Aircraft Engineering and Aerospace TechnologyTable of Contents for Aircraft Engineering and Aerospace Technology. List of articles from the current issue, including Just Accepted (EarlyCite)https://www.emerald.com/insight/publication/issn/0002-2667/vol/96/iss/2?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestAircraft Engineering and Aerospace TechnologyEmerald Publishing LimitedAircraft Engineering and Aerospace TechnologyAircraft Engineering and Aerospace Technologyhttps://www.emerald.com/insight/proxy/containerImg?link=/resource/publication/journal/e565be5aeb8d061fb07464c7b62c3bbb/urn:emeraldgroup.com:asset:id:binary:aeat.cover.jpghttps://www.emerald.com/insight/publication/issn/0002-2667/vol/96/iss/2?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestFatigue crack growth rate and propagation mechanisms of SiC particle reinforced Al alloy matrix compositeshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0010/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to investigate the fatigue crack propagation behavior of SiC particle-reinforced 2124 Al alloy composites under constant amplitude axial loading at a stress ratio of R = 0.1. For this purpose, it is performed experiments and comparatively analyze the results by producing 5, 10, 15 Vol.% SiCp-reinforced composites and unreinforced 2124 Al alloy billets with powder metallurgy (PM) production technique. With the PM production technique, SiCp-reinforced composite and unreinforced 2124 Al alloy billets were produced at 5%, 10%, 15% volume ratios. After the produced billets were extruded and 5 mm thick plates were formed, tensile and fatigue crack propagation compact tensile (CT) samples were prepared. Optical microscope examinations were carried out to determine the microstructural properties of billet and samples. To determine the SiC particle–matrix interactions due to the composite microstructure, unlike the Al alloy, which affects the crack initiation life and crack propagation rate, detailed scanning electron microscopy (SEM) studies have been carried out. Optical microscope examinations for the determination of the microstructural properties of billet and samples showed that although SiC particles were rarely clustered in the Al alloy matrix, they were generally homogeneously dispersed. Fatigue crack propagation rates were determined experimentally. While the highest crack initiation resistance was achieved at 5% SiC volume ratio, the slowest crack propagation rate in the stable crack propagation region was found in the unreinforced 2124 Al alloy. At volume ratios greater than 5%, the number of crack initiation cycles decreases and the propagation rate increases. As a requirement of damage tolerance design, the fatigue crack propagation rate and fatigue behavior of materials to be used in high-tech vehicles such as aircraft structural parts should be well characterized. Therefore, safer use of these materials in critical structural parts becomes widespread. In this study, besides measuring fatigue crack propagation rates, the mechanisms causing crack acceleration or deceleration were determined by applying detailed SEM examinations.Fatigue crack growth rate and propagation mechanisms of SiC particle reinforced Al alloy matrix composites
Adem Karci, Veysel Erturun, Eşref Çakir, Yakup Çam
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.185-192

This study aims to investigate the fatigue crack propagation behavior of SiC particle-reinforced 2124 Al alloy composites under constant amplitude axial loading at a stress ratio of R = 0.1. For this purpose, it is performed experiments and comparatively analyze the results by producing 5, 10, 15 Vol.% SiCp-reinforced composites and unreinforced 2124 Al alloy billets with powder metallurgy (PM) production technique.

With the PM production technique, SiCp-reinforced composite and unreinforced 2124 Al alloy billets were produced at 5%, 10%, 15% volume ratios. After the produced billets were extruded and 5 mm thick plates were formed, tensile and fatigue crack propagation compact tensile (CT) samples were prepared. Optical microscope examinations were carried out to determine the microstructural properties of billet and samples. To determine the SiC particle–matrix interactions due to the composite microstructure, unlike the Al alloy, which affects the crack initiation life and crack propagation rate, detailed scanning electron microscopy (SEM) studies have been carried out.

Optical microscope examinations for the determination of the microstructural properties of billet and samples showed that although SiC particles were rarely clustered in the Al alloy matrix, they were generally homogeneously dispersed. Fatigue crack propagation rates were determined experimentally. While the highest crack initiation resistance was achieved at 5% SiC volume ratio, the slowest crack propagation rate in the stable crack propagation region was found in the unreinforced 2124 Al alloy. At volume ratios greater than 5%, the number of crack initiation cycles decreases and the propagation rate increases.

As a requirement of damage tolerance design, the fatigue crack propagation rate and fatigue behavior of materials to be used in high-tech vehicles such as aircraft structural parts should be well characterized. Therefore, safer use of these materials in critical structural parts becomes widespread. In this study, besides measuring fatigue crack propagation rates, the mechanisms causing crack acceleration or deceleration were determined by applying detailed SEM examinations.

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Fatigue crack growth rate and propagation mechanisms of SiC particle reinforced Al alloy matrix composites10.1108/AEAT-01-2023-0010Aircraft Engineering and Aerospace Technology2023-12-14© 2023 Emerald Publishing LimitedAdem KarciVeysel ErturunEşref ÇakirYakup ÇamAircraft Engineering and Aerospace Technology9622023-12-1410.1108/AEAT-01-2023-0010https://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0010/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Performance analysis of electrical environmental control and ice protection systems in a commercial transport aircrafthttps://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0019/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis paper aims to assess the impact of electrifying the environmental control system (ECS) and ice protection system (IPS), the primary pneumatic system consumers in a conventional commercial transport aircraft, on aircraft weight, range, and fuel consumption. The case study was carried out on Airbus A321-200 aircraft. Design, modelling and analysis processes were carried out on Pacelab SysArc software. Conventional and electrical ECS and IPS architectures were modelled and analysed considering different temperature profiles. The simulation results have shown that the aircraft model with ±270 VDC ECS and IPS architecture is lighter, has a more extended range and has less relative fuel consumption. In addition, the simulation results showed that the maximum range and relative fuel economy of all three aircraft models increased slightly as the temperature increased. Considering the findings in this paper, it is seen that the electrification of the conventional pneumatic system in aircraft has positive contributions in terms of weight, power consumption and fuel consumption. The positive contributions in terms of weight, power consumption and fuel consumption in aircraft will be direct environmental and economic contributions. Apart from the conventional ECS and IPS of the aircraft, two electrical architectures, 230 VAC and ±270 VDC, were modelled and analysed. To see the effects of the three models created in different temperature profiles, analyses were done for cold day, ISA standard day and hot day temperature profiles.Performance analysis of electrical environmental control and ice protection systems in a commercial transport aircraft
Hamdi Ercan, Cüneyt Öztürk, Mustafa Akın
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.193-204

This paper aims to assess the impact of electrifying the environmental control system (ECS) and ice protection system (IPS), the primary pneumatic system consumers in a conventional commercial transport aircraft, on aircraft weight, range, and fuel consumption.

The case study was carried out on Airbus A321-200 aircraft. Design, modelling and analysis processes were carried out on Pacelab SysArc software. Conventional and electrical ECS and IPS architectures were modelled and analysed considering different temperature profiles.

The simulation results have shown that the aircraft model with ±270 VDC ECS and IPS architecture is lighter, has a more extended range and has less relative fuel consumption. In addition, the simulation results showed that the maximum range and relative fuel economy of all three aircraft models increased slightly as the temperature increased.

Considering the findings in this paper, it is seen that the electrification of the conventional pneumatic system in aircraft has positive contributions in terms of weight, power consumption and fuel consumption.

The positive contributions in terms of weight, power consumption and fuel consumption in aircraft will be direct environmental and economic contributions.

Apart from the conventional ECS and IPS of the aircraft, two electrical architectures, 230 VAC and ±270 VDC, were modelled and analysed. To see the effects of the three models created in different temperature profiles, analyses were done for cold day, ISA standard day and hot day temperature profiles.

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Performance analysis of electrical environmental control and ice protection systems in a commercial transport aircraft10.1108/AEAT-01-2023-0019Aircraft Engineering and Aerospace Technology2023-12-18© 2023 Emerald Publishing LimitedHamdi ErcanCüneyt ÖztürkMustafa AkınAircraft Engineering and Aerospace Technology9622023-12-1810.1108/AEAT-01-2023-0019https://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0019/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Long wave infrared signature of swept back leading edges in aircraft frontal aspecthttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0056/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestIn recent years, increased use of all-aspect infrared (IR)-guided missiles based on the long-wave infrared (LWIR; 8–12 µm) band has lowered the probability of aircraft survival in warfare. The lock-on of these highly sensitive missiles is difficult to break, especially from the front. Aerodynamically heated swept-back leading edges (SBLE), because of their high temperature and large area, serve as a prominent LWIR source for aircraft detection from the front. This study aims to report the influence of sweep-back angle (Λ, based on the Mach number [M∞]) on aerodynamic heating and the LWIR signature of SBLE. The temperature along SBLE is obtained numerically as radiation equilibrium temperature (Tw) by discretizing the SBLE length into “n” number of segments, and for each segment, emission based on Tw is evaluated. IR radiance due to reflected external sources (sky-shine and Earthshine) and radiance due to Tw are collectively used to determine the IR contrast between SBLE and its replaced background in the LWIR band (icont-SBLE,LWIR). The results are obtained for low subsonic turboprop aircraft (Λ = 3°, M∞ = 0.44); high subsonic strategic bombers (Λ = 35°, M∞ = 0.8); fifth-generation stealth aircraft (Λ = 40°, M∞ = 1.6); and aircraft with supercruise/supersonic capability (Λ = 50°, M∞ = 2.5). The aircraft with supersonic capability (Λ = 50°, M∞ = 2.5) reports the maximum LWIR signatures and hence the highest visibility from the front. The results obtained are compared with values at Λ = 0° for all cases, which shows that increasing Λ significantly reduces aerodynamic heating and LWIR signatures. The novelty of this study comes from its report on the influence of Λ on the LWIR signatures of aircraft SBLE in the frontal aspect for the first time.Long wave infrared signature of swept back leading edges in aircraft frontal aspect
Kajal Vinayak, Shripad P. Mahulikar
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.205-214

In recent years, increased use of all-aspect infrared (IR)-guided missiles based on the long-wave infrared (LWIR; 8–12 µm) band has lowered the probability of aircraft survival in warfare. The lock-on of these highly sensitive missiles is difficult to break, especially from the front. Aerodynamically heated swept-back leading edges (SBLE), because of their high temperature and large area, serve as a prominent LWIR source for aircraft detection from the front. This study aims to report the influence of sweep-back angle (Λ, based on the Mach number [M]) on aerodynamic heating and the LWIR signature of SBLE.

The temperature along SBLE is obtained numerically as radiation equilibrium temperature (Tw) by discretizing the SBLE length into “n” number of segments, and for each segment, emission based on Tw is evaluated. IR radiance due to reflected external sources (sky-shine and Earthshine) and radiance due to Tw are collectively used to determine the IR contrast between SBLE and its replaced background in the LWIR band (icont-SBLE,LWIR).

The results are obtained for low subsonic turboprop aircraft (Λ = 3°, M = 0.44); high subsonic strategic bombers (Λ = 35°, M = 0.8); fifth-generation stealth aircraft (Λ = 40°, M = 1.6); and aircraft with supercruise/supersonic capability (Λ = 50°, M = 2.5). The aircraft with supersonic capability (Λ = 50°, M = 2.5) reports the maximum LWIR signatures and hence the highest visibility from the front. The results obtained are compared with values at Λ = 0° for all cases, which shows that increasing Λ significantly reduces aerodynamic heating and LWIR signatures.

The novelty of this study comes from its report on the influence of Λ on the LWIR signatures of aircraft SBLE in the frontal aspect for the first time.

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Long wave infrared signature of swept back leading edges in aircraft frontal aspect10.1108/AEAT-02-2023-0056Aircraft Engineering and Aerospace Technology2024-01-18© 2024 Emerald Publishing LimitedKajal VinayakShripad P. MahulikarAircraft Engineering and Aerospace Technology9622024-01-1810.1108/AEAT-02-2023-0056https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0056/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
The effect of global warming and climate changes on aircraft accidents between 2010-2022https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0081/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestOne of the sectors most affected by the variable weather events caused by climate change and global warming is the aviation sector. Especially in aircraft accidents, weather events increasing with climate change and global warming are effective. The purpose of this study is to determine how much the change in weather conditions caused by global warming and climate changes affect the aircraft in the world between the years 2010 and 2022. In this study, it was investigated which weather events were more effective in aircraft crashes by determining the rates of air events and aircraft crashes in aircraft crashes with a passenger capacity of 12 or more that occurred between 2010 and 2022. It is clearly seen that increasing weather conditions with global warming and climate change increase the effect of weather conditions in aircraft crashes. The difference of this study from other studies is the evaluation of the data of the past 12 years, in which the increasing consequences of global warming and climate change have been felt more. It also reveals the necessity of further research on the effects of weather conditions on aircraft.The effect of global warming and climate changes on aircraft accidents between 2010-2022
Tuncer Akay, Cevahir Tarhan
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.215-222

One of the sectors most affected by the variable weather events caused by climate change and global warming is the aviation sector. Especially in aircraft accidents, weather events increasing with climate change and global warming are effective. The purpose of this study is to determine how much the change in weather conditions caused by global warming and climate changes affect the aircraft in the world between the years 2010 and 2022.

In this study, it was investigated which weather events were more effective in aircraft crashes by determining the rates of air events and aircraft crashes in aircraft crashes with a passenger capacity of 12 or more that occurred between 2010 and 2022.

It is clearly seen that increasing weather conditions with global warming and climate change increase the effect of weather conditions in aircraft crashes.

The difference of this study from other studies is the evaluation of the data of the past 12 years, in which the increasing consequences of global warming and climate change have been felt more. It also reveals the necessity of further research on the effects of weather conditions on aircraft.

