Journal Cover Aerospace Science and Technology
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   Hybrid Journal Hybrid journal (It can contain Open Access articles)
   ISSN (Print) 1270-9638
   Published by Elsevier Homepage  [3048 journals]
  • Probe motion compound control for autonomous aerial refueling docking
    • Authors: Zikang Su; Honglun Wang
      Pages: 1 - 13
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Zikang Su, Honglun Wang
      This paper proposed a probe motion compound control scheme for the receiver docking control in the autonomous aerial refueling (AAR). The dynamic equation of the probe is modeled based on the receiver's 6 DOF nonlinear model which considers the influence of the multiple flow disturbance, and these dynamic equations are transformed into the affine nonlinear form for convenient control design. The AAR docking flight controller is divided into several cascade subsystems via back-stepping design technique. The terms which are independent of the virtual control variables in each affine nonlinear subsystem are taken as the “lumped disturbance”, and are accurately estimated for the disturbance compensation in the designed flight controller, by a group of extended states observers (ESO). Then a compound control scheme, which is constituted by back-stepping and ESO, is proposed for AAR docking based on these established receiver dynamics affine nonlinear form and the estimated lumped disturbances by ESOs. In the proposed probe compound scheme, the probe motion is controlled not only via the translational motion of the barycenter but also via the rotational motion of the receiver. The stability of the proposed probe compound control closed-loop system is formally proved by using Lyapunov function technique. Extensive simulations are carried out to verify the effectiveness and improvement with the proposed controller.

      PubDate: 2017-11-05T08:34:20Z
      DOI: 10.1016/j.ast.2017.10.033
      Issue No: Vol. 72 (2017)
       
  • Mission optimisation for a conceptual coaxial rotorcraft for taxi
           applications
    • Authors: J. Enconniere; J. Ortiz-Carretero; V. Pachidis
      Pages: 14 - 24
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): J. Enconniere, J. Ortiz-Carretero, V. Pachidis
      This paper presents the development and an application of a multidisciplinary methodology for the preliminary design assessment of compound coaxial rotorcraft with a counter-rotating rotor system and a rear-mounted propeller. A comprehensive optimisation strategy is deployed to evaluate the environmental and operational benefits of the aforementioned rotorcraft architecture. The code is validated against experimental data prior to the application of the methodology to the evaluation of a conceptual vehicle for intercity taxi applications. Response Surface Models (RSMs) are generated to mimic the rotorcraft performance in order to accelerate the optimisation process. The effects of the defined mission input parameters such as cruise speed, altitude, climb rate or mission length are evaluated. Pareto fronts for fuel burn, N O x emissions and mission duration are obtained. The method was applied to a hypothetical scenario of mission length ranging from 50 to 300 km. Best estimate mission scenario are selected from the Pareto fronts, providing on average 23%, 20%, and 13% simultaneous reductions in mission duration, fuel burn, and N O x emissions when compared to a conventional flight procedure. The picked scenarios coincide with the fuel optimised mission scenarios for each mission length, thus the multi-disciplinary environment was not required. Besides, an “improved” mission procedure is outlined, defining the mission characteristics independently of the mission's length. This procedure yields on average 22%, 14%, and 8% reductions in mission duration, fuel burn, and N O x emissions, respectively.

      PubDate: 2017-11-05T08:34:20Z
      DOI: 10.1016/j.ast.2017.10.031
      Issue No: Vol. 72 (2017)
       
  • Objective quantification of perceived differences between measured and
           synthesized aircraft sounds
    • Authors: Abhishek K. Sahai; Mirjam Snellen; Dick G. Simons
      Pages: 25 - 35
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Abhishek K. Sahai, Mirjam Snellen, Dick G. Simons
      This paper presents an approach with which perceived audible differences in aircraft sounds can be quantified and presented in an objective manner. The objective quantification of the subjectively heard audible differences is intended to serve two primary goals. It can firstly enable developers of auralization technology to make the auralized sounds more realistic by identifying in which aspects the synthesized sounds differ from their real-life counterparts and to what extent. The quantification can secondly provide an improved and more detailed means of distinguishing between aircraft sounds in general, beyond the conventional metrics of A-weighted Sound Pressure level (dBA) or Effective Perceived Noise Level (EPNL) used currently to assess aircraft noise. In this study sound quality metrics are used to quantify the differences in aircraft sounds. These metrics are widely used in other industries such as the automotive sector. Audio files of a reference aircraft, made over identical flight paths at a noise monitoring station in the vicinity of Schiphol airport, are compared in terms of both conventional and sound quality metrics for four measured and four auralized audio files. It is observed from the comparison that differences that may appear small in the conventional metrics can be significant in terms of the sound quality metrics. Significant differences in measured and synthesized sounds are observed for the aircraft considered in this study with regards to the tonal content and fluctuations in amplitude that occur over time. The conventional metrics are seen to capture the overall loudness aspect of aircraft sounds, but give no clear information regarding which spectral or temporal characteristics cause the sounds to be perceived as audibly different.

      PubDate: 2017-11-11T08:45:18Z
      DOI: 10.1016/j.ast.2017.10.035
      Issue No: Vol. 72 (2017)
       
  • Attitude controller design for reusable launch vehicles during reentry
           phase via compound adaptive fuzzy H-infinity control
    • Authors: Qi Mao; Liqian Dou; Qun Zong; Zhengtao Ding
      Pages: 36 - 48
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Qi Mao, Liqian Dou, Qun Zong, Zhengtao Ding
      In this paper, the attitude control problem of reusable launch vehicles (RLVs) during reentry phase is investigated by using compound adaptive fuzzy H-infinity control (CAFHC) strategy in the presence of parameter uncertainties and external disturbances. Firstly, the control-oriented attitude model is established by a model transformation based on the six-degree-of-freedom (6-DoF) dynamic model of the RLV. Secondly, a novel attitude control scheme is developed and the control strategy consists of two parts to achieve a stable and accurate attitude tracking during reentry flight process. An attitude tracking controller is designed utilizing adaptive fuzzy H-infinity control approach combined with an identification model to improve the attitude tracking performance in the interior of fuzzy approximation region of attitude angle. Next, an attitude stabilization controller based on boundary adaptive technique is employed to assure the robustness of the closed-loop system in the exterior of fuzzy approximation region of attitude angle. Furthermore, the stability of the closed-loop system is guaranteed within the framework of Lyapunov theory and the attitude tracking error converges to a small neighborhood around origin. Finally, the simulation results are presented to demonstrate that the effectiveness of the proposed control scheme for reentry RLV, and its tracking performance performs better than the other control method.

      PubDate: 2017-11-11T08:45:18Z
      DOI: 10.1016/j.ast.2017.10.012
      Issue No: Vol. 72 (2017)
       
  • Flow physics and chine control of the water spray generated by an aircraft
           rigid tire rolling on contaminated runways
    • Authors: Kaibin Zhao; Peiqing Liu; Qiulin Qu; Pingchang Ma; Tianxiang Hu
      Pages: 49 - 62
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Kaibin Zhao, Peiqing Liu, Qiulin Qu, Pingchang Ma, Tianxiang Hu
      During the take-off and landing of aircrafts from/on water-contaminated runways, the tire-generated water spray can endanger flight safety, such as engine spray ingestion and spray impingement drag. In this paper, the flow physics and chine control of the spray generated by an aircraft rigid tire are studied by the Smoothed Particle Hydrodynamics (SPH) method. The SPH method is validated by a NASA experiment. The forming and developing progresses of the front and side sprays are described in detail. It is found that the parabolic stripe of initial ejected particles at a given time is the boundary between the spray source region and the non-disturbed region in water film, and the stripes at any time are similar. The effects of tire speed and water film depth on spray angles are evaluated. The chines can effectively control the spray angles by reducing the vertical velocity component of the ejected particles and increasing the lateral component. The arc chines are more effective in controlling the spray than the linear chines.

      PubDate: 2017-11-11T08:45:18Z
      DOI: 10.1016/j.ast.2017.10.036
      Issue No: Vol. 72 (2017)
       
  • A new approach of casing treatment design for high speed compressors
           running at partial speeds with low speed large scale test
    • Authors: Xi Nan; Ning Ma; Feng Lin; Takehiro Himeno; Toshinori Watanabe
      Pages: 104 - 113
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Xi Nan, Ning Ma, Feng Lin, Takehiro Himeno, Toshinori Watanabe
      The instability problems tend to be more severe when high speed compressors operate at partial speeds. This paper proposes an economic approach for casing treatment design that suitable to this situation. Aiming at reducing the expensive and time-consuming high-speed casing treatment experiments, the idea of low-speed similitude of high-speed compressors, which was originally practiced in mid-1980 with the purpose of loss reduction, is now extended to simulate the stability enhancement with casing treatment in this paper. The core idea of this approach is to replace a large portion of design processes for the high-speed compressors (the Prototype) with their equivalent large scale model compressors (the Model). Two different transonic rotors with skewed slots and circumferential grooves casing treatments are conducted as examples to demonstrate this approach. Following the selected similarity rules, the Model is firstly acquired by modeling the near stall point of the Prototype. A variety of casing treatments are designed and assessed on the Model. Then a few more promising configurations can thus be selected via low speed experiments. They are believed to have similar tendency on stall margin improvement on the Prototype. Finally, the selected configurations are converted back to the Prototype with based on the rule of similarity and validated by experimental data. In this paper, principles that guarantee the similitude of the flow field at near stall condition and the effectiveness of the casing treatment are discussed.

