Pub Date : 2024-02-20DOI: 10.1177/09544100241234375
Bin Zhao, Jie Chen, Ruimin Jiang, Min Zhou
Aiming at the problem of cooperative target estimation for multiple strap-down missiles, this paper is concerned with developing a distributed cooperative estimation fusion algorithm considering the random transmission delay of observation data. As a stepping stone, the random measurement delay system with multiple strap-down missiles is transformed into a delay-free system with correlation noise in finite time intervals using the random observation delay model for strap-down missiles. Then, the distributed fusion technique without feedback is proposed based on the design of the best local filter for each strap-down sampling subsystem; To further enhance the robustness and stability of the cooperative target estimation system, the distributed fusion algorithm with feedback is suggested. Simulation in various cases and comparison studies are conducted to verify the effectiveness and robustness of the proposed scheme.
{"title":"Cooperative target estimation of strap-down missiles considering random measurement delay","authors":"Bin Zhao, Jie Chen, Ruimin Jiang, Min Zhou","doi":"10.1177/09544100241234375","DOIUrl":"https://doi.org/10.1177/09544100241234375","url":null,"abstract":"Aiming at the problem of cooperative target estimation for multiple strap-down missiles, this paper is concerned with developing a distributed cooperative estimation fusion algorithm considering the random transmission delay of observation data. As a stepping stone, the random measurement delay system with multiple strap-down missiles is transformed into a delay-free system with correlation noise in finite time intervals using the random observation delay model for strap-down missiles. Then, the distributed fusion technique without feedback is proposed based on the design of the best local filter for each strap-down sampling subsystem; To further enhance the robustness and stability of the cooperative target estimation system, the distributed fusion algorithm with feedback is suggested. Simulation in various cases and comparison studies are conducted to verify the effectiveness and robustness of the proposed scheme.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":1.1,"publicationDate":"2024-02-20","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"139949860","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-02-20DOI: 10.1177/09544100241233320
Hongwei Ma, Zengzeng Wang, Xiang He, Yafei Zhong
Multistage axial compressor first stage rotor blades have occurred non-synchronous vibration (NSV). An experiment, including fluid and structure measurements, is adopted at the NSV occurred conditions to conduct a detailed investigation about NSV and rotating instability. The blade vibration stress was obtained from the strain gauges. High-frequency response two-hole pressure probes captured the flow characteristics at the inlet and outlet of the first stage rotor. NSV has a complex aerodynamic disturbance source. Wavelet and frequency spectrum analyses of pressure and blade stress results were utilized to determine the relationship between the blade vibration and the flow field. The inlet flow angle has been adjusted through the inlet guide vane (IGV) to reduce the vibration intensity. This paper revealed the aerodynamic origin behind this adjustment of IGV angle operation. The yaw angle and the axial direction March number at the first stage rotor inlet at different inlet guide vane (IGV) angles revealed the aerodynamic effect. Three operation points, including the maximum stress point and near surge point, are analyzed to broaden the cognition boundary of the NSV. The time and space characteristics of the rotating instabilities have been studied with the azimuthal mode analysis method. This research confirmed that the enlarged inlet incidence angle and the tip flow worsened caused the first stage rotor non-synchronous vibration in this multistage axial compressor. These experiments help find a straight relationship between the NSV and rotating instability.