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The effect of global warming and climate changes on aircraft accidents between 2010-202210.1108/AEAT-03-2023-0081Aircraft Engineering and Aerospace Technology2023-12-11© 2023 Emerald Publishing LimitedTuncer AkayCevahir TarhanAircraft Engineering and Aerospace Technology9622023-12-1110.1108/AEAT-03-2023-0081https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0081/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Analysis of influencing factors of aircraft fuel tank flammabilityhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0044/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestIt is an indispensable part of airworthiness certification to evaluate the fuel tank flammability exposure time for transport aircraft. There are many factors and complex coupling relationships affecting the fuel tank flammability exposure time. The current work not only lacks a comprehensive analysis of these factors but also lacks the significance of each factor, the interaction relationship and the prediction method of flammability exposure time. The lack of research in these aspects seriously restricts the smooth development of the airworthiness forensics work of domestic large aircraft. This paper aims to clarify the internal relationship between user input parameters and predict the flammability exposure time of fuel tanks for transport aircraft. Based on the requirements of airworthiness certification for large aircraft, an in-depth analysis of the Monte Carlo flammability evaluation source procedures specified in China Civil Aviation Regulation/FAR25 airworthiness regulations was made, the internal relationship between factors affecting the fuel tank flammability exposure time was clarified and the significant effects and interactions of input parameters in the Monte Carlo evaluation model were studied using the response surface method. And the BP artificial neural network training samples with high significance factors were used to establish the prediction model of flammability exposure time. The input parameters in the Monte Carlo program directly or indirectly affect the fuel tank flammability exposure time by means of the influence on the flammability limit or fuel temperature. Among the factors affecting flammability exposure time, the cruising Mach number, balance temperature difference and maximum range are the most significant, and they are all positively correlated with flammability exposure time. Although there are interactions among all factors, the degree of influence on flammability exposure time is not the same. The interaction between maximum range and equilibrium temperature difference is more significant than other factors. The prediction model of flammability exposure time based on multifactor interaction and BP neural network has good accuracy and can be applied to the prediction of fuel tank flammability exposure time. The flammability exposure time prediction model was established based on multifactor interaction and BP neural network. The limited test results were combined with intelligent algorithm to achieve rapid prediction, which saved the test cost and time.Analysis of influencing factors of aircraft fuel tank flammability
Leiming Geng, Ruihua Zhang, Weihua Liu
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.223-232

It is an indispensable part of airworthiness certification to evaluate the fuel tank flammability exposure time for transport aircraft. There are many factors and complex coupling relationships affecting the fuel tank flammability exposure time. The current work not only lacks a comprehensive analysis of these factors but also lacks the significance of each factor, the interaction relationship and the prediction method of flammability exposure time. The lack of research in these aspects seriously restricts the smooth development of the airworthiness forensics work of domestic large aircraft. This paper aims to clarify the internal relationship between user input parameters and predict the flammability exposure time of fuel tanks for transport aircraft.

Based on the requirements of airworthiness certification for large aircraft, an in-depth analysis of the Monte Carlo flammability evaluation source procedures specified in China Civil Aviation Regulation/FAR25 airworthiness regulations was made, the internal relationship between factors affecting the fuel tank flammability exposure time was clarified and the significant effects and interactions of input parameters in the Monte Carlo evaluation model were studied using the response surface method. And the BP artificial neural network training samples with high significance factors were used to establish the prediction model of flammability exposure time.

The input parameters in the Monte Carlo program directly or indirectly affect the fuel tank flammability exposure time by means of the influence on the flammability limit or fuel temperature. Among the factors affecting flammability exposure time, the cruising Mach number, balance temperature difference and maximum range are the most significant, and they are all positively correlated with flammability exposure time. Although there are interactions among all factors, the degree of influence on flammability exposure time is not the same. The interaction between maximum range and equilibrium temperature difference is more significant than other factors. The prediction model of flammability exposure time based on multifactor interaction and BP neural network has good accuracy and can be applied to the prediction of fuel tank flammability exposure time.

The flammability exposure time prediction model was established based on multifactor interaction and BP neural network. The limited test results were combined with intelligent algorithm to achieve rapid prediction, which saved the test cost and time.

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Analysis of influencing factors of aircraft fuel tank flammability10.1108/AEAT-02-2023-0044Aircraft Engineering and Aerospace Technology2023-12-18© 2023 Emerald Publishing LimitedLeiming GengRuihua ZhangWeihua LiuAircraft Engineering and Aerospace Technology9622023-12-1810.1108/AEAT-02-2023-0044https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0044/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Aero-spike and aero-disk effects of on wave drag reduction of supersonic flow past over blunt bodyhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-04-2023-0088/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this study is to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body. Specifically, the study aims to analyze the impact of varying the length of the cylindrical rod in the aerospike (ranging from 0.5 to 2.0 times the diameter of the blunt body) and the diameter of the hemispherical disk (ranging from 0.25 to 0.75 times the blunt body diameter). CFD simulations were conducted at a supersonic Mach number of 2 and a Reynolds number of 2.79 × 106. ICEM CFD and ANSYS CFX solver were used to generate the three-dimensional flow along with its structures. The flow structure and drag coefficient were computed using Reynolds-averaged Navier–Stokes equation model. The drag reduction mechanism was also explained using the idea of dividing streamline and density contour. The performance of the aero spike length and the effect of aero disk size on the drag are investigated. The separating shock is located in front of the blunt body, forming an effective conical shape that reduces the pressure drag acting on the blunt body. It was observed that extending the length of the spike beyond a specific critical point did not impact the flow field characteristics and had no further influence on the enhanced performance. The optimal combination of disk and spike length was determined, resulting in a substantial reduction in drag through the introduction of the aerospike and disk. To predict the accurate results of drag and to reduce the simulation time, a hexa grid with finer mesh structure was adopted in the simulation. The blunt nose structures are primarily employed in the design of rockets, missiles, and re-entry capsules to withstand higher aerodynamic loads and aerodynamic heating. For the optimized size of the aero spike, aero disk is also optimized to use the benefits of both.Aero-spike and aero-disk effects of on wave drag reduction of supersonic flow past over blunt body
Kathiravan Balusamy, Vinothraj A., Suresh V.
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.233-247

The purpose of this study is to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body. Specifically, the study aims to analyze the impact of varying the length of the cylindrical rod in the aerospike (ranging from 0.5 to 2.0 times the diameter of the blunt body) and the diameter of the hemispherical disk (ranging from 0.25 to 0.75 times the blunt body diameter). CFD simulations were conducted at a supersonic Mach number of 2 and a Reynolds number of 2.79 × 106.

ICEM CFD and ANSYS CFX solver were used to generate the three-dimensional flow along with its structures. The flow structure and drag coefficient were computed using Reynolds-averaged Navier–Stokes equation model. The drag reduction mechanism was also explained using the idea of dividing streamline and density contour. The performance of the aero spike length and the effect of aero disk size on the drag are investigated.

The separating shock is located in front of the blunt body, forming an effective conical shape that reduces the pressure drag acting on the blunt body. It was observed that extending the length of the spike beyond a specific critical point did not impact the flow field characteristics and had no further influence on the enhanced performance. The optimal combination of disk and spike length was determined, resulting in a substantial reduction in drag through the introduction of the aerospike and disk.

To predict the accurate results of drag and to reduce the simulation time, a hexa grid with finer mesh structure was adopted in the simulation.

The blunt nose structures are primarily employed in the design of rockets, missiles, and re-entry capsules to withstand higher aerodynamic loads and aerodynamic heating.

For the optimized size of the aero spike, aero disk is also optimized to use the benefits of both.

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Aero-spike and aero-disk effects of on wave drag reduction of supersonic flow past over blunt body10.1108/AEAT-04-2023-0088Aircraft Engineering and Aerospace Technology2024-01-09© 2023 Emerald Publishing LimitedKathiravan BalusamyVinothraj A.Suresh V.Aircraft Engineering and Aerospace Technology9622024-01-0910.1108/AEAT-04-2023-0088https://www.emerald.com/insight/content/doi/10.1108/AEAT-04-2023-0088/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Fuel savings on missed approach through aircraft reinjectionhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2022-0305/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestIn the aerial transportation area, fuel costs are critical to the economic viability of companies, and so urgent measures should be adopted to avoid any unnecessary increase in operational costs. In particular, this paper addresses the case of missed approach manouevres, showing that it is still possible to optimize the usual procedure. The costs involved in a standard procedure following a missed approach are analysed through a simulation model, and they are compared with the improvements achieved with a fast reinjection scheme proposed in a prior work. Experimental results show that, for a standard A320 aircraft, fuel savings ranging from 55% to 90% can be achieved through the reinjection method. To the best of the authors’ knowledge, this work is the first study in the literature addressing the fuel savings benefits obtained by applying a reinjection technique for missed approach manoeuvres.Fuel savings on missed approach through aircraft reinjection
María Carmona, Rafael Casado González, Aurelio Bermúdez, Miguel Pérez-Francisco, Pablo Boronat, Carlos Calafate
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.248-256

In the aerial transportation area, fuel costs are critical to the economic viability of companies, and so urgent measures should be adopted to avoid any unnecessary increase in operational costs. In particular, this paper addresses the case of missed approach manouevres, showing that it is still possible to optimize the usual procedure.

The costs involved in a standard procedure following a missed approach are analysed through a simulation model, and they are compared with the improvements achieved with a fast reinjection scheme proposed in a prior work.

Experimental results show that, for a standard A320 aircraft, fuel savings ranging from 55% to 90% can be achieved through the reinjection method.

To the best of the authors’ knowledge, this work is the first study in the literature addressing the fuel savings benefits obtained by applying a reinjection technique for missed approach manoeuvres.

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Fuel savings on missed approach through aircraft reinjection10.1108/AEAT-11-2022-0305Aircraft Engineering and Aerospace Technology2024-01-22© 2023 María Carmona, Rafael Casado González, Aurelio Bermúdez, Miguel Pérez-Francisco, Pablo Boronat and Carlos Calafate.María CarmonaRafael Casado GonzálezAurelio BermúdezMiguel Pérez-FranciscoPablo BoronatCarlos CalafateAircraft Engineering and Aerospace Technology9622024-01-2210.1108/AEAT-11-2022-0305https://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2022-0305/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 María Carmona, Rafael Casado González, Aurelio Bermúdez, Miguel Pérez-Francisco, Pablo Boronat and Carlos Calafate.http://creativecommons.org/licences/by/4.0/legalcode
Uncertainty quantification of blade geometric deviation on compressor stabilityhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2023-0163/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe geometric parameters of the compressor blade have a noteworthy influence on compressor stability, which should be meticulously designed. However, machining inaccuracies cause the blade geometric parameters to deviate from the ideal design, and the geometric deviation exhibits high randomness. Therefore, the purpose of this study is to quantify the uncertainty and analyze the sensitivity of the impact of blade geometric deviation on compressor stability. In this work, the influence of blade geometric deviation is analyzed based on a subsonic compressor rotor stage, and three-dimensional numerical simulations are used to compute samples with different geometric features. A method of combining Halton sequence and non-intrusive polynomial chaos is adopted to carry out uncertainty quantitative analysis. Sobol’ index and Spearman correlation coefficient are used to analysis the sensitivity and correlation between compressor stability and blade geometric deviation, respectively. The results show that the compressor stability is most sensitive to the tip clearance deviation, whereas deviations in the leading edge radius, trailing edge radius and chord length have minimal impact on the compressor stability. And, the effects of various blade geometric deviations on the compressor stability are basically independent and linearly superimposed. This work provided a new approach for uncertainty quantification in compressor stability analysis. The conclusions obtained in this work provide some reference value for the manufacturing and maintenance of rotor blades.Uncertainty quantification of blade geometric deviation on compressor stability
Tianyuan Ji, Wuli Chu
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.257-264

The geometric parameters of the compressor blade have a noteworthy influence on compressor stability, which should be meticulously designed. However, machining inaccuracies cause the blade geometric parameters to deviate from the ideal design, and the geometric deviation exhibits high randomness. Therefore, the purpose of this study is to quantify the uncertainty and analyze the sensitivity of the impact of blade geometric deviation on compressor stability.

In this work, the influence of blade geometric deviation is analyzed based on a subsonic compressor rotor stage, and three-dimensional numerical simulations are used to compute samples with different geometric features. A method of combining Halton sequence and non-intrusive polynomial chaos is adopted to carry out uncertainty quantitative analysis. Sobol’ index and Spearman correlation coefficient are used to analysis the sensitivity and correlation between compressor stability and blade geometric deviation, respectively.

The results show that the compressor stability is most sensitive to the tip clearance deviation, whereas deviations in the leading edge radius, trailing edge radius and chord length have minimal impact on the compressor stability. And, the effects of various blade geometric deviations on the compressor stability are basically independent and linearly superimposed.

This work provided a new approach for uncertainty quantification in compressor stability analysis. The conclusions obtained in this work provide some reference value for the manufacturing and maintenance of rotor blades.

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Uncertainty quantification of blade geometric deviation on compressor stability10.1108/AEAT-06-2023-0163Aircraft Engineering and Aerospace Technology2023-12-18© 2023 Emerald Publishing LimitedTianyuan JiWuli ChuAircraft Engineering and Aerospace Technology9622023-12-1810.1108/AEAT-06-2023-0163https://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2023-0163/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theoryhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0054/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this study is to establish a thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory. The effects of contact load, inclination angle, fractal dimensional and fractal roughness on thermal contact conductance of rough surfaces were studied using numerical simulation. The results show that the thermal contact conductance of the rough surface increases with the increase of contact load and fractal dimension and decreases with the increase of fractal roughness and inclination angle. The inclination angle of the rough surface has an important influence on the thermal contact conductance of the rough, and it is a factor that cannot be ignored in the study of the thermal contact conductance of rough surfaces. A thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory was established in this study. The achievements of this study provide some theoretical basis for the investigation of the thermal contact conductance of rough surfaces.Thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory
Xianguang Sun
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.265-272

The purpose of this study is to establish a thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory.

The effects of contact load, inclination angle, fractal dimensional and fractal roughness on thermal contact conductance of rough surfaces were studied using numerical simulation.

The results show that the thermal contact conductance of the rough surface increases with the increase of contact load and fractal dimension and decreases with the increase of fractal roughness and inclination angle. The inclination angle of the rough surface has an important influence on the thermal contact conductance of the rough, and it is a factor that cannot be ignored in the study of the thermal contact conductance of rough surfaces.

A thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory was established in this study. The achievements of this study provide some theoretical basis for the investigation of the thermal contact conductance of rough surfaces.

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Thermal contact conductance model of rough surfaces with inclination based on three-dimensional fractal theory10.1108/AEAT-02-2023-0054Aircraft Engineering and Aerospace Technology2024-01-01© 2023 Emerald Publishing LimitedXianguang SunAircraft Engineering and Aerospace Technology9622024-01-0110.1108/AEAT-02-2023-0054https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0054/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Mathematical modelling and PID control system implementation for quadcopter frame Tarot FY650https://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2023-0154/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this study is to find the suitable trajectory path of the Numerical model of the Quadcopter. Quadcopters are widely used in various applications due to their compact size and ease of assembly. Because they are quite unstable, autonomous control systems would be used to overcome this problem. Modelling autonomous control is predominant as the research scope faces challenges because of its highly non-linear, multivariable system with 6 degree of freedom. Quadcopters with antonym systems can operate in an unknown environment by overcoming unexpected disturbances. The first objective when designing such a system is to design an accurate mathematical model to describe the dynamics of the system. Newton’s law of motion was used to build the mathematical model of the system. Establishment of the mathematical model and the physics behind a four propeller drone for the frame TAROT 650 carbon was done. Simulink model was developed based on the mathematical model for simulating the complete dynamics of the drone as well as location and gusts were included to check the stability. The control response of the system was simulated numerically results are discussed. The trajectory path was found. The phases with their own parameters can be used to implement the mathematical model for another type of quadcopter model and achieve quick development.Mathematical modelling and PID control system implementation for quadcopter frame Tarot FY650
Gowtham G., Jagan Raj R.
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.273-284

The purpose of this study is to find the suitable trajectory path of the Numerical model of the Quadcopter. Quadcopters are widely used in various applications due to their compact size and ease of assembly. Because they are quite unstable, autonomous control systems would be used to overcome this problem. Modelling autonomous control is predominant as the research scope faces challenges because of its highly non-linear, multivariable system with 6 degree of freedom.