      PubDate: 2017-11-11T08:45:18Z
      DOI: 10.1016/j.ast.2017.10.032
      Issue No: Vol. 72 (2017)
       
  • A mixed probabilistic–geometric strategy for UAV optimum flight path
           identification based on bit-coded basic manoeuvres
    • Authors: Luciano Blasi; Simeone Barbato; Egidio D'Amato
      Pages: 1 - 11
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Luciano Blasi, Simeone Barbato, Egidio D'Amato
      This paper presents a novel algorithm identifying optimal flight trajectories for Unmanned Aerial Vehicles compliant with environmental constraints. Such constraints are defined in terms of obstacles, fixed way-points and selected destination points. Optimality is evaluated taking the minimum path length as the specific objective function. The proposed path planning strategy is based on an original trajectory modelling coupled with a Particle Swarm optimizer (PSO). Flight paths starting from a specified point and ending at a selected destination point are divided into a finite number of segments made up of circular arcs and straight lines. In the proposed approach such a geometrical sequence is replaced with a finite sequence of binary-coded basic manoeuvres. This novel formulation allows to easily handle the manoeuvres sequence with a fixed number of integer variables taking advantage of PSO capability in handling discrete variables; moreover the use of mixed-type variables provides the optimization procedure a useful flexibility in the “decision making” modelling and operational scenarios definition as well. Specific geometric-based linear obstacle avoidance models have been implemented in addition to suitable penalty functions. The use of these models forces each path to be consistent with the environmental constraints favouring the identification of feasible trajectories with a reduced number of iterations and particles. The path planning model has been developed with particular care devoted to reduce computational effort as well as to improve algorithm capability in handling general-shaped obstacles both in 2-D and 3-D environments. Various applications have been performed in order to test the effectiveness of the proposed flight path generator. Applicability of the proposed optimization model also to vehicles with VTOL and hovering capabilities has been preliminarily assessed.

      PubDate: 2017-09-23T08:17:03Z
      DOI: 10.1016/j.ast.2017.09.007
      Issue No: Vol. 71 (2017)
       
  • Adaptive unscented Kalman filter based on maximum posterior and random
           weighting
    • Authors: Zhaohui Gao; Dejun Mu; Shesheng Gao; Yongmin Zhong; Chengfan Gu
      Pages: 12 - 24
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Zhaohui Gao, Dejun Mu, Shesheng Gao, Yongmin Zhong, Chengfan Gu
      The unscented Kalman filter (UKF) is an effective technique of state estimation for nonlinear dynamic systems. However, its performance depends on prior knowledge on system noise. If the characteristics of system noise are unknown or inaccurate, the filtering solution may be biased or even divergent. This paper presents a new maximum posterior and random weighting based adaptive UKF (MRAUKF) by combining the concepts of maximum posterior and random weighting to overcome this limitation. The proposed MRAUKF computes noise statistics based on the maximum posterior principle, and subsequently adopts the random weighting concept to optimize the obtained maximum posterior estimations by online adjusting the weights on residuals. The maximum posterior and random weighting estimations of noise statistics are established to online estimate and adjust system noise statistics, leading to the improved filtering robustness. Simulation and experimental results demonstrate that the proposed MRAUKF outperforms the classical UKF and adaptive robust UKF in the presence of uncertain system noise statistics.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.08.020
      Issue No: Vol. 71 (2017)
       
  • On the utilisation of nonlinear plasticity models in military aircraft
           fatigue estimation: A preliminary comparison
    • Authors: Dylan Agius; Chris Wallbrink; Weiping Hu; Mladenko Kajtaz; Chun H. Wang; Kyriakos I. Kourousis
      Pages: 25 - 29
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Dylan Agius, Chris Wallbrink, Weiping Hu, Mladenko Kajtaz, Chun H. Wang, Kyriakos I. Kourousis
      Strain-life methodologies are commonly employed for fatigue estimation in military aircraft structures. These methodologies rely on models describing the elastoplastic response of the material under cycling. Despite the numerous advanced plasticity models proposed and utilised in various engineering problems over the past decades, the Masing model remains a popular choice in fatigue analysis software, mainly due to its simplicity. However, in the case of military aircraft load spectra including scattered overloads the Masing model fails to represent adequately transient cyclic phenomena, such as mean stress relaxation and strain ratcheting. In this study, four well-known constitutive plasticity models have been selected as potential substitutes for the Masing model within a defence organisation in-house developed fatigue analysis software. These models assessed were the well-known Multicomponent Armstrong–Frederick Model (MAF) and three of its derivatives: MAF with threshold (MAFT), Ohno–Wang (OW) and MAF with Multiplier (MAFM). The models were calibrated with the use of existing experimental data, obtained from aircraft aluminium alloy tests. Optimisation of the parameters was performed through a genetic algorithm-based commercial software. The models were incorporated in the fatigue analysis software and their performance was evaluated statistically and compared against each other and with the Masing model for a series of different flight load spectra for a military aircraft. The results show that all four models have achieved a drastic improvement in fatigue analysis, with the MAFT model giving a slightly better performance.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.004
      Issue No: Vol. 71 (2017)
       
  • Comprehensive preliminary sizing/resizing method for a fixed wing –
           VTOL electric UAV
    • Authors: Maxim Tyan; Nhu Van Nguyen; Sangho Kim; Jae-Woo Lee
      Pages: 30 - 41
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Maxim Tyan, Nhu Van Nguyen, Sangho Kim, Jae-Woo Lee
      A Fixed Wing (FW) aircraft with Vertical Takeoff and Landing (VTOL) is a new type of aircraft that inherits the hovering, VTOL, and maneuvering properties of multicopters and the power-efficient cruising of an FW aircraft. This paper presents a comprehensive method for FW-VTOL electric UAV sizing and resizing. The method uses newly developed integrated analysis that combines the VTOL propulsion sizing method with modified FW aircraft sizing theories. Performance requirements are specified as a set of functional relations. Several new empirical equations are derived using available data. The required battery capacity and total mass are determined from mission analysis that includes both VTOL and FW mission segments. The design is iteratively resized when the actual components of the propulsion system are selected. A case study of a 3.5-kg FW-VTOL electric UAV is presented in this research. The results of sizing and resizing are compared to parameters of the actual aircraft manufactured. Prediction of most parameters stays within a 10% error threshold.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.008
      Issue No: Vol. 71 (2017)
       
  • A design approach of wide-speed-range vehicles based on the cone-derived
           theory
    • Authors: Tian-tian Zhang; Zhen-guo Wang; Wei Huang; Shi-bin Li
      Pages: 42 - 51
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Tian-tian Zhang, Zhen-guo Wang, Wei Huang, Shi-bin Li
      The hypersonic gliding vehicle is attracting an increasing attention because of its high lift-to-drag ratio and cruising velocity. This kind of vehicle may experience different airspaces as well as different speed environments. Based on the design theory of the cone-derived waverider, a novel design approach of the hypersonic gliding vehicle was proposed in this article, which is accommodated in a wide speed range. The parametric method employed in the ascender line design makes it possible to control the overall configuration of the vehicle, and there are four parameters chosen to describe the ascender line. The numerical approach has been employed to validate the property of this kind of HGV. By analyzing the pressure contour of the vehicles with different Mach numbers, we conclude that this kind of aircraft own good wave-ride properties in the designed speed range. Their aerodynamic performance makes a balance by comparing with the waveriders designed with the two ultimate speeds. Different design Mach number arrangements lead to different aerodynamic properties, and all of them seem to be suitable in the designed speed range. Therefore, this kind of vehicle is worth referring in the aircraft design. The simplified trajectory performance analysis is employed, and the vehicle is assumed to reentry with a given initial condition. With the angle of attack as well as the angle of slide is set to be zero. The CFD method obtained the aerodynamic coefficients of the vehicle at different speeds. The result shows a wavy trajectory with long range, which means that this kind of vehicle is suitable for long-distance transportation.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.010
      Issue No: Vol. 71 (2017)
       
