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Journal Cover Aerospace Science and Technology
  [SJR: 0.816]   [H-I: 49]   [342 followers]  Follow
    
   Hybrid Journal Hybrid journal (It can contain Open Access articles)
   ISSN (Print) 1270-9638
   Published by Elsevier Homepage  [3120 journals]
  • Inverse dynamics particle swarm optimization applied to constrained
           minimum-time maneuvers using reaction wheels
    • Authors: Dario Spiller; Robert G. Melton; Fabio Curti
      Pages: 1 - 12
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): Dario Spiller, Robert G. Melton, Fabio Curti
      The paper deals with the problem of time-optimal spacecraft reorientation maneuvers by means of reaction wheels, with boundary and path constraints. When searching for solutions to optimal attitude-control problems, spacecraft can be easily modeled as controlled by external torques. However, when using actuators such as reaction wheels, conservation of the total angular momentum must be taken into account and the wheel dynamics must be included. A rest-to-rest slew maneuver is considered where an optical sensor cannot be exposed to sources of bright light such as the Earth, the Sun and the Moon. The motion must be constrained to prevent the sensor axis from entering into established keep-out cones. The minimum-time solution is proposed using the Inverse Dynamics Particle Swarm Optimization technique. The attitude and the kinematics of the satellite evolve, leading to the successive attainment of the wheel control input via fixed-step numerical integration. Numerical results are evaluated over different scenarios. It is established that the computation of minimum time maneuvers with the proposed technique leads to near optimal solutions, which fully satisfy all the boundary and path constraints. The ability to converge in a variety of different scenarios always requiring the same computational effort characterizes the proposed technique as a feasible future on-board path-planner.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.038
      Issue No: Vol. 75 (2018)
       
  • An approach and landing guidance design for reusable launch vehicle based
           on adaptive predictor–corrector technique
    • Authors: Maomao Li; Jun Hu
      Pages: 13 - 23
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): Maomao Li, Jun Hu
      In this study, a novel predictor–corrector guidance law based on the all-coefficient adaptive control theory is proposed for the approach and landing phase of an unpowered reusable launch vehicle (RLV). The flight phase includes two portions: the initial gliding flight phase and the final exponential flare one. The equilibrium glide condition and an exponential function of altitude are used to parameterize the guidance commands of two portions and generate the guidance sequence. Based on the first order characteristic model which has the advantages of less characteristic parameters and easy analysis, the all-coefficient adaptive predictor–corrector guidance method is presented. The guidance law has the ability of generating new trajectories online according to the current states and the final conditions of the landing point. Then, the stability and finite-time convergence of the guidance law are analyzed. Considering the process constraints, the fusion guidance law is obtained. Finally, simulation results demonstrate the effectiveness and robustness of the proposed guidance law, with respect to the large initial states errors and parameter uncertainties. The simulations also show that the system under the proposed guidance law converges to a small neighborhood of zero after limited steps.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.037
      Issue No: Vol. 75 (2018)
       
  • Numerical and experimental investigations of single-element and
           double-element injectors using gaseous oxygen/gaseous methane
    • Authors: Nanjia Yu; Yang Zhang; Feng Li; Jian Dai
      Pages: 24 - 34
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): Nanjia Yu, Yang Zhang, Feng Li, Jian Dai
      The objective of this study is to investigate the effects of fuel/oxidizer (F/O) momentum ratio on the mechanism of mixing and combustion of single-element and double-element shear coaxial injectors numerically and experimentally using gaseous oxygen/gaseous methane (GO2/GCH4) and non-intrusive optical diagnostics technique based on planar laser induced fluorescence (PLIF). Instantaneous OH-PLIF images demonstrate that the flame with higher F/O momentum ratio exhibits appearance characteristics with more obvious wrinkled shear layers, extinction and kernels events. Simulation and time-averaged results show that with the increase of F/O momentum ratio at the condition of fixed total mass flow rate, the initial combustion position becomes closer to the front of the chamber. The large mass flow rate of the double-element combustor is beneficial to the thermal protection of the chamber wall. Based on the high-speed images and velocity distribution obtained from numerical data, it is detrimental for the propagation of the flame front at the injection condition of high F/O momentum ratio. In addition, according to the infrared thermal pictures, the flame outside the chamber becomes broader with the increase of F/O momentum ratio, and the plume of double-element combustor will challenge the thermal protection around the equipment. Moreover, new methods are applied in this paper to extract the flame boundary of the OH-PLIF diagrams and to calculate the number of the pixels of the plume from infrared thermal imager results.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.039
      Issue No: Vol. 75 (2018)
       
  • Low energy actuation technique of bistable composites for aircraft
           morphing
    • Authors: F. Nicassio; G. Scarselli; F. Pinto; F. Ciampa; O. Iervolino; M. Meo
      Pages: 35 - 46
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): F. Nicassio, G. Scarselli, F. Pinto, F. Ciampa, O. Iervolino, M. Meo
      Morphing structures for lightweight and energy-efficient aircraft mobile surfaces have been investigated for several years. This paper presents a novel lightweight, passive and low-energy morphing surface concept based on the “lever effect” of a bistable composite plate that can be integrated in aircraft moving surfaces. By using appropriate boundary conditions, it is demonstrated that the magnitude of the activation force on the bistable composite can be tailored to match the differential pressure on the aircraft's airfoil. As a consequence, the bistable laminate can be used as a passive morphing surface. Both numerical simulations and experimental testing are used to prove this concept on a NACA 2412 airfoil structure. The results show that, by choosing proper configuration of constraints, lay-up and aspect ratio of the bistable composite, it is possible to tailor and activate the snap-through mechanism in a passive manner. The proposed concept would save significant weight when compared to an active morphing concept.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.040
      Issue No: Vol. 75 (2018)
       
  • Solid fuel ramjet fuel optimization for maximum thrust to drag ratio and
           impulse density subject to geometric restraints on missile outer mold line
           
    • Authors: Brian McDonald; Jeremy Rice
      Pages: 47 - 57
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): Brian McDonald, Jeremy Rice
      Tactical solid fuel ramjet propulsion system designs are often restricted by a fixed geometric envelope that the vehicle's outer mold line must be contained within. These constraints fix the combustion chamber diameter and length as well as the maximum intake area profile. With the intake area restrained to some maximum value, the maximum air flowrate into the combustor is likewise limited to a specific value for any given design freestream Mach number and altitude. Optimization of fuels for the maximum thrust to drag ratio or specific impulse must consider the linkage of geometric restraints on the selection of fuel materials. In this paper, equations are developed for thrust to drag, net specific impulse, impulse density, and a newly defined performance parameter, the momentum density. These equations are cast in terms of the ratio of the intake area to the aerodynamic reference area. Thermochemical analyses are conducted for a range of common polymers and solid particulate additives. The optimization equations show that as the intake to reference area ratio decreases, maximum performance shifts from materials having the highest heat of combustion per unit mass to materials that maximize the ratio of heat of combustion to the stoichiometric fuel to air ratio, and to operating air to fuel ratios that approach the upper flammability limits. As the intake area to reference area increases, the optimization tends towards fuel that maximize the heating value alone and to operating air to fuel ratios that approach the lower flammability limits. The conclusion of the analysis is that for solid fuel ramjet designs subject to restrictions on inlet area, maximum performance occurs with high density materials that maximizes heat release and minimizes the stoichiometric air to fuel ratio.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.008
      Issue No: Vol. 75 (2018)
       
  • On novel adaptive saturated deployment control of tethered satellite
           system with guaranteed output tracking prescribed performance
    • Authors: Caisheng Wei; Jianjun Luo; Baichun Gong; Mingming Wang; Jianping Yuan
      Pages: 58 - 73
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): Caisheng Wei, Jianjun Luo, Baichun Gong, Mingming Wang, Jianping Yuan
      In this paper, a novel model-free adaptive saturated control approach for deployment of the tethered satellite system (TSS) is developed based on time- and event-triggered mechanisms with consideration of the tether length constraint, actuator saturation and unknown external disturbance. Firstly, an appropriate output prescribed performance function is selected to quantitatively analyze the transient and steady-state performance of the deployment control system. Then, a strict-feedback system is established via an output transformation based on which two new model-free adaptive saturated control schemes are developed. Compared with the existing works, the primary advantage of the proposed approach is that the output tracking performance of the TSS can be guaranteed a priori with excellent disturbance rejection capability. Meanwhile, it is the first time that the event-triggered prescribed performance control scheme is proposed both in theoretic and application views, which dramatically reduces the frequency of controller updating and makes the corresponding designed schemes more applicable in practice. Finally, two group of illustrative examples are employed to validate the effectiveness of the proposed control approach.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.014
      Issue No: Vol. 75 (2018)
       