Quadcopters with antonym systems can operate in an unknown environment by overcoming unexpected disturbances. The first objective when designing such a system is to design an accurate mathematical model to describe the dynamics of the system. Newton’s law of motion was used to build the mathematical model of the system.

Establishment of the mathematical model and the physics behind a four propeller drone for the frame TAROT 650 carbon was done. Simulink model was developed based on the mathematical model for simulating the complete dynamics of the drone as well as location and gusts were included to check the stability.

The control response of the system was simulated numerically results are discussed. The trajectory path was found. The phases with their own parameters can be used to implement the mathematical model for another type of quadcopter model and achieve quick development.

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Mathematical modelling and PID control system implementation for quadcopter frame Tarot FY65010.1108/AEAT-06-2023-0154Aircraft Engineering and Aerospace Technology2024-01-12© 2023 Emerald Publishing LimitedGowtham G.Jagan Raj R.Aircraft Engineering and Aerospace Technology9622024-01-1210.1108/AEAT-06-2023-0154https://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2023-0154/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Design methodology for combustor in advanced gas turbine engines: a reviewhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-12-2022-0351/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestA combustor design is a particularly important and difficult task in the development of gas turbine engines. During studies for accurate and easy combustor design, reasonable design methodologies have been established and used in engine development. The purpose of this paper is to review the design methodology for combustor in development of advanced gas turbine engines. The advanced combustor development task can be successfully achieved in less time and at lower cost by adopting new and superior design methodologies. The review considers the main technical problems (combustion, cooling, fuel injection and ignition technology) in the development of modern combustor design and deals with combustor design methods by dividing it into preliminary design, performance evaluation, optimization and experiment. The advanced combustion and cooling technologies mainly used in combustor design are mentioned in detail. In accordance with the modern combustor design method, the design mechanisms are considered and the methods used in every stage of the design are reviewed technically. The improved performances and strict emission limits of gas turbine engines require the application of advanced technologies when designing combustors. The optimized design mechanism and reasonable performance evaluation methods are very important in reducing experiments and increasing the effectiveness of the design. This paper provides a comprehensive review of the design methodology for the advanced gas turbine engine combustor.Design methodology for combustor in advanced gas turbine engines: a review
Insong Kim, Hakson Jin, Kwangsong Ri, Sunbong Hyon, Cholhui Huang
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.285-296

A combustor design is a particularly important and difficult task in the development of gas turbine engines. During studies for accurate and easy combustor design, reasonable design methodologies have been established and used in engine development. The purpose of this paper is to review the design methodology for combustor in development of advanced gas turbine engines. The advanced combustor development task can be successfully achieved in less time and at lower cost by adopting new and superior design methodologies.

The review considers the main technical problems (combustion, cooling, fuel injection and ignition technology) in the development of modern combustor design and deals with combustor design methods by dividing it into preliminary design, performance evaluation, optimization and experiment. The advanced combustion and cooling technologies mainly used in combustor design are mentioned in detail. In accordance with the modern combustor design method, the design mechanisms are considered and the methods used in every stage of the design are reviewed technically.

The improved performances and strict emission limits of gas turbine engines require the application of advanced technologies when designing combustors. The optimized design mechanism and reasonable performance evaluation methods are very important in reducing experiments and increasing the effectiveness of the design.

This paper provides a comprehensive review of the design methodology for the advanced gas turbine engine combustor.

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Design methodology for combustor in advanced gas turbine engines: a review10.1108/AEAT-12-2022-0351Aircraft Engineering and Aerospace Technology2024-03-01© 2024 Emerald Publishing LimitedInsong KimHakson JinKwangsong RiSunbong HyonCholhui HuangAircraft Engineering and Aerospace Technology9622024-03-0110.1108/AEAT-12-2022-0351https://www.emerald.com/insight/content/doi/10.1108/AEAT-12-2022-0351/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Review on impact, crushing response and applications of re-entrant core sandwich structureshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0122/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestAuxetic sandwich structures are gaining attention because of the negative Poisson’s ratio effect offered by these structures. Re-entrant core was one configuration of the auxetic structures. There is a growing concern about the design and behavior of re-entrant cores in aerospace, marine and protection applications. Several researchers proposed various designs of re-entrant core sandwiches with various materials. The purpose of this study is to review the most recent advances in re-entrant core sandwich structures. This review serves as a guide for researchers conducting further research in this wide field of study. The re-entrant core sandwich structures were reviewed in terms of their design improvements, impact and quasi-static crushing responses. Several design improvements were reviewed including 2D cell, 3D cell, gradient, hierarchical and hybrid configurations. Some common applications of the re-entrant core sandwiches were given at the end of this paper with suggestions for future developments in this field. Generally, the re-entrant configuration showed improved energy absorption and impact response among auxetic structures. The main manufacturing method for re-entrant core manufacturing was additive manufacturing. The negative Poisson’s ratio effect of the re-entrant core provided a wide area of research. Generally, re-entrant cores were mentioned in the review articles as part of other auxetic structures. However, in this review, the focus was solely made on the re-entrant core sandwiches with their mechanics.Review on impact, crushing response and applications of re-entrant core sandwich structures
Mustafa S. Al-Khazraji
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.297-306

Auxetic sandwich structures are gaining attention because of the negative Poisson’s ratio effect offered by these structures. Re-entrant core was one configuration of the auxetic structures. There is a growing concern about the design and behavior of re-entrant cores in aerospace, marine and protection applications. Several researchers proposed various designs of re-entrant core sandwiches with various materials. The purpose of this study is to review the most recent advances in re-entrant core sandwich structures. This review serves as a guide for researchers conducting further research in this wide field of study.

The re-entrant core sandwich structures were reviewed in terms of their design improvements, impact and quasi-static crushing responses. Several design improvements were reviewed including 2D cell, 3D cell, gradient, hierarchical and hybrid configurations. Some common applications of the re-entrant core sandwiches were given at the end of this paper with suggestions for future developments in this field.

Generally, the re-entrant configuration showed improved energy absorption and impact response among auxetic structures. The main manufacturing method for re-entrant core manufacturing was additive manufacturing. The negative Poisson’s ratio effect of the re-entrant core provided a wide area of research.

Generally, re-entrant cores were mentioned in the review articles as part of other auxetic structures. However, in this review, the focus was solely made on the re-entrant core sandwiches with their mechanics.

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Review on impact, crushing response and applications of re-entrant core sandwich structures10.1108/AEAT-05-2023-0122Aircraft Engineering and Aerospace Technology2024-01-26© 2024 Emerald Publishing LimitedMustafa S. Al-KhazrajiAircraft Engineering and Aerospace Technology9622024-01-2610.1108/AEAT-05-2023-0122https://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0122/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Impact time guidance law for arbitrary lead angle using sliding mode controlhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2023-0292/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis paper aims to investigate a novel impact time control guidance (ITCG) law based on the sliding mode control (SMC) for a nonmaneuvering target using the predicted interception point (PIP). To intercept the target with the minimal miss distance and desired impact time, an estimation of time-to-go is introduced. This estimation results in a precise impact time for multimissiles salvo attack the target at the same time. Even for a large lead angle, the desired impact time is achieved by using the sliding mode and Lyapunov stability theory. The singularity issue of the proposed impact time guidance laws is also analyzed to achieve an arbitrary lead angle with the desired impact time. Numerical scenarios with desired impact time are presented to illustrate the performance of the proposed ITCG law. Comparison with the state-of-art impact time guidance laws proves that the guidance law in this paper can enable the missile to intercept the target with minimal miss distance and final impact time error. This method enables multiple missiles to attack the target simultaneously with different distances and arbitrary lead angles. An ITCG law based on sliding mode and Lyapunov stability theory is proposed, and the switching surface is designed based on a novel estimation time-to-go for the missile to intercept the target with minimal miss distance. To intercept the target with initial arbitrary lead angles and desired impact time, the authors analysis the singular issue in SMC to ensure that the missile can intercept the target with arbitrary lead angle. The proposed approach for a nonmaneuvering target using the PIP has simple forms, and therefore, they have the superiority of being implemented easily.Impact time guidance law for arbitrary lead angle using sliding mode control
He Du, Ming Yang, Songyan Wang, Tao Chao
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.307-315

This paper aims to investigate a novel impact time control guidance (ITCG) law based on the sliding mode control (SMC) for a nonmaneuvering target using the predicted interception point (PIP).

To intercept the target with the minimal miss distance and desired impact time, an estimation of time-to-go is introduced. This estimation results in a precise impact time for multimissiles salvo attack the target at the same time. Even for a large lead angle, the desired impact time is achieved by using the sliding mode and Lyapunov stability theory. The singularity issue of the proposed impact time guidance laws is also analyzed to achieve an arbitrary lead angle with the desired impact time.

Numerical scenarios with desired impact time are presented to illustrate the performance of the proposed ITCG law. Comparison with the state-of-art impact time guidance laws proves that the guidance law in this paper can enable the missile to intercept the target with minimal miss distance and final impact time error. This method enables multiple missiles to attack the target simultaneously with different distances and arbitrary lead angles.

An ITCG law based on sliding mode and Lyapunov stability theory is proposed, and the switching surface is designed based on a novel estimation time-to-go for the missile to intercept the target with minimal miss distance. To intercept the target with initial arbitrary lead angles and desired impact time, the authors analysis the singular issue in SMC to ensure that the missile can intercept the target with arbitrary lead angle. The proposed approach for a nonmaneuvering target using the PIP has simple forms, and therefore, they have the superiority of being implemented easily.

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Impact time guidance law for arbitrary lead angle using sliding mode control10.1108/AEAT-11-2023-0292Aircraft Engineering and Aerospace Technology2024-02-02© 2024 Emerald Publishing LimitedHe DuMing YangSongyan WangTao ChaoAircraft Engineering and Aerospace Technology9622024-02-0210.1108/AEAT-11-2023-0292https://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2023-0292/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Model-free low-power observer based robust trajectory tracking control of UAV quadrotor with unknown disturbanceshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-08-2023-0212/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this paper is to design an output-feedback algorithm based on low-power observer (LPO), robust chattering-free controller and nonlinear disturbance observer (DO) to achieve trajectory tracking of quadrotor in the Cartesian plane. To achieve trajectory tracking control, firstly the decoupled rotational and translational model of quadrotor are modified by introducing backstepped state-space variables. In the second step, robust integral sliding mode control is designed based on the proportional-integral-derivative (PID) technique. In the third step, a DO is constructed. In next step, the measurable outputs, i.e. rotational and translational state variables, are used to design the LPO. Finally, in the control algorithm all state variables and its rates are replaced with its estimates obtained using the state-observer. The finding includes output-feedback control (OFC) algorithm designed by using a LPO. A modified backstepping model for rotational and rotational systems is developed prior to the design of integral sliding mode control based on PID technique. Unlike traditional high-gain observers (HGO), this paper used the LPO for state estimation of quadrotor systems to solve the problem of peaking phenomenon in HGO. Furthermore, a nonlinear DO is designed such that it attenuates disturbance with unknown magnitude and frequency. Moreover, a chattering reduction criterion has been introduced to solve the inherited chattering issue of controllers based on sliding mode technique. This paper presents input and output data-driven model-free control algorithm. That is, only input and output of the quadrotor model are required to achieve the trajectory tracking control. Therefore, for practical implementation, the number of on-board sensor is reduced. Although extensive research has been done for designing OFC algorithms for quadrotor, LPO has never been implemented for the rotational and translational state estimations of quadrotor. Furthermore, the mathematical model of rotational and translational systems is modified by using backstepped variables followed by the controller designed using PID and integral sliding mode control technique. Moreover, a DO is developed for attenuation of disturbance with unknown bound, magnitude and frequency.Model-free low-power observer based robust trajectory tracking control of UAV quadrotor with unknown disturbances
Muhammad Nabeel Siddiqui, Xiaolu Zhu, Hanad Rasool, Muhammad Bilal Afzal, Nigar Ahmed
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.316-328

The purpose of this paper is to design an output-feedback algorithm based on low-power observer (LPO), robust chattering-free controller and nonlinear disturbance observer (DO) to achieve trajectory tracking of quadrotor in the Cartesian plane.

To achieve trajectory tracking control, firstly the decoupled rotational and translational model of quadrotor are modified by introducing backstepped state-space variables. In the second step, robust integral sliding mode control is designed based on the proportional-integral-derivative (PID) technique. In the third step, a DO is constructed. In next step, the measurable outputs, i.e. rotational and translational state variables, are used to design the LPO. Finally, in the control algorithm all state variables and its rates are replaced with its estimates obtained using the state-observer.

The finding includes output-feedback control (OFC) algorithm designed by using a LPO. A modified backstepping model for rotational and rotational systems is developed prior to the design of integral sliding mode control based on PID technique. Unlike traditional high-gain observers (HGO), this paper used the LPO for state estimation of quadrotor systems to solve the problem of peaking phenomenon in HGO. Furthermore, a nonlinear DO is designed such that it attenuates disturbance with unknown magnitude and frequency. Moreover, a chattering reduction criterion has been introduced to solve the inherited chattering issue of controllers based on sliding mode technique.

This paper presents input and output data-driven model-free control algorithm. That is, only input and output of the quadrotor model are required to achieve the trajectory tracking control. Therefore, for practical implementation, the number of on-board sensor is reduced.

Although extensive research has been done for designing OFC algorithms for quadrotor, LPO has never been implemented for the rotational and translational state estimations of quadrotor. Furthermore, the mathematical model of rotational and translational systems is modified by using backstepped variables followed by the controller designed using PID and integral sliding mode control technique. Moreover, a DO is developed for attenuation of disturbance with unknown bound, magnitude and frequency.