  • Structural reliability sensitivity analysis based on classification of
           model output
    • Authors: Sinan Xiao; Zhenzhou Lu
      Pages: 52 - 61
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Sinan Xiao, Zhenzhou Lu
      In structural reliability analysis, sensitivity analysis can be used to measure how the input variable influences the failure of structure. In this work, a new reliability sensitivity analysis method is proposed. In the proposed method, the model output is separated into two classes (failure domain and safe domain). The basic idea is that if the failure-conditional probability density function of input variable is significantly different from its unconditional probability density function, then the input variable is sensitive to the failure of structure. The proposed reliability sensitivity indices contain both individual sensitivity index and interaction sensitivity index. The individual sensitivity index can measure the individual effect of input variable on the failure of structure. The asymmetrical interaction sensitivity index can measure how one input variable influences the effect of another input variable on the failure of structure. Additionally, the meanings of the proposed reliability sensitivity indices are also interpreted explicitly, and a data-driven estimation method is also proposed to estimate the proposed reliability sensitivity indices. Finally, a numerical example and two engineering examples are presented to illustrate the rationality of the proposed sensitivity indices and the feasibility of the proposed estimation method.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.009
      Issue No: Vol. 71 (2017)
       
  • Finite-time formation control for multiple flight vehicles with accurate
           linearization model
    • Authors: Enjiao Zhao; Tao Chao; Songyan Wang; Ming Yang
      Pages: 90 - 98
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Enjiao Zhao, Tao Chao, Songyan Wang, Ming Yang
      The finite-time leader–follower formation control problem of Multiple Flight Vehicle (MFV) system with accurate linearization model is considered. Herein, there is only one leader, the interaction topology among the followers is undirected, and the followers are reachable from the leader. Precise feedback linearization based on differential geometry theory is used to linearize the nonlinear motion model of the flight vehicle and the system model with follower track errors is formulated. A distributed formation control protocol based on finite-time control theory is proposed. With the designed control law, the MFV systems can achieve the desired formation in finite time, where the formation configurations can be specified in advance according to the task requirements. Meanwhile, the convergence analysis is proved and the protocol performance is discussed. Finally, simulation results further demonstrate the effectiveness of the proposed method.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.08.018
      Issue No: Vol. 71 (2017)
       
  • Study on the effects of ionization seeds on pulse detonation
           characteristics
    • Authors: Ling Lin; Chunsheng Weng; Qingzhang Chen; Hongyu Jiao
      Pages: 128 - 135
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Ling Lin, Chunsheng Weng, Qingzhang Chen, Hongyu Jiao
      Ionization phenomenon is happened on the wave front due to the presence of high temperature and pressure in the detonation process. Plasma produced in the detonation process can be used as magnetohydrodynamic (MHD) generator or flow controlled by the external magnetic field. However, it is necessary to increase the ionization efficiency by adding metal ions with lower ionization potential owing to the limited amount of plasma produced by detonation. In this paper, a model of pulse detonation engine with ionization seeds was established. The Conservation Element and Solution Element (CE/SE) method was deduced to simulate the interaction between plasma and detonation process. The influence of ionization seed contents on the electrical conductivity and detonation characteristic parameters was analyzed, and the MHD control of detonation process was realized by adding the external electromagnetic field device. The results showed that it had a little influence on the detonation process but a great influence on the generation of detonation plasma by the addition of a certain amount of ionized seed. The ion mass fraction and electrical conductivity in the detonation tube were first increased and then decreased with the increase of ionization seed content, which reached the maximum at the ionization seed mass fraction of 0.05. The acceleration and deceleration process could be achieved by the MHD control.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.015
      Issue No: Vol. 71 (2017)
       
  • Dual-filter transfer alignment for airborne distributed POS based on PVAM
    • Authors: Zhaoxing Lu; Jiancheng Fang; Haojie Liu; Xiaolin Gong; Shicheng Wang
      Pages: 136 - 146
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Zhaoxing Lu, Jiancheng Fang, Haojie Liu, Xiaolin Gong, Shicheng Wang
      Distributed array antenna synthetic aperture radar (SAR), flexible baseline interferometric SAR (InSAR) or integrated multi-task imaging sensors are the most attractive development directions of aerial survey and remote sensing system, and urgently demand a distributed Position and Orientation System (DPOS) to accurately measure multi-nodes time-spatial reference information. However, the traditional transfer alignment (TA) method can't meet the requirements of some high precision interferometric imaging task. To solve the problem, a dual-filter TA method based on position-velocity-attitude matching (PVAM) was proposed. Firstly, the TA method based on attitude matching (AM) is conducted in filter-1 to estimate the flexible angles and the derivatives. Then the TA method based on position-velocity matching (PVM) is conducted in filter-2, and the position and velocity measurements are compensated by the flexible angles and the derivatives estimated by filter-1. Finally, the strapdown solutions of slave IMU are corrected by the corrections from both filter-1 and filter-2. A semi-physical simulation based on airborne DPOS flight experiment has been conducted, verifying that the baseline error has been reduced from 0.0240 m to 0.0083 m and the computation time has been decrease by 3.41%.

      PubDate: 2017-09-30T08:28:22Z
      DOI: 10.1016/j.ast.2017.09.016
      Issue No: Vol. 71 (2017)
       
  • Coverage-based cooperative guidance strategy against highly maneuvering
           target
    • Authors: Wenshan Su; Kebo Li; Lei Chen
      Pages: 147 - 155
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Wenshan Su, Kebo Li, Lei Chen
      Considering highly maneuvering target, a novel cooperative guidance strategy, which aims to make multiple missiles' joint reachable sets cover the target evasion region, is proposed. The assumed engagement scenario is that multiple missiles separate with each other at the beginning of homing phase and then cooperatively intercept the highly maneuvering target which cannot be intercepted by single missile due to insufficient maneuverability. Firstly, for biasing the reachable sets of different missiles to cover different subintervals of target maneuvering range as expected, a biased proportional navigation guidance law (BPN) is designed by introducing the virtual aiming point. Secondly, for the missile-team without communication capability, the minimum relative maneuverability superiority and minimum team size for covering the whole target maneuvering range are derived based on the BPN respectively, and one open-loop cooperative guidance strategy is proposed with an acceptable assumption about the handover errors of midcourse guidance. While for the missile-team with communication capability, a novel coordinating variable is designed and the corresponding closed-loop guidance strategy is proposed. Numerical simulations with different target maneuver modes are presented to verify the guidance performance of the proposed cooperative guidance strategy.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.021
      Issue No: Vol. 71 (2017)
       
  • Swing principle in tether-assisted return mission from an elliptical orbit
    • Authors: Vladimir S. Aslanov; Alexander S. Ledkov
      Pages: 156 - 162
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Vladimir S. Aslanov, Alexander S. Ledkov
      The problem of a tether-assisted payload return from an elliptical orbit is considered in this study. In contrast to the existing works devoted to this issue, the article deals with a tether length control that provides a transfer of the payload into a descent trajectory from the tether rotation mode. Application of the swing principle for the tether control is investigated. The simplified mathematical model of the space tethered system is developed. It is shown that the stable limit cycle could exist under the considered control. The approximate analytical solution for this cycle is obtained. The stability of this solution is studied by the Lyapunov's theorems. The optimal control, which provides transfer of the payload into the descent trajectory with minimum perigee radius, is found as a result of the simulation series. It is shown that the tether should occur several turns before the payload separation. For example, in the YES-2 experiment, it is demonstrated that proposed control makes it possible to perform a payload return using a tether of considerably shorter length. The main conclusion of the paper is that the proposed scheme of the payload deorbit is more effective than the classical static or dynamic tether deployment schemes.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.006
      Issue No: Vol. 71 (2017)
       
  • Optimization of slot geometry in shock wave boundary layer interaction
           phenomenon by using CFD–ANN–GA cycle
    • Authors: M. Karbasizadeh; A.R. Babaei; M. Bazazzadeh; M.D. Menshadi
      Pages: 163 - 171
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): M. Karbasizadeh, A.R. Babaei, M. Bazazzadeh, M.D. Menshadi
      Slot is one of the measures to control the Shock Wave–Boundary Layer Interaction (SBLI) used to avoid strong interference of shock waves with the boundary layer in supersonic flows. In this control measure, the Height of Triple Point (HTP) of λ shock significantly increases, compared to the one without controller, and cause a decline in shock power and pressure drops rate. In this paper, the main focus is on optimization of slot geometry as an influential parameter on the structure of the shock and flow characteristics by using Genetic Algorithm (GA). The averaged implicit Navier–Stokes equations and two equation standard k–ω turbulence models for the numerical simulation of the flow field have been used. The optimization problem is formulated in term of one objective function, namely, height of triple point maximization. Artificial neural network with two hidden layers has been used to achieve objective function based on the numerical simulation of the flow field data base. Root Mean Square Error (RMSE) was calculated for comparison and selecting the best algorithm in sequential steps. In order to simulate and compare the results with data obtained from experimental tests, the Cambridge University's wind tunnel tests and geometry have been used as the base design. The study demonstrates that, the HTP for optimized slot geometry is about 7.6 percentages more than base design.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.019
      Issue No: Vol. 71 (2017)
       