多级轴流压缩机第一级转子叶片发生了非同步振动（NSV）。为了详细研究非同步振动和旋转不稳定性，我们在发生非同步振动的条件下进行了包括流体和结构测量在内的实验。叶片振动应力由应变片获得。高频响应双孔压力探头捕捉了第一级转子入口和出口处的流动特性。NSV 具有复杂的空气动力干扰源。利用压力和叶片应力结果的小波和频谱分析来确定叶片振动与流场之间的关系。通过进气导叶（IGV）调整了进气流角，以降低振动强度。本文揭示了调整 IGV 角度操作背后的空气动力学原理。不同入口导叶（IGV）角度下第一级转子入口的偏航角和轴向马赫数揭示了空气动力学效应。分析了最大应力点和近浪涌点等三个运行点，拓宽了对 NSV 的认知边界。利用方位模态分析方法研究了旋转不稳定性的时间和空间特征。研究证实，在这种多级轴流式压缩机中，入口入射角的增大和顶端流动的恶化导致了第一级转子的非同步振动。这些实验有助于找到非同步振动与旋转不稳定性之间的直接关系。
{"title":"Investigation of inlet flow influenced on the first rotor flow instability at blade non-synchronous vibration state","authors":"Hongwei Ma, Zengzeng Wang, Xiang He, Yafei Zhong","doi":"10.1177/09544100241233320","DOIUrl":"https://doi.org/10.1177/09544100241233320","url":null,"abstract":"Multistage axial compressor first stage rotor blades have occurred non-synchronous vibration (NSV). An experiment, including fluid and structure measurements, is adopted at the NSV occurred conditions to conduct a detailed investigation about NSV and rotating instability. The blade vibration stress was obtained from the strain gauges. High-frequency response two-hole pressure probes captured the flow characteristics at the inlet and outlet of the first stage rotor. NSV has a complex aerodynamic disturbance source. Wavelet and frequency spectrum analyses of pressure and blade stress results were utilized to determine the relationship between the blade vibration and the flow field. The inlet flow angle has been adjusted through the inlet guide vane (IGV) to reduce the vibration intensity. This paper revealed the aerodynamic origin behind this adjustment of IGV angle operation. The yaw angle and the axial direction March number at the first stage rotor inlet at different inlet guide vane (IGV) angles revealed the aerodynamic effect. Three operation points, including the maximum stress point and near surge point, are analyzed to broaden the cognition boundary of the NSV. The time and space characteristics of the rotating instabilities have been studied with the azimuthal mode analysis method. This research confirmed that the enlarged inlet incidence angle and the tip flow worsened caused the first stage rotor non-synchronous vibration in this multistage axial compressor. These experiments help find a straight relationship between the NSV and rotating instability.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":1.1,"publicationDate":"2024-02-20","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"139949717","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2024-02-19DOI: 10.1177/09544100241234374
Saleh Abuhanieh
The capability for solving compressible fluid flows in the rotating frame of reference is added to an existed open-source CFD solver, namely, HiSA solver. The new implementation is explained and validated using the experimental data of the Sikorsky S-76 rotor. A comparison is presented between the moving mesh results obtained from the original HiSA code and the single rotating frame results achieved through the new implementation. The comparison includes an analysis of torque and thrust values, as well as computational costs. The results imply that, for evaluating the performance of an isolated rotor or for shape optimization purposes at the transonic regime, the single rotating frame method, like the one introduced in the current work, can provide accurate results within an acceptable computational budget. Furthermore, the results show that, at least 25 revolutions are required for the transient analysis to reach an acceptable steady-state converged solution like the one obtained by the single rotating frame method.
