Aerospace Science and Technology最新文献

英文中文

Investigation of rapid detonation initiation under injection mixing conditions with downstream low-energy ignition strategies

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-04-01 DOI: 10.1016/j.ast.2025.110178

Xinxin Wang , Xiaodong Cai , Rong Hong , Haorui Liu , Wandong Zhao

This study investigates the deflagration-to-detonation transition (DDT) under injection mixing conditions using high-resolution numerical simulations and a downstream ignition strategy. We analyze the effects of inflow Mach number, fuel equivalence ratio (ER), ignition duration, and ignition location on mixing and detonation initiation. Results show that lower Mach numbers enhance transverse mixing, while higher Mach numbers facilitate faster detonation onset via stronger shear-induced energy deposition. Near-stoichiometric ER reduces initiation distance and boosts detonation velocity, whereas deviations impede flame propagation. Additionally, a new initiation mechanism emerges under low-energy ignition and low internal energy mixtures: a Mach stem forms during flame acceleration, leading to autoignition behind it. Reflected shocks further promote detonation by intensifying interactions with flame fronts. This mechanism demonstrates a pathway to rapid detonation using minimal ignition energy. Furthermore, appropriate ignition duration and location accelerate upstream flame propagation, with suboptimal placement risking poor mixing and detonation failure. These findings provide comprehensive insights into optimizing conditions for efficient detonation initiation under realistic supersonic combustion systems.

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引用次数: 0

Experimental and numerical investigation of a novel strut-mounted roller-screw inerter for helicopter vibration attenuation

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-31 DOI: 10.1016/j.ast.2025.110172

Aykut Tamer , Pierangelo Masarati , Michele Zilletti , Luigi Bottasso

A prominent problem of helicopters is the high vibrational levels due to the high-amplitude excitation forces originating from the main rotor. The ideal solution to reduce vibrations transmitted through the struts is to isolate the fuselage from the main rotor excitation at gearbox struts; therefore the overall vibration attenuation is achieved rather than local solutions. However, the limited available volume around the struts limits the application of existing vibration dampers. To solve the challenge, this work proposes a novel vibration attenuation idea that can effectively perform in confined spaces. Based on the inerter concept of roller-screw type, the axisymmetric design encloses the strut and shares its attachment points, providing a compact solution. The concept is demonstrated through experiments to identify realistic characteristics and rigorous numerical analysis using lumped-parameters and high-fidelity aeroelastic helicopter models to demonstrate vibration mitigation. The results show that the non-linear effects due to friction reduce the effectiveness at low excitation amplitudes; however, satisfactory vibration attenuation levels are achievable at high vibratory loads, a more critical condition for vibration alleviation performance.

{"title":"Experimental and numerical investigation of a novel strut-mounted roller-screw inerter for helicopter vibration attenuation","authors":"Aykut Tamer , Pierangelo Masarati , Michele Zilletti , Luigi Bottasso","doi":"10.1016/j.ast.2025.110172","DOIUrl":"10.1016/j.ast.2025.110172","url":null,"abstract":"<div><div>A prominent problem of helicopters is the high vibrational levels due to the high-amplitude excitation forces originating from the main rotor. The ideal solution to reduce vibrations transmitted through the struts is to isolate the fuselage from the main rotor excitation at gearbox struts; therefore the overall vibration attenuation is achieved rather than local solutions. However, the limited available volume around the struts limits the application of existing vibration dampers. To solve the challenge, this work proposes a novel vibration attenuation idea that can effectively perform in confined spaces. Based on the inerter concept of roller-screw type, the axisymmetric design encloses the strut and shares its attachment points, providing a compact solution. The concept is demonstrated through experiments to identify realistic characteristics and rigorous numerical analysis using lumped-parameters and high-fidelity aeroelastic helicopter models to demonstrate vibration mitigation. The results show that the non-linear effects due to friction reduce the effectiveness at low excitation amplitudes; however, satisfactory vibration attenuation levels are achievable at high vibratory loads, a more critical condition for vibration alleviation performance.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"162 ","pages":"Article 110172"},"PeriodicalIF":5.0,"publicationDate":"2025-03-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143747320","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Flow field analysis of a high agility type aircraft aeroelastic wind tunnel model

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-31 DOI: 10.1016/j.ast.2025.110177