  • Developing an optimal layout design of a satellite system by considering
           natural frequency and attitude control constraints
    • Authors: Mahdi Fakoor; Parviz Mohammad Zadeh; Homa Momeni Eskandari
      Pages: 172 - 188
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Mahdi Fakoor, Parviz Mohammad Zadeh, Homa Momeni Eskandari
      In recent years, there has been a growing research interest in layout design optimization of satellite systems. The layout design optimization of a satellite system is a complex process having a large number of design variables and constraints. This paper presents a hybrid optimization algorithm, which globally explores the design search space using Particle Swarm Optimization (PSO) and gradient-based Sequential Quadratic Programming (SQP) to rapidly locate optimum design point. The majority of the previous research works mainly focused on finding reasonable placement of components in satellite layout design, with some specific requirements, which are essential for the satellite stability, control and performance such as attitude control, non-interference and overlap constraints. In this study, additional requirements such as structural stiffness and natural frequency constraints are also considered. The proposed approach is employed on a simplified international global communication satellite. The obtained results indicate that the consideration of natural frequency and attitude control constraints in the configuration layout design of a satellite system can significantly improve the stability and control of the satellite and thus frequency coupling between satellite and launcher can be prevented. In addition, the results indicate that the proposed method provides an effective way of solving layout design optimization problem of satellite systems.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.012
      Issue No: Vol. 71 (2017)
       
  • Design and analysis on three-dimensional scramjet nozzles with shape
           transition
    • Authors: Zheng Lv; Jinglei Xu; Jianwei Mo
      Pages: 189 - 200
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Zheng Lv, Jinglei Xu, Jianwei Mo
      A trimming method to optimize the configuration of a three-dimensional asymmetric nozzle with shape transition, which aims to obtain good aerodynamic performance and to save weight at the same time, is presented in this paper. Then the effects of the entry shape on the performance of the three-dimensional nozzle are investigated. The streamline tracing involved in an axisymmetric flowfield with optimal thrust is employed to obtain the inviscid contour of the three-dimensional nozzle with shape transition, and the reference temperature method is applied to correct the thickness of the boundary layer. The performance of the designed nozzle is obtained by using computational fluid dynamics. The calculated results show that the trimmed nozzle gains increases in the lift and pitching moment by 427.00% and 10.80%, respectively, with only a 0.76% decrease in the axial thrust coefficient, while the weight can be reduced by as much as 37.51%. For the nozzles with elliptical entrances, as the axial ratio ranges from 1.0 to 2.0, the axial thrust coefficient is increased by 5.38%, while the lift is decreased by 67.74%. When considering the nozzles with rectangular entrances, as the aspect ratio ranges 1.0 to 2.0, the axial thrust coefficient is increased by 3.58%, while the lift and pitching moment are decreased by 82.09% and 16.43%, respectively. Most of the axial thrust is produced on the upper-wall and side-walls in all nozzles, and the contribution of the expansion flow along the side-walls on the thrust generation is pronounced in the nozzle with a relatively smaller entry ratio. However, the majority of the lift and pitching moment are generated on the upper-wall and lower-wall. Besides, the viscosity loss and weight can be reduced by applying the elliptical cross-section in the propulsion system.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.025
      Issue No: Vol. 71 (2017)
       
  • An approach to estimate aircraft touchdown attitudes and control inputs
    • Authors: P. Wu; M. Voskuijl; L.L.M. Veldhuis
      Pages: 201 - 213
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): P. Wu, M. Voskuijl, L.L.M. Veldhuis
      More strict aircraft emission criteria are proposed by the European Union and the United States. Moreover, the rapid development of the global aviation transportation market also asks for more fuel-efficient aircraft. An innovative assisted takeoff and landing technology is proposed in the EU FP-7 project GABRIEL with the aim to improve fuel-efficiency of commercial aircraft. The technology is based on the removal of the conventional landing gears and the introduction of a ground based landing system. This result in a significant reduction of the structural weight and thereby aircraft performance and fuel efficiency are improved. Furthermore, both airport noise and congestion can be decreased as the aircraft can reach a higher take-off speed without the need for longer runways. The ground based system has a yaw degree of freedom and therefore, the conventional de-crab maneuver for crosswind landings can be avoided. As a consequence, the possible touchdown attitudes and control inputs will be different from those seen in conventional landing. The flight attitudes and control inputs of aircraft during touchdown can significantly influence the landing impacts. Generally, these parameters are obtained from flight test or empirical data and they are crucial for landing gear design. However, the use of empirical data is not possible for new innovative designs such as the GABRIEL system. This paper proposes a solution to estimate all possible aircraft touchdown attitudes and control inputs based on flight dynamics simulations and Monte Carlo evaluation. Turbulence is accounted for based on the von Karman turbulence model. The GABRIEL system is used as a test case and 100 sets of stochastic turbulence are implemented. The resulting flight attitudes and control inputs are presented for different control strategies and compared to the results of landing simulations with a conventional landing gear.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.023
      Issue No: Vol. 71 (2017)
       
  • Uncertain reduced-order modeling for unsteady aerodynamics with interval
           parameters and its application on robust flutter boundary prediction
    • Authors: Xianjia Chen; Zhiping Qiu; Xiaojun Wang; Yunlong Li; Ruixing Wang
      Pages: 214 - 230
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Xianjia Chen, Zhiping Qiu, Xiaojun Wang, Yunlong Li, Ruixing Wang
      Computational fluid dynamics based unsteady aerodynamic reduced-order models can significantly improve the efficiency of transonic aeroelastic analysis. In this paper, the concept of the conventional model reduction method based on the system identification theory is extended to aerodynamic subsystems with the consideration of computational fluid dynamics-induced interval uncertainties in simulation to get the aerodynamic reduced-order model as uncertain as the original aerodynamic subsystem. The interval estimation of identified coefficients involved in the uncertain reduced-order model is obtained by utilizing the first-order interval perturbation method. The stability problem of the interval aeroelastic state-space model formulated based on the constructed uncertain aerodynamic reduced-order model is equivalently transformed into a standard interval eigenvalue problem associated with a real non-symmetric interval matrix in which the interval bounds of eigenvalues are evaluated by virtue of the first-order interval matrix perturbation algorithm. A new stability criterion for the interval aeroelastic state matrix is defined to predict the robust flutter boundary of the concerned uncertain aeroelastic system. Two numerical examples with respect to the uncertain aerodynamic ROM constructions and robust flutter boundary predictions of the two-dimensional Isogai wing and the three-dimensional AGARD 445.6 wing in transonic regime are implemented to assess the validity and accuracy of the presented approach. The obtained results are also compared with Monte Carlo simulation solutions as well as numerical and experimental results in the literatures indicating that the proposed method can provide a more robust and conservative prediction on the flutter boundary of an aeroelastic system compared with conventional deterministic aeroelastic analysis approaches.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.018
      Issue No: Vol. 71 (2017)
       
  • Analysis of the crash of a transport aircraft and assessment of
           fuzzy-logic stall recovery
    • Authors: Chuan-Tau Lan; Shawn Keshmiri
      Pages: 231 - 244
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Chuan-Tau Lan, Shawn Keshmiri
      A turboprop transport Flight Data Recorder data indicates that attempt to recover from the stalled conditions has failed, even though the pitch angle is continuously pushed nose-down in accordance with the published stall-recovery technique. The present study is to examine the reasons for the angle of attack being increased in the first place and not being reduced after stall, even though the longitudinal control is set to nose-down. Fuzzy Logic models are used to preserve nonlinear and unsteady aerodynamic effects. It is found that the increasing angle of attack is initially caused by the nose-up pitching moment due to inertial coupling. In the subsequent nose-down control attempt for stall recovery, the stall angles of attack are still increasingly higher because the pitch rate relative to the rotating axes stays mostly positive. The present simulation study is based on a new method to be called “Fuzzy-Logic Dynamic Inversion,” where desired dynamics with reduced Euler angles and angular rates are specified; while the required control inputs in elevator, aileron and rudder are determined, all through Fuzzy Logic models. Results in elevator, aileron and rudder controls are illustrated to demonstrate the possibility of stall recovery and accident prevention. It is shown that if at the first sign of stall in icing and crosswind flight conditions, and the landing is aborted, stall recovery is possible by applying proper control inputs simultaneously about three control axes to reduce the moments of inertial coupling and the roll and pitch angles.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.028
      Issue No: Vol. 71 (2017)
       
  • UGV-to-UAV cooperative ranging for robust navigation in GNSS-challenged
           environments
    • Authors: Victor O. Sivaneri; Jason N. Gross
      Pages: 245 - 255
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Victor O. Sivaneri, Jason N. Gross
      This paper considers cooperative navigation between an Unmanned Aerial Vehicle (UAV) operating in a GNSS-challenged environment with an Unmanned Ground Vehicle (UGV), and focuses on the design of the optimal motion of the UGV to best aide the UAV's navigation solution. Our approach reduces the uncertainty of a UAV's navigation solution through the use of peer-to-peer radio ranging from a cooperative UGV, whose location is designed to improve positioning geometry for the UAV. Two novel cooperative strategies and two different estimation strategies for the UGV to assist a UAV are developed and compared. Through the use of a realistic simulation environment, it is shown that employing UGV-to-UAV cooperative navigation can reduce the positioning error of a UAV that is operating in a GNSS-challenged environment, from approximately 1-meter-level to approximately 10-cm-level 3D positioning error.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.024
      Issue No: Vol. 71 (2017)
       
