Flight Control

A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Aeronautics".

Deadline for manuscript submissions: closed (31 December 2023) | Viewed by 18882

Special Issue Editors


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Guest Editor
Division of Mechanics, Institute of Aeronautics and Applied Mechanics, Faculty of Power and Aeronautical Engineering, Warsaw University of Technology, 00-665 Warsaw, Poland
Interests: flight dynamics; aircraft system identification; optimization methods; modeling and simulation in MATLAB environment
Special Issues, Collections and Topics in MDPI journals

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Guest Editor
Department of Cryogenics and Aeronautical Engineering, Wrocław University of Science and Technology, 50-370 Wroclaw, Poland
Interests: flight mechanics; fluid mechanics; experiments

Special Issue Information

Dear Colleagues, 

Flight control systems play a crucial role in modern aircraft development as they allow us to perform intended tasks and can enhance aircraft capabilities. Their design requires a multidisciplinary approach that incorporates the mechanics of flight modeling, control theory, mathematical optimization, hydraulic and electrical systems analysis, aeronautical regulations, pilot presence, and many others. Integrating those components is a very challenging and time-consuming task. Flight control systems design is thus an evolving area undergoing constant development and innovative changes. This Special Issue aims to present the latest advances in flight control design, which includes (but is not limited to) the following areas: 

  • Adaptative control;
  • Autonomous systems;
  • Guidance, navigation, and control;
  • Neural networks and machine learning;
  • Flight dynamics;
  • Flight testing;
  • Pilot modelling and human–aircraft interaction;
  • Reconfigurable and fault-tolerant control;
  • Risk and safety management;
  • System identification. 

Dr. Piotr Lichota
Dr. Katarzyna Strzelecka
Guest Editors

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Published Papers (13 papers)