  • A fast-prediction surrogate model for large datasets
    • Authors: John T. Hwang; Joaquim R.R.A. Martins
      Pages: 74 - 87
      Abstract: Publication date: April 2018
      Source:Aerospace Science and Technology, Volume 75
      Author(s): John T. Hwang, Joaquim R.R.A. Martins
      Surrogate models approximate a function based on a set of training points and can then predict the function at new points. In engineering, kriging is widely used because it is fast to train and is generally more accurate than other types of surrogate models. However, the prediction time of kriging increases with the size of the dataset, and the training can fail if the dataset is too large or poorly spaced, which limits the accuracy that is attainable. We develop a new surrogate modeling technique—regularized minimal-energy tensor-product splines (RMTS)—that is not susceptible to training failure, and whose prediction time does not increase with the number of training points. The improved scalability with the number of training points is due to the use of tensor-product splines, where energy minimization is used to handle under-constrained problems in which there are more spline coefficients than training points. RMTS scales up to four dimensions with 10–15 spline coefficients per dimension, but scaling beyond that requires coarsening of the spline in some of the dimensions because of the computational cost of the energy minimization step. Benchmarking using a suite of one- to four-dimensional problems shows that while kriging is the most accurate option for a small number of training points, RMTS is the best alternative when a large set of data points is available or a low prediction time is desired. The best-case average root-mean-square error for the 4-D problems is close to 1% for RMTS and just under 10% for kriging.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.030
      Issue No: Vol. 75 (2018)
       
  • A sub-grid scale model with natural near-wall damping
    • Authors: Javad H. Taghinia; Md Mizanur Rahman; Timo Siikonen
      Pages: 1 - 16
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Javad H. Taghinia, Md Mizanur Rahman, Timo Siikonen
      A zero-equation sub-grid scale (SGS) model with a variable eddy-viscosity coefficient C μ is developed for large-eddy simulation. Since C μ is evaluated based on the resolved shear and vorticity parameters accompanied by the hybrid time scale T t (combination of dynamic and Kolmogorov time scales), it is sensitized to non-equilibrium flows, preserving the anisotropic characteristics of turbulence. The current model accounts for the SGS kinetic energy with k s g s = C μ 2 3 ( Δ ¯ S ¯ ) 2 and guarantees the positivity in the energy components. Unlike the original Smagorinsky model, the present SGS model does not need any ad-hoc damping function ( C μ acts as a natural damping function as the wall is approached) or averaging/clipping of the model coefficient for numerical stabilization as required by the dynamic Smagorinsky model (DSM). The model is validated against well-documented flow cases, yielding predictions in good agreement with the direct numerical simulation (DNS) and experimental data. Comparisons indicate some advantages of the new model over the DSM; the current model needs only a single-filtering operation that recovers the numerical stability and computational effort to a greater extent.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.014
      Issue No: Vol. 74 (2018)
       
  • Interval analysis of the wing divergence
    • Authors: Yi Li; Tianhong Wang
      Pages: 17 - 21
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Yi Li, Tianhong Wang
      In this paper the Taylor expansion (TE) is combined with the optimization and anti-optimization problems (OAP) solutions of parameterized interval analysis (PIA) to study the effect of structural uncertainties on the divergence of wing. Through a two-dimensional wing example, the TE + PIA + OAP method developed by this paper is compared with the classic interval analysis (CIA), the results show that the TE + PIA + OAP method can reduce overestimation in the CIA and TE method, and the interval of divergence dynamic pressure predicted by the TE + PIA + OAP method is as narrow as the one from the PIA + OAP methods, but the computation cost of the TE + PIA + OAP method is much lower. In addition an actual engineering example of forward swept wing with uncertain structural parameters is studied by the TE + PIA + OAP method to illustrate the capability to solve the real divergence problem of wing with uncertain structural parameters.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.001
      Issue No: Vol. 74 (2018)
       
  • Global and robust attitude control of a launch vehicle in exoatmospheric
           flight
    • Authors: Fabio Celani
      Pages: 22 - 36
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Fabio Celani
      The goal of this research is to design global and robust attitude control systems for launch vehicles in exoatmospheric flight. An attitude control system is global when it guarantees that the vehicle converges to the desired attitude regardless of its initial condition. Global controllers are important because when large angle maneuvers must be performed, it is simpler to use a single global controller than several local controllers patched together. In addition, the designed autopilots must be robust with respect to uncertainties in the parameters of the vehicle, which means that global convergence must be achieved despite of those uncertainties. Two designs are carried out. In the first one possible delays introduced by the actuator are neglected. The design is performed by using a Lyapunov approach, and the obtained autopilot is a standard proportional-derivative controller. In the second one, the effects of the actuator are considered. Then the design is based on robust backstepping which leads to a memory-less nonlinear feedback of attitude, attitude-rate, and of the engine deflection angle. Both autopilots are validated in a case study.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.016
      Issue No: Vol. 74 (2018)
       
  • Neighboring optimal control for open-time multiburn orbital transfers
    • Authors: Zheng Chen; Shuo Tang
      Pages: 37 - 45
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Zheng Chen, Shuo Tang
      While a parametric approach has been proposed recently to establish the neighboring optimal control (NOC) for multiburn orbital transfer problems with fixed final time, it is not applicable once the final time is free. Indeed, it is quite common for orbital transfer missions to leave their final time free. In this paper, we continue developing the parametric approach and extend it to establishing the NOC for such problems with free final time. The key idea is to construct a new parameterized family of extremals such that every neighboring extremal around a nominal one is totally determined by its final time and a vector. Then, the NOC for free-time scenarios is readily established through deriving the Taylor expansion of the parameterized neighboring extremals; as a by-product, some prerequisite conditions, fundamental for establishing the NOC, are formulated. Moreover, a scheme for performing the NOC is proposed accordingly. Finally, to illustrate the improvement of this paper, a free-time multiburn orbital transfer problem is computed and some Monte Carlo campaigns are tested in the presence of various perturbations, showing that the final condition errors are significantly reduced by employing the NOC developed in the paper.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.003
      Issue No: Vol. 74 (2018)
       
  • An investigation of interface conditions inherent in detached-eddy
           simulation methods
    • Authors: L. Zhou; R. Zhao; W. Yuan
      Pages: 46 - 55
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): L. Zhou, R. Zhao, W. Yuan
      The interfaces where the RANS modeled areas match the LES resolved regions are comparatively investigated with regard to the four popular detached-eddy simulation (DES) variants; namely, one-equation Spalart–Allmaras (SA) and two-equations Menter's SST background DES methods (SA-DES and SST-DES), as well as their respective delayed versions (SA-DDES and SST-DDES). The comparisons are aimed at further interpretation of their performance differences under various flows. Although all four DES variants can consistently predict results in fully separated circular cylinder flow, the SST-DES interface is like the SA-DDES interface around the wall, which indicates that, in this case, the shielding function f d _ cor of SST-DDES is redundant. Moreover, the recalibrated f d _ cor for SST-DDES is found to preserve double the boundary-layer thicknesses in the flat-plate flow, and shown to be too conservative to resolve the unsteady vortex in the cavity-ramp flow. On the other hand, SA-DDES with the shielding function f d shows an advantage by properly balancing the need of reserving the RANS modeled regions for wall boundary layers and generating the unsteady turbulent structures in detached areas.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.005
      Issue No: Vol. 74 (2018)
       
  • Experimental study on ignition characteristics of kerosene–air mixtures
           in V-shaped burner with DC plasma jet igniter
    • Authors: Hua-Lei Zhang; Li-Ming He; Gao-Cheng Chen; Wen-Tao Qi; Jin-Lu Yu
      Pages: 56 - 62
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Hua-Lei Zhang, Li-Ming He, Gao-Cheng Chen, Wen-Tao Qi, Jin-Lu Yu
      In this paper, we investigate the effects of DC plasma jet igniter on ignition characteristics of the kerosene–air mixtures in V-shaped burner. The primary principles of DC plasma igniter design have been put forward to prolong the igniter life and improve ignition performance. The main characteristic of plasma igniter is that the anode arc root is fixed at anode nozzle outside surface other than the anode inner surface in conventional design. The electric characteristics of the plasma jet igniter including voltage–current characteristics, the arc column propagation, emission spectrum and anode arc root movement in a quiescent environment are investigated with high speed CCD together with a recording of the voltage and current. The entire ignition process of the plasma jet igniter is recorded by high speed CCD and the ignition delay times of the kerosene–air mixtures defined by the interval between the onset of the plasma igniter and the establishment of a stable flame are measured with two photomultipliers with one of 10 mm band pass filter centered at 780 mm (the plasma igniter emission spectrum of OI) and the other of 10 mm band pass filter centered at 430 mm (direct flame emission of C H ( A 2 Δ → X 2 Π ) ). Results show that the plasma igniter works in a restrike mode and the anode spot movement is irregular. The plasma igniter jet contains abundance active particles inferred from the emission spectrum. The ignition delay times increase at first and then decrease with the increase of excess air coefficient ranging from 0.91 to 6.38. The delay times increase quickly when the mixtures are extremely lean or rich.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.023
      Issue No: Vol. 74 (2018)
       