  • Numerical study of inflow equivalence ratio inhomogeneity on oblique
           detonation formation in hydrogen–air mixtures
    • Authors: Yishen Fang; Zongmin Hu; Honghui Teng; Zonglin Jiang; Hoi Dick Ng
      Pages: 256 - 263
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Yishen Fang, Zongmin Hu, Honghui Teng, Zonglin Jiang, Hoi Dick Ng
      In this study, numerical simulations using Euler equations with detailed chemistry are performed to investigate the effect of fuel–air composition inhomogeneity on the oblique detonation wave (ODW) initiation in hydrogen–air mixtures. This study aims for a better understanding of oblique detonation wave engine performance under practical operating conditions, among those is the inhomogeneous mixing of fuel and air giving rise to a variation of the equivalence ratio (ER) in the incoming combustible flow. This work focuses primarily on how a variable equivalence ratio in the inflow mixture affects both the formation and characteristic parameters of the oblique detonation wave. In this regard, the present simulation imposes initially a lateral linear distribution of the mixture equivalence ratio within the initiation region. The variation is either from fuel-lean or fuel-rich to the uniform stoichiometric mixture condition above the oblique shock wave. The obtained numerical results illustrate that the reaction surface is distorted in the cases of low mixture equivalence ratio. The so-called “V-shaped” flame is observed but differed from previous results that it is not coupled with any compression or shock wave. Analyzing the temperature and species density evolution also shows that the fuel-lean and fuel-rich inhomogeneity have different effects on the combustion features in the initiation region behind the oblique shock wave. Two characteristic quantities, namely the initiation length and the ODW surface position, are defined to describe quantitatively the effects of mixture equivalence ratio inhomogeneity. The results show that the initiation length is mainly determined by the mixture equivalence ratio in the initiation region. Additional computations are performed by reversing ER distribution, i.e., with the linear variation above the initiation region of uniform stoichiometric condition and results also demonstrate that the ODW position is effectively determined by the ER variation before the ODW, which has in turn only negligible effect on the initiation length.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.027
      Issue No: Vol. 71 (2017)
       
  • A cancelling method based on Nth-order SSC algorithm for solving active
           cancellation problems
    • Authors: Mingxu Yi; Jun Huang; Zhijun Meng
      Pages: 264 - 271
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Mingxu Yi, Jun Huang, Zhijun Meng
      In this paper, a mathematical model of cancelling signal on the basis of Nth-order SSC (Spectrum Spread and Compression) algorithm is established for the Nonlinear Frequency Modulated (NLFM) signal. The explicit expressions of these cancelling signals are derived. Simulation results show that the cancelling signal could reduce the target gain when the time delay is relatively large. The proposed method is compared with the existing methods in the literature and satisfactory results obtained. Comparison between active cancellation technology and passive stealth technology is implemented in detail. The discussions demonstrate that active cancellation technology is more applicable to solve the low-frequency stealth problem.

      PubDate: 2017-10-08T08:50:02Z
      DOI: 10.1016/j.ast.2017.09.038
      Issue No: Vol. 71 (2017)
       
  • Finite-time tracking control of hypersonic vehicle with input saturation
    • Authors: Jing-Guang Sun; Sheng-Li Xu; Shen-Min Song; Xi-Jun Dong
      Pages: 272 - 284
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Jing-Guang Sun, Sheng-Li Xu, Shen-Min Song, Xi-Jun Dong
      This paper studies the finite-time tracking control problem of hypersonic vehicle in presence of model parameter uncertainties, external disturbance and input saturation. Firstly, to cope with the unknown upper bound disturbance of the system, an adaptive fast terminal sliding mode controller is designed based on a non-homogeneous disturbance observer (NHDO) which can alleviate the chattering and make the proposed controller strongly robust. Meanwhile, a new saturation function is introduced to solve the singular problems of the controller. Secondly, to further solve the problem of input saturation, the hyperbolic tangent function and auxiliary system are introduced to design an anti-saturation fast adaptive terminal sliding mode controller which not only can satisfy the requirements of the actuator's physical limitation, but also guarantees that the sliding mode manifold is finite-time stable. Finally, Lyapunov theory is used to prove the stability of the designed controller strictly, and the numerical simulations of the longitudinal model of the hypersonic vehicle are carried out, which further confirm the robustness and effectiveness of the two designed controllers.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.036
      Issue No: Vol. 71 (2017)
       
  • Monitoring multi-axial vibrations of flexible rockets using
           sensor-instrumented reference strain structures
    • Authors: Natsuki Tsushima; Weihua Su; Hector Gutierrez; Michael G. Wolf; Edwin D. Griffin; Jarrod T. Whittaker; Marie P. Dumoulin
      Pages: 285 - 298
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Natsuki Tsushima, Weihua Su, Hector Gutierrez, Michael G. Wolf, Edwin D. Griffin, Jarrod T. Whittaker, Marie P. Dumoulin
      Strain sensors (e.g., fiber optic strain sensors) can be used to measure the deformation of flexible rockets during launches, in order to monitor and control rocket flight attitude. In this paper, strain sensors are instrumented on multi-axial reference strain structures for a convenient monitor of rocket bending vibrations. Reference strain structures are attached longitudinally along the outer surface of thin-walled flexible rockets. As the medium between the sensors and rocket, the structural design of reference strain structures, as well as the sensor spacing along them, is optimized using an integrated multi-objective optimization approach, which ensures that the reference strain structures will accurately track the deformation of the rocket surface. In addition, kinematic equations are developed to allow for an accurate prediction of the bending deflection of the rocket center axis by using the strain data measured on the rocket surface. Finally, the performance of the optimal reference strain structure is evaluated using different numerical simulations of the flexible rocket.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.026
      Issue No: Vol. 71 (2017)
       
  • Influences of shield ratio on the infrared signature of serpentine nozzle
    • Authors: Wen Cheng; Zhanxue Wang; Li Zhou; Xiaolin Sun; Jingwei Shi
      Pages: 299 - 311
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Wen Cheng, Zhanxue Wang, Li Zhou, Xiaolin Sun, Jingwei Shi
      The criteria for designing the serpentine nozzle on the aspect of infrared suppression is undocumented currently. The aim of this paper is to investigate the influences of the shield ratio on the infrared signature of the serpentine nozzle. Firstly, a validation study was conducted to get the reliable computational method of the infrared signature. Then, the infrared signatures of the serpentine nozzles with variable shield ratios were investigated numerically, and the comparison between the serpentine nozzle and the circular nozzle was also conducted. Results show that 28.9% of the infrared radiation intensity, on average, can be reduced by the single serpentine nozzle, as compare with the circular nozzle. The shield ratio has little effect on the infrared signature of gas outside of the serpentine nozzle exit. And the completely shielded serpentine nozzle cannot bring the greatest benefit to the infrared stealth. The single serpentine nozzle with shield ratio of 0.75 and the double serpentine nozzle with shield ratio of 0.5 have better infrared stealth performance than the others. The visible area ratio can be used as an indicator to evaluate the infrared signature of the serpentine nozzle. In order to provide an effective infrared suppression, the visible area ratio of the single serpentine nozzle and the double serpentine nozzle should be restricted under 0.15 and 0.35, respectively.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.001
      Issue No: Vol. 71 (2017)
       
  • Multi-objective optimization for re-entry spacecraft conceptual design
           using a free-form shape generator
    • Authors: Antonio Viviani; Luigi Iuspa; Andrea Aprovitola
      Pages: 312 - 324
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Antonio Viviani, Luigi Iuspa, Andrea Aprovitola
      In this paper we developed a multi-disciplinary, multi-objective optimization procedure for the shape generation of re-entry spacecrafts performing conventional landing from a Low Earth Orbit return mission. A special free-form parametric model, able to define complex vehicle shapes with no explicit support surfaces, was defined and used for this purpose. Model capabilities have been preliminary validated by emulating the HOPE-X vehicle prototype and computing the aerodynamic coefficients at Mach number 2 and 10. Multi-objective optimization has been performed by considering a multidisciplinary approach comprising aerodynamic analysis, trajectory estimation, and heating analysis starting by fixed waypoints along the descent path. A Pareto front based on mass and cross range objective functions was generated, highlighting the existence of several design scenarios: minimum mass, maximum winglet, maximum cross range. The existing trade-offs between the objective functions were related mainly to bank angle values and vehicle length, featuring main design trends.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.030
      Issue No: Vol. 71 (2017)
       