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Model-free low-power observer based robust trajectory tracking control of UAV quadrotor with unknown disturbances10.1108/AEAT-08-2023-0212Aircraft Engineering and Aerospace Technology2024-02-13© 2024 Emerald Publishing LimitedMuhammad Nabeel SiddiquiXiaolu ZhuHanad RasoolMuhammad Bilal AfzalNigar AhmedAircraft Engineering and Aerospace Technology9622024-02-1310.1108/AEAT-08-2023-0212https://www.emerald.com/insight/content/doi/10.1108/AEAT-08-2023-0212/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Comparison of drilling of Inconel 625 by AWJM and WEDMhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0068/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to comparatively analyze the cut parts obtained as a result of cutting the Ni-based Inconel 625 alloy, which is widely used in the aerospace industry, with the wire electro-discharge machining (WEDM) and abrasive water jet machining (AWJM) methods in terms of macro- and microanalyses. In this study, calipers, Mitutoyo SJ-210, Nikon SMZ 745 T, scanning electron microscope and energy dispersive X-ray were used to determine kerf, surface roughness and macro- and microanalyses. Considering the applications in the turbine industry, it has been determined that the WEDM method is suitable to meet the standards for the machinability of Inconel 625 alloy. In contrast, the AWJM method does not meet the standards. Namely, while the kerf angle was formed because the hole entrance diameters of the holes obtained with AWJM were larger than the hole exit diameters, the equalization of the hole entry and exit dimensions, thanks to the perpendicularity and tension sensitivity of the wire electrode used in the holes drilled with WEDM ensured that the kerf angle was not formed. It is known that the surface roughness of the parts used in the turbine industry is accepted at Ra = 0.8 µm. In this study, the average roughness value obtained from the successful drilling of Inconel 625 alloy with the WEDM method was 0.799 µm, and the kerf angle was obtained as zero. In the cuts made with the AWJM method, thermal effects such as debris, microcracks and melted materials were not observed; an average surface roughness of 2.293 µm and a kerf of 0.976° were obtained.Comparison of drilling of Inconel 625 by AWJM and WEDM
Ferhat Ceritbinmez, Ali Günen
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.329-336

This study aims to comparatively analyze the cut parts obtained as a result of cutting the Ni-based Inconel 625 alloy, which is widely used in the aerospace industry, with the wire electro-discharge machining (WEDM) and abrasive water jet machining (AWJM) methods in terms of macro- and microanalyses.

In this study, calipers, Mitutoyo SJ-210, Nikon SMZ 745 T, scanning electron microscope and energy dispersive X-ray were used to determine kerf, surface roughness and macro- and microanalyses.

Considering the applications in the turbine industry, it has been determined that the WEDM method is suitable to meet the standards for the machinability of Inconel 625 alloy. In contrast, the AWJM method does not meet the standards. Namely, while the kerf angle was formed because the hole entrance diameters of the holes obtained with AWJM were larger than the hole exit diameters, the equalization of the hole entry and exit dimensions, thanks to the perpendicularity and tension sensitivity of the wire electrode used in the holes drilled with WEDM ensured that the kerf angle was not formed.

It is known that the surface roughness of the parts used in the turbine industry is accepted at Ra = 0.8 µm. In this study, the average roughness value obtained from the successful drilling of Inconel 625 alloy with the WEDM method was 0.799 µm, and the kerf angle was obtained as zero. In the cuts made with the AWJM method, thermal effects such as debris, microcracks and melted materials were not observed; an average surface roughness of 2.293 µm and a kerf of 0.976° were obtained.

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Comparison of drilling of Inconel 625 by AWJM and WEDM10.1108/AEAT-03-2023-0068Aircraft Engineering and Aerospace Technology2024-02-02© 2024 Emerald Publishing LimitedFerhat CeritbinmezAli GünenAircraft Engineering and Aerospace Technology9622024-02-0210.1108/AEAT-03-2023-0068https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0068/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Multiobjective optimization of process parameters of AZ91D/AgNPs/TiO composite fabricated by friction stir processing using response surface methodology and desirabilityhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2023-0196/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis paper aims to determine the optimum parametric settings for yielding superior mechanical properties, namely, ultimate tensile strength (UTS), yield strength (YS) and percentage elongation (EL) of AZ91D/AgNPs/TiO2 hybrid composite fabricated by friction stir processing. An empirical model has been developed to govern crucial influencing parameters, namely, rotation speed (RS), tool transverse speed (TS), number of passes (NPS) and reinforcement fraction (RF) or weight percentage. Box Behnken design (BBD) with four input parameters and three levels of each parameter was used to design the experimental work, and analysis of variance (ANOVA) was used to check the acceptability of the developed model. Desirability function analysis (DFA) for a multiresponse optimization approach is integrated with response surface methodology (RSM). The individual desirability index (IDI) was calculated for each response, and a composite desirability index (CDI) was obtained. The optimal parametric settings were determined based on maximum CDI values. A confirmation test is also performed to compare the actual and predicted values of responses. The relationship between input parameters and output responses (UTS, YS, and EL) was investigated using the Box-Behnken design (BBD). Silver nanoparticles (AgNPs) and nano-sized titanium dioxide (TiO2) enhanced the ultimate tensile strength and yield strength. It was observed that the inclusion of AgNPs led to an increase in ductility, while the increase in the weight fraction of TiO2 resulted in a decrease in ductility. AZ91D/AgNPs/TiO2 hybrid composite finds enormous applications in biomedical implants, aerospace, sports and aerospace industries, especially where lightweight materials with high strength are critical. In terms of optimum value through desirability, the experimental trials yield the following results: maximum value of UTS (318.369 MPa), maximum value of YS (200.120 MPa) and EL (7.610) at 1,021 rpm of RS, 70 mm/min of TS, 4 NPS and level 3 of RF.Multiobjective optimization of process parameters of AZ91D/AgNPs/TiO composite fabricated by friction stir processing using response surface methodology and desirability
Ram Niwas, Vikas Kumar
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.337-347

This paper aims to determine the optimum parametric settings for yielding superior mechanical properties, namely, ultimate tensile strength (UTS), yield strength (YS) and percentage elongation (EL) of AZ91D/AgNPs/TiO2 hybrid composite fabricated by friction stir processing.

An empirical model has been developed to govern crucial influencing parameters, namely, rotation speed (RS), tool transverse speed (TS), number of passes (NPS) and reinforcement fraction (RF) or weight percentage. Box Behnken design (BBD) with four input parameters and three levels of each parameter was used to design the experimental work, and analysis of variance (ANOVA) was used to check the acceptability of the developed model. Desirability function analysis (DFA) for a multiresponse optimization approach is integrated with response surface methodology (RSM). The individual desirability index (IDI) was calculated for each response, and a composite desirability index (CDI) was obtained. The optimal parametric settings were determined based on maximum CDI values. A confirmation test is also performed to compare the actual and predicted values of responses.

The relationship between input parameters and output responses (UTS, YS, and EL) was investigated using the Box-Behnken design (BBD). Silver nanoparticles (AgNPs) and nano-sized titanium dioxide (TiO2) enhanced the ultimate tensile strength and yield strength. It was observed that the inclusion of AgNPs led to an increase in ductility, while the increase in the weight fraction of TiO2 resulted in a decrease in ductility.

AZ91D/AgNPs/TiO2 hybrid composite finds enormous applications in biomedical implants, aerospace, sports and aerospace industries, especially where lightweight materials with high strength are critical.

In terms of optimum value through desirability, the experimental trials yield the following results: maximum value of UTS (318.369 MPa), maximum value of YS (200.120 MPa) and EL (7.610) at 1,021 rpm of RS, 70 mm/min of TS, 4 NPS and level 3 of RF.

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Multiobjective optimization of process parameters of AZ91D/AgNPs/TiO composite fabricated by friction stir processing using response surface methodology and desirability10.1108/AEAT-07-2023-0196Aircraft Engineering and Aerospace Technology2024-02-28© 2024 Emerald Publishing LimitedRam NiwasVikas KumarAircraft Engineering and Aerospace Technology9622024-02-2810.1108/AEAT-07-2023-0196https://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2023-0196/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Helicopter airfoil aerodynamic characteristics and rotor trim in sandy environmentshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2023-0308/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe sandy environment is one of the typical environments in which helicopters operate. Air-sand two-phase flow in sandy environments may be an important factor affecting flight safety. Taking a typical example, this paper aims to investigate the aerodynamic and rotor trim characteristics of the UH-60 helicopter in sandy environments. A computational study is conducted to simulate the air-sand flow over airfoils based on the Euler–Lagrange framework. The simulation uses the S-A turbulence model and the two-way momentum coupling methodology. Additionally, the trim characteristics of the UH-60 rotor are calculated based on the isolated rotor trim algorithm. The simulation results show that air-sand flow significantly affects the aerodynamic characteristics of the SC1095 airfoil and the SC1094R8 airfoil. The presence of sand particles leads to a decrease in lift and an increase in drag. The calculation results of the UH-60 helicopter rotor indicate that the thrust decreases and the torque increases in the sandy environment. To maintain a steady forward flight in sandy environments, it is necessary to increase the collective pitch and the longitudinal cyclic pitch. In this paper, the aerodynamic characteristics of airfoils and the trim characteristics in the air-sand flow of the UH-60 helicopter are discussed, which might be a new view to analyse the impact of sandy environments on helicopter safety and manoeuvring.Helicopter airfoil aerodynamic characteristics and rotor trim in sandy environments
Zejian Huang, Yihua Cao, Yanyang Wang
Aircraft Engineering and Aerospace Technology, Vol. 96, No. 2, pp.348-355

The sandy environment is one of the typical environments in which helicopters operate. Air-sand two-phase flow in sandy environments may be an important factor affecting flight safety. Taking a typical example, this paper aims to investigate the aerodynamic and rotor trim characteristics of the UH-60 helicopter in sandy environments.

A computational study is conducted to simulate the air-sand flow over airfoils based on the Euler–Lagrange framework. The simulation uses the S-A turbulence model and the two-way momentum coupling methodology. Additionally, the trim characteristics of the UH-60 rotor are calculated based on the isolated rotor trim algorithm.

The simulation results show that air-sand flow significantly affects the aerodynamic characteristics of the SC1095 airfoil and the SC1094R8 airfoil. The presence of sand particles leads to a decrease in lift and an increase in drag. The calculation results of the UH-60 helicopter rotor indicate that the thrust decreases and the torque increases in the sandy environment. To maintain a steady forward flight in sandy environments, it is necessary to increase the collective pitch and the longitudinal cyclic pitch.

In this paper, the aerodynamic characteristics of airfoils and the trim characteristics in the air-sand flow of the UH-60 helicopter are discussed, which might be a new view to analyse the impact of sandy environments on helicopter safety and manoeuvring.

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Helicopter airfoil aerodynamic characteristics and rotor trim in sandy environments10.1108/AEAT-11-2023-0308Aircraft Engineering and Aerospace Technology2024-02-22© 2024 Emerald Publishing LimitedZejian HuangYihua CaoYanyang WangAircraft Engineering and Aerospace Technology9622024-02-2210.1108/AEAT-11-2023-0308https://www.emerald.com/insight/content/doi/10.1108/AEAT-11-2023-0308/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Aeroengine gas trajectory prediction using time-series analysis auto regressive integrated moving averagehttps://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0018/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to propose the use of time series autoregressive integrated moving average (ARIMA) models to predict gas path performance in aero engines. The gas path is a critical component of an aero engine and its performance is essential for safe and efficient operation of the engine. The study analyzes a data set of gas path performance parameters obtained from a fleet of aero engines. The data is preprocessed and then fitted to ARIMA models to predict the future values of the gas path performance parameters. The performance of the ARIMA models is evaluated using various statistical metrics such as mean absolute error, mean squared error and root mean squared error. The results show that the ARIMA models can accurately predict the gas path performance parameters in aero engines. The proposed methodology can be used for real-time monitoring and controlling the gas path performance parameters in aero engines, which can improve the safety and efficiency of the engines. Both the Box-Ljung test and the residual analysis were used to demonstrate that the models for both time series were adequate. To determine whether or not the two series were stationary, the Augmented Dickey–Fuller unit root test was used in this study. The first-order ARIMA models were selected based on the observed autocorrelation function and partial autocorrelation function. Further, the authors find that the trend of predicted values and original values are similar and the error between them is small.Aeroengine gas trajectory prediction using time-series analysis auto regressive integrated moving average
M. Mary Victoria Florence, E. Priyadarshini
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to propose the use of time series autoregressive integrated moving average (ARIMA) models to predict gas path performance in aero engines. The gas path is a critical component of an aero engine and its performance is essential for safe and efficient operation of the engine.

The study analyzes a data set of gas path performance parameters obtained from a fleet of aero engines. The data is preprocessed and then fitted to ARIMA models to predict the future values of the gas path performance parameters. The performance of the ARIMA models is evaluated using various statistical metrics such as mean absolute error, mean squared error and root mean squared error. The results show that the ARIMA models can accurately predict the gas path performance parameters in aero engines.

The proposed methodology can be used for real-time monitoring and controlling the gas path performance parameters in aero engines, which can improve the safety and efficiency of the engines. Both the Box-Ljung test and the residual analysis were used to demonstrate that the models for both time series were adequate.

To determine whether or not the two series were stationary, the Augmented Dickey–Fuller unit root test was used in this study. The first-order ARIMA models were selected based on the observed autocorrelation function and partial autocorrelation function.

Further, the authors find that the trend of predicted values and original values are similar and the error between them is small.

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Aeroengine gas trajectory prediction using time-series analysis auto regressive integrated moving average10.1108/AEAT-01-2023-0018Aircraft Engineering and Aerospace Technology2023-08-01© 2023 Emerald Publishing LimitedM. Mary Victoria FlorenceE. PriyadarshiniAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2023-08-0110.1108/AEAT-01-2023-0018https://www.emerald.com/insight/content/doi/10.1108/AEAT-01-2023-0018/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
An easy-to-produce HIS-based MIMO radio altimeter antenna design for aircrafthttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0034/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this study is to design a radio altimeter antenna whose production process is facilitated and can work with multiple-input multiple-output (MIMO) properties to provide space gain on the aircraft. To create an easy-to-produce MIMO, a two-storied structure consisting of a reflector and a top antenna was designed. The dimensions of the reflector were prevented to get smaller to supply easy production. The unit cell nearly with the same dimensions of a lower frequency was protected through the original cell design. The co-planar structure with the use of a via connection was modified and a structure was achieved with no need to via for easy production, too. Finally, the antennas were placed side by side and the distance between them was optimized to achieve a MIMO operation. As a result, an easy-to-produce, compact and successful radio altimeter antenna was obtained with high antenna parameters such as 10.14 dBi gain and 10.55 dBi directivity, and the conical pattern along with proper MIMO features, through original reflector surface and top antenna system. Since radio altimeter antennas require high radiation properties, the microstrip antenna structure is generally used in literature. This paper contributes by presenting the radio altimeter application with antenna-reflective structure participation. The technical solutions were developed during the design, focusing on an easy manufacturing process for both the reflective surface and the upper antenna. Also, the combination of International Telecommunication Union’s recommended features that require high antenna properties was achieved, which is challenging to reach. In addition, by operating the antenna as a successful MIMO, two goals of easy production and space gain on aircraft have been attained at the same time.An easy-to-produce HIS-based MIMO radio altimeter antenna design for aircraft
Serap Kiriş, Muharrem Karaaslan
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this study is to design a radio altimeter antenna whose production process is facilitated and can work with multiple-input multiple-output (MIMO) properties to provide space gain on the aircraft.