{"title":"Development of a steady-state computational fluid dynamics solver for transonic rotorcraft flows","authors":"Saleh Abuhanieh","doi":"10.1177/09544100241234374","DOIUrl":"https://doi.org/10.1177/09544100241234374","url":null,"abstract":"The capability for solving compressible fluid flows in the rotating frame of reference is added to an existed open-source CFD solver, namely, HiSA solver. The new implementation is explained and validated using the experimental data of the Sikorsky S-76 rotor. A comparison is presented between the moving mesh results obtained from the original HiSA code and the single rotating frame results achieved through the new implementation. The comparison includes an analysis of torque and thrust values, as well as computational costs. The results imply that, for evaluating the performance of an isolated rotor or for shape optimization purposes at the transonic regime, the single rotating frame method, like the one introduced in the current work, can provide accurate results within an acceptable computational budget. Furthermore, the results show that, at least 25 revolutions are required for the transient analysis to reach an acceptable steady-state converged solution like the one obtained by the single rotating frame method.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":1.1,"publicationDate":"2024-02-19","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"139949661","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-11-14DOI: 10.1177/09544100231207558
Zhaolin Chen, XiaoHui Wei, Tianhang Xiao, Ning Qin
A 2-D airfoil shape optimization in transonic low-Reynolds number regime is conducted. A Navier–Stokes flow solver with a transition model (k-ω SST γ-Re θ ) is used to evaluate the fitness function. Single-point and multi-point formulations of the optimization results are compared. In addition, the effects of Mach number and angles of attack on aerodynamic characteristics of the optimized airfoils are investigated under low Reynolds number (Re = 17,000) and high-subsonic-flow ([Formula: see text]) conditions. The results show that the corresponding drag divergence Mach number curves of the conventional airfoil present almost a parallel shifting at the entire Mach number range. By contrast, the unconventional airfoil starts showing a significant drag reduction when Mach number is greater than 0.75. Besides, the maximum lift-to-drag ratio is highly influenced by the Mach number because of the formation, movement, type, and strength of a shock wave. In addition, the distinguishing difference in the conclusion between two airfoils is that the lift fluctuation of the conventional airfoil amplifies with the increase of the Mach number. However, the unconventional airfoil shows an opposite trend.
{"title":"Optimization of transonic low-Reynolds number airfoil based on genetic algorithm","authors":"Zhaolin Chen, XiaoHui Wei, Tianhang Xiao, Ning Qin","doi":"10.1177/09544100231207558","DOIUrl":"https://doi.org/10.1177/09544100231207558","url":null,"abstract":"A 2-D airfoil shape optimization in transonic low-Reynolds number regime is conducted. A Navier–Stokes flow solver with a transition model (k-ω SST γ-Re θ ) is used to evaluate the fitness function. Single-point and multi-point formulations of the optimization results are compared. In addition, the effects of Mach number and angles of attack on aerodynamic characteristics of the optimized airfoils are investigated under low Reynolds number (Re = 17,000) and high-subsonic-flow ([Formula: see text]) conditions. The results show that the corresponding drag divergence Mach number curves of the conventional airfoil present almost a parallel shifting at the entire Mach number range. By contrast, the unconventional airfoil starts showing a significant drag reduction when Mach number is greater than 0.75. Besides, the maximum lift-to-drag ratio is highly influenced by the Mach number because of the formation, movement, type, and strength of a shock wave. In addition, the distinguishing difference in the conclusion between two airfoils is that the lift fluctuation of the conventional airfoil amplifies with the increase of the Mach number. However, the unconventional airfoil shows an opposite trend.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-14","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134957168","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-11-14DOI: 10.1177/09544100231203408
Mohamed Arif Mohamed, Zhen Wei Teo, Tze How New, Bing Feng Ng
The ability to predict the characteristics of engine exhaust is important to determining heat signatures on various components of the aircraft body. This is often obtained through numerical means such as RANS, which however relies closely on the choice of the turbulence model for accurate predictions. In this study, predictions of exhaust plumes from four turbulence models are compared against results from particle image velocimetry of a scaled helicopter engine exhaust. These models include the standard k − ϵ, realizable k − ϵ, shear-stress transport (SST) k − ω, and Durbin’s turbulence models. All four turbulence models managed to capture the general shape of the exhaust when analyzed through the velocity contours at two measurement windows. However, the comparisons of velocity contours fail to describe the shift of the predicted plume from the experiments, which is important for fuselage/tail impingement. To obtain further insights to the shifts, a visual correlation in the form of a confidence ellipse through principal component analysis (PCA) is introduced and plotted for the predicted plumes. All the models’ plume predictions showed quantifiable shifts in their mean-centers when compared to the measurements. In terms of matching the measurements’ statistical maximum variance of the plume distribution at the furthermost plane, it was found that the realizable k − ϵ performed the best among the models. On the other hand, the SST k − ω and Durbin’s model performed the best in predicting bivariate ( x and y coordinates) distribution of the plume at the furthermost plane.