  • Conceptual design, performance and stability analysis of a 200 passengers
           Blended Wing Body aircraft
    • Authors: Sami Ammar; Clément Legros; Jean-Yves Trépanier
      Pages: 325 - 336
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Sami Ammar, Clément Legros, Jean-Yves Trépanier
      The Blended Wing Body (BWB) is a type of innovative aircraft, based on the flying wing concept. For this aircraft, the literature has reported performance improvements compared to conventional aircraft: economy of fuel, reduction of the weight of the structure, increased payload capacity and less impact on the environment. However, most BWB studies have focused on large aircraft and it is not sure whether the gains are the same for smaller aircraft. The main objective of this study is to perform the conceptual design of a 200 passengers BWB and compare its performance against an equivalent conventional A320 aircraft in terms of payload and range. Moreover, an emphasis will be placed on obtaining a stable aircraft, with the analysis of static and dynamic stability over its flight envelope. This kind of aircraft has a lack of stability due to the absence of vertical tail. Most studies of stability were already realized on reduced size models of BWB, but there is no study on a 200 passengers BWB. The design of the BWB was realized with the platform called Computerized Environment for Aircraft Synthesis and Integrated Optimization Methods (CEASIOM). The airplane, the engines and the control surfaces were obtained in the geometrical module AcBuilder. This design platform, suitable for conventional aircraft design, has been modified and additional tools have been integrated in order to achieve the aerodynamic, performance and stability analysis of the BWB aircraft. The aerodynamic coefficients are calculated from Tornado program. The BWB flight envelope was created based on aeronautical data of A320 aircraft. From this flight envelope, we have got back several thousand possible points of flight. The static and dynamic stability was studied using the longitudinal and lateral matrices of stability and the Flying Qualities Requirements for every flight point.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.037
      Issue No: Vol. 71 (2017)
       
  • Spiral coning manoeuvre for in-orbit low thrust characterisation in
           CubeSats
    • Authors: A. Macario-Rojas; K.L. Smith
      Pages: 337 - 346
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): A. Macario-Rojas, K.L. Smith
      The ability to accurately measure the level of thrust during in-orbit operations is fundamental to the characterisation of emerging propulsion systems for nanosatellites. Many new CubeSat missions use propulsion systems with thrust levels in the order of few micro-Newtons. Whilst laboratory sensing resources are able to resolve such low thrust values, in complementary in-orbit characterisation are limited and in the main not compatible with the standard CubeSat mission. Additionally, typical in-orbit assessment of micro-thrust is generally carried out through body angular speed changes, the effectiveness of which is drastically reduced when external perturbations and sensor noise approach or exceed the thruster action on the CubeSat. This investigation sets out to improve in-orbit micro-thrust characterisation via changes in body angular velocity periodicity due to off-centred thrust action in nearly axisymmetric CubeSats. Unlike traditional methods that rely on determining angular acceleration this method employs a frequency analysis of the transversal component of the angular velocity signal with the aim of reducing measurement error. Numerical simulations support the feasibility and adequacy of the proposed low-thrust gauging method, particularly for weak and noisy sensor signals. The robustness of the method allows for interchangeable analysed signal and enables the use of simple commercial-off-the-shelf rate sensors in fine micro-thrust characterisation.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.035
      Issue No: Vol. 71 (2017)
       
  • Effect of elastic deformation on flight dynamics of projectiles with large
           slenderness ratio
    • Authors: RuHao Hua; ZhengYin Ye; Jie Wu
      Pages: 347 - 359
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): RuHao Hua, ZhengYin Ye, Jie Wu
      The elastic deformation of modern projectiles with large slenderness ratio cannot be ignored with the increasing of flight speed and maneuverability. Unsteady Reynolds-averaged Navier–Stokes (URANS) Equations are solved through CFD technique in this paper. Based on the frame of unstructured mesh, techniques of rigid-motion mesh and inverse-distance-weighted (IDW) morphing mesh are adopted to treat the rigid motion caused by flight dynamics and flexible structure deformation due to aeroelasticity, respectively. Moreover, the six degree of freedom (SDOF) dynamic equations and static aeroelastic equation are solved through the aerodynamic coupling. Numerical results of both free flight case and aeroelastic case calculated by the in-house code agree well with the experimental data, validating the numerical method. A projectile model with X–X configuration is constructed to investigate the effect of elastic deformation on the flight dynamics. Comparison results show that the longitudinal oscillation is more affected by the elastic deformation than the centroid motion, and the oscillation cycle of the orientation angle increases. Furthermore, the trajectories of rigid models with various centroid locations are simulated, illustrating that the elastic deformation could move the aerodynamic center forward and weaken the margin of the static stability margin. In the end, detailed analysis and comparison of the pressure distribution indicates the mechanism by which the elastic deformation leads to the movement of the aerodynamic center and changes the flight dynamic characteristics of the flexible projectile.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.029
      Issue No: Vol. 71 (2017)
       
  • New approach to investigate nonlinear dynamic response and vibration of
           imperfect functionally graded carbon nanotube reinforced composite double
           curved shallow shells subjected to blast load and temperature
    • Authors: Dinh Duc Nguyen; Quoc Quan Tran; Dinh Khoa Nguyen
      Pages: 360 - 372
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Dinh Duc Nguyen, Quoc Quan Tran, Dinh Khoa Nguyen
      This paper presents a new approach – using analytical solution to investigate nonlinear dynamic response and vibration of imperfect functionally graded carbon nanotube reinforced composite (FG-CNTRC) double curved shallow shells. The double curved shallow shells are reinforced by single-walled carbon nanotubes (SWCNTs) which vary according to the linear functions of the shell thickness. The shells are resting on elastic foundations and subjected to blast load and temperature. The shell's effective material properties are assumed to depend on temperature and estimated through the rule of mixture. By applying higher order shear theory, Galerkin method and fourth-order Runge–Kutta method and the Airy stress function, nonlinear dynamic response and natural frequency for thick imperfect FG-CNTRC double curved shallow shells are determined. In numerical results, the influences of geometrical parameters, elastic foundations, initial imperfection, temperature increment and nanotube volume fraction on the nonlinear vibration of the FG-CNTRC double curved shallow shells are investigated. The proposed results are validated by comparing with those of other authors.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.031
      Issue No: Vol. 71 (2017)
       
  • Adaptive optimal gliding guidance independent of QEGC
    • Authors: Jianwen Zhu; Shengxiu Zhang
      Pages: 373 - 381
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Jianwen Zhu, Shengxiu Zhang
      A novel multiple constrained adaptive gliding guidance method which is independent of quasi-equilibrium gliding condition (QEGC) and standard trajectory is proposed in this paper. The gliding guidance task is decomposed into longitudinal and lateral directions. In longitudinal direction, an altitude control model is established independent of QEGC, a hierarchical adaptive guidance strategy is introduced to control the vehicle to achieve equilibrium flight state and to meet the terminal altitude and flight-path angle constraints. In lateral direction, a heading error control model is constructed and the optimal control is employed to eliminate the heading error in real time with minimum energy consumption. In addition, the terminal velocity magnitude is predicted and corrected analytically based on lift–drag ratio, and the coordination strategy between guidance and velocity control is proposed to realize multi-constraint gliding guidance. This algorithm can generate angle-of-attack and bank angle commands which can meet the given terminal constraints with high precision based on the current flight states analytically, and has strong robustness to the initial deviation and environmental deviation.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.033
      Issue No: Vol. 71 (2017)
       
  • An investigation on the effect of pitchwise endwall design in a turbine
           cascade at different incidence angles
    • Authors: K.N. Kiran; S. Anish
      Pages: 382 - 391
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): K.N. Kiran, S. Anish
      This paper describes the effects of non-axisymmetric endwall profiling on the aerodynamic performance of a linear turbine cascade at different incidence angles. The sinusoidal profiling is carried out with constant profile curvature along the mean streamline path. Three different profiles, with varying hump to dip height, are analyzed numerically and the performances are compared with the planar profile. Reynolds Averaged Navier Stokes (RANS) equations are solved in their conservative form using Finite Volume Method with SST turbulence model. The calculated results indicate that the profiled endwall minimizes the lateral movement of weaker boundary layer fluid from the hub-pressure side corner. In comparison with planar case, the flow deviations are largely contained with endwall profiling but closer to the endwall it enhances the overturning and secondary flow kinetic energy. The reduction in loss coefficient is estimated to be 1.3%, 8.7% and 38% for incidence angles of −10°, nominal and +15° respectively. The sinusoidal profiling has brought down the pitch averaged flow deviation and secondary flow kinetic energy at nominal and positive incidence angles but the impact is insignificant at negative incidence. Profiling minimizes the rolling up of the passage vortex and makes the passage vortex to migrate closer to the endwall. This flow modification brings down the losses in the core flow but enhances the losses near the endwall.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.032
      Issue No: Vol. 71 (2017)
       
