in the context of usefulness of conducted researches.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-25T04:25:41Z
      DOI: 10.1177/0954410020934012
       
  • Numerical investigation of mutual interaction between a pusher propeller
           and a fuselage
    • Authors: Marcin Figat, Paulina Piątkowska
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This article presents the numerical analysis of the aerodynamic impact of the fuselage on propeller performance in the pusher configuration and the propeller impact on the flow around the fuselage. The main aim of the presented investigation was to find the magnitude of the interaction between the propeller and the fuselage. This effect was evaluated based on the analysis of the change of the fuselage drag and the propeller thrust according to the change of the propeller's geometry. All obtained results allowed to prepare the methodology for choosing the best propeller geometry for the aircraft in the pusher configuration. During the investigation, the impact of the propeller geometry on the results was analysed. First of all, the change of the blade pitch ratio and the propeller radius was tested. Computation was made for numerous flight conditions and propeller rotation rate. As a main result, the relation between the propeller performance and the fuselage in the pusher configuration was found. Especially, the significant influence on the propeller thrust caused by the fuselage and the influence on the fuselage drag caused by the propeller were observed. Finally, all the obtained outcomes were used to create a knowledge base, which was next used to select the best propeller geometry to satisfy the required power condition for a level flight for the newly designed aircraft PW-Chimera.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-08T11:54:21Z
      DOI: 10.1177/0954410020932796
       
  • Large aircraft reliability study as important aspect of the aircraft
           systems' design changes and improvements
    • Authors: Włodzimierz Balicki, Paweł Głowacki, Leszek Loroch
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Along with the increase in air traffic, the number of reported aviation events also increases. The authors have performed processing of the data included in the European Coordination Centre for Aviation Incident Reporting Systems (ECCAIRS) analyzing large aircraft reliability and safety of their operations considering events according to ICAO aviation occurrence categories. The airframe systems are the biggest contributor to the total number of reported events which occurred on the Polish registered large aircraft in the years 2008–2018. Failures of these systems caused more than 23% of total reported aviation events. Based on the ECCAIRS data, determination of occurrence categories essential for aviation safety was performed with the assessment of their safety risk level. Also, detailed airframe systems' reliability study was carried out in order to assess the real reason of their failures. Airframe systems' faults were assigned to the specific ATA chapters and then to each of their sections. The most frequently occurring defects of each unit of the particular airframe system were identified. The results of this analysis may support the decisions of supervisory authorities in the areas where security threats are most important. They can also help aircraft operators with identification of the airframe units which require special attention, and as support in improvement of safety analysis. Identification of significant parts due to the frequency of malfunctions of the system components can be a support for designers who have the information needed to improve the design of specific airframe system.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-08T11:54:02Z
      DOI: 10.1177/0954410020925626
       
  • Numerical multidisciplinary optimization of aircraft with flight dynamic
           stability constraints
    • Authors: Jacek Mieloszyk
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Classical approach to conceptual and preliminary design in aerospace sciences reaches limits. To go further and achieve better, competitive results’ use of optimization methods becomes mandatory. The trend is clearly visible in professional software for simulations equipped with optimization tools, which was not standard just decade ago. Many examples of multidisciplinary optimization were shown, especially with coupled aerodynamics and structure analyses, but only few of them consider dynamic stability effects in the conceptual and preliminary design. The article shows successful example of multidisciplinary design and optimization of Vertical Takeoff and Landing aircraft, which includes coupling of aerodynamics, mass analyses and innovative approach of constraints from flight dynamic stability. Presented optimization framework revealed potential for more efficient conceptual and preliminary design with the result of known dynamic stability characteristics of the aircraft. Obtaining the dynamic stability characteristics is not common on the early stages of the design yet crucial for the aircraft flight performance.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-01T03:56:32Z
      DOI: 10.1177/0954410020926659
       