Julius Stegmüller, Andreas Molz, Paloma García-Guillén, Christian Breitsamter

Modern high-agility aircraft are often affected by the consequences of tail buffeting effects at subsonic speeds and medium to high angles of attack. High pressure fluctuations with distinct frequency contents characterize the flow field downstream of vortex breakdown and are often responsible for the dynamic structural response, which can result in heavy structural damage and degraded handling qualities. For analyzing the flow field and the frequency content of the pressure fluctuations over a modular full-span wind tunnel model with either rigid or aeroelastically scaled double-delta wings and horizontal and vertical tailplanes, stereoscopic particle image velocimetry measurements and measurements with a fast-response aerodynamic pressure probe are performed. The aeroelastically scaled components are 3D-printed from polylactide whereas the rigid components are made of aluminum and serve as a reference case. When comparing the rigid and aeroelastically scaled configurations, significant differences in the axial vortex core velocities in some measurement planes can be detected, while the power spectral densities of the flow field pressure fluctuations show similar characteristics with slight differences in the amplitudes.

{"title":"Flow field analysis of a high agility type aircraft aeroelastic wind tunnel model","authors":"Julius Stegmüller, Andreas Molz, Paloma García-Guillén, Christian Breitsamter","doi":"10.1016/j.ast.2025.110177","DOIUrl":"10.1016/j.ast.2025.110177","url":null,"abstract":"<div><div>Modern high-agility aircraft are often affected by the consequences of tail buffeting effects at subsonic speeds and medium to high angles of attack. High pressure fluctuations with distinct frequency contents characterize the flow field downstream of vortex breakdown and are often responsible for the dynamic structural response, which can result in heavy structural damage and degraded handling qualities. For analyzing the flow field and the frequency content of the pressure fluctuations over a modular full-span wind tunnel model with either rigid or aeroelastically scaled double-delta wings and horizontal and vertical tailplanes, stereoscopic particle image velocimetry measurements and measurements with a fast-response aerodynamic pressure probe are performed. The aeroelastically scaled components are 3D-printed from polylactide whereas the rigid components are made of aluminum and serve as a reference case. When comparing the rigid and aeroelastically scaled configurations, significant differences in the axial vortex core velocities in some measurement planes can be detected, while the power spectral densities of the flow field pressure fluctuations show similar characteristics with slight differences in the amplitudes.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"162 ","pages":"Article 110177"},"PeriodicalIF":5.0,"publicationDate":"2025-03-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143747321","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"OA","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Large-Eddy Simulations of kerosene spray combustion in a supersonic jet flow

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-27 DOI: 10.1016/j.ast.2025.110164

Florian Kissel, Guillaume Ribert, Pascale Domingo

High-speed reactive two-phase flows, relevant for the development of future scramjet engines, are studied by Large-Eddy Simulations. The present configuration is inspired by the supersonic Cheng's burner replacing hydrogen with kerosene in the gaseous or liquid phase. The sonic kerosene injection is surrounded by a supersonic co-flow of hot vitiated air, ensuring a jet flame's stabilization. The global equivalence of the burner is set to 0.5. A 5-species global mechanism and a more detailed mechanism from the literature involving 19 species and 54 reactions are compared with a new optimized chemistry involving 18 species and 29 reactions. The impact of ignition delay and droplet size on flame stabilization dynamics and combustion modes is discussed. The necessity of a sub-grid closure for the source terms is assessed. Whatever the kerosene phase, the flame structure is dominated by the non-premixed combustion regime, even if a non-negligible contribution of the rich premixed regime is detected.

引用次数: 0

Research on error compensation methods of wing deflection reconstruction based on FBG strain sensors

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-26 DOI: 10.1016/j.ast.2025.110167

Yiheng Yuan, Wang Sun, Yingjie Zhang, Pinghui Wang, Song Han

Fiber Bragg Grating (FBG) sensors have the characteristics of compact size, multiplexing capability, and high sensitivity, making them ideal for structural health monitoring of unmanned aerial vehicles (UAVs). However, FBG strain Sensors measure wing strain rather than wing deflection, necessitating the conversion of strain into deflection through a deflection reconstruction algorithm. The deflection reconstruction process is influenced by the number and placement of sensors, as well as the reconstruction algorithm, which may result in measurement errors.