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Research

14 pages, 4727 KiB  
Article
Derivation and Flight Test Validation of Maximum Rate of Climb during Takeoff for Fixed-Wing UAV Driven by Propeller Engine
by Katsumi Watanabe, Takuma Shibata and Masazumi Ueba
Aerospace 2024, 11(3), 233; https://doi.org/10.3390/aerospace11030233 - 15 Mar 2024
Viewed by 735
Abstract
In recent years, the use of fixed-wing Unmanned Aerial Vehicles (UAVs) has expanded, and the use of fixed-wing UAVs is expected to expand due to their usefulness for long-range operations. Different from manned aircraft, no provision is required regarding climb angle at takeoff [...] Read more.
In recent years, the use of fixed-wing Unmanned Aerial Vehicles (UAVs) has expanded, and the use of fixed-wing UAVs is expected to expand due to their usefulness for long-range operations. Different from manned aircraft, no provision is required regarding climb angle at takeoff for fixed-wing UAVs. Therefore, fixed-wing UAVs can take off by taking advantage of their performance. In addition, propeller engines are the propulsion device currently used by most fixed-wing UAVs. However, the thrust force generated by a propeller engine decreases as its airspeed increases. In such circumstances, this paper describes how to derive a maximum rate of climb in which the characteristics of the propeller engine are taken into account, with the aim of reducing takeoff time by maximizing the rate of climb during takeoff. The derivation uses optimization problems with a dependency of the thrust force on the airspeed. After the derivation of the maximum rate of climb, we first checked whether the maximum rate of climb obtained for the mass system was feasible for takeoff at the rate of climb by using a 6-DOF flight simulation, and then confirmed its validity through flight experiments. Full article
(This article belongs to the Special Issue Flight Control)
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21 pages, 7402 KiB  
Article
A Decentralized Voting and Monitoring Flight Control Actuation System for eVTOL Aircraft
by Ruichen He, Florian Holzapfel, Johannes Bröcker, Yi Lai and Shuguang Zhang
Aerospace 2024, 11(3), 195; https://doi.org/10.3390/aerospace11030195 - 29 Feb 2024
Viewed by 921
Abstract
The emergence of eVTOL (electrical Vertical Takeoff and Landing) aircraft necessitates the development of safe and efficient systems to meet stringent certification and operational requirements. The primary state-of-the-art technology for flight control actuation in eVTOL aircraft is electro-mechanical actuators (EMAs), which heavily rely [...] Read more.
The emergence of eVTOL (electrical Vertical Takeoff and Landing) aircraft necessitates the development of safe and efficient systems to meet stringent certification and operational requirements. The primary state-of-the-art technology for flight control actuation in eVTOL aircraft is electro-mechanical actuators (EMAs), which heavily rely on multiple redundancies of critical components to achieve fault tolerance. However, challenges persist in terms of insufficient reliability, immaturity, and a lack of a measurable evaluation method. This research addresses these issues by elucidating the design requirements for EMAs in eVTOL aircraft and proposing a systematic design and evaluation approach for EMA architecture. A key enhancement involves the incorporation of decentralized voting and monitoring (VoDeMo) mechanisms within the Electronic Control Units (ECUs) to improve the overall safety of the EMA. The paper introduces an innovative triple-dual redundant architecture for aircraft control effectors, comprising three dissimilar lanes of ECUs and two similar redundant parallel channels of power electronics and motors. The design is synergistically supported by a comprehensive evaluation that incorporates quantifiable Model-Based Safety Assessment (MBSA), utilizing both physical simulation and logical safety models. Hardware-In-the-Loop (HIL) tests are conducted on a constructed prototype to validate the proposed architecture. Full article
(This article belongs to the Special Issue Flight Control)
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15 pages, 6259 KiB  
Article
Study of Paired Approach Wake Separation Based on Crosswinds
by Weijun Pan, Yanqiang Jiang, Junjie Zhou, Wei Ye and Yuqin Zhang
Aerospace 2024, 11(2), 146; https://doi.org/10.3390/aerospace11020146 - 09 Feb 2024
Viewed by 847
Abstract
The effect of crosswinds on paired approach (PA) procedures for Closely Spaced Parallel Runways (CSPR) is investigated in this paper by fully utilizing the crosswind environment to implement a more efficient PA and increase runway capacity. An improved wake dissipation model is used [...] Read more.