  • Flat-upper-surface wing for experimental high altitude unmanned aerial
           vehicle
    • Authors: Cezary Galiński, Magdalena Gronowska, Wieńczysław Stalewski, Konrad Gumowski
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Technology of photovoltaic cells and lithium batteries is developing rapidly. As a result, more and more attempts are made to build solar high altitude long endurance airplanes. Unfortunately, data on altitude impact on photovoltaic cells and batteries performance are not easily available. Moreover, acquisition cost of cells is still high. As a result, high altitude long endurance airplanes design is expensive and risky. Therefore, a tool for inexpensive testing of cells is needed. A small and very light unmanned aerial vehicle can be used for this purpose. It could fly as high as the envisaged high altitude long endurance airplane with a small number of cells and batteries, providing valuable information on them. The weight of such an experimental unmanned aerial vehicle could be minimized because long endurance would not be required, so heavy load of lithium batteries could be minimized, reducing also weights of other components. Wings of this unmanned aerial vehicle should enable installation of various types of photovoltaic cells including rigid ones. Therefore, it would be advantageous to apply an airfoil with a flat-upper-surface as large as possible. Unfortunately, flat-upper-surface airfoils are not popular in airfoils catalogs. Therefore, an attempt was undertaken to design an airfoil with 75% flat upper surface. The research focused on maximization of the lift-to-drag ratio and power factor assuming low Reynolds numbers conditions since it was designed for a small unmanned aerial vehicle for photovoltaic cells testing. This paper contains description of design methodology, design assumptions, and the obtained results. Moreover, the authors describe the experiment undertaken to verify the design. The wind tunnel and a semi-span model used for this experiment are presented together with the obtained results. The model has a similar structure to the envisaged structure of unmanned aerial vehicle, so flexibility of the wing is taken into account.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-26T12:26:24Z
      DOI: 10.1177/0954410020925642
       
  • Optimization of the three-segment aerofoil arrangement based on wind
           tunnel tests and numerical analysis
    • Authors: Wojciech Grendysa, Bartosz Olszański
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper presents the optimization of multi-element aerofoil for the LAR-3 Puffin -- STOL light transport aircraft concept proposal. Based on the geometry and aerodynamic characteristics of the well-known and proven in flight three-segment NACA 63A416 aerofoil, the authors explore the possibility of enhancing its high-lift performance by the movement of slot and flap position in extended (deployed) aerodynamic configuration. In order to determine the optimum positions of aerofoil segments (elements), a multi-step optimization approach was developed. It combines computational fluid dynamics simulations that were used for design space screening and preliminary optimization together with low-turbulence wind tunnel tests which yielded certain results. To decrease the numerical cost of the computer simulation campaign, Design of experiment methods (optimal space-filling design among others) were employed instead of exhausting full factorial (parametric) design. Response surface models of major aerodynamic coefficients (lift, drag, pitching moment) at predicted maximum lift coefficient (CL max) point allowed to narrow down search space and identify several candidates for optimal configuration to be checked experimentally. Wind tunnel tests campaign confirmed the major trends observed in computational fluid dynamics derived response surface contour plots. For the optimum aerodynamic configuration, chosen experimental CL max is over 3.9, which is a 10% increase over the baseline (initial slat and flap positions) case. In parallel, the maximum lift-to-drag ratio gain at that point was almost 19%. The research outlined in this paper was conducted on behalf of the aircraft production company and its results will be applied in a newly designed transport aircraft.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-20T01:23:46Z
      DOI: 10.1177/0954410020926028
       
  • 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
       
  • Influence of tubercles’ spanwise distribution on swept wings for
           unmanned aerial vehicles
    • Authors: Charalampos Papadopoulos, Vasilis Katsiadramis, Kyros Yakinthos
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      In this work, a 3D numerical study on the influence of the spanwise distribution of tubercles on a unmanned aerial vehicle wing is presented. The idea of using tubercles in aeronautics comes from the humpback whale (Megaptera novaeangliae) which has a characteristic flipper, with a spanwise scalloped leading edge, creating an almost sinusoidal shape, consisting of bumps called tubercles. The whale uses this layout in order to achieve high underwater maneuverability. Early experimental research showed a great potential in enhancing the 3D aerodynamic characteristics of a wing. Most of the existing experimental results concern infinite wings (2D) models and are accompanied with substantial loss in lift and increase in drag in pre-stall region. On the other hand, 3D finite models have displayed a better overall aerodynamic performance (increased lift and moment, but also, decreased drag). At a range of Reynolds number between 500,000 and 1,000,000 (based on the mean chord of the flipper), tubercles act as virtual fences, introducing a pair of counter rotating vortices that delays the stall of the flipper, a phenomenon that the whales exploit to perform sharp turns and catch their prey. The aforementioned Reynolds number range is the same as the operational Reynolds number for typical unmanned aerial vehicles. To assess the influence of the tubercles installation on UAV wings, a full 3D computational study is carried-out with the use of CFD tools which at a first phase are validated and calibrated with the available literature experimental data. Then, computations are performed for different spanwise tubercles distributions. The results show that there is a noticeable potential on controlling the flow on the wings of a UAV operating in a Reynolds number range between 500,000 and 1,000,000 (based on UAV’s wing mean chord), which can lead to an aerodynamic performance and efficiency increase.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-04-17T06:53:18Z
      DOI: 10.1177/0954410020919583
       