To reduce reconstruction errors in wing deflection, this study proposes two approaches: the optimization of sensor placement and the application of an error compensation function.

Firstly, a method for optimizing the location and number of FBG sensors based on a genetic algorithm is proposed. By optimizing the sensor installation location and number to be more consistent with the actual strain distribution characteristics of the wing, the error of wing deflection measurement based on deflection reconstruction algorithm is reduced.

Additionally, a novel error compensation function based on Fourier Series Fitting is developed. Static loading test validated that this approach achieves high accuracy and robustness under varying load conditions, making it suitable for improving the precision of wing deflection reconstruction.

Finally, the wing deflection reconstruction system was implemented on a vertical take-off and landing (VTOL) UAV with a wingspan of 4 meters, achieving real-time measurement of wing deflection and twist angle during flight, verifying the effectiveness of both the reconstruction system and the error compensation function through flight testing.

{"title":"Research on error compensation methods of wing deflection reconstruction based on FBG strain sensors","authors":"Yiheng Yuan, Wang Sun, Yingjie Zhang, Pinghui Wang, Song Han","doi":"10.1016/j.ast.2025.110167","DOIUrl":"10.1016/j.ast.2025.110167","url":null,"abstract":"<div><div>Fiber Bragg Grating (FBG) sensors have the characteristics of compact size, multiplexing capability, and high sensitivity, making them ideal for structural health monitoring of unmanned aerial vehicles (UAVs). However, FBG strain Sensors measure wing strain rather than wing deflection, necessitating the conversion of strain into deflection through a deflection reconstruction algorithm. The deflection reconstruction process is influenced by the number and placement of sensors, as well as the reconstruction algorithm, which may result in measurement errors.</div><div>To reduce reconstruction errors in wing deflection, this study proposes two approaches: the optimization of sensor placement and the application of an error compensation function.</div><div>Firstly, a method for optimizing the location and number of FBG sensors based on a genetic algorithm is proposed. By optimizing the sensor installation location and number to be more consistent with the actual strain distribution characteristics of the wing, the error of wing deflection measurement based on deflection reconstruction algorithm is reduced.</div><div>Additionally, a novel error compensation function based on Fourier Series Fitting is developed. Static loading test validated that this approach achieves high accuracy and robustness under varying load conditions, making it suitable for improving the precision of wing deflection reconstruction.</div><div>Finally, the wing deflection reconstruction system was implemented on a vertical take-off and landing (VTOL) UAV with a wingspan of 4 meters, achieving real-time measurement of wing deflection and twist angle during flight, verifying the effectiveness of both the reconstruction system and the error compensation function through flight testing.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"161 ","pages":"Article 110167"},"PeriodicalIF":5.0,"publicationDate":"2025-03-26","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143746875","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Thermo-chemical non-equilibrium flows and aerodynamic loads of Type III shock/shock interaction on blunted lip of a Mach 12 inlet

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-26 DOI: 10.1016/j.ast.2025.110168

Dengke Li , Bo Sun , Chunliang Dai , Xiong Chen , Yanjin Man

The aerothermal challenges are often accentuated by shock/shock interactions (SSI) in the hypervelocity flow. Due to the small scale, the blunted inlet lip is likely to encounter an extremely harsh thermal environment among the components of an airbreathing vehicle. In the present study, an open-source solver named hy2Foam is used to obtain the unexplored flow characteristics of small-scale Type IIIa SSI at Mach 12 with incident shock angle

β

12^{\circ}

15^{\circ}

and

18^{\circ}

. The transition from Type III to Type IV SSI is observed with the increase of dimensionless intercept

I_{r}

. Different from the results in the previous study, the shear layer in Type IIIa SSI doesn't directly attach but deflects twice before the final attachment. This phenomenon stems from the coupling between the shear layer attachment process and the shock wave/shear layer interaction. As

I_{r}

increases, the Mach reflection (MR) within Type IIIa SSI can transform to regular reflection (RR) owing to the decrease of the first deflection angle of shear layer, and the flows remain steady during the transition process from Type IIIa SSI to Type IV SSI. The flow within a valley region between the shear layer and the cylinder surface is the closest to thermal equilibrium and chemical equilibrium in the whole flow field. Due to the strong recompression shock generated by the final attachment of the shear layer, the peak of pressure and heat transfer at the critical state transitioning from Type IIIa to Type IV SSI is the highest for Type IIIa SSI, achieving 30.8 and 40.08 times to the stagnation values in undisturbed flow, respectively. Finally, a new power-law correlation between pressure and heat transfer intensifications for Type III SSI on the blunted lip of a Mach 12 inlet is obtained. Those insights in this research are poised to provide a reference for the thermal protection design of high Mach number inlets.