To create an easy-to-produce MIMO, a two-storied structure consisting of a reflector and a top antenna was designed. The dimensions of the reflector were prevented to get smaller to supply easy production. The unit cell nearly with the same dimensions of a lower frequency was protected through the original cell design. The co-planar structure with the use of a via connection was modified and a structure was achieved with no need to via for easy production, too. Finally, the antennas were placed side by side and the distance between them was optimized to achieve a MIMO operation.

As a result, an easy-to-produce, compact and successful radio altimeter antenna was obtained with high antenna parameters such as 10.14 dBi gain and 10.55 dBi directivity, and the conical pattern along with proper MIMO features, through original reflector surface and top antenna system.

Since radio altimeter antennas require high radiation properties, the microstrip antenna structure is generally used in literature. This paper contributes by presenting the radio altimeter application with antenna-reflective structure participation. The technical solutions were developed during the design, focusing on an easy manufacturing process for both the reflective surface and the upper antenna. Also, the combination of International Telecommunication Union’s recommended features that require high antenna properties was achieved, which is challenging to reach. In addition, by operating the antenna as a successful MIMO, two goals of easy production and space gain on aircraft have been attained at the same time.

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An easy-to-produce HIS-based MIMO radio altimeter antenna design for aircraft10.1108/AEAT-02-2023-0034Aircraft Engineering and Aerospace Technology2023-11-28© 2023 Emerald Publishing LimitedSerap KirişMuharrem KaraaslanAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2023-11-2810.1108/AEAT-02-2023-0034https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0034/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Experimental study of the operating parameters on the performance of a single-stage Stirling cryocooler cooling infrared sensor for space applicationhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0051/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this study is to achieve lower and lower temperature as infrared sensors works faster and better used for space application. For getting good quality images from space, the infrared sensors are need to keep in cryogenic temperature. Cooling to cryogenic temperatures is necessary for space-borne sensors used for space applications. Infrared sensors work faster or better at lower temperatures. It is the need for time to achieve lower and lower temperatures. This study presents the investigation of the critical Stirling cryocooler parameters that influence the cold end temperature. In the paper, the design approach, the dimensions gained through thermal analysis, experimental procedure and testing results are discussed. The effect of parameters such as multilayer insulation, helium gas charging pressure, compressor input voltage and cooling load was investigated. The performance of gold-plated and aluminized multilayer insulation is checked. The tests were done with multilayer insulation covering inside and outside the Perspex cover. By using aluminized multilayer insulation inside and outside the Perspex cover, the improvement of 16 K in cool-down temperature was achieved. The cryocooler is charged with helium gas. The pressure varies between 14 and 18 bar. The optimum cooling is obtained for 17 bar gas pressure. The piston stroke increased as the compressor voltage increased, resulting in total helium gas compression. The optimum cool-down temperature was attained at 85 V. The cryocooler is designed to achieve the cool-down temperature of 2 W cooling load at 100 K. The lowest cool-down temperature recorded was 105 K at a 2 W cooling load. Multilayer insulation is the major item that keeps the thermal radiation from the sun from reaching the copper tip.Experimental study of the operating parameters on the performance of a single-stage Stirling cryocooler cooling infrared sensor for space application
Fayaz Kharadi, Karthikeyan A, Virendra Bhojwani, Prachi Dixit, Nand Jee Kanu, Nidhi Jain
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this study is to achieve lower and lower temperature as infrared sensors works faster and better used for space application. For getting good quality images from space, the infrared sensors are need to keep in cryogenic temperature. Cooling to cryogenic temperatures is necessary for space-borne sensors used for space applications. Infrared sensors work faster or better at lower temperatures. It is the need for time to achieve lower and lower temperatures.

This study presents the investigation of the critical Stirling cryocooler parameters that influence the cold end temperature. In the paper, the design approach, the dimensions gained through thermal analysis, experimental procedure and testing results are discussed.

The effect of parameters such as multilayer insulation, helium gas charging pressure, compressor input voltage and cooling load was investigated. The performance of gold-plated and aluminized multilayer insulation is checked. The tests were done with multilayer insulation covering inside and outside the Perspex cover.

By using aluminized multilayer insulation inside and outside the Perspex cover, the improvement of 16 K in cool-down temperature was achieved. The cryocooler is charged with helium gas. The pressure varies between 14 and 18 bar. The optimum cooling is obtained for 17 bar gas pressure. The piston stroke increased as the compressor voltage increased, resulting in total helium gas compression. The optimum cool-down temperature was attained at 85 V.

The cryocooler is designed to achieve the cool-down temperature of 2 W cooling load at 100 K. The lowest cool-down temperature recorded was 105 K at a 2 W cooling load. Multilayer insulation is the major item that keeps the thermal radiation from the sun from reaching the copper tip.

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Experimental study of the operating parameters on the performance of a single-stage Stirling cryocooler cooling infrared sensor for space application10.1108/AEAT-02-2023-0051Aircraft Engineering and Aerospace Technology2023-06-13© 2023 Emerald Publishing LimitedFayaz KharadiKarthikeyan AVirendra BhojwaniPrachi DixitNand Jee KanuNidhi JainAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2023-06-1310.1108/AEAT-02-2023-0051https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0051/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Role of spirulina microalgae blends in the micro gas turbine on engine performance and emission characteristicshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0052/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestDiesel has traditionally been considered the best-suited and most widely used fuel in various sectors, including manufacturing industries, power production, automobiles and transportation. However, with the ongoing crisis of fossil fuel inadequacy, the search for alternative fuels and their application in these sectors has become increasingly important. One particularly interesting and beneficial alternative fuel is biodiesel derived from bio sources. In this research, an attempt was made to use biodiesel in an unconventional micro gas turbine engine. It will remove the concentric use of diesel engines for power production by improving fuel efficiency as well as increasing the power production rate. Before the fuel is used enormously, it has to be checked in many ways such as performance, emission and combustion analysis experimentally. In this paper, a detailed experimental study was made for the use of Spirulina microalgae biodiesel in a micro gas turbine. A small-scale setup with the primary micro gas turbine and secondary instruments such as a data acquisition system and AVL gas analyser. The reason for selecting the third-generation microalgae is due to its high lipid and biodiesel production rate. For the conduction of experimental tests, certain conditions were followed in addition that the engine rotating rpm was varied from 4,000, 5,000 and 6,000 rpm. The favourable and predicted results were obtained with the use of microalgae biodiesel. The performance and combustion results were not exactly equal or greater for biodiesel blends but close to the values of pure diesel; however, the reduction in the emission of CO was at the appreciable level for the used spirulina microalgae biodiesel. The emission of nitrogen oxides and carbon dioxide was a little higher than the use of pure diesel. This experimental analysis results proved that the use of spirulina microalgae biodiesel is both economical and effective replacement for fossil fuel.Role of spirulina microalgae blends in the micro gas turbine on engine performance and emission characteristics
Gokulnath R., Booma Devi
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

Diesel has traditionally been considered the best-suited and most widely used fuel in various sectors, including manufacturing industries, power production, automobiles and transportation. However, with the ongoing crisis of fossil fuel inadequacy, the search for alternative fuels and their application in these sectors has become increasingly important. One particularly interesting and beneficial alternative fuel is biodiesel derived from bio sources.

In this research, an attempt was made to use biodiesel in an unconventional micro gas turbine engine. It will remove the concentric use of diesel engines for power production by improving fuel efficiency as well as increasing the power production rate. Before the fuel is used enormously, it has to be checked in many ways such as performance, emission and combustion analysis experimentally.

In this paper, a detailed experimental study was made for the use of Spirulina microalgae biodiesel in a micro gas turbine. A small-scale setup with the primary micro gas turbine and secondary instruments such as a data acquisition system and AVL gas analyser. The reason for selecting the third-generation microalgae is due to its high lipid and biodiesel production rate. For the conduction of experimental tests, certain conditions were followed in addition that the engine rotating rpm was varied from 4,000, 5,000 and 6,000 rpm. The favourable and predicted results were obtained with the use of microalgae biodiesel.

The performance and combustion results were not exactly equal or greater for biodiesel blends but close to the values of pure diesel; however, the reduction in the emission of CO was at the appreciable level for the used spirulina microalgae biodiesel. The emission of nitrogen oxides and carbon dioxide was a little higher than the use of pure diesel. This experimental analysis results proved that the use of spirulina microalgae biodiesel is both economical and effective replacement for fossil fuel.

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Role of spirulina microalgae blends in the micro gas turbine on engine performance and emission characteristics10.1108/AEAT-02-2023-0052Aircraft Engineering and Aerospace Technology2023-10-09© 2023 Emerald Publishing LimitedGokulnath R.Booma DeviAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2023-10-0910.1108/AEAT-02-2023-0052https://www.emerald.com/insight/content/doi/10.1108/AEAT-02-2023-0052/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Flow of visco elastic fluid through a porous medium with chemical reactionhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2022-0094/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestChemical reaction effects are added to the governing equation. This paper aims to get the solution by converting the partial differential equation into an ordinary differential equation and solve using a perturbation scheme and applying the boundary conditions. In this paper, the authors discussed the chemical reaction effects of heat and mass transfer on megnato hydro dynamics free convective rotating flow of a visco-elastic incompressible electrically conducting fluid past a vertical porous plate through a porous medium with suction and heat source. The authors analyze the effect of time dependent fluctuating suction on a visco-elastic fluid flow. Using variable parameters of the fluid, the velocity, temperature and concentration of the fluid are analyzed through graphs. The velocity profile reduces by increasing the values of thermal Grashof number (Gr), mass Grashof number (Gc) and the magnetic parameter (M). On the other hand, the velocity profile gets increased by increasing the permeability parameter (K). The temperature profile decreases by raising the value of Prandtl number (Pr) and frequency of oscillation parameter (ω). However, the source parameter (S) has the opposite effect on the temperature profile. The concentration profile reduces in all points by raising the chemical reaction parameter Kl, Schmidt number Sc, frequency of oscillation ω and the time t.Flow of visco elastic fluid through a porous medium with chemical reaction
Niranjana N., Vidhya M., Govindarajan A., Rajesh K.
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

Chemical reaction effects are added to the governing equation. This paper aims to get the solution by converting the partial differential equation into an ordinary differential equation and solve using a perturbation scheme and applying the boundary conditions.

In this paper, the authors discussed the chemical reaction effects of heat and mass transfer on megnato hydro dynamics free convective rotating flow of a visco-elastic incompressible electrically conducting fluid past a vertical porous plate through a porous medium with suction and heat source. The authors analyze the effect of time dependent fluctuating suction on a visco-elastic fluid flow.

Using variable parameters of the fluid, the velocity, temperature and concentration of the fluid are analyzed through graphs.

The velocity profile reduces by increasing the values of thermal Grashof number (Gr), mass Grashof number (Gc) and the magnetic parameter (M). On the other hand, the velocity profile gets increased by increasing the permeability parameter (K). The temperature profile decreases by raising the value of Prandtl number (Pr) and frequency of oscillation parameter (ω). However, the source parameter (S) has the opposite effect on the temperature profile. The concentration profile reduces in all points by raising the chemical reaction parameter Kl, Schmidt number Sc, frequency of oscillation ω and the time t.

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Flow of visco elastic fluid through a porous medium with chemical reaction10.1108/AEAT-03-2022-0094Aircraft Engineering and Aerospace Technology2022-08-01© 2022 Emerald Publishing LimitedNiranjana N.Vidhya M.Govindarajan A.Rajesh K.Aircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2022-08-0110.1108/AEAT-03-2022-0094https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2022-0094/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2022 Emerald Publishing Limited
Research on vibration characteristics of rocket engine flow pipelinehttps://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0078/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestAbnormal vibrations often occur in the liquid oxygen kerosene transmission pipelines of rocket engines, which seriously threaten their safety. Improper handling can result in failed rocket launches and significant economic losses. Therefore, this paper aims to examine vibrations in transmission pipelines. In this study, a three-dimensional high-pressure pipeline model composed of corrugated pipes, multi-section bent pipes, and other auxiliary structures was established. The fluid–solid coupling method was used to analyse vibration characteristics of the pipeline under various external excitations. The simulation results were visualised using MATLAB, and their validity was verified via a thermal test. In this study, the vibration mechanism of a complex high-pressure pipeline was examined via a visualisation method. The results showed that the low-frequency vibration of the pipe was caused by fluid self-excited pressure pulsation, whereas the vibration of the engine system caused a high-frequency vibration of the pipeline. The excitation of external pressure pulses did not significantly affect the vibrations of the pipelines. The visualisation results indicated that the severe vibration position of the pipeline thermal test is mainly concentrated between the inlet and outlet and between the two bellows. The results of this study aid in understanding the causes of abnormal vibrations in rocket engine pipelines. The causes of different vibration frequencies in the complex pipelines of rocket engines and the propagation characteristics of external vibration excitation were obtained.Research on vibration characteristics of rocket engine flow pipeline
Su Yong, Gong Wu-Qi
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

Abnormal vibrations often occur in the liquid oxygen kerosene transmission pipelines of rocket engines, which seriously threaten their safety. Improper handling can result in failed rocket launches and significant economic losses. Therefore, this paper aims to examine vibrations in transmission pipelines.

In this study, a three-dimensional high-pressure pipeline model composed of corrugated pipes, multi-section bent pipes, and other auxiliary structures was established. The fluid–solid coupling method was used to analyse vibration characteristics of the pipeline under various external excitations. The simulation results were visualised using MATLAB, and their validity was verified via a thermal test.

In this study, the vibration mechanism of a complex high-pressure pipeline was examined via a visualisation method. The results showed that the low-frequency vibration of the pipe was caused by fluid self-excited pressure pulsation, whereas the vibration of the engine system caused a high-frequency vibration of the pipeline. The excitation of external pressure pulses did not significantly affect the vibrations of the pipelines. The visualisation results indicated that the severe vibration position of the pipeline thermal test is mainly concentrated between the inlet and outlet and between the two bellows.

The results of this study aid in understanding the causes of abnormal vibrations in rocket engine pipelines.

The causes of different vibration frequencies in the complex pipelines of rocket engines and the propagation characteristics of external vibration excitation were obtained.