  • Optimality considerations for propulsive fuselage power savings
    • Authors: Arne Seitz, Anaïs Luisa Habermann, Martijn van Sluis
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The paper discusses optimality constellations for the design of boundary layer ingesting propulsive fuselage concept aircraft under special consideration of different fuselage fan power train options. Therefore, a rigorous methodical approach for the evaluation of the power saving potentials of propulsive fuselage concept aircraft configurations is provided. Analytical formulation for the power-saving coefficient metric is introduced, and, the classic Breguet–Coffin range equation is extended for the analytical assessment of boundary layer ingesting aircraft fuel burn. The analytical formulation is applied to the identification of optimum propulsive fuselage concept power savings together with computational fluid dynamics numerical results of refined and optimised 2D aero-shapings of the bare propulsive fuselage concept configuration, i.e. fuselage body including the aft–fuselage boundary layer ingesting propulsive device, obtained during the European Union-funded DisPURSAL and CENTRELINE projects. A common heuristic for the boundary layer ingesting efficiency factor is derived from the best aero-shaping cases of both projects. Based thereon, propulsive fuselage concept aircraft design optimality is parametrically analysed against variations in fuselage fan power train efficiency, systems weight impact and fuselage-to-overall aircraft drag ratio in cruise. Optimum power split ratios between the fuselage fan and the underwing main fans are identified. The paper introduces and discusses all assumptions necessary in order to apply the presented evaluation approach. This includes an in-depth explanation of the adopted system efficiency definitions and drag/thrust bookkeeping standards.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-04-08T11:51:49Z
      DOI: 10.1177/0954410020916319
       
  • Lightweight unmanned aerial vehicle for emergency medical service –
           Synthesis of the layout
    • Authors: Tomasz Goetzendorf-Grabowski, Andrzej Tarnowski, Marcin Figat, Jacek Mieloszyk, Bogdan Hernik
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The article presents the innovative unmanned aerial vehicle project for emergency medical services. Designed unmanned aerial vehicle combines vertical takeoff and landing characteristics with fast forward flight capability that are vital to perform such an emergency medical mission. The main purpose of the designed unmanned aerial vehicle is to deliver the necessary medical package to the place where access is difficult, and estimated arrival time of conventional ambulance is too long. The cost of the support of such unmanned aerial vehicle could be significantly lower than in case of medical helicopter, which is not necessary in some cases. Designed unmanned aerial vehicle can also be used for fast delivery of essential medical substances (e.g. blood). The selection of configuration was the first and crucial step of the design. After analysis of many different copter configurations, together with selected crash reports analysis, the coaxial quadcopter configuration crossed with conventional airplane was selected. All power units for VTOL capability are electric, and they are doubled for redundancy purposes, with maximum T/W ratio about 2.0. Such configuration allows to sustain a stable flight (vertical phases) in case of one motor failure. Two versions of the vehicle are designed: fully electric (power units for the forward flight and vertical takeoff and landing are electric) and mixed where forward flight unit is a small piston engine. The final layout was the result of conceptual investigation and preliminary research, MDO and trade-off analysis, where as many aspects as possible were considered. The main problem was to meet the vertical takeoff and landing capabilities, relatively long range and endurance, expected payload (3 kg) and the requirement not to exceed 25 kg of maximum take-off weight. Paper presents the design process from initial requirement to the final configuration accepted to be manufactured.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-03-06T09:49:18Z
      DOI: 10.1177/0954410020910584
       