  • Immersion and invariance-based control of novel moving-mass flight
           vehicles
    • Authors: Changsheng Gao; Jianqing Li; Yidi Fan; Wuxing Jing
      Pages: 63 - 71
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Changsheng Gao, Jianqing Li, Yidi Fan, Wuxing Jing
      This paper proposes a novel configuration of the moving mass trim control system with a single moving mass and jet thrusters, which increases the control authority of control system while decreasing its volume and weight requirement. As a stepping stone, dynamics model of a moving-mass flight vehicle and control problem of attitude are formulated. Then, an adaptive controller based on the immersion and invariance approach is designed to guarantee asymptotic stability of the closed-loop system, and a slight improvement is developed to simplify the process. Furthermore, the estimator, which estimates aerodynamic coefficients and unmeasurable terms, is designed without certainty equivalence or linear parameterization. Instead, a procedure is provided to add cross terms between the parameter estimates and the plant states. Also, the associated stability proof is constructive and accomplished by the development of a Lyapunov function. In addition to stability proofs, numerical simulation results in different case are presented to illustrate the performance of the proposed schemes.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.017
      Issue No: Vol. 74 (2018)
       
  • Effect of cowl shock on restart characteristics of simple ramp type
           hypersonic inlets with thin boundary layers
    • Authors: Lianjie Yue; Yinan Jia; Xiao Xu; Xinyu Zhang; Peng Zhang
      Pages: 72 - 80
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Lianjie Yue, Yinan Jia, Xiao Xu, Xinyu Zhang, Peng Zhang
      The effect of cowl angle on the restart characteristics of simple ramp type hypersonic inlets was experimentally investigated in shock tunnel equipped with schlieren imagery and static pressure measurement. The cowl shock strength is found to be a key factor that determines the inlet restart and makes the restart contraction ratios significantly deviate from the Kantrowitz criterion. Stronger cowl shock tends to degrade the inlet restart capability by causing larger separation bubble and higher pressure loss during the restarting process. In particular, a sensitive range of the cowl angles, within which the restart contraction ratio decreases rapidly, was identified. A design concept of multiple noncoalesced cowl shocks was thus proposed and proven to significantly improve the inlet restart capability.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.018
      Issue No: Vol. 74 (2018)
       
  • Numerical study of non-reacting flowfields of a swirling trapped vortex
           ramjet combustor
    • Authors: Song Chen; Dan Zhao
      Pages: 81 - 92
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Song Chen, Dan Zhao
      In this work, 3D numerical investigations of a trapped vortex combustor operated in different swirling flow conditions are performed by solving Reynolds-averaged Navier–Stokes equations with Reynolds-stress model. Emphasis is placed on the non-reacting flowfield characteristics and the stability of the locked vortex. Validation is performed first by comparing the present results with experimental data available. It shows that the Reynolds-stress model can provide good predictions for flows with a swirl number up to 0.98. It is also found that the cavity vortex can be trapped well in flows with different swirl numbers. To further study the “locked” vortices, flow disturbances are introduced to the trapped vortex combustor via suddenly increasing swirl number from 0.6 to 0.98. The transient simulation results reveal that the cavity vortex is highly resistant to the flow disturbances and is still well trapped in the cavity, while vortex shedding of the conventional breakdown vortex is observed in the presence of the flow disturbances. Turbulence intensity and kinetic energy are found to be significantly increased by approximately 300%, which indicates that the fuel–air mixing can be dramatically improved. This study shows that the swirling trapped vortex combustor is an alternative promising robust and efficient combustor concept.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.006
      Issue No: Vol. 74 (2018)
       
  • Distributed adaptive synchronization for multiple spacecraft formation
           flying around Lagrange point orbits
    • Authors: Wei Wang; Giovanni Mengali; Alessandro A. Quarta; Jianping Yuan
      Pages: 93 - 103
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Wei Wang, Giovanni Mengali, Alessandro A. Quarta, Jianping Yuan
      This paper presents a distributed adaptive control framework for multiple spacecraft formation flying around Lagrange point orbits, which account for unmeasurable velocities and (spacecraft) mass uncertainties. The nominal trajectory for the formation system is a halo orbit parameterized by Fourier series expansions. Such an explicit, albeit approximate, description of the nominal trajectory facilitates each spacecraft in formation to include the relative state information into a cooperative feedback control system design, so that the relative motion can be driven towards a desired trajectory while maintaining a group synchronization during the maneuver. The developed distributed control strategies rely on the protocols formulated on an undirected topology with mutual information interactions, utilizing every available neighbor-to-neighbor communication data couplings, in order to improve the reliability of the formation. Numerical simulations show that the proposed adaptive control laws guarantee global asymptotic convergence and system robustness.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.007
      Issue No: Vol. 74 (2018)
       
  • Optimal nozzle Mach number for maximizing altitude of sounding rocket
    • Authors: Sang-Hyeon Lee
      Pages: 104 - 111
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Sang-Hyeon Lee
      A pseudo-analytic approach is applied to determine the optimal nozzle Mach number for maximizing the altitude of a sounding rocket flying in a standard atmosphere. The one-dimensional rocket momentum equation including thrust, gravitational force and aerodynamic drag is considered, for which it is impossible to obtain an analytic solution in a general form. In this work, a piecewise pseudo-analytic approach with a constant parameter introduced to make the velocity integral in the governing equation analytic is applied. The rocket flight in the standard atmosphere is analyzed by dividing the entire range into small intervals where the drag parameter and the gravitational acceleration can be treated as a constant in each interval. A characteristic equation exists and provides accurate predictions of the optimal nozzle Mach number for maximizing the altitude at burn-out state or at apogee.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.019
      Issue No: Vol. 74 (2018)
       
  • Gust response of rigid and elastically mounted airfoils at a transitional
           Reynolds number
    • Authors: Caleb J. Barnes; Miguel R. Visbal
      Pages: 112 - 119
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Caleb J. Barnes, Miguel R. Visbal
      This article explores the evolution of the unsteady flow structure for rigid and elastically mounted NACA0012 airfoils subject to a parallel vortical gust disturbance at a Reynolds number of Re = 150 , 000 using implicit large-eddy simulation coupled with a structural dynamics model. A Taylor vortex is supplied upstream of the airfoil and shown to be successful at eliciting laminar separation flutter in the elastically mounted system under conditions which normally require an artificial disturbance. Much of the gust-induced moment responsible for flutter excitation is supplied by a two-stage flow transition process on the undersurface of the wing after the gust passes the airfoil. The gust response triggers transition of the separated lower-surface boundary layer into spanwise coherent vortices followed by a laminar separation bubble accompanied by a secondary emergence of flow transition. These events appear to be analogous in some ways to the processes that appear during fully developed flutter but occur over a shorter timescale.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.025
      Issue No: Vol. 74 (2018)
       
  • Investigations of self-excited vibration in splitter plate with a cavity
           in the supersonic mixing layer
    • Authors: Hao Li; Jian-guo Tan; Ju-wei Hou; Dong-dong Zhang
      Pages: 120 - 131
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Hao Li, Jian-guo Tan, Ju-wei Hou, Dong-dong Zhang
      The combustion efficiency strongly depends on the mixing process of fuel–air for a combined cycle engine. For an efficient mixing process, self-excited vibration of splitter plate is more practical than that of forced vibration. We have carried out a comprehensive numerical and experimental study of self-excited vibration. Finite element method (FEM) is applied to analyze the natural modal of vibration. Large eddy simulations (LES) are employed for the mechanistic study of self-excited vibration and their influence factors are evaluated from the experimental studies. Displacements are accurately measured by a non-intrusive laser vibrometer. It is found that the first-order vibration shape is quasi-two-dimensional and responsible for improved mixing. We found that self-excited vibration of the splitter plate is motivated by a cavity due to the acoustic self-oscillation. Moreover, self-excited vibration depends on the gap distance between fixed rod and splitter plate, static pressure difference between upper and lower nozzle outlet, length to depth ratio (K) and aft wall angle (θ). At zero gap distance, the frequency is up to 2.5 times higher than that of 3 mm gap distance. So, reduction in gap distance can efficiently increase the frequency of self-excited vibration, which is an encouraging point for the future study.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.009
      Issue No: Vol. 74 (2018)
       
