Maximum Take-off Weight Research Articles

Automated Design Process of a Fixed Wing UAV Maximizing Endurance

Unmanned aerial vehicle (UAV) design necessitates significant effort in prototyping, testing, and design iterations. To reduce design time and improve wing performance, an automated design and optimization framework is proposed utilizing open-source software, including OpenVSP: VSPAERO & Parasite Drag Tool, XFOIL, and Python. This study presents a preliminary UAV wing design methodology, emphasizing weight estimation, drag analysis, stall prediction, and endurance optimization. The maximum takeoff weight of the UAV was calculated after estimating the empty weight using a linear regression from data from 20 existing similar UAVs. The wing and engine sizing were determined using the matching plot technique. A solver with low-fidelity models, combining the Vortex Lattice Method (VLM) and analytical expressions, was used to predict the drag coefficient and maximum lift coefficient of the designed wing. An optimization process using a genetic algorithm was applied to maximize endurance while satisfying requirements such as rate of climb, stall, and maximum speeds. The optimized wing was analyzed with computational fluid dynamics (CFD), and its aerodynamic characteristics were compared with those obtained using VLM and the suggested aerodynamic solver. According to the CFD results, the proposed aerodynamic solver estimated the drag coefficient at zero angle of attack with an error of 17.2% compared to 63.1% using the VLM classic method. The error on the maximum lift coefficient estimation was limited to 5.3%. In terms of optimization, the framework showed an increase in the endurance ratio of up to 2% compared to the Artificial Neural Network method coupled with XFLR5. The primary advantage of the suggested framework is the utilization of open-source software, giving a cost-effective and accessible solution for small and medium-sized startups to design and optimize UAVs to achieve mission objectives.

Концепция использования современных технических средств при проведении обследования помещений, ограничивающего конституционное право человека и гражданина на неприкосновенность жилища

The article provides a systematic analysis of current legislation, generally accepted judicial practice and various doctrinal approaches to the legal regulation of public inspection of premises, which restricts the constitutional right of a person and citizen to inviolability of the home, and suggests ways to improve current legislation aimed at eliminating gaps in the legal regulation of the procedure for conducting this operational search event. As a result of the study, special powers of employees of operational units were noted when drawing up a protocol for the inspection of a home and using video and audio recordings, film and photography, as well as other technical and other means during this operational search event. The paper describes the legal regime and classification of unmanned aerial vehicles depending on the maximum take-off weight and type of ownership. Taking into account modern scientific and technical achievements, it is proposed to use a quadcopter using the following innovative methods when conducting a housing survey: video shooting, panoramic, nodal and detailed photography, “online broadcast”, three-dimensional modeling, modeling using artificial intelligence, digital protocol compilation and methods: photo and (or) video shooting using infrared and ultraviolet detectors, video shooting using gas analysis detectors, X-ray photography. The use of modern technical means in conducting a home survey will ensure objectivity, increase efficiency and guarantee high efficiency of this operational investigative measures.

Granular Bait Applications for Management of Rangeland Grasshoppers Using a Remotely Piloted Aerial Application System

Rangeland grasshoppers are an endemic species that play an essential role in the rangeland ecosystem but can cause severe economic damage when populations reach outbreak levels. Remotely piloted aerial application systems (RPAASs) offer an alternative method to carry out aerial insecticide applications in relatively small areas. The objective of this study was to investigate the efficacy of a granular bait, 2% Sevin (with the active ingredient carbaryl), applied by an RPAAS. The bait was applied on four replicated 4.05-hectare (10-acre) plots at a rate of 2.27 kg/ha (5 lbs/acre) with an RPAAS on a private ranch in New Mexico. Applications resulted in a normalized population reduction of 70.32% ± 16.54% standard error of the mean (SEM) of bait-susceptible species. Although some of the observed reduction in population may be attributed to aging, the net effect was most likely due to the ingestion of bait based on field observations of rapid mortality after ingestion and other factors, like past experience with carbaryl bait treatments on grasshoppers. Plots required at least two flights due to the Federal Aviation Administration’s (FAA) maximum takeoff weight requirement for small RPAASs. Combined, these results indicate that RPAASs can provide treatment capabilities in relatively small areas, i.e., population hotspots, preferably before outbreak levels are reached.