{"title":"Thermo-chemical non-equilibrium flows and aerodynamic loads of Type III shock/shock interaction on blunted lip of a Mach 12 inlet","authors":"Dengke Li , Bo Sun , Chunliang Dai , Xiong Chen , Yanjin Man","doi":"10.1016/j.ast.2025.110168","DOIUrl":"10.1016/j.ast.2025.110168","url":null,"abstract":"<div><div>The aerothermal challenges are often accentuated by shock/shock interactions (SSI) in the hypervelocity flow. Due to the small scale, the blunted inlet lip is likely to encounter an extremely harsh thermal environment among the components of an airbreathing vehicle. In the present study, an open-source solver named hy2Foam is used to obtain the unexplored flow characteristics of small-scale Type IIIa SSI at Mach 12 with incident shock angle <span><math><mi>β</mi></math></span> of <span><math><msup><mn>12</mn><mo>∘</mo></msup></math></span>, <span><math><msup><mn>15</mn><mo>∘</mo></msup></math></span> and <span><math><msup><mn>18</mn><mo>∘</mo></msup></math></span>. The transition from Type III to Type IV SSI is observed with the increase of dimensionless intercept <span><math><msub><mi>I</mi><mi>r</mi></msub></math></span>. Different from the results in the previous study, the shear layer in Type IIIa SSI doesn't directly attach but deflects twice before the final attachment. This phenomenon stems from the coupling between the shear layer attachment process and the shock wave/shear layer interaction. As <span><math><msub><mi>I</mi><mi>r</mi></msub></math></span> increases, the Mach reflection (MR) within Type IIIa SSI can transform to regular reflection (RR) owing to the decrease of the first deflection angle of shear layer, and the flows remain steady during the transition process from Type IIIa SSI to Type IV SSI. The flow within a valley region between the shear layer and the cylinder surface is the closest to thermal equilibrium and chemical equilibrium in the whole flow field. Due to the strong recompression shock generated by the final attachment of the shear layer, the peak of pressure and heat transfer at the critical state transitioning from Type IIIa to Type IV SSI is the highest for Type IIIa SSI, achieving 30.8 and 40.08 times to the stagnation values in undisturbed flow, respectively. Finally, a new power-law correlation between pressure and heat transfer intensifications for Type III SSI on the blunted lip of a Mach 12 inlet is obtained. Those insights in this research are poised to provide a reference for the thermal protection design of high Mach number inlets.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"161 ","pages":"Article 110168"},"PeriodicalIF":5.0,"publicationDate":"2025-03-26","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143746876","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Supersonic combustion flow field reconstruction in a scramjet based on deep learning method

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-25 DOI: 10.1016/j.ast.2025.110169

Shicai Huang , Ye Tian , Xue Deng , Maotao Yang , Erda Chen , Hua Zhang

Efficiently predicting combustion flow in scramjet engines enhances early state awareness, which is crucial for facilitating active flow control and ensuring reliable and stable engine operation. Aiming at the problem that the traditional optical test method of ground wind tunnel test is limited by complex and narrow space and difficult to obtain data, the study introduces a supersonic flow field reconstruction model that utilizes local-global feature grouping and fusion to realize fast reconstruction of flow field schlieren image based on sparse pressure data obtained by tiny wall pressure sensor. To prevent the loss of many shallow gradients in the transmission process, a two-layer gradient calculation strategy is designed to preserve the shallow gradient features to the maximum extent. To reduce the number of model parameters and enhance the information crafty interaction between different flow fields, a binary segmentation mask strategy is designed to transform the image dimensionality and block. To assess the effectiveness of the proposed model, experiments were conducted using hydrogen fuel combustion test data obtained from a pulse combustion wind tunnel with an inflow Mach number of 2.5. When compared to other models, our model demonstrated a significant 14.37 % improvement in structural similarity and an 8.35 % improvement in peak signal-to-noise ratio indicators. Most notably, these improvements were achieved while maintaining the lowest computational complexity.