  • Numerical simulation for the decomposition of DT-3 in a monopropellant
           thruster
    • Authors: Zi-guang Gao; Guo-xiu Li; Tao Zhang; Xu-hui Liu; Zi-huan Wang; Xin Liu
      Pages: 132 - 144
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Zi-guang Gao, Guo-xiu Li, Tao Zhang, Xu-hui Liu, Zi-huan Wang, Xin Liu
      DT-3 is a high performance low-freezing monopropellant developed in the 1980s in China, DT-3 is also the world's first practical use of low freezing point propellant. As a monopropellant for pose control, it is necessary to precisely control the thrust output. In this paper, a three-dimensional simulation is conducted, mass flow rate, preheating temperature and the size of the catalytic bed are studied for the controlling of thrust and start-up process. The results indicate that: When the mass flow is around 92.8 g/s, the thrust force reaches the designed thrust. Increasing mass flow, will lead to the temperature rise in the catalytic bed, and thus slightly accelerate the reaction process, eventually obtaining a relatively small propellant remaining at the catalytic bed exit position. Secondly, the increase in the preheat temperature facilitates faster entry of the thruster into the desired output state, while the impact on the final thrust is limited. Thirdly, the length of the posterior catalytic bed has an optimal value, the short catalytic bed causes the components to not react completely, excessive catalytic bed will affect the thrust output, and the best length is around 32 mm.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.035
      Issue No: Vol. 74 (2018)
       
  • RISE and disturbance compensation based trajectory tracking control for a
           quadrotor UAV without velocity measurements
    • Authors: Xingling Shao; Qinxiao Meng; Jun Liu; Honglun Wang
      Pages: 145 - 159
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Xingling Shao, Qinxiao Meng, Jun Liu, Honglun Wang
      This paper presents a velocity-free robust trajectory tracking control for a quadrotor unmanned aerial vehicle (UAV) with consideration of parametric uncertainties and external disturbances by effectively integrating robust integral of the sign of the error (RISE) feedback control with extended state observer (ESO). The original cascaded dynamics of quadrotor UAV is first derived in a strict form with lumped disturbances. Then, the robust RISE partial state feedback controllers with disturbance compensation are respectively synthesized in position and attitude loop, where the unmeasurable velocity states and disturbance compensation terms are estimated by ESO, the synthesized RISE control law is then accounted for attenuating the residual estimation error to achieve enhanced robustness against uncertainties. The major feature of proposed method is that fundamentally different anti-disturbance mechanisms of disturbance suppression-based RISE control and disturbance observer-based control are combined to handle the lumped disturbances simultaneously, which preserves their theoretical advantages while overcoming their performance limitations. Moreover, the proposed controller theoretically guarantees that the tracking error converges to a small neighborhood around the origin. The effectiveness and superiority of proposed control method are investigated in simulations against disturbances due to parametric uncertainties, wind gust and bounded perturbations.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.029
      Issue No: Vol. 74 (2018)
       
  • Design method of internal waverider inlet under non-uniform upstream for
           inlet/forebody integration
    • Authors: Guoping Huang; Fengyuan Zuo; Wenyou Qiao
      Pages: 160 - 172
      Abstract: Publication date: March 2018
      Source:Aerospace Science and Technology, Volume 74
      Author(s): Guoping Huang, Fengyuan Zuo, Wenyou Qiao
      A novel Bump-integrated three-dimensional internal waverider inlet (IWI) design method is presented for high-speed inlet/forebody integration. The low-kinetic-energy (boundary layer) flow generated by a blunted leading-edge and forebody boundary layer represents an extreme challenge in the integration of aircraft forebody and inlet. In this method, such an inlet's flowpath is divided into the entrance shockwave segment, the isentropic compression segment and the isolator. First, a three-dimensional inverse method of characteristics (3D-IMOC) is developed to obtain a compression surface that can generate a requested entrance shock wave in non-uniform upstream flow. This configuration realizes the integration of IWI and aircraft fuselage by incorporating a Bump to remove most of the boundary layer flow. This is followed by a three-dimensional, isentropic compression flow-path with cross sectional areas conforming to the specified Mach number distribution. Finally, a new three-dimensional Bump-integrated IWI was tested in M = 6 wind tunnel, under a rather thick boundary layer upstream flow (37% height of inlet entrance). Both of the experimental data and numerical simulation results show that, the new method of IWI and Bump can overcome serious boundary layer flow problems and improve the inlet performance.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.012
      Issue No: Vol. 74 (2018)
       
  • Optimal fuel consumption finite-thrust orbital hopping of aeroassisted
           spacecraft
    • Authors: Runqi Chai; Al Savvaris; Antonios Tsourdos; Senchun Chai; Yuanqing Xia
      Abstract: Publication date: Available online 3 February 2018
      Source:Aerospace Science and Technology
      Author(s): Runqi Chai, Al Savvaris, Antonios Tsourdos, Senchun Chai, Yuanqing Xia
      In the paper, the problem of minimum-fuel aeroassisted spacecraft regional reconnaissance (orbital hopping) is considered. A new nonlinear constrained optimal control formulation is designed and constructed so as to describe this mission scenario. This formulation contains multiple exo-atmospheric and atmospheric flight phases and correspondingly, two sets of flight dynamics. The constructed continuous-time optimal control system is then discretized via a multi-phase global collocation technique. The resulting discrete-time system is optimized using a newly proposed gradient-based optimization algorithm. Several comparative simulations are carried out and the obtained optimal results indicate that it is effective and feasible to use the proposed multi-phase optimal control design for achieving the aeroassisted vehicle orbital hopping mission.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.026
       
  • Distributed almost global finite-time attitude consensus of multiple
           spacecraft without velocity measurements
    • Authors: Haichao Gui; George Vukovich
      Abstract: Publication date: Available online 3 February 2018
      Source:Aerospace Science and Technology
      Author(s): Haichao Gui, George Vukovich
      This paper addresses the attitude consensus problem of multiple rigid bodies in terms of the unit quaternion parameterization. By employing Lyapunov theory and homogeneous techniques, distributed finite-time attitude consensus laws are proposed for leader-following and leaderless multi-agent systems, with full-state (i.e., attitude plus angular velocity) or attitude-only measurements. Specifically, sliding mode observers are used to estimate the leader's information in finite time for followers without direct access to the leader. The so-called “separation principle” is then established between the observers and the consensus controllers. In addition, quaternion filtering systems are constructed to inject the necessary damping into the closed-loop system when angular velocity measurements are absent. In all scenarios, the proposed methods ensure almost global finite-time convergence, avoid the unwinding problem, and yield continuous control torques with a priori known bounds. Numerical examples demonstrate the effectiveness of the proposed methods.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.015
       
  • Spacecraft attitude fault-tolerant control based on iterative learning
           observer and control allocation
    • Authors: Qinglei Hu; Guanglin Niu; Chenliang Wang
      Abstract: Publication date: Available online 3 February 2018
      Source:Aerospace Science and Technology
      Author(s): Qinglei Hu, Guanglin Niu, Chenliang Wang
      In this paper, an observer-based fault-tolerant control scheme is proposed for the attitude stabilization of rigid spacecraft in the presence of actuator fault, configuration misalignment, input saturation and even external disturbances simultaneously. More specifically, an iterative learning observer is firstly developed to estimate the torque deviation and steer the estimation errors into some small residual sets. And also the detailed derivations of the observer are provided, along with a thorough analysis for the associated ultimate bounded stability and estimation error convergence property. Then, an integral-type sliding mode control law is designed to produce the three-axis virtual control signals with the desired performance for being distributed among the individual actuators. Under this, a robust control allocation algorithm is developed to map the virtual control demand onto individual actuator in an optimal manner, which takes into account the estimation uncertainties and ensures some fault-tolerant ability. The key feature of the proposed strategies is that the whole closed-loop fault tolerant control system can be guaranteed theoretically to be stable by the development of Lyapunov methodology. Numerical simulation results are presented to illustrate and highlight the fine performance benefits obtained using the proposed schemes.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.031
       