{"title":"On the numerical assessment of engine exhaust plumes with a scaled cowling","authors":"Mohamed Arif Mohamed, Zhen Wei Teo, Tze How New, Bing Feng Ng","doi":"10.1177/09544100231203408","DOIUrl":"https://doi.org/10.1177/09544100231203408","url":null,"abstract":"The ability to predict the characteristics of engine exhaust is important to determining heat signatures on various components of the aircraft body. This is often obtained through numerical means such as RANS, which however relies closely on the choice of the turbulence model for accurate predictions. In this study, predictions of exhaust plumes from four turbulence models are compared against results from particle image velocimetry of a scaled helicopter engine exhaust. These models include the standard k − ϵ, realizable k − ϵ, shear-stress transport (SST) k − ω, and Durbin’s turbulence models. All four turbulence models managed to capture the general shape of the exhaust when analyzed through the velocity contours at two measurement windows. However, the comparisons of velocity contours fail to describe the shift of the predicted plume from the experiments, which is important for fuselage/tail impingement. To obtain further insights to the shifts, a visual correlation in the form of a confidence ellipse through principal component analysis (PCA) is introduced and plotted for the predicted plumes. All the models’ plume predictions showed quantifiable shifts in their mean-centers when compared to the measurements. In terms of matching the measurements’ statistical maximum variance of the plume distribution at the furthermost plane, it was found that the realizable k − ϵ performed the best among the models. On the other hand, the SST k − ω and Durbin’s model performed the best in predicting bivariate ( x and y coordinates) distribution of the plume at the furthermost plane.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-14","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134954555","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-11-10DOI: 10.1177/09544100231205154
Tianhong Jiang, Yaolong Liu, Yao Zheng, Ali Elham
Due to the high power and energy demands of large aircraft and low energy density of batteries, parallel hybrid-electric propulsion (HEP) is an in-between solution to reduce carbon emissions instead of full-electric propulsion. Considering the long in-service time, parallel HEP retrofit of existing large commercial or transport aircraft is favorable for technical risks and economic values. To enhance the reliability of the analysis, a comprehensive parallel HEP retrofit study of a narrow-body airliner is carried out in this paper. At first, efficient and accurate aerodynamic and engine deck surrogate models are built to include multidisciplinary impacts. To capture the off-design performance, a detailed flight mission is modeled based on exact flight dynamic equations, which is solved in a time-stepping manner. Then, the hybridization levels and battery usage strategy are thoroughly investigated. Besides, fuel burn, carbon emissions, and total energy consumption performance is studied for fixed take-off weight and selected flight range scenarios. The results show that the total energy in flight can be reduced by 7.9% and fuel combustion and carbon emissions can be reduced by 13.2% for the 2000 km flight distance case compared to the reference case. The greenhouse gas (GHG) emission reductions are 3059.8 kg CO2, 1220.0 kg H 2 O, 0.1937 kg SO 4 , and 0.0387 kg soot. As our methods have incorporated reliable multidisciplinary data and off-design features, the results and conclusions on the parallel HEP retrofit of large aircraft for carbon emission reductions and sustainable aviation goals are more relevant.