{"title":"Supersonic combustion flow field reconstruction in a scramjet based on deep learning method","authors":"Shicai Huang , Ye Tian , Xue Deng , Maotao Yang , Erda Chen , Hua Zhang","doi":"10.1016/j.ast.2025.110169","DOIUrl":"10.1016/j.ast.2025.110169","url":null,"abstract":"<div><div>Efficiently predicting combustion flow in scramjet engines enhances early state awareness, which is crucial for facilitating active flow control and ensuring reliable and stable engine operation. Aiming at the problem that the traditional optical test method of ground wind tunnel test is limited by complex and narrow space and difficult to obtain data, the study introduces a supersonic flow field reconstruction model that utilizes local-global feature grouping and fusion to realize fast reconstruction of flow field schlieren image based on sparse pressure data obtained by tiny wall pressure sensor. To prevent the loss of many shallow gradients in the transmission process, a two-layer gradient calculation strategy is designed to preserve the shallow gradient features to the maximum extent. To reduce the number of model parameters and enhance the information crafty interaction between different flow fields, a binary segmentation mask strategy is designed to transform the image dimensionality and block. To assess the effectiveness of the proposed model, experiments were conducted using hydrogen fuel combustion test data obtained from a pulse combustion wind tunnel with an inflow Mach number of 2.5. When compared to other models, our model demonstrated a significant 14.37 % improvement in structural similarity and an 8.35 % improvement in peak signal-to-noise ratio indicators. Most notably, these improvements were achieved while maintaining the lowest computational complexity.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"161 ","pages":"Article 110169"},"PeriodicalIF":5.0,"publicationDate":"2025-03-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143725414","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Prescribed-time fault-tolerant tracking control for quadrotor UAV with guaranteed performance

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-24 DOI: 10.1016/j.ast.2025.110162

Miaomiao Tian , Nan Wang , Zhikai Wang , Zhumu Fu , Fazhan Tao

During the actual flight of unmanned aerial vehicle (UAV), there are widespread actuator faults, system uncertainties and external disturbances, which bring significant threats to its steady flight. This paper studies the problem of prescribed-time fault-tolerant tracking control for quadrotor UAV with guaranteed performance. By integrating error reconstruction and dynamic adjustment strategies, this paper ensures that the position and attitude tracking errors converge to the preset performance range within the prescribed time. Furthermore, this paper utilizes the disturbance observer failure compensation method to simultaneously handle the multiplicative and additive failures of the actuator simultaneously, and presents a nonlinear observer-based adaptive strategy to estimate additive faults and unknown external disturbances acting on the UAV. Simulation results show that the position and attitude tracking errors converge to the preset accuracy within the prescribed time without actuator failures. When actuator failures occur, the attitude tracking error still converges on time, and the convergence rate of position error is minimally affected. Thus, the proposed control algorithm for quadrotor UAV is validated.

{"title":"Prescribed-time fault-tolerant tracking control for quadrotor UAV with guaranteed performance","authors":"Miaomiao Tian , Nan Wang , Zhikai Wang , Zhumu Fu , Fazhan Tao","doi":"10.1016/j.ast.2025.110162","DOIUrl":"10.1016/j.ast.2025.110162","url":null,"abstract":"<div><div>During the actual flight of unmanned aerial vehicle (UAV), there are widespread actuator faults, system uncertainties and external disturbances, which bring significant threats to its steady flight. This paper studies the problem of prescribed-time fault-tolerant tracking control for quadrotor UAV with guaranteed performance. By integrating error reconstruction and dynamic adjustment strategies, this paper ensures that the position and attitude tracking errors converge to the preset performance range within the prescribed time. Furthermore, this paper utilizes the disturbance observer failure compensation method to simultaneously handle the multiplicative and additive failures of the actuator simultaneously, and presents a nonlinear observer-based adaptive strategy to estimate additive faults and unknown external disturbances acting on the UAV. Simulation results show that the position and attitude tracking errors converge to the preset accuracy within the prescribed time without actuator failures. When actuator failures occur, the attitude tracking error still converges on time, and the convergence rate of position error is minimally affected. Thus, the proposed control algorithm for quadrotor UAV is validated.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"161 ","pages":"Article 110162"},"PeriodicalIF":5.0,"publicationDate":"2025-03-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143704200","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Fuel stratification effects on thermoacoustic instability for swirl spray flame in an aero-engine centrally-staged combustor