  • Time Efficient Sliding Mode Controller based on Bang-Bang Logic for
           Satellite Attitude Control
    • Authors: You Li; Dong Ye; Zhaowei Sun
      Abstract: Publication date: Available online 3 February 2018
      Source:Aerospace Science and Technology
      Author(s): You Li, Dong Ye, Zhaowei Sun
      In order to improve the convergence rate of standard sliding mode controller, time efficient controllers based on Bang-Bang logic for satellite attitude stabilization and tracking control are developed in this paper. The time efficient open-loop control algorithm Bang-Bang control is combined with closed-loop sliding mode control to improve system robustness. A two-stage structure sliding mode with a fixed angular velocity stage and a fixed deceleration stage is proposed in this paper. The sliding mode parameter is real-time updating hence the modified sliding mode could have Bang-Bang character. The system inertia matrix uncertainty and disturbance torque is discussed and the controller proposed in this paper is robust to the perturbation. The control torque constraint is also discussed and the constraint on control parameters is given. The performance of the controller is demonstrated by numerical simulation.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.042
       
  • Multidimensional material response simulations of a full-scale tiled
           ablative heatshield
    • Authors: Jeremie B.E. Meurisse; Jean Lachaud; Francesco Panerai; Chun Tang; Nagi N. Mansour
      Abstract: Publication date: Available online 2 February 2018
      Source:Aerospace Science and Technology
      Author(s): Jeremie B.E. Meurisse, Jean Lachaud, Francesco Panerai, Chun Tang, Nagi N. Mansour
      The Mars Science Laboratory (MSL) was protected during Mars atmospheric entry by a 4.5 meter diameter heatshield, which was constructed by assembling 113 thermal tiles made of NASA's flagship porous ablative material, Phenolic Impregnated Carbon Ablator (PICA). Analysis and certification of the tiles thickness were based on a one-dimensional model of the PICA response to the entry aerothermal environment. This work provides a detailed three-dimensional heat and mass transfer analysis of the full-scale MSL tiled heatshield. One-dimensional and three-dimensional material response models are compared at different locations of the heatshield. The three-dimensional analysis is made possible by the use of the Porous material Analysis Toolbox based on OpenFOAM (PATO) to simulate the material response. PATO solves the conservation equations of solid mass, gas mass, gas momentum and total energy, using a volume-averaged formulation that includes production of gases from the decomposition of polymeric matrix. Boundary conditions at the heatshield forebody surface were interpolated in time and space from the aerothermal environment computed with the Data Parallel Line Relaxation (DPLR) code at discrete points of the MSL trajectory. A mesh consisting of two million cells was created in Pointwise, and the material response was performed using 840 processors on NASA's Pleiades supercomputer. The present work constitutes the first demonstration of a three-dimensional material response simulation of a full-scale ablative heatshield with tiled interfaces. It is found that three-dimensional effects are pronounced at the heatshield outer flank, where maximum heating and heat loads occur for laminar flows.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.013
       
  • On the influence of optimization algorithm and initial design on wing
           aerodynamic shape optimization
    • Authors: Yin Yu; Zhoujie Lyu; Zelu Xu; Joaquim R.R.A. Martins
      Abstract: Publication date: Available online 2 February 2018
      Source:Aerospace Science and Technology
      Author(s): Yin Yu, Zhoujie Lyu, Zelu Xu, Joaquim R.R.A. Martins
      Aerodynamic shape optimization is a useful tool in wing design, but the impact of the choice of optimization algorithm and the multimodality of the design space in wing design optimization is still poorly understood. To address this, we benchmark both gradient-based and gradient-free optimization algorithms for computational fluid dynamics based aerodynamic shape optimization problems based on the Common Research Model wing geometry. The aerodynamic model solves the Reynolds-averaged Navier–Stokes equations with a Spalart–Allmaras turbulence model. The drag coefficient is minimized subject to lift, pitching moment, and geometry constraints, with up to 720 shape variables and 11 twist variables for two mesh sizes. We benchmark six gradient-based and three gradient-free algorithms by comparing both the accuracy of the optima and the computational cost. Most of the optimizers reach similar optima, but the gradient-based methods converge to more accurate solutions at a much lower computational cost. Since multimodality and nonsmoothness of the design space are common arguments for the use of gradient-free methods, we investigate these issues by solving the same optimization problem starting from a series of randomly generated initial geometries, as well as a wing based on the NACA 0012 airfoil with zero twist and constant thickness-to-chord ratio. All the optimizations consistently converge to practically identical results, where the differences in drag are within 0.05%, and the shapes and pressure distributions are very similar. Our overall conclusion is that the design space for wing design optimization with a fixed planform is largely convex, with a very small flat region that is multimodal because of numerical errors. However, this region is so small, and the differences in drag so minor, that the design space can be considered unimodal for all practical purposes.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2018.01.016
       
  • Scramjet performance for ideal combustion processes
    • Authors: Tristan Vanyai; Mathew Bricalli; Stefan Brieschenk; Russell R. Boyce
      Abstract: Publication date: Available online 2 February 2018
      Source:Aerospace Science and Technology
      Author(s): Tristan Vanyai, Mathew Bricalli, Stefan Brieschenk, Russell R. Boyce
      Models used to analyse scramjet engine cycles are typically either very simple or focused on specific combustor geometries. This paper uses a quasi-one-dimensional, inviscid chemical equilibrium solver to examine scramjet engines defined by predetermined combustion processes. This solver is capable of rapidly analysing combustion processes without requiring predefined geometries and is validated against the Hyshot II flight experiment and a constant area combustor ground test. Combustion occurs using constant area, constant pressure, constant Mach number or constant temperature processes. Constant area combustion typically produces the highest specific impulse for given combustor entrance conditions. Maximum pressure and temperature limitations were introduced, and multiple engines with combinations of combustion process were examined. It was found that engines with a combination of all four combustion processes can have lower maximum values of pressure and temperature whilst maintaining high performance compared with constant area combustor engines.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.ast.2017.12.021
       
  • Low-complexity prescribed performance control for spacecraft attitude
           stabilization and tracking
    • Authors: Jianjun Luo; Caisheng Wei; Honghua Dai; Zeyang Yin; Xing Wei; Jianping Yuan
      Abstract: Publication date: Available online 2 February 2018
      Source:Aerospace Science and Technology
      Author(s): Jianjun Luo, Zeyang Yin, Caisheng Wei, Jianping Yuan
      In this paper, the attitude stabilization and tracking control problem is investigated for spacecraft with consideration of unknown dynamics and external disturbance. A low-complexity prescribed performance attitude control scheme is presented to solve this problem. Different from existing works, there are threefold prominent advantages. The first one is the transient performance and the steady performance of the system is guaranteed by a user-defined function rather than depending on repeated adjustment of controller parameters. The second is that no information of the system and external disturbance is necessary in the developed control scheme, which means the method is model-free. Moreover, the developed low-complexity controller is calculated without any time-consuming iterative operations; thus it's significantly advantageous in engineering applications. It is proved that the state variables converge to the prescribed region at a prescribed exponential rate under the proposed control scheme. Four groups of numerical simulations are organized to validate the effectiveness of the method.

      PubDate: 2018-02-05T12:25:54Z
      DOI: 10.1016/j.isatra.2018.01.016
       
  • Three-dimensional salvo attack guidance considering communication delay
    • Authors: Shaoming He; Mingu Kim; Tao Song; Defu Lin
      Pages: 1 - 9
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Shaoming He, Mingu Kim, Tao Song, Defu Lin
      This paper investigates the problem of salvo attack guidance law design for multiple missiles against a stationary target considering communication delay. The proposed guidance law requires no information on time-to-go and consists of two guidance stages. More specifically, a simple decentralized midcourse guidance law is proposed to provide desired initial conditions for the latter phase. As for the second guidance phase, all missiles are switched to classical pure proportional navigation guidance law. Considering the fact that the communication delay is inevitable in real applications, the corresponding allowed maximum upper bound of the communication delay among all interceptors is also derived analytically. Nonlinear numerical simulations clearly confirm the effectiveness of the proposed formulation.