  • Nozzle dimension design for aircraft engine infrared signature and thrust
           active control using MOEA/D
    • Authors: Yu Zhao, Shijie Zheng
      First page: 2133
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Aircraft infrared signature is one of the most important properties for the military aircraft survivability. In terms of military aircraft, the exhaust system is the most significant infrared radiation source. The exhaust system accounts for more than 90% of the aircraft infrared radiation, and that the exhaust nozzle contributes the most significant infrared radiation of the whole radiation energy provided by the exhaust system from the rear aspect. Low detectionable feature for military aircraft has attracted more importance to promote aircraft survivability via reducing infrared signature. The alteration of nozzle exit area affects an aircraft engine performance; meanwhile, it severely influences the engine infrared signature radiation from the rear side. The present paper is mainly focused on searching an appropriate group of nozzle exit diameter and throat to exit diameter ratio, which can reduce infrared signature radiation while cutting down the loss of thrust. Hence, objectives involve two aspects: one is minimum infrared signature level, and the other is minimum thrust loss. The multi-objective evolutionary algorithm based on decomposition has been employed to solve this bi-objective optimization problem. The optimization results illustrate that dimension selection range and throat to exit diameter ratio exert more important effect on the thrust loss and infrared signature level. Furthermore, the thrust plays significant role for deciding nozzle exit diameter and throat diameter.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-11T11:15:38Z
      DOI: 10.1177/0954410020932805
       
  • Direct impact angle control guidance for passive homing missiles
    • Authors: Dongsoo Cho, Seung-Hwan Kim
      First page: 2139
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      A guidance problem for impact angle control applicable to a missile equipped with a passive seeker is investigated. The proposed guidance law based on the sliding mode control consists of the proportional navigation guidance term with a varying navigation constant and the additional biased term which is a function of the novel sliding surface. The guidance law directly uses the impact angle error but does not require any range information to be estimated by the additional filter loop of the passive seeker. The asymmetric limit of the passive seeker’s field-of-view is also taken into account, which is not violated by introducing a continuous switching function. The finite time convergence properties are analyzed, and nonlinear simulations are performed to investigate the effectiveness of the proposed guidance law.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-11T11:09:38Z
      DOI: 10.1177/0954410020930488
       
  • Validation case for supersonic boundary layer and turbulent aero-optical
           investigation in high-Reynolds-number freestream by WCNS-E-5
    • Authors: Xi-Wan Sun, Wei Liu
      First page: 2153
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Turbulent flat-plate boundary layer flows have been widely employed for numerical validation in the aero-optical field. In present study, the laminar-to-turbulent evolution induced by wavy roughness in high-Reynolds-number supersonic freestream is investigated using a numerical technique based on the fifth-order weighted compact nonlinear scheme (WCNS-E-5). The computational procedure and post-processing method are described in detail, and the acquired instantaneous flow structure and statistical data are compared with other theoretical, experimental, and numerical results to demonstrate the feasibility of predicting turbulence using WCNS-E-5. Further, to reduce the computational resources required to simulate turbulent flow, a velocity correlation function is introduced to decrease the computational domain in the spanwise direction. Additionally, the effects of different grid sizes on the simulation results are examined by reducing the number of cells in the streamwise, wall-normal, and spanwise directions. Finally, the authors conduct a tentative investigation into the aero-optical effects of the laminar-to-turbulent flowfield using a ray-tracing method, considering both the feasibility of aero-optical detection and the effect of grid scale on the time-averaged imaging quality, as well as a deeper probe into the characteristic structures reflected by aero-optical frequency spectrum. The results elucidated that the wall-normal grid number has the strongest influence on the transitional location, and undoubtedly affects wavefront aberrations. However, different gird scales lead to similar aero-optical spectrum, and revealed the Kolmogorov-type turbulence at small-scale regime. As a prelude to further aero-optical simulations of wall-bounded flows, the current study provides some reference for the code validation process and aero-optical interrogation.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-12T11:24:05Z
      DOI: 10.1177/0954410020932015
       
  • Simulation and experimental results of an unmanned towed system
    • Authors: Nghia Huynh, Carlos Montalvo
      First page: 2167
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This report investigates the flight dynamics of a small-scale (2 ft) towed system using a quadcopter and actively controlled payload. A towed system includes a main driver to propel the system forward connected to a payload via a tether. The towed system here is unique, and in that the driver is a scratch built quadcopter while the payload is also a scratch built actively controlled aircraft. The payload is designed to carry a small instrument that must be sufficiently far away from all interferences created by a quadcopter. A fully non-linear full state model is created and utilized to reveal that oscillations in the payload are decreased with the introduction of a PD controller on the payload. An experimental setup is built to validate simulation results. Experiments show that an actively controlled payload can decrease the attitude oscillations of the payload.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-19T10:17:38Z
      DOI: 10.1177/0954410020934059
       