Conversion of a Coaxial Rotorcraft to a UAV—Lessons Learned

A coaxial helicopter with a maximum take-off weight of 600 kg was converted to an unmanned aerial vehicle. A minimally invasive robotic actuator system was developed, which can be retrofitted onto the copilot seat of the rotorcraft in a short period of time to enable automatic flight. The automatic flight control robot includes electromechanical actuators, which are connected to the cockpit inceptors and control the helicopter. Most of the sensors and avionic components were integrated into the modular robotic system for faster integration into the rotorcraft. The mechanical design of the control system, the development of the robot control software, and the control system architecture are described in this paper. Furthermore, the multi-body simulation of the robotic system and the estimation of the linear low-order actuator models from hover-frame flight test data are discussed. The developed technologies in this study are not specific to a coaxial helicopter and can be applied to the conversion of any crewed flight vehicle with mechanical controls to unmanned or fly-by-wire. This agile development of a full-size flying test-bed can accelerate the testing of advanced flight control laws, as well as advanced air mobility-related functions.

Aerodynamic and Structural Design Procedures Supported by Fabrication and Flight Testing of a Small Unmanned Helicopter

In this work, typical design, production, and testing procedures for a small unmanned helicopter are explained and performed. In doing so, preliminary sizing of the helicopter and three main disciplines are conducted: aerodynamic analytical and numerical simulations, power calculations, and structure analysis assessment. First, a thorough survey is implemented to obtain the trends for the maximum take-off weight versus some design constraints such as rotor diameter, motor power, payload, and empty weight. Performance calculation results are obtained to figure out all aspects that correspond to the specified mission. The designed rotor geometry along with the aerodynamic characteristics and flight performance variables is then validated using the blade element theory and numerical simulations. Second, based on the power curves obtained for different flight regimes, an electric brushless motor is selected. The numerical simulations (Computational Fluid Dynamics) analysis is used to enhance the selection which implies that the motor power should be greater than 5.4 kW to overcome the drag forces. The motor power selection corresponds to a maximum rotor pitch angle of 15∘ and a maximum rotor speed of 1450 RPM. Then, the aerodynamic loads are used as an input for the structural analysis using one-way coupling of fluid–structure interaction (FSI) and consequently designing the internal structure of the blade. Eventually, the internal structure manufactured using carbon fiber-reinforced polymer (CFRP) by applying a combined technique between wet layup and compression molding. The blade is statically tested compared with numerical finite element model results. The fuselage structure along with hub and tail units is manufactured and assembled with the existing on-shelf components to examine the helicopter lift capability with different payloads up to 9 kg. The results show that the detailed design process is significant for manufacturing such blades and the helicopter is capable of lifting off the ground with various payloads depending on the rotor pitch angles (8∘, 12∘, and 15∘) at a constant rotor speed of 1450 RPM.

Aircraft Takeoff and Landing Weight Estimation from Surveillance Data

Aircraft weight estimation is a common problem facing researchers working with aircraft surveillance data. Although knowledge of an aircraft’s weight and thrust is required for many types of analyses, such as those evaluating aircraft acoustic noise, fuel burn, and emissions, these parameters are typically not available from surveillance sources. Instead, researchers generally only have access to basic aircraft states: lateral position, groundspeed, and altitude. Therefore, methods for estimating the weight of aircraft from these basic states become necessary in cases where aircraft performance is a key component of the analysis. This paper introduces two weight estimation models: one for the estimation of aircraft takeoff weight from departure data, and another for the estimation of aircraft landing weight from arrival data. The models are mathematically simple but grounded in knowledge of aircraft certification, airline operations, and aircraft flight management system logic. The landing weight estimation model proposed is shown to have a mean absolute error equivalent to 2.66% of maximum takeoff weight and a standard deviation of 3.35% of maximum takeoff weight when validated using onboard data recordings from 240 Airbus A320 flights. Similarly, the proposed takeoff weight estimation model is shown to have a mean absolute error of 2.83% of the maximum takeoff weight and a standard deviation of 3.55% of the maximum takeoff weight when applied to the same validation dataset.