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Research on vibration characteristics of rocket engine flow pipeline10.1108/AEAT-03-2023-0078Aircraft Engineering and Aerospace Technology2024-03-11© 2024 Emerald Publishing LimitedSu YongGong Wu-QiAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-1110.1108/AEAT-03-2023-0078https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0078/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Adaptive sensor fault tolerant control with prescribed performance for unmanned autonomous helicopter based on neural networkshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0080/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis paper aims to study a neural network-based fault-tolerant controller to improve the tracking control performance of an unmanned autonomous helicopter with system uncertainty, external disturbances and sensor faults, using the prescribed performance method. To ensure that the tracking error satisfies the prescribed performance, the authors adopt an error transformation function method. A control scheme based on the neural network and high-order disturbance observer is designed to guarantee the boundedness of the closed-loop system. A simulation is performed to prove the validity of the control scheme. The developed adaptive fault-tolerant control method makes the system with sensor fault realize tracking control. The error transformation function method can effectively handle the prescribed performance requirements. Sensor fault can be regarded as a type of system uncertainty. The uncertainty can be approximated accurately using neural networks. A high-order disturbance observer can effectively suppress compound disturbances. The tracking performance requirements of unmanned autonomous helicopter system are considered in the design of sensor fault-tolerant control. The inequality constraint that the output tracking error must satisfy is transformed into an unconstrained problem by introducing an error transformation function. The fault state of the velocity sensor is considered as the system uncertainty, and a neural network is used to approach the total uncertainty. Neural network estimation errors and external disturbances are treated as compound disturbances, and a high-order disturbance observer is constructed to compensate for them.Adaptive sensor fault tolerant control with prescribed performance for unmanned autonomous helicopter based on neural networks
Min Wan, Mou Chen, Mihai Lungu
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This paper aims to study a neural network-based fault-tolerant controller to improve the tracking control performance of an unmanned autonomous helicopter with system uncertainty, external disturbances and sensor faults, using the prescribed performance method.

To ensure that the tracking error satisfies the prescribed performance, the authors adopt an error transformation function method. A control scheme based on the neural network and high-order disturbance observer is designed to guarantee the boundedness of the closed-loop system. A simulation is performed to prove the validity of the control scheme.

The developed adaptive fault-tolerant control method makes the system with sensor fault realize tracking control. The error transformation function method can effectively handle the prescribed performance requirements. Sensor fault can be regarded as a type of system uncertainty. The uncertainty can be approximated accurately using neural networks. A high-order disturbance observer can effectively suppress compound disturbances.

The tracking performance requirements of unmanned autonomous helicopter system are considered in the design of sensor fault-tolerant control. The inequality constraint that the output tracking error must satisfy is transformed into an unconstrained problem by introducing an error transformation function. The fault state of the velocity sensor is considered as the system uncertainty, and a neural network is used to approach the total uncertainty. Neural network estimation errors and external disturbances are treated as compound disturbances, and a high-order disturbance observer is constructed to compensate for them.

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Adaptive sensor fault tolerant control with prescribed performance for unmanned autonomous helicopter based on neural networks10.1108/AEAT-03-2023-0080Aircraft Engineering and Aerospace Technology2024-03-29© 2024 Emerald Publishing LimitedMin WanMou ChenMihai LunguAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-2910.1108/AEAT-03-2023-0080https://www.emerald.com/insight/content/doi/10.1108/AEAT-03-2023-0080/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Trend in tropopause warming and its influence aircraft performancehttps://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2022-0128/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestTemperature anomalies in the upper troposphere have become a reality as a result of global warming, which has a noticeable impact on aircraft performance. The purpose of this study is to investigate the total air temperature (TAT) anomaly observed during the cruise level and its impact on engine parameter variations. Empirical methodology is used in this study, and it is based on measurements and observations of anomalous phenomena on the tropopause. The primary data were taken from the Boeing 747-8F's enhanced flight data recorder, which refers to the quantitative method, while the qualitative method is based on a literature review and interviews. The GEnx Integrated Vehicle Health Management system was used for the study's evaluation of engine performance to support the complete range of operational priorities throughout the entire engine lifecycle. The study's findings indicate that TAT and SAT anomalies, which occur between 270- and 320-feet flight level, have a substantial impact on aircraft performance at cruise altitude and, as a result, on engine parameters, specifically an increase in fuel consumption and engine exhaust gas temperature values. The TAT and Ram Rise anomalies were the focus of the atmospheric deviations, which were assessed as major departures from the International Civil Aviation Organizations–defined International Standard Atmosphere, which is obvious on a positive tendency and so goes against the norms. Necessary fixed flight parameters gathered from the aircraft's enhanced airborne flight recorder (EAFR) via Aeronautical Radio Incorporated (ARINC) 664 Part 7 at a certain velocity and altitude interfacing with the diagnostic program direct parameter display (DPD), allow for analysis of aircraft performance in a real-time frame. Thus, processed data transmits to the ground maintenance infrastructure for future evaluation and for proper maintenance solutions. A real-time analysis of aircraft performance is possible using the diagnostic program DPD in conjunction with necessary fixed flight parameters obtained from the aircraft's EAFR via ARINC 664 Part 7 at a specific speed and altitude. Thus, processed data is transmitted to the ground infrastructure for maintenance to be evaluated in the future and to find the best maintenance fixes.Trend in tropopause warming and its influence aircraft performance
Mehmet Necati Cizrelioğullari, Tapdig Veyran Imanov, Tugrul Gunay, Aliyev Shaiq Amir
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

Temperature anomalies in the upper troposphere have become a reality as a result of global warming, which has a noticeable impact on aircraft performance. The purpose of this study is to investigate the total air temperature (TAT) anomaly observed during the cruise level and its impact on engine parameter variations.

Empirical methodology is used in this study, and it is based on measurements and observations of anomalous phenomena on the tropopause. The primary data were taken from the Boeing 747-8F's enhanced flight data recorder, which refers to the quantitative method, while the qualitative method is based on a literature review and interviews. The GEnx Integrated Vehicle Health Management system was used for the study's evaluation of engine performance to support the complete range of operational priorities throughout the entire engine lifecycle.

The study's findings indicate that TAT and SAT anomalies, which occur between 270- and 320-feet flight level, have a substantial impact on aircraft performance at cruise altitude and, as a result, on engine parameters, specifically an increase in fuel consumption and engine exhaust gas temperature values. The TAT and Ram Rise anomalies were the focus of the atmospheric deviations, which were assessed as major departures from the International Civil Aviation Organizations–defined International Standard Atmosphere, which is obvious on a positive tendency and so goes against the norms.

Necessary fixed flight parameters gathered from the aircraft's enhanced airborne flight recorder (EAFR) via Aeronautical Radio Incorporated (ARINC) 664 Part 7 at a certain velocity and altitude interfacing with the diagnostic program direct parameter display (DPD), allow for analysis of aircraft performance in a real-time frame. Thus, processed data transmits to the ground maintenance infrastructure for future evaluation and for proper maintenance solutions.

A real-time analysis of aircraft performance is possible using the diagnostic program DPD in conjunction with necessary fixed flight parameters obtained from the aircraft's EAFR via ARINC 664 Part 7 at a specific speed and altitude. Thus, processed data is transmitted to the ground infrastructure for maintenance to be evaluated in the future and to find the best maintenance fixes.

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Trend in tropopause warming and its influence aircraft performance10.1108/AEAT-05-2022-0128Aircraft Engineering and Aerospace Technology2023-02-15© 2023 Emerald Publishing LimitedMehmet Necati CizrelioğullariTapdig Veyran ImanovTugrul GunayAliyev Shaiq AmirAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2023-02-1510.1108/AEAT-05-2022-0128https://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2022-0128/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2023 Emerald Publishing Limited
Optimization of landing gear under consideration of vibration comfort for civil aircrafthttps://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0130/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to investigate the design and optimization of landing gear buffers to improve the landing-phase comfort of civil aircraft. The vibration comfort during the landing and taxiing phases is calculated and evaluated based on the flight-testing data for a type of civil aircraft. The calculation and evaluation are under the guidance of the vibration comfort standard of GB/T13441.1-2007 and related files. The authors establish here a rigid-flexible coupled multibody dynamics finite element model of one full-size aircraft. Furthermore, the authors also implement a dynamic simulation for the landing and taxiing processes. Also, an analysis of how the main parameters of the buffers affect the vibration comfort is presented. Finally, the optimization of the single-chamber and double-chamber buffers in the landing gear is performed considering vibration comfort. The double-chamber buffer with optimized parameters in landing gear can improve the vibration comfort of the aircraft during the landing and taxiing phases. Moreover, the comfort index can be increased by 25.6% more than that of a single-chamber type. To the best of the authors’ knowledge, this study first investigates the evaluation methods and evaluation indexes on the aircraft vibration comfort, then further conducts the optimization of the parameters of landing gear buffer with different structures, so as to improve the comfort of aircraft passengers during landing process. Most of the current studies on aircraft landing gear have focused on the strength and safety of the landing gear, with very limited research on cabin vibration comfort during landing and subsequent taxiing because of the coupling of runway surface unevenness and airframe vibration.Optimization of landing gear under consideration of vibration comfort for civil aircraft
Shuowen Yan, Pu Xue, Long Liu, M.S. Zahran
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to investigate the design and optimization of landing gear buffers to improve the landing-phase comfort of civil aircraft.

The vibration comfort during the landing and taxiing phases is calculated and evaluated based on the flight-testing data for a type of civil aircraft. The calculation and evaluation are under the guidance of the vibration comfort standard of GB/T13441.1-2007 and related files. The authors establish here a rigid-flexible coupled multibody dynamics finite element model of one full-size aircraft. Furthermore, the authors also implement a dynamic simulation for the landing and taxiing processes. Also, an analysis of how the main parameters of the buffers affect the vibration comfort is presented. Finally, the optimization of the single-chamber and double-chamber buffers in the landing gear is performed considering vibration comfort.

The double-chamber buffer with optimized parameters in landing gear can improve the vibration comfort of the aircraft during the landing and taxiing phases. Moreover, the comfort index can be increased by 25.6% more than that of a single-chamber type.

To the best of the authors’ knowledge, this study first investigates the evaluation methods and evaluation indexes on the aircraft vibration comfort, then further conducts the optimization of the parameters of landing gear buffer with different structures, so as to improve the comfort of aircraft passengers during landing process. Most of the current studies on aircraft landing gear have focused on the strength and safety of the landing gear, with very limited research on cabin vibration comfort during landing and subsequent taxiing because of the coupling of runway surface unevenness and airframe vibration.

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Optimization of landing gear under consideration of vibration comfort for civil aircraft10.1108/AEAT-05-2023-0130Aircraft Engineering and Aerospace Technology2024-03-12© 2024 Emerald Publishing LimitedShuowen YanPu XueLong LiuM.S. ZahranAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-1210.1108/AEAT-05-2023-0130https://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0130/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Formation fault-tolerant control for multiple UAVs with external disturbanceshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0148/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this paper is to construct an event-triggered finite-time fault-tolerant formation tracking controller, which can achieve a time-varying formation control for multiple unmanned aerial vehicles (UAVs) during actuator failures and external perturbations. First, this study developed the formation tracking protocol for each follower using UAV formation members, defining the tracking inaccuracy of the UAV followers’ location. Subsequently, this study designed the multilayer event-triggered controller based on the backstepping method framework within finite time. Then, considering the actuator failures, and added self-adaptive thought for fault-tolerant control within finite time, the event-triggered closed-loop system is subsequently shown to be a finite-time stable system. Furthermore, the Zeno behavior is analyzed to prevent infinite triggering instances within a finite time. Finally, simulations are conducted with external disturbances and actuator failure conditions to demonstrate formation tracking controller performance. It achieves improved performance in the presence of external disturbances and system failures. Combining limited-time adaptive control and event triggering improves system stability, increase robustness to disturbances and calculation efficiency. In addition, the designed formation tracking controller can effectively control the time-varying formation of the leader and followers to complete the task, and by adding a fixed-time observer, it can effectively compensate for external disturbances and improve formation control accuracy. A formation-following controller is designed, which can handle both external disturbances and internal actuator failures during formation flight, and the proposed method can be applied to a variety of formation control scenarios and does not rely on a specific type of UAV or communication network.Formation fault-tolerant control for multiple UAVs with external disturbances
Ziyuan Ma, Huajun Gong, Xinhua Wang
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this paper is to construct an event-triggered finite-time fault-tolerant formation tracking controller, which can achieve a time-varying formation control for multiple unmanned aerial vehicles (UAVs) during actuator failures and external perturbations.

First, this study developed the formation tracking protocol for each follower using UAV formation members, defining the tracking inaccuracy of the UAV followers’ location. Subsequently, this study designed the multilayer event-triggered controller based on the backstepping method framework within finite time. Then, considering the actuator failures, and added self-adaptive thought for fault-tolerant control within finite time, the event-triggered closed-loop system is subsequently shown to be a finite-time stable system. Furthermore, the Zeno behavior is analyzed to prevent infinite triggering instances within a finite time. Finally, simulations are conducted with external disturbances and actuator failure conditions to demonstrate formation tracking controller performance.

It achieves improved performance in the presence of external disturbances and system failures. Combining limited-time adaptive control and event triggering improves system stability, increase robustness to disturbances and calculation efficiency. In addition, the designed formation tracking controller can effectively control the time-varying formation of the leader and followers to complete the task, and by adding a fixed-time observer, it can effectively compensate for external disturbances and improve formation control accuracy.

A formation-following controller is designed, which can handle both external disturbances and internal actuator failures during formation flight, and the proposed method can be applied to a variety of formation control scenarios and does not rely on a specific type of UAV or communication network.

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Formation fault-tolerant control for multiple UAVs with external disturbances10.1108/AEAT-05-2023-0148Aircraft Engineering and Aerospace Technology2024-03-13© 2024 Emerald Publishing LimitedZiyuan MaHuajun GongXinhua WangAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-1310.1108/AEAT-05-2023-0148https://www.emerald.com/insight/content/doi/10.1108/AEAT-05-2023-0148/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Effect of Soret on MHD rotating fluid flow through a porous medium with heat and mass transferhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2022-0162/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this paper to analyze the effect of Soret with heat and mass transfer on an unsteady two-dimensional Magnetohydrodynamics flow through a porous medium under the influence of the uniform transverse magnetic field in a rotating parallel plate is considered. A mathematical model was developed using the slip conditions under unsteady state situations. Analytical expressions for the velocity, temperature and concentration profiles, wall shear stress, rates of heat and mass transfer and volumetric flow rate were obtained and computationally discussed with respect to the non-dimensional parameters. Further, the velocity reduces with increasing Hartmann number M and increases with Grashof number Gr and permeability parameter K. It is observed that temperature reduces with an increase in Prandtl number Pr and ω. It is noted that the thermal radiation increases with increase in Soret number Sr, Schmidt number Sc, Prandtl number pr and ω. Concentration decreases with an increase in radiation parameter R and chemical reaction parameter Kc.Effect of Soret on MHD rotating fluid flow through a porous medium with heat and mass transfer
Lawanya T., Vidhya M., Govindarajan A.
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this paper to analyze the effect of Soret with heat and mass transfer on an unsteady two-dimensional Magnetohydrodynamics flow through a porous medium under the influence of the uniform transverse magnetic field in a rotating parallel plate is considered.