  • Design and experimental validation of a specialized pressure-measuring
           
    • Authors: Xin Xu, Dawei Liu, Keming Cheng, Dehua Chen
      First page: 2186
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The internal drag of the unconventional inner flow channel of blended wing body aircraft must be measured accurately to correct the air intake effect of the blended wing body flow-through model in wind tunnel tests. In this study, the pressure distribution of the inner flow channel under the interaction of internal and external flows was obtained through numerical simulation. A specialized pressure-measuring rake was designed based on the numerical results, and a validation test was conducted in a 2.4 m × 2.4 m transonic wind tunnel. Compared with the flow in traditional inlets/nozzles, the flow in the unconventional inner channel in the current research is asymmetric, the distortion index is higher, and the internal drag is more sensitive to flow changes. The wind tunnel test results have a good correlation with the numerical results, and the repeatability of the test results is satisfactory, indicating that the measurement accuracy and precision of the pressure-measuring rake are acceptable. The design method of the specialized rake is feasible, and it can be used to guide the measurement of complex flow in unconventional inner flow channels of blended wing body aircraft.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-06T05:20:31Z
      DOI: 10.1177/0954410020938971
       
  • Adaptive robust sliding mode simultaneous control of spacecraft attitude
           and micro-vibration based on magnetically suspended control and sensitive
           gyro
    • Authors: Yuan Ren, Xiaocen Chen, Yuanwen Cai, Weijie Wang, Qiang Liu
      First page: 2197
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      This paper puts forward realizing the synchronous control of spacecraft attitude and micro-vibration by taking advantage of magnetic suspension rotor deflection of magnetically suspended control and sensitive gyro. A disturbance-observer is designed to estimate the complex micro vibration which is hard to measure, and an adaptive robust sliding mode controller is designed as an integrated controller and its stability is proven by Lyapunov stability criterion. The band-pass filters are introduced to separate low frequency and high frequency control signals and takes them as input commands of gyro frame and magnetic suspension rotor, respectively. The semi-physical simulation results testify that the proposed method has good vibration suppression effect as well as improves the attitude convergence rate.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-10T11:26:05Z
      DOI: 10.1177/0954410020939790
       
  • Remaining useful life prediction for auxiliary power unit based on
           particle filter
    • Authors: Jiachen Guo, Jing Cai, Heng Jiang, Xin Li
      First page: 2211
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Auxiliary power unit is one of the indispensable systems for civil aviation aircraft but the traditional planned maintenance cannot meet the actual needs of airlines. In this work, the key performance parameters of the auxiliary power unit are selected by using recursive feature elimination method. With the selected parameters, the remaining useful flight cycle of the auxiliary power unit is predicted by applying particle filter techniques. Some improved algorithms such as Gaussian particle filter and auxiliary particle filter are also compared. The experimental results demonstrate that the particle filter-based method has high prediction accuracy and engineering application value.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-07-14T02:36:29Z
      DOI: 10.1177/0954410020940882
       
  • Numerical parametric study on three-dimensional rectangular counter-flow
           thrust vectoring control
    • Authors: Kexin Wu, Guang Zhang, Tae Ho Kim, Heuy Dong Kim
      First page: 2221
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Recently, fluidic thrust vectoring control is popular for micro space launcher propulsion systems due to its several advantages, such as fast dynamic responsiveness, better control effectiveness, and no moving mechanical equipment. Counter-flow thrust vectoring control is an especially effective technique by utilizing less suction flow to control the mainstream deflection flexibly. In the current work, theoretical and numerical analyses are performed together to elaborate on the performance of the three-dimensional rectangular counter-flow thrust vectoring control system. A new propulsion nozzle of Mach 2.5 is devised by method of characteristics. To testify the feasibility and accuracy of the present research methodology, numerical results were validated against experimental data from the open literature. The computational result using the standard k-epsilon turbulence model reveals a good match with experimentally measured static pressure values along the centerline of the upper suction collar. The influence of several key parameters on vectoring performance is investigated herein, including the mainstream temperature, collar radius, horizontal collar length, and gap height. Critical parameters have been quantitatively analyzed, such as static pressure distribution along the centerline of the upper suction collar, pitching angle, suction mass flow ratio, and thrust coefficient. Furthermore, the flow-field features are qualitatively expounded based on the static pressure contour, streamline, iso-turbulent kinetic energy contour, and iso-Mach number contour. Some important conclusions are offered for further studies. The mainstream temperature mainly affects different dynamic characteristics of the mixing shear layer, including the convective Mach number of the shear layer, density ratio, and flow velocity ratio. The collar radius influences the pressure gradient near the suction collar surface significantly. The pitching angle increases rapidly with the increasing collar radius. As the horizontal collar length increases, the systematic deflection angle initially increases then decreases. The highest pitching angle is obtained for L/H = 3.53. With regard to the gap height, a larger gap height can achieve a higher pitching angle.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-22T11:08:48Z
      DOI: 10.1177/0954410020925602
       