  • A multi-source information fusion fault diagnosis for aviation hydraulic
           pump based on the new evidence similarity distance
    • Authors: Chuanqi Lu; Shaoping Wang; Xingjian Wang
      Pages: 392 - 401
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Chuanqi Lu, Shaoping Wang, Xingjian Wang
      Aviation hydraulic pump is one of the key components in aircraft hydraulic system, therefore, high-precision fault diagnosis is essential to improve the reliability and performance of hydraulic pump. A novel multi-source information fusion fault diagnosis method is proposed based on the Dempster–Shafer (D–S) evidence theory, which utilizes the three-level signals from pump level, hydraulic power system level and hydraulic actuation system level. The feature vectors of these three levels are extracted as three bodies of evidences (BOEs) and the fuzzy membership function is employed to construct the basic probability assignments (BPAs) of three BOEs. In order to solve the issue of combining the conflicting evidences, the D–S evidence theory based on the new evidence similarity distance is developed to combine the obtained BPAs. Finally, the making-decision rules are given to diagnose the faults. The diagnosis results validate that the proposed method not only can increase significantly the belief level of supporting the diagnosis target, but also has the ability to diagnose fault of pump correctly even if a sensor is faulty.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.040
      Issue No: Vol. 71 (2017)
       
  • On the thermo-structural response of a composite closeout in a
           regeneratively cooled thrust chamber
    • Authors: M. Ferraiuolo; W. Petrillo; A. Riccio
      Pages: 402 - 411
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): M. Ferraiuolo, W. Petrillo, A. Riccio
      Composite materials could be very useful when applied to structural rocket engine components since they can allow significant weight savings thanks to their high specific strength and high specific stiffness. In the present work, a carbon fiber reinforced composite has been adopted to replace the typical heavy metallic closeout structure of a regeneratively cooled thrust chamber of a liquid rocket engine. The composite structure has been considered wrapped over the inner liner of the thrust chamber, made of copper alloys, and provides hoop strength for withstanding the fuel/coolant pressure in the cooling channels. The main aim of the paper is to investigate the influence of the geometry and the thermo-mechanical load on the structural response of the analyzed composite closeout. This study is expected to provide a better understanding of the physical phenomena occurring during the service life of the chamber together with an effective identification of the sizing loads that should be considered in the design phase of the closeout structure.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.041
      Issue No: Vol. 71 (2017)
       
  • Multidisciplinary optimisation of single-stage sounding rockets using
           solid propulsion
    • Authors: Adam Okninski
      Pages: 412 - 419
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Adam Okninski
      Existing sounding rockets are based on earlier proven designs and often utilize surplus military solid rocket motors. Therefore commonly non-optimal, in terms of performance for a given payload, configurations are utilized. This paper presents a methodology for finding close-to-optimal, in terms of launch mass minimization, design configurations for small unguided sounding rockets. A numerical, multidisciplinary approach is used. During the optimization process vehicle sizing and corresponding aerodynamics modelling is done. The implemented flight simulation module is simplified due to unknown, during the conceptual design phase, rocket mass distributions along vehicle major axes. Special attention is given to propulsion system sizing and thrust level selection. This paper presents optimization of sounding rockets with lift capabilities equivalent to sending small payloads above the Von Karman line. The ultimate aim of this paper is to present methods to improve sounding rocket performance at an early stage of design, to enable conducting more efficient microgravity research. Various concepts, such as using different expansion ratio nozzles for different payload envelopes and masses, are discussed. Optimization results for maximizing the apogee of a small sounding rocket are presented. Due to the lack of published corresponding research, guidelines for future sounding rocket developments, based on numerical investigations, are given. The significance of the study is due to the emergence of new sounding rocket designs, without use of surplus motors, and the possibility to improve vehicle efficiency after a few decades of little alteration.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.039
      Issue No: Vol. 71 (2017)
       
  • Preliminary integrated analysis for modeling and optimizing space stations
           at conceptual level
    • Authors: Kaiqiang Wang; Bainan Zhang; Tao Xing
      Pages: 420 - 431
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Kaiqiang Wang, Bainan Zhang, Tao Xing
      Key disciplines at the conceptual design stage for space station are introduced, which are configuration, dynamics and control, and power disciplines. The main variables and parameters in the three disciplines are presented, and the relevant disciplinary analysis models are developed. The integrated analysis framework of the space station is obtained afterward. Then, the multidisciplinary optimization for solar array configuration is taken as an example of the space station optimization based on the integrated analysis model. The optimization problem is modeled with the use of the collaborative optimization (CO). The system-level and three disciplinary optimization models are introduced. In the optimization process, MATLAB is utilized for simulation, and the adaptive genetic algorithm (AGA) is applied as the basic optimization algorithm. It is shown that the optimization problem is effectively solved with the use of the CO and AGA. Moreover, using the integrated analysis framework, the parameters of space station are successfully calculated with high computational efficiency at the conceptual design stage.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.048
      Issue No: Vol. 71 (2017)
       
  • Nussbaum-based fuzzy adaptive nonlinear fault-tolerant control for
           hypersonic vehicles with diverse actuator faults
    • Authors: Chaofang Hu; Xianpeng Zhou; Binghan Sun; Wenjing Liu; Qun Zong
      Pages: 432 - 440
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Chaofang Hu, Xianpeng Zhou, Binghan Sun, Wenjing Liu, Qun Zong
      Stability is a challenging problem for control system of the hypersonic vehicle when partial loss of effectiveness fault and stuck fault happen on elevators and engine. In this paper, a fuzzy adaptive nonlinear fault-tolerant control (FTC) method based on Nussbaum gain technique is proposed for hypersonic vehicles with diverse faults. The cases that one of elevators is lock-in-place and another elevator or the engine is partial loss of effectiveness are addressed. The longitudinal model is transformed into the strict feedback formation. The baseline controllers for altitude and velocity commands tracking are designed using dynamic surface control (DSC) and dynamic inversion. The partial loss of effectiveness faults of elevators and engine are combined in the control gain functions, and Nussbaum approach is introduced to avoid singularity of controllers. The unknown nonlinear functions involving the stuck fault and original uncertain nonlinear items are approximated by fuzzy logic systems. The norm estimation technique is utilized to reduce the number of fuzzy adaptive parameters. This greatly alleviates the calculation burden. It is guaranteed that all signals of the closed-loop FTC system are semi-global uniformly ultimately bounded. Finally, the numerical simulations involving diverse faults demonstrate the effectiveness of the proposed FTC approach.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.10.002
      Issue No: Vol. 71 (2017)
       
  • Two-dimensional heliocentric dynamics approximation of an electric sail
           with fixed attitude
    • Authors: Lorenzo Niccolai; Alessandro A. Quarta; Giovanni Mengali
      Pages: 441 - 446
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Lorenzo Niccolai, Alessandro A. Quarta, Giovanni Mengali
      This work analyzes an approximate solution of the equations of motion for a spacecraft propelled by an Electric Solar Wind Sail with a fixed attitude. The peculiarity of such a propulsion system is that its thrust scales as the inverse heliocentric distance. This represents a substantial difference from a classical solar sail, whose propelling force is known to be proportional to inverse square distance from the Sun. Assuming a heliocentric, two-dimensional mission scenario, the polar form of the spacecraft trajectory equation is obtained for a closed parking orbit of given characteristics by means of an asymptotic expansion procedure. The proposed approach significantly improves the existing results as presented in the literature. A suitable choice of propulsion system parameters and parking orbit characteristics provides interesting similarities with recent solutions obtained for a solar sail-based spacecraft in a heliocentric, two-dimensional, mission scenario.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.045
      Issue No: Vol. 71 (2017)
       
  • Enhancement of quality of modal test results of an unmanned aerial vehicle
           wing by implementing a multi-objective genetic algorithm optimization
    • Authors: Nima Pedramasl; Melin Şahin; Erdem Acar
      Pages: 447 - 463
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Nima Pedramasl, Melin Şahin, Erdem Acar
      Due the fact that aircraft structures work in an environment with lots of dynamic forces, it is of vital importance to perform a dynamic analysis to understand dynamic characteristics of aircraft in that specific environment. These characteristics are usually obtained using numerical methods (finite element analysis) or experimental methods (classical modal analysis). In classical modal analysis, quality of test equipment plays a critical role in final results' accuracy and completeness. There is another important factor which is expertise of a test engineer. Test engineer uses his/her experience to find sufficient/optimum numbers, types and locations of transducers. This process sometimes would be time consuming and exhausting which results in degradation of test results quality. In this paper an algorithm is developed and implemented to find numbers, types and locations of transducers in a modal test which will make results of test more reliable. In this study, an unmanned aerial vehicle used as dummy structure to test functionality of developed algorithm. This algorithm utilized two toolboxes from MATLAB (multi-objective genetic algorithm toolbox and parallel computing toolbox) and MSC© NASTRAN finite element solver. A genetic algorithm based optimization is performed in which MSC© NASTRAN was used to calculate dynamic characteristics of UAV wing. Since this was a time and resource consuming process a parallel computing cluster is also utilized which decreased run times at least fourfold. In algorithm it was tried to find optimum numbers, types and locations of transducers which will result in minimum cost and error in test results. Error was defined as a summation of mode shape observability error, mass loading error and optimum driving point error. At the end of study optimization results are presented and validated by classical modal analysis.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.042
      Issue No: Vol. 71 (2017)
       