The effect of crosswinds on paired approach (PA) procedures for Closely Spaced Parallel Runways (CSPR) is investigated in this paper by fully utilizing the crosswind environment to implement a more efficient PA and increase runway capacity. An improved wake dissipation model is used to quickly predict the change in the wake velocity field for the PA procedures. The change in the width of the hazard zone is explored in detail using the roll moment coefficient as a determination index. The calculation method for the hazard zone of a wake encounter in a PA is designed considering the influence of crosswind, turbulence, and ground effect. The results show the diffusion rate of the hazard zone and a decrease in the width of the maximum hazard zone under a breezeless environment with increases in the turbulence intensity. The maximum hazard zone width decreases with an increase in crosswind speed. Favorable crosswinds can reduce wake separation and improve the efficiency of a PA. Lower turbulence intensity has a better crosswind effect under a normal PA. The 3-degree offset PA can accommodate larger unfavorable crosswinds, with a higher turbulence intensity having a better crosswind effect. The 3-degree offset PA can substantially increase the proportion of time when no wake affects the PA procedure. Full article
(This article belongs to the Special Issue Flight Control)
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25 pages, 9658 KiB  
Article
Impact-Angle Constraint Guidance and Control Strategies Based on Deep Reinforcement Learning
by Junfang Fan, Denghui Dou and Yi Ji
Aerospace 2023, 10(11), 954; https://doi.org/10.3390/aerospace10110954 - 13 Nov 2023
Viewed by 925
Abstract
In this study, two different impact-angle-constrained guidance and control strategies using deep reinforcement learning (DRL) are proposed. The proposed strategies are based on the dual-loop and integrated guidance and control types. To address comprehensive flying object dynamics and the control mechanism, a Markov [...] Read more.
In this study, two different impact-angle-constrained guidance and control strategies using deep reinforcement learning (DRL) are proposed. The proposed strategies are based on the dual-loop and integrated guidance and control types. To address comprehensive flying object dynamics and the control mechanism, a Markov decision process is used to solve the guidance and control problem, and a real-time impact-angle error in the state vector is used to improve the model applicability. In addition, a reasonable reward mechanism is designed based on the state component which reduces both the miss distance and the impact-angle error and solves the problem of sparse rewards in DRL. Further, to overcome the negative effects of unbounded distributions on bounded action spaces, a Beta distribution is used instead of a Gaussian distribution in the proximal policy optimization algorithm for policy sampling. The state initialization is then realized using a sampling method adjusted to engineering backgrounds, and the control strategy is adapted to a wide range of operational scenarios with different impact angles. Simulation and Monte Carlo experiments in various scenarios show that, compared with other methods mentioned in the experiment in this paper, the proposed DRL strategy has smaller impact-angle errors and miss distance, which demonstrates the method’s effectiveness, applicability, and robustness. Full article
(This article belongs to the Special Issue Flight Control)
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33 pages, 14464 KiB  
Article
An Extension Algorithm of Regional Eigenvalue Assignment Controller Design for Nonlinear Systems
by Ahmet Çağrı Arıcan, Engin Hasan Çopur, Gokhan Inalhan and Metin Uymaz Salamci
Aerospace 2023, 10(10), 893; https://doi.org/10.3390/aerospace10100893 - 19 Oct 2023
Viewed by 1169
Abstract
This paper provides a new method to nonlinear control theory, which is developed from the eigenvalue assignment method. The main purpose of this method is to locate the pointwise eigenvalues of the linear-like structure built by freezing the nonlinear systems at a given [...] Read more.
This paper provides a new method to nonlinear control theory, which is developed from the eigenvalue assignment method. The main purpose of this method is to locate the pointwise eigenvalues of the linear-like structure built by freezing the nonlinear systems at a given time instant in a desired disk region. Since the control requirements for the transient response characteristics are the major constraints on the selection of the disk centre and radius, two different update algorithms are also developed to reshape the disk region by changing the disk centre and radius at each time step. The effectiveness of the proposed methods is tested in both simulations and experiments. A validated three-DOF laboratory helicopter is used for experiments. Full article
(This article belongs to the Special Issue Flight Control)
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20 pages, 2418 KiB  
Article
Cooperative Guidance Law for High-Speed and High-Maneuverability Air Targets
by Fırat Yılmaz Cevher and Mehmet Kemal Leblebicioğlu
Aerospace 2023, 10(2), 155; https://doi.org/10.3390/aerospace10020155 - 08 Feb 2023
Cited by 1 | Viewed by 1727
Abstract
In this paper, a novel cooperative and predictive guidance law is proposed to intercept high-speed and high-maneuverability targets with inferior interceptors. The purpose of guidance is cooperatively covering the most-probable locations where the target may be in the future. To fulfill this purpose, [...] Read more.