  • Integrated celestial navigation for spacecraft using interferometer and
           Earth sensor
    • Authors: Kai Xiong, Chunling Wei
      First page: 2248
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      An integrated celestial navigation scheme for spacecrafts based on an optical interferometer and an ultraviolet Earth sensor is presented in this paper. The optical interferometer is adopted to measure the change in inter-star angles due to stellar aberration, which provides information on the velocity of the spacecraft in the plane perpendicular to the direction of the observed star. In order to enhance the navigation performance, the measurements obtained from the ultraviolet Earth sensor is used to eliminate the unfavorable effect caused by the gravitational deflection of starlight. As the prior knowledge about the optical path delay bias of the optical interferometer may be ambiguous, a Q-learning extended Kalman filter is derived to fuse the two types of measurements, and estimate the kinematic state together with the optical path delay bias. The solution of the autonomous navigation system consists of position, velocity and attitude of the spacecraft. Numerical simulation shows that an evident improvement in navigation accuracy can be achieved by introducing the ultraviolet Earth sensor into the navigation system. In addition, it is shown that the Q-learning extended Kalman filter performs better than the traditional extended Kalman filter.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-26T12:25:02Z
      DOI: 10.1177/0954410020927522
       
  • A novel adaptive second-order nonsingular terminal sliding mode guidance
           law design
    • Authors: Shuai Xu, Min Gao, Dan Fang, Yi Wang, Baochen Li
      First page: 2263
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Aiming at the problem of missile attacking ground target in pitch plane, combined with a composite fast nonsingular terminal sliding mode, a new adaptive finite-time stable guidance law with attack angle constraint is designed based on the second-order sliding mode control. The improved extended state observer is used to estimate the uncertainties and compensate the control quantity, and the dynamic control gains are designed to avoid the problem about “excessive estimation” of the parameter upper limit. According to the Lyapunov stability theory, it is proved that the system states can converge into a small neighborhood near the equilibrium point in a finite time. Monte Carlo simulation is carried out by randomly generating initial conditions, which proves that the guidance law has strong adaptability to different initial conditions and has good guidance precision.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-26T12:23:41Z
      DOI: 10.1177/0954410020926769
       
  • Robust proportional incremental nonlinear dynamic inversion control of a
           flying-wing tailsitter
    • Authors: Yunjie Yang, Xiangyang Wang, Jihong Zhu, Xiaming Yuan, Xiaojun Zhang
      First page: 2274
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Tailsitter unmanned aerial vehicles are utilized extensively nowadays since they merge advantages of both fixed-wing unmanned aerial vehicles and rotary-wing unmanned aerial vehicles. However, their attitude control suffers from unknown nonlinearities and disturbances due to the wide flight envelope. To solve the problems, a robust attitude controller based on a newly designed flying-wing tailsitter is proposed in this paper. By employing the angular acceleration feedback to compensate unmodeled dynamics, the proportional incremental nonlinear dynamic inversion control law is first developed. The proportional incremental nonlinear dynamic inversion strengthens the conventional nominal gain incremental nonlinear dynamic inversion with a proportional term to reflect the change of the angular acceleration more directly. Therefore, the tailsitter has a quicker response and performs better in suppressing model uncertainties and external disturbances. Since the angular acceleration is difficult to measure in practice, an angular acceleration estimation method is then established to provide accurate signals for the proportional incremental nonlinear dynamic inversion. The signals are derived as complementary results of model prediction method and direct differential method. Theoretical analysis and systematic simulations are conducted to corroborate the effectiveness of the developed estimation method as well as the robustness of the proposed controller.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-26T12:30:03Z
      DOI: 10.1177/0954410020926657
       