A mathematical model was developed using the slip conditions under unsteady state situations. Analytical expressions for the velocity, temperature and concentration profiles, wall shear stress, rates of heat and mass transfer and volumetric flow rate were obtained and computationally discussed with respect to the non-dimensional parameters. Further, the velocity reduces with increasing Hartmann number M and increases with Grashof number Gr and permeability parameter K.

It is observed that temperature reduces with an increase in Prandtl number Pr and ω. It is noted that the thermal radiation increases with increase in Soret number Sr, Schmidt number Sc, Prandtl number pr and ω.

Concentration decreases with an increase in radiation parameter R and chemical reaction parameter Kc.

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Effect of Soret on MHD rotating fluid flow through a porous medium with heat and mass transfer10.1108/AEAT-06-2022-0162Aircraft Engineering and Aerospace Technology2022-09-19© 2022 Emerald Publishing LimitedLawanya T.Vidhya M.Govindarajan A.Aircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2022-09-1910.1108/AEAT-06-2022-0162https://www.emerald.com/insight/content/doi/10.1108/AEAT-06-2022-0162/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2022 Emerald Publishing Limited
Externally applied magnetic force effect on the Couette flow of a viscous incompressible fluid with heat source/sinkhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2022-0187/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to investigate the effect of externally applied magnetic force and heat transfer with a heat source/sink on the Couette flow with viscous dissipation in a horizontal rotating channel. The magnetic force is added to the governing equations. The effects of fluid flow parameters are observed under the applied magnetic force. In this system, the magnetic force is applied perpendicular to the plane of the fluid flow. In recent years, the magnetic field has renewed interest in aerospace technology. The physical and theoretical approach in the multidisciplinary field of magneto fluid dynamics (MFD) is applied in the field of aerospace vehicle design. Authors use the perturbation method to solve and find the approximate solutions of differential equations. First, convert the partial differential equation to ordinary differential equation and calculate the approximate solutions in two cases. The first solution got by assuming heat generating in the fluid and the second one got when heat absorbing. After applying the external magnetic force, the effects of various fluid parameters velocity, temperature, skin friction coefficient and Nusselt number are found and discussed using tables and graphs. It is found that the velocity of the fluid has decreased tendency when the rotation of the fluid and magnetic force on the fluid increases. The temperature of the fluid, Prandtl value and Eckert number increased when the heat source generated heat. When heat absorbs the heat, sink parameter increases and the temperature of the fluid decreases. Also, while heat absorbs, the temperature increases when the Prandtl value and Eckert number increase. The skin friction coefficient on the surface increases, when the rotation parameter and the magnetic force parameter of the fluid increase. In the case of heat generating, the Nusselt number increased, while the Eckert number and Prandtl numbers increased. Also, the Nusselt number has larger values when the heat source parameter has near the constant temperature, and it has smaller values when the temperature varies. In the case of heat-absorbing, the Nusselt number decreased when the Eckert and Prandtl numbers increased. Also, the Nusselt number varies up and down while the heat absorbing parameter increases.Externally applied magnetic force effect on the Couette flow of a viscous incompressible fluid with heat source/sink
Sheeba Juliet S., Vidhya M., Govindarajan A.
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to investigate the effect of externally applied magnetic force and heat transfer with a heat source/sink on the Couette flow with viscous dissipation in a horizontal rotating channel. The magnetic force is added to the governing equations. The effects of fluid flow parameters are observed under the applied magnetic force. In this system, the magnetic force is applied perpendicular to the plane of the fluid flow. In recent years, the magnetic field has renewed interest in aerospace technology. The physical and theoretical approach in the multidisciplinary field of magneto fluid dynamics (MFD) is applied in the field of aerospace vehicle design.

Authors use the perturbation method to solve and find the approximate solutions of differential equations. First, convert the partial differential equation to ordinary differential equation and calculate the approximate solutions in two cases. The first solution got by assuming heat generating in the fluid and the second one got when heat absorbing. After applying the external magnetic force, the effects of various fluid parameters velocity, temperature, skin friction coefficient and Nusselt number are found and discussed using tables and graphs.

It is found that the velocity of the fluid has decreased tendency when the rotation of the fluid and magnetic force on the fluid increases. The temperature of the fluid, Prandtl value and Eckert number increased when the heat source generated heat. When heat absorbs the heat, sink parameter increases and the temperature of the fluid decreases. Also, while heat absorbs, the temperature increases when the Prandtl value and Eckert number increase.

The skin friction coefficient on the surface increases, when the rotation parameter and the magnetic force parameter of the fluid increase. In the case of heat generating, the Nusselt number increased, while the Eckert number and Prandtl numbers increased. Also, the Nusselt number has larger values when the heat source parameter has near the constant temperature, and it has smaller values when the temperature varies. In the case of heat-absorbing, the Nusselt number decreased when the Eckert and Prandtl numbers increased. Also, the Nusselt number varies up and down while the heat absorbing parameter increases.

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Externally applied magnetic force effect on the Couette flow of a viscous incompressible fluid with heat source/sink10.1108/AEAT-07-2022-0187Aircraft Engineering and Aerospace Technology2022-10-28© 2022 Emerald Publishing LimitedSheeba Juliet S.Vidhya M.Govindarajan A.Aircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2022-10-2810.1108/AEAT-07-2022-0187https://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2022-0187/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2022 Emerald Publishing Limited
Effects of hydrogen and chicken waste blends in the internal combustion engines for superior engine performance and emission characteristics assisted with graphite oxidehttps://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2022-0201/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe demand for energy is increasing massively due to urbanization and industrialization. Due to the massive usage of diesel engines in the transportation sector, global warming is increasing rapidly. The purpose of this paper is to use hydrogen as the potential alternative for diesel engine. A series of tests conducted in the twin cylinder four stroke diesel engine at various engine speeds. In addition to the hydrogen, the ultrasonication is applied to add the nanoparticles to the neat diesel. The role of nanoparticles on engine performance is effective owing to its physicochemical properties. Here, neat diesel mixed 30% of biodiesel along with the hydrogen at the concentration of 10%, 20% and 30% and 50 ppm of graphite oxide to form the blends DNH10, DNH20 and DNH30. Inclusion of both hydrogen and nanoparticles increases the brake power and brake thermal efficiency (BTE) of the engine with relatively less fuel consumption. Compared to all blends, the maximum BTE of 33.3% has been reported by 30% hydrogen-based fuel. On the contrary, the production of the pollutants also reduces as the hydrogen concentration increases. Majority of the pollutants such as carbon monoxide, carbon dioxide and hydrocarbon were dropped massively compared to diesel. On the contrary, there is no reduction in nitrogen of oxides (NOx). Highest production of NOx was witnessed for 30% hydrogen fuel due to the premixed combustion phase and cylinder temperatures.Effects of hydrogen and chicken waste blends in the internal combustion engines for superior engine performance and emission characteristics assisted with graphite oxide
Dinesh R., Stanly Jones Retnam, Dev Anand M., Edwin Raja Dhas J.
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The demand for energy is increasing massively due to urbanization and industrialization. Due to the massive usage of diesel engines in the transportation sector, global warming is increasing rapidly. The purpose of this paper is to use hydrogen as the potential alternative for diesel engine.

A series of tests conducted in the twin cylinder four stroke diesel engine at various engine speeds. In addition to the hydrogen, the ultrasonication is applied to add the nanoparticles to the neat diesel. The role of nanoparticles on engine performance is effective owing to its physicochemical properties. Here, neat diesel mixed 30% of biodiesel along with the hydrogen at the concentration of 10%, 20% and 30% and 50 ppm of graphite oxide to form the blends DNH10, DNH20 and DNH30.

Inclusion of both hydrogen and nanoparticles increases the brake power and brake thermal efficiency (BTE) of the engine with relatively less fuel consumption. Compared to all blends, the maximum BTE of 33.3% has been reported by 30% hydrogen-based fuel. On the contrary, the production of the pollutants also reduces as the hydrogen concentration increases.

Majority of the pollutants such as carbon monoxide, carbon dioxide and hydrocarbon were dropped massively compared to diesel. On the contrary, there is no reduction in nitrogen of oxides (NOx). Highest production of NOx was witnessed for 30% hydrogen fuel due to the premixed combustion phase and cylinder temperatures.

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Effects of hydrogen and chicken waste blends in the internal combustion engines for superior engine performance and emission characteristics assisted with graphite oxide10.1108/AEAT-07-2022-0201Aircraft Engineering and Aerospace Technology2022-09-23© 2022 Emerald Publishing LimitedDinesh R.Stanly Jones RetnamDev Anand M.Edwin Raja Dhas J.Aircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2022-09-2310.1108/AEAT-07-2022-0201https://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2022-0201/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2022 Emerald Publishing Limited
Developing conceptual model for safety risk management in aviation as building capacity and resilience strategyhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2023-0178/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to develop an inclusive, multidisciplinary, flexible and organizationally adaptable safety risk management framework, including diversity management, that will be implemented to ensure safety is and remains at the desired level. If the number of incidents and potential incidents that could lead to accidents and their impact rates are to be reduced operationally and administratively, aviation safety risks and sources of risk must be better understood, sources of risk identified, and the safety risk management framework designed in an organization-specific and organization-wide sustainable way. At this point, it is necessary to draw the conceptual framework well and to define the boundaries of the concepts well. In this study, a framework model that can be adapted to the organization is proposed to optimize the management of risks and provide both efficient and effective resource allocation and organizational structure design in its operations and management functions. The qualitative research method – triple techniques – was deemed appropriate for this study, which aims to identify, examine, interpret and develop the situations of safety management models. In this context, document analysis, business process modeling technique and Delphi techniques from qualitative research methods were used via integration as the methodology of this research. To manage dynamic civil aviation management activities and business processes effectively and efficiently, the risk management process is the building block of the “Proposed Process Model” that supports the decision-making processes of aviation organizations and managers. This “Framework Conceptual Model” building block also helps build capacity and resilience by enabling continuous development, organizational learning, and flexible structuring. This research is limited to air transportation and aviation safety management issues. This research is limited specifically to a safety-based risk management framework for the aviation industry. This research may have social implications as source saving, optimum resource use and capacity building will make a contribution to society and add value besides operational and practical implementation. This research may contribute to more safe operations and functions in the aviation industry. Management and academia may gain considerable support from this research to manage their safety risks via a corporate-tailored risk management framework, both improving resilience and developing corporate capacity. With this model presented, decision-makers will have a guiding structure that can optimally manage the main risk types that may be encountered in the safety risk in the fields of suppliers, manufacturers, demand changes, logistics, information management, environmental, legal and regulatory. Existing studies in the literature are generally in the form of algorithms and cannot be used as a decision-making support tool. This model aims to fill the gap in the literature. In addition, added value may be created by applying this model to optimum management safety risks in the real aviation industry and its related sectors.Developing conceptual model for safety risk management in aviation as building capacity and resilience strategy
Ayse KUCUK YILMAZ, Konstantinos N. MALAGAS, Triant G. FLOURIS
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to develop an inclusive, multidisciplinary, flexible and organizationally adaptable safety risk management framework, including diversity management, that will be implemented to ensure safety is and remains at the desired level. If the number of incidents and potential incidents that could lead to accidents and their impact rates are to be reduced operationally and administratively, aviation safety risks and sources of risk must be better understood, sources of risk identified, and the safety risk management framework designed in an organization-specific and organization-wide sustainable way. At this point, it is necessary to draw the conceptual framework well and to define the boundaries of the concepts well. In this study, a framework model that can be adapted to the organization is proposed to optimize the management of risks and provide both efficient and effective resource allocation and organizational structure design in its operations and management functions.

The qualitative research method – triple techniques – was deemed appropriate for this study, which aims to identify, examine, interpret and develop the situations of safety management models. In this context, document analysis, business process modeling technique and Delphi techniques from qualitative research methods were used via integration as the methodology of this research.

To manage dynamic civil aviation management activities and business processes effectively and efficiently, the risk management process is the building block of the “Proposed Process Model” that supports the decision-making processes of aviation organizations and managers. This “Framework Conceptual Model” building block also helps build capacity and resilience by enabling continuous development, organizational learning, and flexible structuring.

This research is limited to air transportation and aviation safety management issues. This research is limited specifically to a safety-based risk management framework for the aviation industry. This research may have social implications as source saving, optimum resource use and capacity building will make a contribution to society and add value besides operational and practical implementation.

This research may contribute to more safe operations and functions in the aviation industry.

Management and academia may gain considerable support from this research to manage their safety risks via a corporate-tailored risk management framework, both improving resilience and developing corporate capacity. With this model presented, decision-makers will have a guiding structure that can optimally manage the main risk types that may be encountered in the safety risk in the fields of suppliers, manufacturers, demand changes, logistics, information management, environmental, legal and regulatory. Existing studies in the literature are generally in the form of algorithms and cannot be used as a decision-making support tool. This model aims to fill the gap in the literature. In addition, added value may be created by applying this model to optimum management safety risks in the real aviation industry and its related sectors.

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Developing conceptual model for safety risk management in aviation as building capacity and resilience strategy10.1108/AEAT-07-2023-0178Aircraft Engineering and Aerospace Technology2024-03-20© 2024 Emerald Publishing LimitedAyse KUCUK YILMAZKonstantinos N. MALAGASTriant G. FLOURISAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-2010.1108/AEAT-07-2023-0178https://www.emerald.com/insight/content/doi/10.1108/AEAT-07-2023-0178/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Autonomous flight performance optimization of fixed-wing unmanned aerial vehicle with morphing wingtiphttps://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2022-0262/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this paper is to improve autonomous flight performance of a fixed-wing unmanned aerial vehicle (UAV) via simultaneous morphing wingtip and control system design conducted with optimization, computational fluid dynamics (CFD) and machine learning approaches. The main wing of the UAV is redesigned with morphing wingtips capable of dihedral angle alteration by means of folding. Aircraft dynamic model is derived as equations depending only on wingtip dihedral angle via Nonlinear Least Squares regression machine learning algorithm. Data for the regression analyses are obtained by numerical (i.e. CFD) and analytical approaches. Simultaneous perturbation stochastic approximation (SPSA) is incorporated into the design process to determine the optimal wingtip dihedral angle and proportional-integral-derivative (PID) coefficients of the control system that maximizes autonomous flight performance. The performance is defined in terms of trajectory tracking quality parameters of rise time, settling time and overshoot. Obtained optimal design parameters are applied in flight simulations to test both longitudinal and lateral reference trajectory tracking. Longitudinal and lateral autonomous flight performances of the UAV are improved by redesigning the main wing with morphing wingtips and simultaneous estimation of PID coefficients and wingtip dihedral angle with SPSA optimization. This paper originally discusses the simultaneous design of innovative morphing wingtip and UAV flight control system for autonomous flight performance improvement. The proposed simultaneous design idea is conducted with the SPSA optimization and a machine learning algorithm as a novel approach.Autonomous flight performance optimization of fixed-wing unmanned aerial vehicle with morphing wingtip
Tugrul Oktay, Yüksel Eraslan
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this paper is to improve autonomous flight performance of a fixed-wing unmanned aerial vehicle (UAV) via simultaneous morphing wingtip and control system design conducted with optimization, computational fluid dynamics (CFD) and machine learning approaches.