Selection of unmanned aerial vehicle systems for border monitoring using the MPSI-SPOTIS method

PurposeThe present paper assesses the decision problem of selecting Unmanned Aerial Vehicle Systems (SARP) by the hybrid MPSI-SPOTIS approach for deployment in border control and transborder illicit combat.Design/methodology/approachBy the hybrid MCDA MPSI-SPOTIS approach, and from the database available in Gettinger (2019), models were filtered by Endurance, Range, Maximum Take-Off Weight (MTOW), and Payload, fitting within the classification of Categories EB 0 and 2. Category EB 1 was not considered in this study due to the limited number of models in the data source.FindingsThe use of the Multi-Criteria Decision Analysis (MCDA) tool MPSI-SPOTIS allowed the determination of weights by stochastic criteria, applied in a ranking method resistant to reverse ordering. The application of the method identified the Raybird-3 (Cat EB 0) and Searcher (Mk3) (Cat EB 2) models as the best alternatives. From a proposed clustering, other selection possibilities with close performance in the evaluation were presented. The cost criterion was not taken into consideration due to the absence of information in the data source employed. Future studies are suggested to include criteria related to the life cycle and acquisition cost of the models.Research limitations/implicationsThe cost criterion was not taken into consideration due to the absence of information in the data source used. Future studies are suggested to include criteria related to the life cycle and acquisition cost of the models.Originality/valueThis paper aims to propose a technology selection method applied to complex defense acquisitions when multiple factors influence the decision makers and it is hard to obtain a major optimum solution in multitask and multi-mission platform.

Aeronautical Infrastructure Optimization to Enhance the Strategic and Operational Business Performance of Small Airports

Airport business management literature commonly evaluates the performance of these organizations based on the number of passengers and cargo processed, aircraft movements, number of runways, number of employees, and passengers’ terminal dimension, among other factors concerning commercial aviation, commonly applied to large structures, with gaps being identified in the literature of small airports. The identification and analysis of the main factors that interfere with the performance of airport business, and the strategic and operational management of small public airports have been analyzed from 24 small Brazilian airports. The study employed the Data Envelopment Analysis method based on the Charnes, Cooper, and Rhodes model, as well as the Banker, Charnes, and Cooper model, adapted for Visual Basic Analysis. The main results demonstrate that infrastructural characteristics, such as runway width, pavement classification number, runway length, buldings' age, maximum take-off weight of the critical aircraft, passenger terminal size, and air cargo and passengers processed, show significant relevance to the operational performance of small airports. The main theoretical implications of the study are the definition of strategic and operational parameters that impact the performance of small airports, which can be used as a reference by air transport decision-makers.

Sensitivity Analysis for Design Parameters of Electric Tilt-Rotor Aircraft

In recent years, there has been rapid development in electric aircraft, particularly electric vertical takeoff and landing (eVTOL) aircraft, as part of efforts to promote green aviation. During the conceptual design stage, it is crucial to select appropriate values for key parameters and conduct sensitivity analysis on these parameters. This study focuses on an electric tilt-rotor aircraft and proposes a performance analysis method for electric aircraft while developing a general design tool specifically for this type of aircraft. Subsequently, the impact of wing incidence angle, sweep angle, span, propeller solidity, battery-specific energy, and battery mass on range, maximum takeoff weight, and hover power are analyzed. The results show that the battery mass, wingspan, and wingtip chord length have great effects on the maximum takeoff weight; among these, battery mass had the greatest influence. In terms of range, the battery energy density has a great positive effect on range, while the increase in wing angle of incidence, wingtip chord length and battery mass have some negative effects on range.

A Drone Based Image Dataset Generation Methodology for Single Image Super Resolution

The advancements in drone technologies, digital imaging, computer vision techniques, and the liberalized laws related to drone flying have opened up drone-based applications such as the delivery of supplies, search and rescue, aerial surveillance, and so on. The drones, especially the nano/micro/small drones, may be mounted with only low-resolution camera(s) due to their maximum takeoff weight limitations. The low-resolution images generated by the cameras, if used for landing, can result in faulty detection unless the photos are taken from a very close distance to the point of interest. Detection and recognition of the point(s) of interest as early as possible is required to ensure sufficient response time for safe maneuvering. Hence, the images are to be captured at greater heights or distances from the point(s) of interest, and obtaining the high-resolution images from the captured low-resolution images is crucial. The High Resolution (HR) and the Low Resolution (LR) image pairs for training super-resolution models in the works presented in literature are generated using two different cameras or the HR images are captured by the camera and LR images are generated by degrading the HR images. As both methods are not appropriate for small/micro/nano category drones, we propose a novel method based on Ground Sampling Distance (GSD) to capture the LR and HR images. In this paper, we have presented the designed methodology for the creation of a dataset using drone-mounted cameras covering a broad spectrum of views of the target(s) suitable for training and testing of the Single Image Super-Resolution (SISR) models. We also present a methodology for selecting an appropriate target for imaging that enables the visual quality assessment of the developed super-resolution model.