{"title":"Investigating the potential of parallel hybrid-electric retrofit of narrow-body airliner for emission reduction","authors":"Tianhong Jiang, Yaolong Liu, Yao Zheng, Ali Elham","doi":"10.1177/09544100231205154","DOIUrl":"https://doi.org/10.1177/09544100231205154","url":null,"abstract":"Due to the high power and energy demands of large aircraft and low energy density of batteries, parallel hybrid-electric propulsion (HEP) is an in-between solution to reduce carbon emissions instead of full-electric propulsion. Considering the long in-service time, parallel HEP retrofit of existing large commercial or transport aircraft is favorable for technical risks and economic values. To enhance the reliability of the analysis, a comprehensive parallel HEP retrofit study of a narrow-body airliner is carried out in this paper. At first, efficient and accurate aerodynamic and engine deck surrogate models are built to include multidisciplinary impacts. To capture the off-design performance, a detailed flight mission is modeled based on exact flight dynamic equations, which is solved in a time-stepping manner. Then, the hybridization levels and battery usage strategy are thoroughly investigated. Besides, fuel burn, carbon emissions, and total energy consumption performance is studied for fixed take-off weight and selected flight range scenarios. The results show that the total energy in flight can be reduced by 7.9% and fuel combustion and carbon emissions can be reduced by 13.2% for the 2000 km flight distance case compared to the reference case. The greenhouse gas (GHG) emission reductions are 3059.8 kg CO2, 1220.0 kg H 2 O, 0.1937 kg SO 4 , and 0.0387 kg soot. As our methods have incorporated reliable multidisciplinary data and off-design features, the results and conclusions on the parallel HEP retrofit of large aircraft for carbon emission reductions and sustainable aviation goals are more relevant.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-10","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135136520","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-11-06DOI: 10.1177/09544100231212651
Sabah Saib, Yahia Achour, Tarak Ghennam, Khoudir Marouani, Nassim Rizoug
This study aims for endurance improvement and energy saving by addressing, on the one hand, the design optimization (DO) of a hybrid electric source (HES) consisting of a fuel cell (FC) and a battery, intended for a fixed-wing electric UAV (EUAV) and, on the other hand, the power management (PM) of the sources. The DO employed the particle swarm optimization (PSO) algorithm, which determines the FC power and the battery capacity that offer the sources the least weight; the findings resulted in a weight reduction of 8%; this gain saves a total energy of 90.28 kJ that can extend the “endurance stage” by more than 26 min. Three power management strategies (PMS) are examined: frequency separation (FS), which gives good performance in PM and protects the FC from dehydration by smoothing its response; power limitation (PL) which protects the FC from exceeding its maximum power, and fuzzy logic (FL), which allows for more flexible power distribution between the sources and has the main advantage of controlling the battery state of charge (SOC). A fourth strategy was developed by combining the benefits of these three strategies. This hybrid PMS (HPMS) saves 2% of FC energy compared to the FL strategy. FS and PL save, respectively, 41% and 44% of FC energy, but the SOC can drop significantly, especially for longer missions. The experimental findings, achieved using a laboratory test bench, validate the simulated results obtained from MATLAB/Simulink software, confirming the validity and performance of the proposed HPMS.
{"title":"Design optimization and power management of a fuel cell-battery fixed-wing electric UAV","authors":"Sabah Saib, Yahia Achour, Tarak Ghennam, Khoudir Marouani, Nassim Rizoug","doi":"10.1177/09544100231212651","DOIUrl":"https://doi.org/10.1177/09544100231212651","url":null,"abstract":"This study aims for endurance improvement and energy saving by addressing, on the one hand, the design optimization (DO) of a hybrid electric source (HES) consisting of a fuel cell (FC) and a battery, intended for a fixed-wing electric UAV (EUAV) and, on the other hand, the power management (PM) of the sources. The DO employed the particle swarm optimization (PSO) algorithm, which determines the FC power and the battery capacity that offer the sources the least weight; the findings resulted in a weight reduction of 8%; this gain saves a total energy of 90.28 kJ that can extend the “endurance stage” by more than 26 min. Three power management strategies (PMS) are examined: frequency separation (FS), which gives good performance in PM and protects the FC from dehydration by