In this paper, a novel cooperative and predictive guidance law is proposed to intercept high-speed and high-maneuverability targets with inferior interceptors. The purpose of guidance is cooperatively covering the most-probable locations where the target may be in the future. To fulfill this purpose, predicted target states in the form of a probability density function were obtained using limited target information, i.e., noisy position data for one case and maneuverability limits for the second case, at first. Next, the likelihood of the reachable set of interceptors was computed over the predicted target state. Then, to increase the probability of interception in a finite time, the interceptors’ trajectories were adjusted collaboratively depending on the likelihood values. An extensive Monte Carlo study, with practically applicable simulation parameters, was used to demonstrate the effectiveness of the proposed methods against targets in challenging maneuvering modes. Full article
(This article belongs to the Special Issue Flight Control)
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0 pages, 11197 KiB  
Article
Influence of the Projectile Rotation on the Supersonic Fluidic Element
by Yufang Wang and Nannan Wang
Aerospace 2023, 10(1), 35; https://doi.org/10.3390/aerospace10010035 - 31 Dec 2022
Cited by 2 | Viewed by 1147
Abstract
The effects of projectile rotation on the internal and external flow fields of the supersonic fluidic element are numerically studied using sliding grid technique and the RNG k-ε turbulence model. The effects of rotating speed on internal and external flow fields, switching time [...] Read more.
The effects of projectile rotation on the internal and external flow fields of the supersonic fluidic element are numerically studied using sliding grid technique and the RNG k-ε turbulence model. The effects of rotating speed on internal and external flow fields, switching time and output characteristics are studied. The results show that: for the external flow field, there is no obvious change in the flow field structure at low angular velocity; when the angular velocity increases to 20 r/s, the flow field structure becomes obviously asymmetric due to the Coriolis force; the flow field far away from the surface of the projectile body (more than 0.3 m) is much more affected than the flow field near the surface of the projectile body. The influence of projectile rotation on the internal flow field is much weaker than on the external flow field, and the change of internal flow field is not obvious when the rotational speed is less than 20 r/s. The switching time decreases with the increase in angular velocity, and within normal range of the angular velocity, the deviation of switching time from that without rotation is within 5%. The change of thrust distribution is not obvious when the rotational speed is less than 20 r/s. However, when the rotational speed reaches 50 r/s, the thrust of the middle part of the right nozzle increases by about 20 N. Full article
(This article belongs to the Special Issue Flight Control)
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22 pages, 6498 KiB  
Article
Multiple Constraints-Based Adaptive Three-Dimensional Back-Stepping Sliding Mode Guidance Law against a Maneuvering Target
by Qingli Shi, Hua Wang and Hao Cheng
Aerospace 2022, 9(12), 796; https://doi.org/10.3390/aerospace9120796 - 05 Dec 2022
Cited by 2 | Viewed by 1154
Abstract
This paper addresses the issue of a complex three-dimension (3-D) terminal guidance process that is used against maneuvering targets while considering both the terminal impact angle (TIA) and field-of-view (FOV) angle constraints. According to the highly coupled and nonlinear 3-D terminal guidance model, [...] Read more.
This paper addresses the issue of a complex three-dimension (3-D) terminal guidance process that is used against maneuvering targets while considering both the terminal impact angle (TIA) and field-of-view (FOV) angle constraints. According to the highly coupled and nonlinear 3-D terminal guidance model, an adaptive back-stepping sliding-mode guidance law algorithm is proposed in order to guarantee the stability and robustness of the guidance system. Considering the explicit expression of the line-of-sight (LOS) angle in the kinematics and dynamics of the terminal guidance process, the TIA constraint is transformed into an LOS constraint based on their well-known relationship. In view of the challenges in obtaining the motion information of maneuvering targets, an adaptive law design is introduced in order to estimate and compensate for external disturbances caused by the maneuvering of the target and modeling uncertainty. In addition, because the FOV angle represented by the overall leading angle is not a state variable in the sliding-mode guidance system, it is decoupled into two partial leading angles based on a specific transformation relation, so the 3-D terminal guidance control problem is converted into separate tracking system control issues in the pitch and yaw planes. Then, the Lyapunov stability theory is utilized to substantiate the stability of the guidance system, where the Lyapunov functions in both of the subsystems consist of the LOS and partial FOV state error terms. Finally, a series of simulations of various motion states of maneuvering targets under different terminal cases were carried out. It was proved that the terminal guidance design based on the strategies presented above was able to obtain the desired LOS constraints with satisfying the FOV limitation, and the simulation results verified the effectiveness, universality, and significance for practical applications of the proposed guidance design method. Full article