      PubDate: 2017-12-12T17:56:50Z
      DOI: 10.1016/j.ast.2017.11.019
      Issue No: Vol. 73 (2017)
       
  • Characterisation of turbine behaviour for an engine overspeed prediction
           model
    • Authors: Lucas Pawsey; David John Rajendran; Vassilios Pachidis
      Pages: 10 - 18
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Lucas Pawsey, David John Rajendran, Vassilios Pachidis
      This paper focuses on the characterisation of turbine overspeed behaviour to be integrated into an engine overspeed model capable of predicting the terminal speed of the high pressure turbine (HPT) in the event of a high pressure shaft failure. The engine considered in this study features a single stage HPT with a shrouded contra-rotating rotor with respect to the single stage intermediate pressure turbine (IPT). The HPT performance is characterised in terms of torque and mass flow function for a range of expansion ratios at various non-dimensional rotational speeds (NH), up to 200% of the design value. Additionally, for each HPT expansion ratio and NH, the change in capacity of the downstream IPT, for different IPT non-dimensional rotational speeds (NI), also needs to be characterised due to the extremely positive incidence angle of the flow from the upstream rotor. An automated toolkit is developed to generate these characteristic maps for both the HPT and IPT. An unlocated high pressure shaft failure will result in rearward movement of the rotor sub-assembly. This causes changes in the rotor tip and rim seal regions, and in the rim seal leakage flow properties. Therefore, in the present work, a high fidelity characterisation of turbine behaviour with the inclusion of tip and rim seals is carried out at three different displacement locations, 0 mm, 10 mm and 15 mm, to improve terminal speed estimation. Furthermore, there is a possibility of damage to the tip seal fins of the HPT rotor due to unbalance in the spool that may result in contact between the rotor aerofoil tip and the casing. Consequently, another set of characteristics are generated with damaged tip fins at each displacement location. It is observed from the characteristics that the torque of the HPT rotor decreases with increasing NH. The HPT mass flow function initially decreases and then increases with an increase in NH. The IPT mass flow function initially remains similar and then decreases with increase in NH above values of 150%. The HPT rotor torque and IPT mass flow function decrease with rearward movement of the HPT rotor sub-assembly for all values of NH. With worn tip seal fins all parameters mentioned previously are lower than in the nominal undamaged case. The high fidelity characterisation of turbines that follows the sequence of events after a shaft failure, as described in this work, can provide accurate predictions of terminal speed and thus act as a tool for testing design modifications that can result in better management and control of the over-speed event.

      PubDate: 2017-12-12T17:56:50Z
      DOI: 10.1016/j.ast.2017.11.037
      Issue No: Vol. 73 (2017)
       
  • Development of the ILR-33 “Amber” sounding rocket for
           microgravity experimentation
    • Authors: Blazej Marciniak; Adam Okninski; Bartosz Bartkowiak; Michal Pakosz; Kamil Sobczak; Wojciech Florczuk; Damian Kaniewski; Jan Matyszewski; Pawel Nowakowski; Dawid Cieslinski; Grzegorz Rarata; Pawel Surmacz; Dominik Kublik; Damian Rysak; Jaromir Smetek; Piotr Wolanski
      Pages: 19 - 31
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Blazej Marciniak, Adam Okninski, Bartosz Bartkowiak, Michal Pakosz, Kamil Sobczak, Wojciech Florczuk, Damian Kaniewski, Jan Matyszewski, Pawel Nowakowski, Dawid Cieslinski, Grzegorz Rarata, Pawel Surmacz, Dominik Kublik, Damian Rysak, Jaromir Smetek, Piotr Wolanski
      This paper gives an overview of the development of the ILR-33 ”Amber” sounding rocket designated for microgravity experiments, that is under development at Institute of Aviation in Warsaw, Poland. The lack of an easily accessible and affordable platform for this kind of research was one of the key reasons for this work. The proposed design enables performing experiments in microgravity for almost 150 seconds with an apogee over 100 km. Combining these results with a relatively low price per launch and short deployment time gives a possibility to establish a firm position on the dynamic market. This article describes also the rocket structure and the vehicle's capabilities. The proposed design utilizes a hybrid rocket motor with High Test Peroxide as an oxidizer along with two reusable solid rocket boosters. The early phase analysis of the rocket configuration and propellant considerations are also presented in the paper. Furthermore, there have been already several on-ground test performed such as: wind tunnel research and motor firings. The proposed design is considered as an introduction to small launch vehicle technology.

      PubDate: 2017-12-12T17:56:50Z
      DOI: 10.1016/j.ast.2017.11.034
      Issue No: Vol. 73 (2017)
       
  • Conceptual design of a Blended Wing Body MALE UAV
    • Authors: P. Panagiotou; S. Fotiadis-Karras; K. Yakinthos
      Pages: 32 - 47
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): P. Panagiotou, S. Fotiadis-Karras, K. Yakinthos
      The current work is an aerodynamic design study of a Blended Wing Body (BWB) Medium-Altitude-Long-Endurance (MALE) Unmanned-Aerial-Vehicle (UAV). Using a combined approach of presizing tools and computational simulations, a step-by-step layout design study was conducted to define the key layout characteristics and select the optimal airframe-engine combination. Trade studies were also carried out to optimize the aerodynamic performance and stability. The traditional sizing and aerodynamic estimation methods were adopted to incorporate the characteristics of the BWB platform, whereas CFD computations were employed in order to calculate the aerodynamic and stability coefficients, during the layout comparison and trade studies. Drawings and tables are provided to show the progression of the design study at each stage. The performance specifications are also compared with a conventional UAV platform to point out the main advantages and disadvantages of the BWB for MALE UAV applications.

      PubDate: 2017-12-12T17:56:50Z
      DOI: 10.1016/j.ast.2017.11.032
      Issue No: Vol. 73 (2017)
       
  • Fault detection in operating helicopter drivetrain components based on
           support vector data description
    • Authors: V. Camerini; G. Coppotelli; S. Bendisch
      Pages: 48 - 60
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): V. Camerini, G. Coppotelli, S. Bendisch
      The objective of the paper is to develop a vibration-based automated procedure dealing with early detection of mechanical degradation of helicopter drive train components using Health and Usage Monitoring Systems (HUMS) data. An anomaly-detection method devoted to the quantification of the degree of deviation of the mechanical state of a component from its nominal condition is developed. This method is based on an Anomaly Score (AS) formed by a combination of a set of statistical features correlated with specific damages, also known as Condition Indicators (CI), thus the operational variability is implicitly included in the model through the CI correlation. The problem of fault detection is then recast as a one-class classification problem in the space spanned by a set of CI, with the aim of a global differentiation between normal and anomalous observations, respectively related to healthy and supposedly faulty components. In this paper, a procedure based on an efficient one-class classification method that does not require any assumption on the data distribution, is used. The core of such an approach is the Support Vector Data Description (SVDD), that allows an efficient data description without the need of a significant amount of statistical data. Several analyses have been carried out in order to validate the proposed procedure, using flight vibration data collected from a H135, formerly known as EC135, servicing helicopter, for which micro-pitting damage on a gear was detected by HUMS and assessed through visual inspection. The capability of the proposed approach of providing better trade-off between false alarm rates and missed detection rates with respect to individual CI and to the AS obtained assuming jointly-Gaussian-distributed CI has been also analysed.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.043
      Issue No: Vol. 73 (2017)
       
  • “Wasted performance” minimization of the multi-purpose mini-satellite
           platform for an EO mission using a dynamic simulation-based model
    • Authors: Asad Saghari; Shima Rahmani; Amirreza Kosari; Masoud Ebrahimi
      Pages: 61 - 77
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Asad Saghari, Shima Rahmani, Amirreza Kosari, Masoud Ebrahimi
      With growing interest in low-cost, high performance and short time-to-flight satellites, the idea of using multi-purpose satellite platforms has drawn many attentions. However, under certain conditions in which the required capability of a mission is slightly more than the capabilities of a specified variant of a multi-purpose platform, it must jump to the next higher capability variant. Correspondingly, lots of unnecessary performance or capability might be imposed. In this paper, the main goal is to minimize the wasted performance in the design of an Earth Observation (EO) satellite based on a multi-purpose platform. To this end, a dynamic simulation-based model has been developed with the capability of simulating the satellite performance characteristics, EO mission requirements and supportable performance by the platform under various circumstances and throughout the satellite lifetime. To minimize the wasted performance we face with an optimization problem containing a complicated non-linear, non-convex and multi-modal design space which is enveloped by constraints of the mission requirements and supportable performance of the platform. A novel criterion as an objective function has been introduced aiming at the incompatibility reduction between the capabilities of a certain multi-purpose platform variant and the EO mission requirements. This objective function considers all the performance variations during the entire mission lifetime rather than solely the worst case. Through this, the mission and orbital characteristics have been determined by which there would be no need to jump to the higher performance variant in order to satisfy the EO mission requirements.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.038
      Issue No: Vol. 73 (2017)
       