  • Optimal control of rapid cooperative spacecraft rendezvous with multiple
           specific-direction thrusts
    • Authors: Piaoyi Su, Weiming Feng, Yang Kun, Zhao Junfeng
      First page: 2296
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Focusing on the long-distance rapid cooperative rendezvous of two spacecraft under finite continuous thrust, this paper proposes a practical strategy for space operation using multiple specific-direction thrusts. Based on the orbital dynamic theory and Pontryagin’s maximum principle, the dynamic equations and optimal control equations for radial, circumferential, and normal thrust are determined. The optimization method is a hybrid algorithm. The initial costate variables for the fuel-optimal rapid cooperative rendezvous problem are obtained using quantum particle swarm optimization and subsequently set as the initial values in the sequence quadratic programming to search for the exact convergent solution. The elliptical and near-circular mission orbital rendezvous for spacecraft with multiple specific-direction thrusts are simulated and optimized. Numerical examples verifying the proposed method are provided. The results facilitate easier realization of rapid spacecraft maneuvering under continuous thrust conditions.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-29T12:53:35Z
      DOI: 10.1177/0954410020926768
       
  • The compound bowing design in a highly loaded linear cascade with large
           turning angle
    • Authors: Xingxu Xue, Songtao Wang, Lei Luo, Xun Zhou
      First page: 2323
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      Numerical simulation was carried out to study the influences of blade-bowing designs based on a highly loaded cascade with large turning angle, while the compound bowing design showed much lower endwall loss than the conventional design in this study. Generally, it showed that the increased turning angle would strengthen the adverse pressure gradient on the suction surface, so the side effect of negative blade bowing angle would be enhanced because of the reduced flow filed stability near suction–endwall corner. However, the positive corner bowing angle that applied in the compound bowing design would enhance the flow field stability near the suction–endwall corner by adjusting spanwise pressure gradient and velocity triangle, so the side effect of negative blade bowing angle would be suppressed and lead to weaker secondary flow. In detail, the blade bowing angle (as well as the corner bowing angle in the conventional bowed cascades) was varied from −5° to −30° in this study, while the reductions of the loss coefficient in the compound bowed cascades were about 0.662.16 times higher (the absolute differences were about 0.0067 0.0097) than the corresponding conventional bowed cascades. Moreover, the Reynolds number and Mach number at the outlet plane were kept at 2.4 × 105 and 0.6, respectively, during the bowing design to ensure the comparability.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-06-01T03:54:52Z
      DOI: 10.1177/0954410020926658
       
  • Research on probabilistic risk assessment of aeroengine rotor failure
    • Authors: Yuanyuan Guo, Youchao Sun, Longbiao Li
      First page: 2337
      Abstract: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering, Ahead of Print.
      The failure of aeroengine rotor will cause a great threat to the flight safety. A new risk assessment model to assess the possibility and severity of failure of aeroengine rotor is presented addressing the problems that its failure samples are not sufficient and that early potential failures are not easy to recognize. The key point of this model is to determine the risk mechanism with diverse failures, failure phase, failure process, historical failure time, and check interval. The risk mechanism is to determine the failure is in the state of relevant failure or not. The failure phases are divided from initial operational state to potential failure and eventually developed to functional failure. The failure process is to confirm possible cascading failure of parts at the same level or of systems. Historical failure time is useful to describe the tendency of failure in the future by Weibull distribution, which is very suitable to depict the rule of mechanical parts failure. Considering check intervals makes the model more complete. These factors to be considered in risk modeling are complete. The relationship has been simulated between failure process and check interval of engine rotor. The risk assessment flowchart of engine rotor has been established after determining failure correlation including nonrelevant failure and relevant failure to another part. The primary failure probability has been predicted through Monte Carlo simulation. In the case of aircraft bursting into flames due to fuel tank breakdown resulted from turbine disk debris, the probability and process model of relevant turbine debris failure have been established through the cartridge receiver, airfoil, fuel tank in sequence. The relevant failure risk of engine part has been evaluated to ensure the safe operation of aeroengine. The aeroengine rotor failure risk model will have great significance in eliminating potential failure and reducing sudden failure.
      Citation: Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering
      PubDate: 2020-05-20T01:26:08Z
      DOI: 10.1177/0954410020926662
       
 
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