smoothing its response; power limitation (PL) which protects the FC from exceeding its maximum power, and fuzzy logic (FL), which allows for more flexible power distribution between the sources and has the main advantage of controlling the battery state of charge (SOC). A fourth strategy was developed by combining the benefits of these three strategies. This hybrid PMS (HPMS) saves 2% of FC energy compared to the FL strategy. FS and PL save, respectively, 41% and 44% of FC energy, but the SOC can drop significantly, especially for longer missions. The experimental findings, achieved using a laboratory test bench, validate the simulated results obtained from MATLAB/Simulink software, confirming the validity and performance of the proposed HPMS.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-06","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135684835","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-11-06DOI: 10.1177/09544100231212647
Zhenyu Wei, Kai Chen, Zhijiang Shao
Traditional trajectory optimization methods for the parafoil system set fixed-point landing as the objective. However, in recent payload fairing recovery missions, the recovery system comprising two parafoils is collaboratively recovered by a mobile vehicle, posing new challenges to the current trajectory optimization technique. In order to recover two parafoil systems autonomously with an unmanned surface vessel, this paper presents a trajectory optimization framework composed of three following component processes consecutively. Firstly, a feasibility judgment algorithm based on reachable boundary estimation is designed to determine the possibility of recovering two parafoil systems. Secondly, the decoupled-then-simultaneous strategy is proposed to enhance the convergence of solving the collaborative recovery problem. Thirdly, the finite-element collocation approach is utilized to convert the formulated trajectory optimization problems into nonlinear programming (NLP) problems, which are solved by a highly efficient NLP solver. Simulation results show that the proposed trajectory optimization framework can efficiently generate the optimal trajectory for recovering two parafoil systems with a vessel.
{"title":"Trajectory optimization for collaborative recovery of parafoil systems using unmanned vessel","authors":"Zhenyu Wei, Kai Chen, Zhijiang Shao","doi":"10.1177/09544100231212647","DOIUrl":"https://doi.org/10.1177/09544100231212647","url":null,"abstract":"Traditional trajectory optimization methods for the parafoil system set fixed-point landing as the objective. However, in recent payload fairing recovery missions, the recovery system comprising two parafoils is collaboratively recovered by a mobile vehicle, posing new challenges to the current trajectory optimization technique. In order to recover two parafoil systems autonomously with an unmanned surface vessel, this paper presents a trajectory optimization framework composed of three following component processes consecutively. Firstly, a feasibility judgment algorithm based on reachable boundary estimation is designed to determine the possibility of recovering two parafoil systems. Secondly, the decoupled-then-simultaneous strategy is proposed to enhance the convergence of solving the collaborative recovery problem. Thirdly, the finite-element collocation approach is utilized to convert the formulated trajectory optimization problems into nonlinear programming (NLP) problems, which are solved by a highly efficient NLP solver. Simulation results show that the proposed trajectory optimization framework can efficiently generate the optimal trajectory for recovering two parafoil systems with a vessel.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-11-06","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135637304","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-10-31DOI: 10.1177/09544100231206568
Najeeba Murtuzapurwala, Malaikannan G
The Thermal Protection System, or TPS, is the layer that shields spacecraft from the extreme aerodynamic heating that occurs during re-entry and launch. We proposed an optimized design of an Integrated Thermal Protection System (ITPS) in this paper. An ITPS provides thermal protection as well as structural load-bearing capability. Thermo-chemical simulations are carried out with the help of the Fully Implicit Ablation and Thermal response program (FIAT) for better understanding of the thermo-chemical behavior of the system under the extreme conditions of a space vehicle. Similarly, a structural analysis is performed to determine the structural integrity of the proposed ITPS design. Both the thermo-chemical and structural analyses show that the proposed corrugated-core sandwich structure is viable in terms of recession, pyrolysis, and structural integrity.