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-24 DOI: 10.1016/j.ast.2025.110163

Ziyu Qin , Heng Song , Xiao Han , Lei Li , Yuzhen Lin , Liangliang Xu

The centrally-staged combustor features a flexible radial fuel stratification to achieve well-organized energy utilization but still faces serious thermoacoustic instability. In this study, we conducted experiments with four stratification ratios (SRs) under the realistic stratified swirl spray flame configuration. The stratification effect on thermoacoustic instability is investigated using the 20 kHz simultaneous PIV and CH₂O-PLIF optical diagnostic. The combustor is prone to severe thermoacoustic oscillation at lower SR, in which the pressure fluctuation substantially follows the heat release dynamic with a short time delay. At higher SR, the relative phase between them becomes irregular accompanied by intermittent oscillations. Two types of flame dynamics are found. The heat release region features an axial mode for lower SR, while it presents the transverse oscillation which destroys the thermoacoustic coupling for higher SR. Further, a coupling metric quantifying the similarity between the vortex probability distribution and flame surface density is proposed to measure the vortex-flame coupling. The highest value for the lowest SR is attributed to the strong flame-flow interaction downstream of the primary recirculation zone, which contributes to the unsteady heat release. Finally, it is found that the local flame stabilization is similar to the counterflow flame.

{"title":"Fuel stratification effects on thermoacoustic instability for swirl spray flame in an aero-engine centrally-staged combustor","authors":"Ziyu Qin , Heng Song , Xiao Han , Lei Li , Yuzhen Lin , Liangliang Xu","doi":"10.1016/j.ast.2025.110163","DOIUrl":"10.1016/j.ast.2025.110163","url":null,"abstract":"<div><div>The centrally-staged combustor features a flexible radial fuel stratification to achieve well-organized energy utilization but still faces serious thermoacoustic instability. In this study, we conducted experiments with four stratification ratios (SRs) under the realistic stratified swirl spray flame configuration. The stratification effect on thermoacoustic instability is investigated using the 20 kHz simultaneous PIV and CH<sub>2</sub>O-PLIF optical diagnostic. The combustor is prone to severe thermoacoustic oscillation at lower SR, in which the pressure fluctuation substantially follows the heat release dynamic with a short time delay. At higher SR, the relative phase between them becomes irregular accompanied by intermittent oscillations. Two types of flame dynamics are found. The heat release region features an axial mode for lower SR, while it presents the transverse oscillation which destroys the thermoacoustic coupling for higher SR. Further, a coupling metric quantifying the similarity between the vortex probability distribution and flame surface density is proposed to measure the vortex-flame coupling. The highest value for the lowest SR is attributed to the strong flame-flow interaction downstream of the primary recirculation zone, which contributes to the unsteady heat release. Finally, it is found that the local flame stabilization is similar to the counterflow flame.</div></div>","PeriodicalId":50955,"journal":{"name":"Aerospace Science and Technology","volume":"161 ","pages":"Article 110163"},"PeriodicalIF":5.0,"publicationDate":"2025-03-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"143746874","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":1,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}

引用次数: 0

Composite anti-unwinding sliding mode attitude tracking control for rigid spacecraft: A prescribed-time method

IF 5 1区工程技术 Q1 ENGINEERING, AEROSPACE

Aerospace Science and Technology

Pub Date : 2025-03-24 DOI: 10.1016/j.ast.2025.110158

Haibin Sun , Yi Xu , Linlin Hou , Dong Yang , Ticao Jiao

This study addresses the prescribed-time anti-unwinding attitude tracking control problem for a rigid spacecraft. First, a prescribed-time sliding mode disturbance observer is designed to estimate the disturbances. To ensure prescribed-time convergence and avoid the unwinding phenomenon, a new prescribed-time sliding mode surface is established by introducing a hyperbolic sine function such that the designed sliding surface includes two equilibrium points. Subsequently, an attitude tracking control law with a disturbance compensation term is designed to guarantee that the system states converge to the sliding surface in a prescribed-time. The simulation results reveal that anti-unwinding prescribed-time attitude tracking of a rigid spacecraft can be achieved by using the developed attitude control law.

引用次数: 0

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