Design and Development of Quadcopter’s Frame

The purpose of this research is to obtain a design and prototype of quadcopter’s frame, which be developed further as a quadcopter that capable of carrying payload of 1.5 kg with flight time of more than 20 minutes. The design begins with estimating the values of Maximum Take-Off Weight (MTOW), required thrust, propeller dimensions, and wheelbase dimensions of the quadcopter. The results show the MTOW of the quadcopter was 5 kg, with each motor requiring of 2.5 kg of thrust per arm. The wheelbase dimension was 790 mm, and the propeller diameter of 17 inches. A wheelbase dimension was utilized to develop three conceptual frame designs for the quadcopter. The three conceptual designs were selected using a decision matrix table. The selected design was calculated for its structural strength by applying a load of 2.5 kg on each motor mounting. The results show the maximum stress value of 21.17 MPa, the maximum deflection of 3.5 mm, and safety factor of 22.44. Then, the prototype of quadcopter’s frame was manufactured. Therefore, the producing of prototype was measured the mass and deflection. Based on the measurements conducted, the quadcopter frame has an actual mass of 595 grams and a maximum deflection of 3.7 mm. The actual deflection value and the actual mass are close to the deflection and mass obtained from the calculation.

Design methodology and simulation analysis of hybrid fuel cell and battery systems for powering unmanned aircraft systems

Hybrid fuel cell and battery systems are currently among the most promising options for increasing the flight autonomy of unmanned aircraft systems. The ratio between the size of the battery and the fuel cell is a delicate issue in power system design. Both devices are necessary; the battery, with a quick response, to respond to demand peaks, and the fuel cell, with a slow dynamic response, but greater energy density, to increase its autonomy. One of the most important tasks in designing the hybrid system is selecting the appropriate size of each component so that it can respond to peak demands, extend its autonomy as far as possible and reduce its weight.The purpose of this research was to establish a design procedure for hybrid power systems for unmanned aerial systems, which means selecting the best size for each component. The hybrid system's modeling and the dynamic simulation of its response to demand curves are the foundation of this procedure. The simulation results include the consumption of the fuel cell as a function of time, the battery’s charge status and an estimate of its autonomy for each demand curve. The analysis of the simulation results guides the choice of component size, adapting each one to achieve minimum fuel consumption for the characteristic demand curves for which the unmanned aircraft system is designed. After modeling the components of the hybrid system and describing the simulation procedure, this contribution proposes an example of a design for an unmanned aircraft system with a maximum takeoff weight of 20 kg. In this case, the most appropriate size for the fuel cell was 2000 W and a 3000 mAh 10S Lithium-Polymer battery. Fitted with a 12-liter hydrogen tank, stored at a maximum pressure of 300 bar, an estimated autonomy of 1.5 h can be expected.

Multi-Physics Digital Model of an Aluminum 2219 Liquid Hydrogen Aircraft Tank

Future liquid hydrogen-powered aircraft requires the design and optimization of a large number of systems and subsystems, with cryogenic tanks being one of the largest and most critical. Considering previous space applications, these tanks are usually stiffened by internal members such as stringers, frames, and stiffeners resulting in a complex geometry that leads to an eventual reduction in weight. Cryogenic tanks experience a variety of mechanical and thermal loading conditions and are usually constructed out of several different materials. The complexity of the geometry and the loads highlights the necessity for a computational tool in order to conduct analysis. In this direction, the present work describes the development of a multi-physics finite element digital simulation, conducting heat transfer and structural analysis in a fully parametric manner in order to be able to support the investigation of different design concepts, materials, geometries, etc. The capabilities of the developed model are demonstrated by the design process of an independent-type aluminum 2219 cryogenic tank for commuter aircraft applications. The designed tank indicates a potential maximum take-off weight reduction of about 8% for the commuter category and demonstrates that aluminum alloys are serious candidate materials for future aircraft.