(This article belongs to the Special Issue Flight Control)
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20 pages, 805 KiB  
Article
Inverse Optimal Zero Effort Miss Guidance Based on Disturbance Observer
by Biao Ma, Mou Chen, Yaohua Shen and Mihai Lungu
Aerospace 2022, 9(12), 767; https://doi.org/10.3390/aerospace9120767 - 28 Nov 2022
Cited by 3 | Viewed by 1481
Abstract
To intercept a maneuvering target in a two-dimensional plane, the inverse optimal guidance law based on zero effort miss (ZEM) and disturbance observer (DO) is studied in this paper. Firstly, the relative kinematics equation is simplified to obtain the missile-target ZEM and its [...] Read more.
To intercept a maneuvering target in a two-dimensional plane, the inverse optimal guidance law based on zero effort miss (ZEM) and disturbance observer (DO) is studied in this paper. Firstly, the relative kinematics equation is simplified to obtain the missile-target ZEM and its dynamics. In order to enhance the robustness of the inverse optimal guidance law, the integral of the ZEM is introduced as a new state to form an augmented system with the original system based on the idea of proportional integral (PI) control. Then, the target maneuver acceleration is assumed as the unknown external disturbance of the guidance augmented system, which is estimated by the DO. Based on the estimated value of DO and the backstepping method, the inverse optimal guidance law is designed to reduce the adverse effect of the disturbance on the guidance system. Finally, simulations are designed to verify the effectiveness of the inverse optimal guidance method based on DO. Full article
(This article belongs to the Special Issue Flight Control)
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24 pages, 3890 KiB  
Article
Sliding Mode Backstepping Control for the Ascent Phase of Near-Space Hypersonic Vehicle Based on a Novel Triple Power Reaching Law
by Shutong Huang, Ju Jiang and Ouxun Li
Aerospace 2022, 9(12), 755; https://doi.org/10.3390/aerospace9120755 - 26 Nov 2022
Cited by 1 | Viewed by 1050
Abstract
This paper presents a novel sliding mode backstepping control scheme for the ascent phase of a near-space hypersonic vehicle (NSHV) base on a triple power reaching law (TPRL). A new model transformation is proposed for NSHV with uncertain parameters subject to uncertainties during [...] Read more.
This paper presents a novel sliding mode backstepping control scheme for the ascent phase of a near-space hypersonic vehicle (NSHV) base on a triple power reaching law (TPRL). A new model transformation is proposed for NSHV with uncertain parameters subject to uncertainties during ascent phase. To shorten the reaching time and reduce the chattering in sliding mode scheme, TPRL is proposed. Then, based on TPRL, a sliding mode backstepping control scheme is proposed, which is combined with new adaptive laws to further reduce the adverse impact of uncertainties. Simulation results demonstrate that TPRL is effective, and the proposed controller for the ascent phase of NSHV is robust with respect to uncertainties. Full article
(This article belongs to the Special Issue Flight Control)
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24 pages, 5858 KiB  
Article
Longitudinal Aerodynamic Parameter Identification for Blended-Wing-Body Aircraft Based on a Wind Tunnel Virtual Flight Test
by Lixin Wang, Shang Tai, Ting Yue, Hailiang Liu, Yanling Wang and Chen Bu
Aerospace 2022, 9(11), 689; https://doi.org/10.3390/aerospace9110689 - 04 Nov 2022
Cited by 2 | Viewed by 1670
Abstract
The wind tunnel virtual flight test realizes the dynamic semi-free flight of the model in the wind tunnel through the deflections of the control surface and uses the test data to identify the aerodynamic derivatives. The difference in dynamics between the wind tunnel [...] Read more.
The wind tunnel virtual flight test realizes the dynamic semi-free flight of the model in the wind tunnel through the deflections of the control surface and uses the test data to identify the aerodynamic derivatives. The difference in dynamics between the wind tunnel virtual flight and the free flight leads to discrepancies between the identification and theoretical results. To solve the problems, a step-by-step identification and correction method for aerodynamic derivatives is established based on the difference between the equations of motion of wind tunnel virtual flight and free flight to identify and correct the lift, drag derivatives, pitch moment derivatives, and velocity derivatives, respectively. To establish an aerodynamic parameter identification model, the flight dynamics equation is expressed as a decoupled form of the free flight force and the influence of the test support frame force on the model’s motions through linearization. To ensure the identification accuracy of each aerodynamic derivative, an excitation signal design method based on amplitude–frequency characteristic analysis is proposed. The longitudinal aerodynamic parameter identification results of a blended-wing-body aircraft show that identification results with higher accuracy can be obtained by adopting the proposed identification and correction method. Full article