  • Inertia parameters identification for cellular space robot through
           interaction
    • Authors: Haitao Chang; Panfeng Huang; Zhenyu Lu; Yizhai Zhang; Zhongjie Meng; Zhengxiong Liu
      Pages: 464 - 474
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Haitao Chang, Panfeng Huang, Zhenyu Lu, Yizhai Zhang, Zhongjie Meng, Zhengxiong Liu
      Most of the technologies are in high-speed evolution nowadays. But the spacecraft, however, is still high-priced and takes years to construct. Besides that, it is hardly to service since the conventional spacecraft are not serviceable designed. Facing those challenges, the concept of cellular space robot is presented in this paper for both spacecraft system construction and on-orbit service. As a typical on-orbit service task, the non-cooperative target takeover control is considered in this paper. Specifically, the inertia parameters identification for takeover control is studied in this paper. Because the cells in the cellular space robot are interconnected and networked, an interactive parameter identification algorithm is presented to solve the parameter identification problem by cells interaction. The algorithm is distributed and both synchronous and asynchronous interactions are supported. The algorithm is validated and analyzed by numerical simulations.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.09.044
      Issue No: Vol. 71 (2017)
       
  • Deployment strategies for planar multi-tethered satellite formation
    • Authors: Guang Zhai; Fei Su; Jingrui Zhang; Bin Liang
      Pages: 475 - 484
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Guang Zhai, Fei Su, Jingrui Zhang, Bin Liang
      This paper presents the deployment strategies for planar multi-tethered satellite formation that spins in orbit plane. By using Lagrange principles, the deployment dynamics, which treat the parent satellite as a finite sized rigid body, are established under gravity gradient perturbation. Comparing with the simplified dynamics that take the parent satellite as mass point, the model in this work enables the investigation on dynamical coupling between parent satellite and tethers. To achieve successful deployment, typical strategies are developed firstly with active gravity gradient compensation, both the tether deployment rate and parent satellite spinning profile are derived under specific motion constraints, after that, the deployment strategy capable of compensating the gravity gradient perturbation is also developed. For the fully deployed system, the minimum spinning rate that ensures the configuration stabilization is mathematically analyzed by employing Jacobi Integrator. Finally, series numerical simulations are performed to validate the proposed deployment strategies.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.10.009
      Issue No: Vol. 71 (2017)
       
  • A prescribed performance control approach guaranteeing small overshoot for
           air-breathing hypersonic vehicles via neural approximation
    • Authors: Xiangwei Bu; Yu Xiao; Ke Wang
      Pages: 485 - 498
      Abstract: Publication date: December 2017
      Source:Aerospace Science and Technology, Volume 71
      Author(s): Xiangwei Bu, Yu Xiao, Ke Wang
      This paper investigates a new prescribed performance control (PPC) methodology for the longitudinal dynamic model of an air-breathing hypersonic vehicle via neural approximation. To release the restriction on traditional PPC that the initial tracking errors have to be known in advance for control design, a novel performance function is exploited. Moreover, the devised controller is capable of guaranteeing prescribed performance on the velocity and altitude tracking errors. Neural networks (NNs) are employed to approximate the unknown vehicle dynamics and a minimal-learning parameter scheme is utilized to update the norm of NN's weight vector. Hence, a low computational burden design is achieved without using back-stepping. The semi-globally uniform boundedness of all the closed-loop signals is insured by Lyapunov synthesis. Finally, simulation results are presented to validate the efficacy of the proposed control approach.

      PubDate: 2017-10-14T08:55:28Z
      DOI: 10.1016/j.ast.2017.10.005
      Issue No: Vol. 71 (2017)
       
  • Database self-expansion based on artificial neural network: An approach in
           aircraft design
    • Authors: Shuyue Wang; Gang Sun Wanchun Chen Yongjian Zhong
      Abstract: Publication date: January 2018
      Source:Aerospace Science and Technology, Volume 72
      Author(s): Shuyue Wang, Gang Sun, Wanchun Chen, Yongjian Zhong
      Aircraft design today requires large amount of CFD calculation. For example when Natural Laminar Flow technique is applied to reduce aircraft skin friction drag by extending laminar length over surface, flowfield calculation related with airfoil laminar transition is computationally intense. Situations like this make iterative trial-and-error approach very inefficient. In order to improve this, this paper aims to exploit airfoil database of geometry and aerodynamic performance (from accumulated experiment and CFD calculation results) based on Artificial Neural Network to develop the approach of database self-expansion. It can generate airfoils with better aerodynamic performance from original database, so that the new airfoils can be applied to improve local aerodynamic performance of aircraft. The motive of the approach is to utilize the resource of accumulated optimization products in order to aid aircraft design. In this paper, we will discuss its application in laminar length extension over the surface of nacelle and wing. Geometry description in preparation of database establishment, configuration of network training, and workflow will be described in the paper.

      PubDate: 2017-11-11T08:45:18Z
       
  • Attitude acquisition from an arbitrary tumbling state using two skewed
           reaction wheels
    • Authors: H.Sh. Ousaloo
      Abstract: Publication date: Available online 3 November 2017
      Source:Aerospace Science and Technology
      Author(s): H.Sh. Ousaloo
      A novel technique has been designed that creates rapid nutation damping and accurate spin rate control for a spacecraft with arbitrary inertia ratio. In this approach the satellite incorporates two symmetrically inclined reaction wheels in a V configuration and stabilization is achieved by simultaneously controlling the angular velocity of the satellite and the wheels. The method furnishes gyroscopic stiffness and steers interchange of momentum between the wheels and the satellite main body. A Monte Carlo type approach is used to verify stability and it is shown that the controller provides automatically logical recovery of the desired spin for any initial state and inertia ratio. Moreover, results of single wheel simulations demonstrate the efficacy of the proposed concept.

      PubDate: 2017-11-05T08:34:20Z
      DOI: 10.1016/j.ast.2017.10.040
       
  • Ground effects on the stability of separated flow around a NACA 4415
           airfoil at low Reynolds numbers
    • Authors: Wei He; Peng Yu; Larry K.B. Li
      Abstract: Publication date: Available online 3 November 2017
      Source:Aerospace Science and Technology
      Author(s): Wei He, Peng Yu, Larry K.B. Li
      We perform a linear BiGlobal modal stability analysis on the separated flow around a NACA 4415 airfoil at low Reynolds numbers ( R e = 300 –1000) and a high angle of attack ( α = 20 ° ), with a focus on the effect of the airfoil's proximity to two different types of ground: a stationary ground and a moving ground. The results show that the most dominant perturbation is a Kelvin–Helmholtz mode, which gives rise to a supercritical Hopf bifurcation to a global mode, leading to large-scale vortex shedding at a periodic limit cycle. As the airfoil approaches the ground, this mode can become more unstable or less unstable, depending on the specific type of ground: introducing a stationary ground to an otherwise groundless system is destabilizing but introducing a moving ground is stabilizing, although both effects weaken with increasing Re. By performing a Floquet analysis, we find that short-wavelength secondary instabilities are damped by a moving ground but are amplified by a stationary ground. By contrast, long-wavelength secondary instabilities are relatively insensitive to ground type. This numerical–theoretical study shows that the ground can have an elaborate influence on the primary and secondary instabilities of the separated flow around an airfoil at low Re. These findings could be useful for the design of micro aerial vehicles and for improving our understanding of natural flyers such as insects and birds.

      PubDate: 2017-11-05T08:34:20Z
      DOI: 10.1016/j.ast.2017.10.039
       
  • Anti-disturbance fault tolerant initial alignment for inertial navigation
           system subjected to multiple disturbances
    • Authors: Songyin Cao; Lei Guo; Wenhua Chen
      Abstract: Publication date: Available online 3 November 2017
      Source:Aerospace Science and Technology
      Author(s): Songyin Cao, Lei Guo, Wenhua Chen
      Modeling error, stochastic error of inertial sensor, measurement noise and environmental disturbance affect the accuracy of an inertial navigation system (INS). In addition, some unpredictable factors, such as system fault, directly affect the reliability of INSs. This paper proposes a new anti-disturbance fault tolerant alignment approach for a class of INSs subjected to multiple disturbances and system faults. Based on modeling and error analysis, stochastic error of inertial sensor, measurement noise, modeling error and environmental disturbance are formulated into different types of disturbances described by a Markov stochastic process, Gaussian noise and a norm-bounded variable, respectively. In order to improve the accuracy and reliability of an INS, an anti-disturbance fault tolerant filter is designed. Then, a mixed dissipative/guarantee cost performance is applied to attenuate the norm-bounded disturbance and to optimize the estimation error. Slack variables and dissipativeness are introduced to reduce the conservatism of the proposed approach. Finally, compared with the unscented Kalman filter (UKF), simulation results for self-alignment of an INS are provided based on experimental data. It can be shown that the proposed method has an enhanced disturbance rejection and attenuation performance with high reliability.

      PubDate: 2017-11-05T08:34:20Z
      DOI: 10.1016/j.ast.2017.10.041
       
 
 
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