OPTIMIZATION INTERNATIONAL AIRPORT CAPACITY I GUSTI NGURAH RAI BALI

This study discuss about optimization capacity on land new to I Gusti Airport International Ngurah Rai, Bali with review evaluation capacity moment this . Development Airport Ngurah Rai consists of of 2 ( two ) stages . Stage One consists of from development international terminal capacity 189,000 meters square . For Step second , development in the form of international terminal expansion to 265,891 meters. Development process stage I requires reclamation land area of 47 ( four twenty seven ) hectares , and when it's 35 ( three twenty five) hectares has reclaimed . Whereas the rest still in the licensing process and targeted done mid December 2018. Development process Phase II will need reclamation land area of 71 ( seven twenty one ) hectare . Based on plan development Stages I and II are the addition of an apron covering an area of 47.9 hectares so need conducted calculation optimization capacity airport Ngurah Rai with consider adequate condition of the apron and runway given by PT Angkasa Pura I considering that condition expansion more carry on will difficult because cost development consequence crisis global pandemic and vulnerable damage surroundings _ _ waters airport that can damage area . Study this aim for knowing amount movement the plane that will wearing a new apron To use serve movement aircraft in 2022. Analysis _ started with needs runway length in Maximum Take - off Weight (MTOW) for type aircraft largest operating _ in 2019 (Boeing 747 – 400), next analyze amount runway capacity and prediction year 2022 in operation mix (Take – off/Landing), and do analysis linear regression for estimate amount movement aircraft optimally for _ 2022. Continued _ with determine score rmonth , Rday , and the later Rhour multiplied with score results regression for get score Nmonth , Nday , and Nhour . Nhour value in year planning . Analysis result show evaluation condition existing runway 3000 m to runway needs with aircraft Boeing 747 – 400ER plan , then results correct to influence height from sea level (Fe) is 3323.4 m, while correct to influence temperature air (Ft) obtained 3620.4 m. amount movement aircraft in 2022 results _ analysis operating runway capacity arrival just is 18 operations per hour, for operation departure just is 44 operations per hour, whereas for operation mixture is 58 operations per hour. The result of parking stand requirements is obtained in class of aircraft 3 C as many as 34 parking stands and 4C as many as 45 parking stands. From result determination new apron capacity in 2022 with _ total dimensions 122130.4 m 2 . Next recommended for conducted analysis more carry on in the form of arrangement of parking stands on land reclamation stages I and II. for 2022 .

Material Analysis of a Sample Airless Wheel That Can Be Used in The Landing Gear of Small UAVs for Smart Agriculture Using the Finite Element Method

Unmanned Aerial Vehicles (UAV) in categories 1 and 2 are small in size and have a maximum gross take-off weight between 5 - 40 N and 40 - 250 N, with a normal operating altitude of 120 m above ground level. One of the major concerns for UAVs is the weight of aerial vehicles, which is aided primarily by material changes. Tires for UAVs are primarily made of ABS, rubber silicone, and nylon to reduce flight problems. These tires add to the dead weight and drag during flight.
 In response to these issues, a rigorous analysis was performed to select the best possible configuration using various materials such as Acrylonitrile Butadiene Styrene (ABS), rubber silicone, and nylon materials. UAV wheel manufacturing is commonly known as the production of high-strength light-weighing parts. It also provides a broader range of design options and favors an iterative design approach, so it is important the method of production. To achieve the best results, the iterative design approach was used in the analysis application. The study incorporated a design for airless tires with 40 spokes, leveraging Autodesk Inventor Pro for modeling. Subsequently, the designs underwent Finite Element Analysis to assess static radial strength. Bending and torsion stresses during landing were deemed negligible owing to the lightweight nature of categories 1 and 2 UAVs and were thus excluded from the analysis.
 The modeled wheel design in Autodesk Inventor Pro analysis based on Finite Element Analysis of the airless wheels revealed that wheels made of rubber silicone exhibited superior shock absorption, with a deformation of only 2.314 mm at 625 N upon impact during UAV landing maneuvers.

Toward Estimating Noise–Power–Distance Curves for Propeller-Powered Zero-Emission Hydrogen Aircraft

As part of the UK Research and Innovation project New Aviation, Propulsion, Knowledge and Innovation Network (NAPKIN), a high-level framework was developed for the assessment of the noise impact of the proposed regional-sized hydrogen-powered aircraft. This study consists of the methodology used to generate the industry-standard noise–power–distance (NPD) curves from individual component noise analysis, specifically propeller tonal noise. The model is based on an asymptotic analysis of a frequency domain propeller tonal noise model combined with a linear approximation, taking advantage of the logarithmic nature of noise. An error analysis on the linear approximation assumption proves that the relative error between predicted and actual values of the noise remains below 10% for appropriately chosen baseline points. Verification of the framework was achieved through a bench-marking procedure that compared predictions of departure NPD curves for current technology regional aircraft against published ones over a range of operational power settings. Finally, departure and approach NPD predictions for three of the NAPKIN hydrogen concept aircraft are presented. Concepts featuring a larger, slower-rotating propeller with an increased number of blades relative to the reference aircraft showed benefits over the reference aircraft, despite, in some cases, increases in maximum takeoff weight.