(This article belongs to the Special Issue Flight Control)
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30 pages, 4004 KiB  
Article
Backstepping- and Sliding Mode-Based Automatic Carrier Landing System with Deck Motion Estimation and Compensation
by Mihai Lungu, Mou Chen and Dana-Aurelia Vîlcică (Dinu)
Aerospace 2022, 9(11), 644; https://doi.org/10.3390/aerospace9110644 - 24 Oct 2022
Cited by 4 | Viewed by 1779
Abstract
This paper addresses the automatic carrier landing problem in the presence of deck motion, carrier airwake disturbance, wind shears, wind gusts, and atmospheric turbulences. By transforming the 6-DOF aircraft model into an affine dynamic with angle of attack controlled by thrust, the equations [...] Read more.
This paper addresses the automatic carrier landing problem in the presence of deck motion, carrier airwake disturbance, wind shears, wind gusts, and atmospheric turbulences. By transforming the 6-DOF aircraft model into an affine dynamic with angle of attack controlled by thrust, the equations associated to the resultant disturbances are deduced; then, a deck motion prediction block (based on a recursive-least squares algorithm) and a tracking differentiator-based deck motion compensation block are designed. After obtaining the aircraft reference trajectory, the backstepping control method is employed to design a novel automatic carrier landing system with three functional parts: a guidance control system, an attitude control system, and an approach power compensation system. The design of the attitude subsystem involves the flight path control, the control of the attitude angles, and the control of the angular rates. To obtain convergence performance for the closed-loop system, the backstepping technique is combined with sliding mode-based command differentiators for the computation of the virtual commands and extended state observers for the estimation of the disturbances. The global stability of the closed-loop architecture is analyzed by using the Lyapunov theory. Finally, simulation results verify the effectiveness of the proposed carrier landing system, the aircraft reference trajectory being accurately tracked. Full article
(This article belongs to the Special Issue Flight Control)
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19 pages, 3053 KiB  
Article
Field-to-View Constrained Integrated Guidance and Control for Hypersonic Homing Missiles Intercepting Supersonic Maneuvering Targets
by Zhibing Li, Quanlin Dong, Xiaoyue Zhang, Huanrui Zhang and Feng Zhang
Aerospace 2022, 9(11), 640; https://doi.org/10.3390/aerospace9110640 - 24 Oct 2022
Cited by 4 | Viewed by 1495
Abstract
An integrated guidance and control (IGC) scheme considering the field-of-view (FOV) constraint is proposed in this paper for hypersonic skid-to-turn (STT) missiles with a strapdown seeker intercepting a high-speed maneuvering target, which is based on the backstepping control (BC), barrier Lyapunov function (BLF), [...] Read more.
An integrated guidance and control (IGC) scheme considering the field-of-view (FOV) constraint is proposed in this paper for hypersonic skid-to-turn (STT) missiles with a strapdown seeker intercepting a high-speed maneuvering target, which is based on the backstepping control (BC), barrier Lyapunov function (BLF), sliding mode control (SMC), dynamic surface control (DSC), and reduced-order extended state observer (ESO). First, a fifth-order strict feedback IGC model considering the rudder delay dynamics is derived, which also considers the drag effect on the axial velocity. Second, the missile guidance control system based on the BC consists of seeker, guidance, angle-of-attack, attitude, and rudder subsystems. The seeker subsystem was designed based on the BLF, and the other four subsystems were designed based on the SMC. The system-lumped disturbances, including unknown target maneuvers, unmodeled parts, perturbations caused by aerodynamic parameter variations, and external disturbances, were estimated and compensated for using the reduced-order ESO. The DSC prevented the “differential explosion” caused by virtual control commands introduced by the BC. Subsequently, the stability of the closed-loop system was strictly proven using the Lyapunov theory, and the boundedness of the FOV angle was strictly derived. Finally, the simulation results demonstrated the effectiveness and robustness of the proposed IGC scheme. Full article
(This article belongs to the Special Issue Flight Control)
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