  • Ignition delay kinetic model of boron particle based on bidirectional
           diffusion mechanism
    • Authors: Daolun Liang; Jianzhong Liu; Yunan Zhou; Junhu Zhou; Kefa Cen
      Pages: 78 - 84
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Daolun Liang, Jianzhong Liu, Yunan Zhou, Junhu Zhou, Kefa Cen
      Ignition delay time of B particles is one of the key factors that influence their burnout ratio in the afterburner. In this study, the micro/nanofabricated slice measurement of a combustion residue particle of B was carried out. By combining the experimental results obtained with previous experimental results, the surface diffusion mechanism of a single B particle was completely verified. Then, an ignition delay kinetic model of B particle was developed using the principles of semiempirical models. By ensuring the initial ignition temperature, the ignition delay of a single B particle can be divided into two stages: (i) heat transfer stage and (ii) low-temperature oxidation stage. The existence of both O2 diffusion and (BO) n diffusion (bidirectional diffusion) was confirmed during the low-temperature oxidation stage. Only heat transfer between the B particle and surroundings occurred during the heat transfer stage, whereas both heat transfer and oxidation occurred during the low-temperature oxidation stage. The oxidation involves four global reactions: (i) evaporation of B2O3, (ii) diffusion of O2, (iii) diffusion of (BO) n , and (iv) reaction of H2O. The final ignition delay time of a B particle is equal to the sum of the lasting times of heat transfer stage and low-temperature oxidation stage. The results of computed ignition delay time obtained by the model are consistent with the previous experimental data under O2/H2O atmosphere. According to the prediction of the model, the increase in the initial particle size will prolong the ignition delay time of a B particle. The ignition delay of a small B particle is dominated by the low-temperature oxidation stage, whereas the ignition delay of a large B particle is dominated by the heat transfer stage.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.044
      Issue No: Vol. 73 (2017)
       
  • Civil turbofan engine exhaust aerodynamics: Impact of bypass nozzle
           after-body design
    • Authors: Ioannis Goulos; Tomasz Stankowski; David MacManus; Philip Woodrow; Christopher Sheaf
      Pages: 85 - 95
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Ioannis Goulos, Tomasz Stankowski, David MacManus, Philip Woodrow, Christopher Sheaf
      It is envisaged that the next generation of civil large turbofan engines will be designed for greater bypass ratios when compared to contemporary architectures. The underlying motivation is to reduce specific thrust and improve propulsive efficiency. Concurrently, the aerodynamic performance of the exhaust system is anticipated to play a key role in the success of future engine architectures. The transonic flow topology downstream of the bypass nozzle can be significantly influenced by the after-body geometry. This behavior is further complicated by the existence of the air-flow vent on the nozzle after-body which can have an impact on the performance of the exhaust system. This paper aims to investigate the aerodynamics associated with the geometry of the bypass nozzle after-body and to establish guidelines for the design of separate-jet exhausts with respect to future large turbofan engines. A parametric geometry definition has been derived based on Class-Shape Transformation (CST) functions for the representation of post-nozzle-exit components such as after-bodies, plugs, and air-flow vents. The developed method has been coupled with an automatic mesh generation and a Reynolds Averaged Navier–Stokes (RANS) flow solution method, thus devising an integrated aerodynamic design tool. A cost-effective optimization strategy has been implemented consisting of methods for Design Space Exploration (DSE), Response Surface Modeling (RSM), and Genetic Algorithms (GAs). The combined approach has been deployed to explore the aerodynamic design space associated with the bypass nozzle after-body geometry for a Very High Bypass Ratio (VHBR) turbofan engine with separate-jet exhausts. A detailed investigation has been carried out to expose the transonic flow mechanisms associated with the effect of after-body curvature combined with the impact of the air-flow vent. A set of optimum curved after-body geometries has been obtained, with each subsequently compared against their respective conical representation. The obtained results suggest that no significant performance improvements can be obtained through curving the nozzle after-body relative to the case of a conical design. However, it is shown that the application of surface curvature has the potential to unlock new parts in the design space that allow analysts to reduce the required after-body length without any loss in aerodynamic performance. The developed approach complements the existing tool-set of enabling technologies for the design and optimization of future large aero-engines, consequently leading to increased thrust and reduced Specific Fuel Consumption (SFC).

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.09.002
      Issue No: Vol. 73 (2017)
       
  • An analytical study of the vibroacoustic response of a ribbed plate
    • Authors: Tao Fu; Zhaobo Chen; Hongying Yu; Chengfei Li; Xiaoxiang Liu
      Pages: 96 - 104
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Tao Fu, Zhaobo Chen, Hongying Yu, Chengfei Li, Xiaoxiang Liu
      An analytical model is developed to investigate the sound transmission loss from orthogonally rib-stiffened plates structure under diffuse acoustic field excitation. The validity and feasibility of the model are verified by comparing the present theoretical predictions with numerical and experimental results published previously. The influences of modal coupling terms, boundary condition and stiffener spacing on sound power and sound transmission loss are subsequently presented. Numerical discussion of the model demonstrates the significant influence of both boundary conditions and stiffener spacing upon the mode shape, sound power and sound transmission loss for stiffened plate, wherein sound power decreases and sound transmission loss increases as the amount of constraint increases.

      PubDate: 2017-12-12T17:56:50Z
      DOI: 10.1016/j.ast.2017.11.047
      Issue No: Vol. 73 (2017)
       
  • A combined criteria-based method for hypersonic three-dimensional boundary
           layer transition prediction
    • Authors: Ling Zhou; Rui Zhao; Renfu Li
      Pages: 105 - 117
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Ling Zhou, Rui Zhao, Renfu Li
      In this paper, a new method for predicting hypersonic three-dimensional (3D) boundary layer transition is developed. It is based on the Re θ / M e criterion for streamwise instability and on the Re cf , new criterion for crossflow instability. An intermittency function is also formulated and applied to combine laminar and turbulent flows. Additionally, a computational grid pretreating method, compatible with modern computational fluid dynamics (CFD) techniques based on parallel execution is adopted in order to obtain the boundary layer parameters. Four criteria are compared to define the boundary layer edge. A HIFiRE-5 elliptic cone at different Reynolds numbers is adopted to validate the performance of the criteria-based transition model and the effectiveness of the four criteria for boundary layer edge definition. The results show that the boundary layer edge of complex hypersonic 3D flows could be obtained properly with the computational grid pretreating method and the combination of h 0 / h 0 , ∞ = 0.995 and ( d u / d H ) / ( d u / d H ) w = 0.1 criteria. Moreover, the computed Re θ / M e and Re cf , new distributions in the region between the leading edge and the centerline are similar to the N-factor for streamwise and crossflow instabilities from linear parabolized stability equation (PSE) methods. The shape and trend of the transition onsets predicted by the criteria-based transition model between the centerline and leading edge of HIFiRE-5 agree well with the experiment. However, as for the transition on the centerline, which is dominated by the inflection point in streamwise velocity profiles, using the criteria-based transition model related to boundary layer thickness would predict the transition onset prematurely.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.002
      Issue No: Vol. 73 (2017)
       
  • Improved rotor aeromechanics predictions using a fluid structure
           interaction approach
    • Authors: Younghyun You; Deokhwan Na; Sung N. Jung
      Pages: 118 - 128
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Younghyun You, Deokhwan Na, Sung N. Jung
      The measured HART (Higher harmonic control Aeroacoustic Rotor Test) I data in a descending flight condition is validated using various numerical approaches including CFD (Computational Fluid Dynamics)–CSD (Computational Structural Dynamics) coupled analyses with isolated rotor model and rotor-fuselage model. A CSD-alone approach is also conducted for reference purpose. A three-dimensional (3D) compressible RANS (Reynolds Averaged Navier Stokes) flow solver is employed for the CFD code. Good convergence behavior is found for both coupling analyses. It is observed that the rotor-fuselage model improves the correlation significantly as compared with the measured data. Specifically, the highly oscillating section normal forces signals marked in the advancing and retreating sides of the rotor are captured accurately. Detailed harmonic analysis and the gradient of the airloads signals are observed to prove the validity of the prediction model. The upwash induced due to a fuselage as well as the increased vorticity over the rotor flow fields are attributed to the enhanced correlation. The predicted blade elastic motions and structural moments also indicate improvements with the present rotor-fuselage model.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.041
      Issue No: Vol. 73 (2017)
       
  • First-order shear deformation theory for orthotropic doubly-curved shells
           based on a modified couple stress elasticity
    • Authors: Farajollah Zare Jouneghani; Payam Mohammadi Dashtaki; Rossana Dimitri; Michele Bacciocchi; Francesco Tornabene
      Pages: 129 - 147
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Farajollah Zare Jouneghani, Payam Mohammadi Dashtaki, Rossana Dimitri, Michele Bacciocchi, Francesco Tornabene
      This paper investigates the micro- and nano-mechanical behavior of orthotropic doubly-curved shells by considering the New Modified Couple Stress Theory (NMCST). The higher order continuum assumed by the NMCST includes three material length scale parameters in order to capture the size-effect of anisotropic and orthotropic materials. The governing equations of the problem are based on the First-order Shear Deformation Theory (FSDT). According to the proposed NMCST, the expressions of the physical components for the strain and curvature tensors are obtained in an orthogonal curvilinear coordinate system. Then, the governing differential equations and boundary conditions are derived by applying the energy method and Hamilton's principle. A comparative investigation between our numerical results and the ones available in the literature proves the capability of the proposed formulation in predicting the micro- and nano-mechanical behavior of orthotropic doubly-curved shells.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.045
      Issue No: Vol. 73 (2017)
       