Parametric Analysis for Hybrid–Electric Regional Aircraft Conceptual Design and Development

This paper proposes a conceptual analysis of the limitations related to the development (and integration) of hybrid–electric propulsion on regional transport aircraft, with the aim to identify a feasibility space for this innovative aircraft concept. Hybrid–electric aircraft have attracted the interest of aeronautical research as these have the potential to reduce fuel consumption and, thus, the related greenhouse gas emissions. Nevertheless, considering the development of such an aircraft configuration while keeping the constraints deriving from technological and/or operating aspects loose could lead to the analysis of concepts that are unlikely to be realised. In this paper, specifically to outline the boundaries constraining the actual development of such aircraft, the influence on overall aircraft design and performance of the main technological, operating, and design factors characterising the development of such a configuration is analysed and discussed at a conceptual level. Specifically, the current achievable gravimetric battery energy density (BED) is identified as the main limiting factor for the development of regional hybrid–electric aircraft, and a sensitivity analysis shows the correlation of this important technological parameter with aircraft performance in terms of both fuel consumption and energy efficiency. In this context, minimum technological development thresholds are therefore identified to enable the effective development of this type of aircraft; namely, a minimum of BED = 500 Wh/kg at battery pack level is identified as necessary to provide tangible benefits. From an operating point of view, flight distance is the most limiting design requirement, and a proper assessment of the design range is necessary if a hybrid–electric aircraft is to be designed to achieve lower emissions than the state of the art; flight ranges equal to or lower than 600 nm are to be considered for this type of aircraft. As a bridging of both of the previous constraints, a change in the design paradigm with respect to established practices for state-of-the-art aircraft is necessary. More specifically, penalisations in maximum take-off weight and overall aircraft energy efficiency may be necessary if the aim is to reduce direct in-flight consumption by means of integration of hybrid–electric powertrains.

Design and fabrication of a fixed-wing Unmanned Aerial Vehicle (UAV)

Unmanned Aerial Vehicles (UAVs) have been widely used both in military and civil across the world in recent years. Nevertheless, their design always involves complex design optimization variables and decisions. Therefore, this paper aims to guide through the designing, manufacturing, and testing of an electrically powered radio-controlled aircraft for achieving a take-off, cruise, safe landing and carrying the highest payload possible. The whole process involves several phases, design phase, structural analysis, performance analysis, materials used, manufacturing, and finally aircraft testing. The final aircraft was designed with an empty weight and maximum take-off weight of 15.43 Ibs and 33.07 Ibs respectively while the wingspan, cruising speed and maximum speed were 70.1 in., 46 ft/s and 78 ft/sec respectively with a total take-off distance of 100 ft.

Design optimization to minimize wake of wide-body transport aircraft

A cutting-edge sustainable airplane can improve performance and enhance profitability are the most significant future concerns for the aviation industry. This research aims to analyse and assess aircraft performance of mid-size transport aircraft through retrofit designs. An optimal design was employed using a non-linear optimization approach under two cases fuselage alone (case 1) and a combination of fuselage-wing-empennage (case 2). These two retrofit concepts were evaluated and compared with the baseline A310–200 aircraft. To compute the objective function for various cases, non-linear solver was used from MATLAB optimization tool box. Additionally, the parametric studies of aerodynamics, structural and operational performance analysis are carried out by using Advanced Aircraft Analysis (AAA) 5.0 software. Sensitivity analyses are performed to assess the overall effectiveness of the optimal design. The key findings of this paper obtained the following parameters as Maximum take-off Weight, empty weight, fuel weight, parasite drag coefficient, rate of climb, take-off distance and landing distance are improved by 0.36%, 26.02%, 10.12%, 4.48%, 27.24%, 13.13% and 9.2% respectively. At the end, the paper provides a clear statement on the managerial perceptions and future insights for the upcoming challenges towards sustainable design.

World experience in the use of unmanned aerial vehicles in agriculture

The study of unmanned aerial vehicles in agriculture is limited by regulations. Each country has its own restrictions in the form of flight altitude, maximum takeoff weight, and the required package of documents. The article discusses the main regulatory restrictions for the use of drones in various countries of the world.