{"title":"Thermo-chemical and structural analysis of integrated thermal protection system for a space vehicle","authors":"Najeeba Murtuzapurwala, Malaikannan G","doi":"10.1177/09544100231206568","DOIUrl":"https://doi.org/10.1177/09544100231206568","url":null,"abstract":"The Thermal Protection System, or TPS, is the layer that shields spacecraft from the extreme aerodynamic heating that occurs during re-entry and launch. We proposed an optimized design of an Integrated Thermal Protection System (ITPS) in this paper. An ITPS provides thermal protection as well as structural load-bearing capability. Thermo-chemical simulations are carried out with the help of the Fully Implicit Ablation and Thermal response program (FIAT) for better understanding of the thermo-chemical behavior of the system under the extreme conditions of a space vehicle. Similarly, a structural analysis is performed to determine the structural integrity of the proposed ITPS design. Both the thermo-chemical and structural analyses show that the proposed corrugated-core sandwich structure is viable in terms of recession, pyrolysis, and structural integrity.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-10-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135871686","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Pub Date : 2023-10-23DOI: 10.1177/09544100231205671
Kain Dipendrasingh, Srinivas MVV, Arun Kumar R
The present study numerically investigates the blade tip leakage flow (TLF) over a gas turbine blade cascade at low speed ( M ex = 0.1) and transonic ( M ex = 0.92) blade exit Mach numbers, with and without shroud motion. Various blade tip flow phenomena like flow separation and reattachment, flow choking, shock-boundary layer interaction (SBLI), and tip leakage vortex (TLV) are studied, and their influence on the heat transfer coefficient (HTC) distributions over the blade tip and near tip blade suction surface is investigated. The study has found some new zones of heat transfer due to the interaction of horseshoe vortices (HSV) with the blade tip. In addition to the primary TLV, a secondary TLV is formed, which increases near tip blade suction surface heat transfer. The flow transitions to supersonic speed in the aft portion of the blade tip for the transonic case resulting in complex heat transfer distribution due to SBLI, compared to near-uniform distribution for the low speed case. This supersonic flow results in a smaller flow separation zone near the aft portion of the pressure side tip edge for the transonic case compared to the low speed case. The qualitative nature of HTC near the leading edge of the blade tip resembles for low speed and transonic cases. However, near the trailing edge of the blade tip, the qualitative nature of HTC for the transonic case shows a substantial difference due to supersonic flow compared to the low speed case.
{"title":"Blade tip leakage flow and heat transfer characteristics over a gas turbine blade at subsonic and transonic exit conditions","authors":"Kain Dipendrasingh, Srinivas MVV, Arun Kumar R","doi":"10.1177/09544100231205671","DOIUrl":"https://doi.org/10.1177/09544100231205671","url":null,"abstract":"The present study numerically investigates the blade tip leakage flow (TLF) over a gas turbine blade cascade at low speed ( M ex = 0.1) and transonic ( M ex = 0.92) blade exit Mach numbers, with and without shroud motion. Various blade tip flow phenomena like flow separation and reattachment, flow choking, shock-boundary layer interaction (SBLI), and tip leakage vortex (TLV) are studied, and their influence on the heat transfer coefficient (HTC) distributions over the blade tip and near tip blade suction surface is investigated. The study has found some new zones of heat transfer due to the interaction of horseshoe vortices (HSV) with the blade tip. In addition to the primary TLV, a secondary TLV is formed, which increases near tip blade suction surface heat transfer. The flow transitions to supersonic speed in the aft portion of the blade tip for the transonic case resulting in complex heat transfer distribution due to SBLI, compared to near-uniform distribution for the low speed case. This supersonic flow results in a smaller flow separation zone near the aft portion of the pressure side tip edge for the transonic case compared to the low speed case. The qualitative nature of HTC near the leading edge of the blade tip resembles for low speed and transonic cases. However, near the trailing edge of the blade tip, the qualitative nature of HTC for the transonic case shows a substantial difference due to supersonic flow compared to the low speed case.","PeriodicalId":54566,"journal":{"name":"Proceedings of the Institution of Mechanical Engineers Part G-Journal of Aerospace Engineering","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-10-23","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135366606","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
京公网安备 11010802042870号
京ICP备2023020795号-1
扫码关注我们
求助内容：
应助结果提醒方式：