  • An efficient single-loop strategy for reliability-based multidisciplinary
           design optimization under non-probabilistic set theory
    • Authors: Xiaojun Wang; Ruixing Wang; Lei Wang; Xianjia Chen; Xinyu Geng
      Pages: 148 - 163
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Xiaojun Wang, Ruixing Wang, Lei Wang, Xianjia Chen, Xinyu Geng
      Non-probabilistic reliability based multidisciplinary design optimization (NRBMDO) offers a powerful tool for making reliable decisions with the consideration of uncertain-but-bounded uncertainties for complex engineering systems. However, the prohibitive computation and convergence difficulties caused by the directly coupling of uncertainty based multidisciplinary analysis (UMDA), non-probabilistic reliability analysis (NRA) and MDO would seriously hamper the application of NRBMDO. In this paper, an efficient single loop strategy for NRBMDO (SLS_NRBMDO) is developed to decouple the nested issue and thus improve the computational efficiency. The key idea of the proposed strategy is decoupling NRBMDO with several cycles of sequential MDO, UMDA, NRA and translating distance calculation (TDC). For UMDA, three methods, i.e., the first order interval Taylor expansion method, the interval vertex theorem, the direct optimization approach are formulated. Besides, NRA is conducted on the basis of the expanded non-probabilistic stress–strength interference model and the volume ratio thought, which provides a clear and definite assessment criterion for the structural safety with uncertain-but-bounded parameters. Furthermore, the translating strategy based on the performance measure approach is proposed to shift and update the constraints, and the expression of the translating distance is mathematically derived to accelerate the design procedure. Eventually, the effectiveness and efficiency of the proposed method are illustrated with one numerical case and one practical supersonic wing optimization design problem.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.11.046
      Issue No: Vol. 73 (2017)
       
  • An experimental study of surface wettability effects on dynamic ice
           accretion process over an UAS propeller model
    • Authors: Yang Liu; Linkai Li; Haixing Li; Hui Hu
      Pages: 164 - 172
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Yang Liu, Linkai Li, Haixing Li, Hui Hu
      An experimental study was conducted to evaluate the effects of surface wettability on the dynamic ice accretion process over the surface of a rotating Unmanned-Aerial-System (UAS) propeller model and the resultant aerodynamic performance degradation due to the ice accretion. A propeller model was installed in an Icing Research Tunnel at Iowa State University (i.e., ISU-IRT) with its surface wettability changed significantly (i.e., hydrophilic surface versus superhydrophobic surface). In addition to acquiring “phase-locked” images to reveal the dynamic ice accretion process over the rotating propeller surface, the thrust generation and the required power input to drive the propeller model to operate at a constant rotation speed were also measured during the ice accretion process. The dynamic ice accretion process over the rotating propeller surface was found to vary remarkably with changes to the propeller surface wettability. By making the propeller surface superhydrophobic, the detrimental effects of the ice accretion on the aerodynamic performance of the propeller model were found to be mitigated greatly with much less ice accretion over the propeller surface, significant reduction of the thrust loss and less demand for extra power consumption due to the ice accretion, in comparison with the case with the propeller surface being hydrophilic.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.003
      Issue No: Vol. 73 (2017)
       
  • UAV collision avoidance exploitation for noncooperative trajectory
           modification
    • Authors: Pietro Pierpaoli; Amir Rahmani
      Pages: 173 - 183
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Pietro Pierpaoli, Amir Rahmani
      Distributed collision-free trajectories are generally obtained through a continuous sharing of information between vehicles. With the intent of investigating possible sources of vulnerability in autonomous frameworks, we formalize a procedure malicious players can follow to influence other. In this paper we propose a strategy for steering a UAV towards predetermined targets. The strategy described here relies on the existence of a flight information sharing protocol (i.e. ADS-B) and predictable collision avoidance algorithms. A model predictive controller is applied to the switching system representing a pair of UAVs coupled by the presence of an imminent collision. As showed by means of numerical simulations and robot experiments, the result is a loss of autonomy on the UAV. Our results suggest the need to include the subject of our study in the discussion on safe automated airspace.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.008
      Issue No: Vol. 73 (2017)
       
  • Maximum likelihood principle and moving horizon estimation based adaptive
           unscented Kalman filter
    • Authors: Bingbing Gao; Shesheng Gao; Gaoge Hu; Yongmin Zhong; Chengfan Gu
      Pages: 184 - 196
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Bingbing Gao, Shesheng Gao, Gaoge Hu, Yongmin Zhong, Chengfan Gu
      The classical unscented Kalman filter (UKF) requires prior knowledge on statistical characteristics of system noises for state estimation of a nonlinear dynamic system. If the statistical characteristics of system noises are unknown or inaccurate, the UKF solution will be deteriorated or even divergent. This paper presents a novel adaptive UKF by combining the maximum likelihood principle (MLP) and moving horizon estimation (MHE) to overcome this limitation. This method constructs an optimization based estimation of system noise statistics according to MLP. Subsequently, it further establishes a moving horizon strategy to improve the computational efficiency of the MLP based optimization estimation. Based on above, a new expectation maximization technique is developed to iteratively compute the MLP and MHE based noise statistic estimation by replacing complex smoothed estimates with filtering estimates for further improvement of the computational efficiency. The proposed method can achieve the online estimation of system noise statistic and enhance the robustness of the classical UKF. The efficacy of the proposed adaptive UKF is demonstrated through simulations and practical experiments in the INS/GPS integrated navigation.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.007
      Issue No: Vol. 73 (2017)
       
  • Fault-tolerant adaptive finite-time attitude synchronization and tracking
           control for multi-spacecraft formation
    • Authors: Chengxi Zhang; Jihe Wang; Dexin Zhang; Xiaowei Shao
      Pages: 197 - 209
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Chengxi Zhang, Jihe Wang, Dexin Zhang, Xiaowei Shao
      This paper addresses the attitude synchronization and tracking control (ASTC) for multi-spacecraft formation system (MFS) under undirected and directed graph. First, a new adaptive nonsingular fast terminal sliding mode surface (ANFTSMS) is developed. It has both the merits of the NFTSM avoiding singularity and the adaptive method regulating the relative weighting of parameters. This provides designers a new way to improve the control performance. Second, by applying ANFTSMS, the proposed ANFTSM-controllers (ANFTSMCs) provide high precision finite-time convergence, robust to time-varying disturbances, uncertainties and accommodate to actuator faults, limited inputs. Moreover, the ANFTSMCs also achieve simple structure, inexpensive computations and chattering-free for continuous design. Few studies have addressed these problems simultaneously. Finally, effectiveness of the algorithms are verified via simulations.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.004
      Issue No: Vol. 73 (2017)
       
  • Matching error of the iterative closest contour point algorithm for
           terrain-aided navigation
    • Authors: Kedong Wang; Tongqian Zhu; Yujie Qin; Rui Jiang; Yong Li
      Pages: 210 - 222
      Abstract: Publication date: February 2018
      Source:Aerospace Science and Technology, Volume 73
      Author(s): Kedong Wang, Tongqian Zhu, Yujie Qin, Rui Jiang, Yong Li
      The algorithm used for terrain-aided navigation (TAN) is crucial for high performance of such a totally autonomous and long-duration underwater vehicle navigation technique. This paper presents results of an investigation into the matching errors of the revised iterative closest contour point (ICCP) algorithm for underwater TAN. In particular the quantitative relationship of the matching errors with terrain features is studied in this paper. Among 10 terrain factors, 6 of them have been shown to have the most influence on the accuracy of the revised ICCP algorithm. Three statistical methods, including multiple regression, logistic regression, and discriminant analysis, are applied to mathematically derive the different relationships between the terrain factors and the matching errors. Each formula uses no more than three terrain factors to fit the matching errors. This paper also studies the effect of other factors, including the initial error of the inertial navigation system (INS), the INS accuracy, the map resolution, the vehicle speed, the matching path length, and the sonar accuracy, on the matching errors.

      PubDate: 2017-12-27T04:58:02Z
      DOI: 10.1016/j.ast.2017.12.010
      Issue No: Vol. 73 (2017)
       
 
 
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