The main wing of the UAV is redesigned with morphing wingtips capable of dihedral angle alteration by means of folding. Aircraft dynamic model is derived as equations depending only on wingtip dihedral angle via Nonlinear Least Squares regression machine learning algorithm. Data for the regression analyses are obtained by numerical (i.e. CFD) and analytical approaches. Simultaneous perturbation stochastic approximation (SPSA) is incorporated into the design process to determine the optimal wingtip dihedral angle and proportional-integral-derivative (PID) coefficients of the control system that maximizes autonomous flight performance. The performance is defined in terms of trajectory tracking quality parameters of rise time, settling time and overshoot. Obtained optimal design parameters are applied in flight simulations to test both longitudinal and lateral reference trajectory tracking.

Longitudinal and lateral autonomous flight performances of the UAV are improved by redesigning the main wing with morphing wingtips and simultaneous estimation of PID coefficients and wingtip dihedral angle with SPSA optimization.

This paper originally discusses the simultaneous design of innovative morphing wingtip and UAV flight control system for autonomous flight performance improvement. The proposed simultaneous design idea is conducted with the SPSA optimization and a machine learning algorithm as a novel approach.

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Autonomous flight performance optimization of fixed-wing unmanned aerial vehicle with morphing wingtip10.1108/AEAT-09-2022-0262Aircraft Engineering and Aerospace Technology2024-03-29© 2024 Emerald Publishing LimitedTugrul OktayYüksel EraslanAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-2910.1108/AEAT-09-2022-0262https://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2022-0262/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Fuel consumption estimation for jet aircraft: a mathematical model developmenthttps://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2023-0237/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestReducing aircraft fuel consumption has become a paramount research area, focusing on optimizing operational parameters like speed and altitude during the cruise phase. However, the existing methods for fuel reduction often rely on complex experimental calculations and data extraction from embedded systems, making practical implementation challenging. To address this, this study aims to devise a simple and accessible approach using available information. In this paper, a novel analytic method to estimate and optimize fuel consumption for aircraft equipped with jet engines is proposed, with a particular emphasis on speed and altitude parameters. The dynamic variations in weight caused by fuel consumption during flight are also accounted for. The derived fuel consumption equation was rigorously validated by applying it to the Boeing 737–700 and comparing the results against the fuel consumption reference tables provided in the Boeing manual. Remarkably, the equation yielded closely aligned outcomes across various altitudes studied. In the second part of this paper, a pioneering approach is introduced by leveraging the particle swarm optimization algorithm (PSO). This novel application of PSO allows us to explore the equation’s potential in finding the optimal altitude and speed for an actual flight from Algiers to Brussels. The results demonstrate that using the main findings of this study, including the innovative equation and the application of PSO, significantly simplifies and expedites the process of determining the ideal parameters, showcasing the practical applicability of the approach. The suggested methodology stands out for its simplicity and practicality, particularly when compared to alternative approaches, owing to the ready availability of data for utilization. Nevertheless, its applicability is limited in scenarios where zero wind effects are a prevailing factor. The research opens up new possibilities for fuel-efficient aviation, with a particular focus on the development of a unique fuel consumption equation and the pioneering use of the PSO algorithm for optimizing flight parameters. This study’s accessible approach can pave the way for more environmentally conscious and economical flight operations.Fuel consumption estimation for jet aircraft: a mathematical model development
Amar Benkhaled, Amina Benkhedda, Braham Benaouda Zouaoui, Soheyb Ribouh
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

Reducing aircraft fuel consumption has become a paramount research area, focusing on optimizing operational parameters like speed and altitude during the cruise phase. However, the existing methods for fuel reduction often rely on complex experimental calculations and data extraction from embedded systems, making practical implementation challenging. To address this, this study aims to devise a simple and accessible approach using available information.

In this paper, a novel analytic method to estimate and optimize fuel consumption for aircraft equipped with jet engines is proposed, with a particular emphasis on speed and altitude parameters. The dynamic variations in weight caused by fuel consumption during flight are also accounted for. The derived fuel consumption equation was rigorously validated by applying it to the Boeing 737–700 and comparing the results against the fuel consumption reference tables provided in the Boeing manual. Remarkably, the equation yielded closely aligned outcomes across various altitudes studied. In the second part of this paper, a pioneering approach is introduced by leveraging the particle swarm optimization algorithm (PSO). This novel application of PSO allows us to explore the equation’s potential in finding the optimal altitude and speed for an actual flight from Algiers to Brussels.

The results demonstrate that using the main findings of this study, including the innovative equation and the application of PSO, significantly simplifies and expedites the process of determining the ideal parameters, showcasing the practical applicability of the approach.

The suggested methodology stands out for its simplicity and practicality, particularly when compared to alternative approaches, owing to the ready availability of data for utilization. Nevertheless, its applicability is limited in scenarios where zero wind effects are a prevailing factor.

The research opens up new possibilities for fuel-efficient aviation, with a particular focus on the development of a unique fuel consumption equation and the pioneering use of the PSO algorithm for optimizing flight parameters. This study’s accessible approach can pave the way for more environmentally conscious and economical flight operations.

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Fuel consumption estimation for jet aircraft: a mathematical model development10.1108/AEAT-09-2023-0237Aircraft Engineering and Aerospace Technology2024-03-18© 2024 Emerald Publishing LimitedAmar BenkhaledAmina BenkheddaBraham Benaouda ZouaouiSoheyb RibouhAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-1810.1108/AEAT-09-2023-0237https://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2023-0237/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Hovering performance analysis of helicopter rotor blades using supercritical airfoilhttps://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2023-0244/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to find the characteristics of supercritical airfoil in helicopter rotor blades for hovering phase using numerical analysis and the validation using experimental results. Using numerical analysis in the forward phase of the helicopter, supercritical airfoil is compared with the conventional airfoil for the aerodynamic performance. The multiple reference frame method is used to produce the results for rotational analysis. A grid independence test was carried out, and validation was obtained using benchmark values from NASA data. From the analysis results, a supercritical airfoil in hovering flight analysis proved that the NASA SC rotor produces 25% at 5°, 26% at 12° and 32% better thrust at 8° of collective pitch than the HH02 rotor. Helicopter performance parameters are also calculated based on momentum theory. Theoretical calculations prove that the NASA SC rotor is better than the HH02 rotor. The results of helicopter performance prove that the NASA SC rotor provides better aerodynamic efficiency than the HH02 rotor. The novelty of the paper is it proved the aerodynamic performance of supercritical airfoil is performing better than the HH02 airfoil. The results are validated with the experimental values and theoretical calculations from the momentum theory.Hovering performance analysis of helicopter rotor blades using supercritical airfoil
Inamul Hasan, Mukesh R., Radha Krishnan P., Srinath R., Boomadevi P.
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to find the characteristics of supercritical airfoil in helicopter rotor blades for hovering phase using numerical analysis and the validation using experimental results.

Using numerical analysis in the forward phase of the helicopter, supercritical airfoil is compared with the conventional airfoil for the aerodynamic performance. The multiple reference frame method is used to produce the results for rotational analysis. A grid independence test was carried out, and validation was obtained using benchmark values from NASA data.

From the analysis results, a supercritical airfoil in hovering flight analysis proved that the NASA SC rotor produces 25% at 5°, 26% at 12° and 32% better thrust at 8° of collective pitch than the HH02 rotor. Helicopter performance parameters are also calculated based on momentum theory. Theoretical calculations prove that the NASA SC rotor is better than the HH02 rotor. The results of helicopter performance prove that the NASA SC rotor provides better aerodynamic efficiency than the HH02 rotor.

The novelty of the paper is it proved the aerodynamic performance of supercritical airfoil is performing better than the HH02 airfoil. The results are validated with the experimental values and theoretical calculations from the momentum theory.

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Hovering performance analysis of helicopter rotor blades using supercritical airfoil10.1108/AEAT-09-2023-0244Aircraft Engineering and Aerospace Technology2024-01-25© 2024 Emerald Publishing LimitedInamul HasanMukesh R.Radha Krishnan P.Srinath R.Boomadevi P.Aircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-01-2510.1108/AEAT-09-2023-0244https://www.emerald.com/insight/content/doi/10.1108/AEAT-09-2023-0244/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Viewing electromagnetic scattering characteristics on air-brake of a stealth planehttps://www.emerald.com/insight/content/doi/10.1108/AEAT-10-2023-0269/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThis study aims to learn the dynamic radar cross-section (RCS) of a deflection air brake. The aircraft model with delta wing, V-shaped tail and blended wing body is designed, and high-precision unstructured grid technology is used to deal with the surface of air brake and fuselage. The calculation method based on multiple tracking and dynamic scattering is presented to calculate RCS. The fuselage has a low scattering level, and the opening air brake will bring obvious dynamic RCS effects to itself and the whole machine. The average indicator of air brake RCS can be lower than –0.6 dBm2 under the tail azimuth, while that of forward and lateral direction is lower. The mean RCS of fuselage is obviously higher than that of air brake, while the deflected air brake and its cabin can still provide strong scattering sources at some azimuths. When the air brake is opening, the change amplitude of the aircraft forward RCS can exceed 19.81 dBm2. This research has practical significance for the dynamic electromagnetic scattering analysis and stealth design of the air brake. The calculation method for aircraft RCS considering air brake dynamic deflection has been established.Viewing electromagnetic scattering characteristics on air-brake of a stealth plane
Zeyang Zhou, Jun Huang
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

This study aims to learn the dynamic radar cross-section (RCS) of a deflection air brake.

The aircraft model with delta wing, V-shaped tail and blended wing body is designed, and high-precision unstructured grid technology is used to deal with the surface of air brake and fuselage. The calculation method based on multiple tracking and dynamic scattering is presented to calculate RCS.

The fuselage has a low scattering level, and the opening air brake will bring obvious dynamic RCS effects to itself and the whole machine. The average indicator of air brake RCS can be lower than –0.6 dBm2 under the tail azimuth, while that of forward and lateral direction is lower. The mean RCS of fuselage is obviously higher than that of air brake, while the deflected air brake and its cabin can still provide strong scattering sources at some azimuths. When the air brake is opening, the change amplitude of the aircraft forward RCS can exceed 19.81 dBm2.

This research has practical significance for the dynamic electromagnetic scattering analysis and stealth design of the air brake.

The calculation method for aircraft RCS considering air brake dynamic deflection has been established.

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Viewing electromagnetic scattering characteristics on air-brake of a stealth plane10.1108/AEAT-10-2023-0269Aircraft Engineering and Aerospace Technology2024-04-01© 2024 Emerald Publishing LimitedZeyang ZhouJun HuangAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-04-0110.1108/AEAT-10-2023-0269https://www.emerald.com/insight/content/doi/10.1108/AEAT-10-2023-0269/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited
Global fast terminal sliding mode control for automatic carrier landing with environmental disturbanceshttps://www.emerald.com/insight/content/doi/10.1108/AEAT-10-2023-0273/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatestThe purpose of this paper is to solve the challenging problem of automatic carrier landing with the presence of environmental disturbances. Therefore, a global fast terminal sliding mode control (GFTSMC) method is proposed for automatic carrier landing system (ACLS) to achieve safe carrier landing control. First, the framework of ACLS is established, which includes flight glide path model, guidance model, approach power compensation system and flight controller model. Subsequently, the carrier deck motion model and carrier air-wake model are presented to simulate the environmental disturbances. Then, the detailed design steps of GFTSMC are provided. The stability analysis of the controller is proved by Lyapunov theorems and LaSalle’s invariance principle. Furthermore, the arrival time analysis is carried out, which proves the controller has fixed time convergence ability. The numerical simulations are conducted. The simulation results reveal that the proposed method can guarantee a finite convergence time and safe carrier landing under various conditions. And the superiority of the proposed method is further demonstrated by comparative simulations and Monte Carlo tests. The GFTSMC method proposed in this paper can achieve precise and safe carrier landing with environmental disturbances, which has important referential significance to the improvement of ACLS controller designs.Global fast terminal sliding mode control for automatic carrier landing with environmental disturbances
Zhuoer Yao, Zi Kan, Daochun Li, Haoyuan Shao, Jinwu Xiang
Aircraft Engineering and Aerospace Technology, Vol. ahead-of-print, No. ahead-of-print, pp.-

The purpose of this paper is to solve the challenging problem of automatic carrier landing with the presence of environmental disturbances. Therefore, a global fast terminal sliding mode control (GFTSMC) method is proposed for automatic carrier landing system (ACLS) to achieve safe carrier landing control.

First, the framework of ACLS is established, which includes flight glide path model, guidance model, approach power compensation system and flight controller model. Subsequently, the carrier deck motion model and carrier air-wake model are presented to simulate the environmental disturbances. Then, the detailed design steps of GFTSMC are provided. The stability analysis of the controller is proved by Lyapunov theorems and LaSalle’s invariance principle. Furthermore, the arrival time analysis is carried out, which proves the controller has fixed time convergence ability.

The numerical simulations are conducted. The simulation results reveal that the proposed method can guarantee a finite convergence time and safe carrier landing under various conditions. And the superiority of the proposed method is further demonstrated by comparative simulations and Monte Carlo tests.

The GFTSMC method proposed in this paper can achieve precise and safe carrier landing with environmental disturbances, which has important referential significance to the improvement of ACLS controller designs.

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Global fast terminal sliding mode control for automatic carrier landing with environmental disturbances10.1108/AEAT-10-2023-0273Aircraft Engineering and Aerospace Technology2024-03-29© 2024 Emerald Publishing LimitedZhuoer YaoZi KanDaochun LiHaoyuan ShaoJinwu XiangAircraft Engineering and Aerospace Technologyahead-of-printahead-of-print2024-03-2910.1108/AEAT-10-2023-0273https://www.emerald.com/insight/content/doi/10.1108/AEAT-10-2023-0273/full/html?utm_source=rss&utm_medium=feed&utm_campaign=rss_journalLatest© 2024 Emerald Publishing Limited