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Journal articles on the topic 'Payload sizing'
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1
Haley, J. G., T. P. McCall, I. W. Maynard, and B. Chudoba. "A sizing-based approach to evaluate hypersonic demonstrators: demonstrator-carrier constraints." Aeronautical Journal 124, no. 1279 (April 17, 2020): 1318–49. http://dx.doi.org/10.1017/aer.2020.30.
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ABSTRACTThe objective of this study is to identify, evaluate, and provide recommendations towards the realisation of near-term hypersonic flight hardware through the consideration of carrier vehicle constraints. The current rush of available funds for hypersonic research cannot cause a program to ignore growth potential for future missions. The prior NB-52 carrier vehicles, famous for the X-15 and X-43A missions, are retired. Next generation hypersonic demonstrator requirements will necessitate a substitution of carrier vehicle capability. Flight vehicle configuration, technology requirements, and recommendations are arrived at by constructing and evaluating a hypersonic technology demonstrator design matrix. This multi-disciplinary parametric sizing investigation of hypersonic vehicle demonstrators focuses on the evaluation of the combined carrier platform, booster, and hypersonic cruiser solution space topography. Promising baseline configurations are evaluated against operational requirements by trading fuel type, endurance cruise time, and payload weight. The multi-disciplinary study results are constrained with carrier payload mass and geometry limitations. The multi-disciplinary results provide physical insights into near-term hypersonic demonstrator payload and cruise time requirements that will stretch the capability of existing carrier aircraft. Any growth in hypersonic research aircraft size or capability will require new carrier vehicle investments.
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Avanzini, Giulio, Emanuele L. de Angelis, Fabrizio Giulietti, and Edmondo Minisci. "Optimal Sizing of Electric Multirotor Configurations." MATEC Web of Conferences 233 (2018): 00028. http://dx.doi.org/10.1051/matecconf/201823300028.
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A sizing tool for the definition of the configuration of electrically powered multirotor platforms is developed, which accounts for a realistic battery discharge model. The tool is developed to provide the community with the possibility of deriving the best configuration for performing a given task, while accounting for specific constraints and performance requirements. An evolutionary algorithm is used for searching the design space and to identify feasible designs with optimal performance in terms of maximum hovering time on the target and payload weight fraction.
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ONEL, Alexandru-Iulian, and Teodor-Viorel CHELARU. "Weights and sizing assessment in the context of small launcher design." INCAS BULLETIN 12, no. 3 (September 1, 2020): 137–50. http://dx.doi.org/10.13111/2066-8201.2020.12.3.11.
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The paper presents mathematical models that can be used to quickly define preliminary key aspects regarding the sizing and weight characteristics of studied small launchers. The tool developed based on the proposed mathematical models can be used for standalone liquid propelled stage design or it can be integrated in an iterative multidisciplinary optimisation design scheme (MDO) for a preliminary small launcher design, able to insert the desired payload into a predefined orbit.
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Kumar, A., S. C. Sati, and A. K. Ghosh. "Design, Testing, and Realisation of a Medium Size Aerostat Envelope." Defence Science Journal 66, no. 2 (March 23, 2016): 93. http://dx.doi.org/10.14429/dsj.66.9291.
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<p>The design, testing and realisation aspects during the development of a medium size aerostat envelope in the present work. The payload capacity of this aerostat is 300 kg at 1 km above mean sea level. The aerostat envelope is the aerodynamically shaped fabric enclosure part of the aerostat which generally uses helium for lifting useful payloads to a specified height. The envelope volume estimation technique is discussed which provides the basis for sizing. The design, material selection, testing and realisation aspects of this aerostat envelope are also discussed. The empirical formulas and finite element analysis are used to estimate the aerodynamic, structural and other design related parameters of the aerostat. Equilibrium studies are then explained for balancing forces and moments in static conditions. The tether profile estimation technique is discussed to estimate blow by distance and tether length. A comparison of estimated and measured performance parameters during trials has also been discussed.</p>
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Chávez, Javier Enrique Orna, Otto Fernando Balseca Sampedro, Jorge Isaías Caicedo Reyes, Diego Fernando Mayorga Pérez, Edwin Fernando Viteri Núñez, and Catalina Margarita Verdugo Bernal. "Análisis Y Diseño De Una Aeronave No Tripulada Para Uso Agrícola." European Scientific Journal, ESJ 13, no. 6 (February 28, 2017): 135. http://dx.doi.org/10.19044/esj.2017.v13n6p135.
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The present research proposes to diversify the use of unmanned aerial vehicles (UAV) with rotating wings for applications in the agricultural sector. For this we have analyzed and designed an unmanned aircraft. In first instance the applications of this type of aircraft in this sector were reviewed to determine possible design conditions that would aid in the sizing and design of the aircraft. Once the requirements had been determined, aerodynamic analysis was carried out to size up and launch the required power output for the craft. This in order to optimize the weight and autonomous fight time to finally design an aircraft prototype built in carbon fiber with the aid of fault theories as applied to composite materials. At the end of the research, an unmanned aircraft of 6 rotors, each with an installed power supply of 700W was designed. The aircraft has an autonomous flight time of 40 minutes without a payload, 20 minutes with a payload of 3Kg, and 8 minutes with a payload of 5Kg. The commercial application of these aircraft are the monitoring of land and fumigation in inaccessible areas.
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Patil, Ankur S., and Emily J. Arnold. "Sensor-Driven Preliminary Wing Ground Plane Sizing Approach and Applications." International Journal of Aerospace Engineering 2018 (July 2, 2018): 1–15. http://dx.doi.org/10.1155/2018/6378635.
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Structurally integrated antenna arrays provide synergies allowing the integration of large apertures onto airborne platforms. However, the surrounding airframe can greatly impact the performance of the antenna array. This paper presents a sensor-driven preliminary wing ground plane sizing approach to provide insight into the implications of design decisions on payload performance. The improvement of a wing-integrated antenna array that utilizes the wing as a ground plane motivated this study. Relationships for wing span, wing chord, and thickness are derived from extensive parametric electromagnetic simulations based on optimum antenna performance. It is expected that these equations would be used after an initial wing-loading design point has been selected to provide the designer guidance into how various wing parameters might affect the integrated antenna performance.
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Behroo, Mahan, Afshin Banazadeh, and Andisheh Rahimi Golkhandan. "Design Methodology and Preliminary Sizing of an Unmanned Mars Exploration Plane (UMEP)." Applied Mechanics and Materials 332 (July 2013): 15–20. http://dx.doi.org/10.4028/www.scientific.net/amm.332.15.
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This paper discusses the mission requirements and design constraints for an Unmanned Martian research aircraft based on a tailor-made classical airplane design methodology. First, the exploration mission is described using the information from previous real-world experiences and the desired payload is proposed accordingly. The environmental conditions that dictate severe constraints to the design space are characterized afterwards. The conventional airplane design cycle is modified to address the lack of statistical data and to define a proper design recycling criteria. Eventually, the outcome is presented in the form of a novel configuration that is well suited to carry out the specified exploration mission, flying low and slow over the Martian surface.
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da Silva, José Roberto Cândido, and Gefeson Mendes Pacheco. "An Extended Methodology for Sizing Solar Unmanned Aerial Vehicles: Theory and Development of a Python Framework for Design Assist." Sensors 21, no. 22 (November 12, 2021): 7541. http://dx.doi.org/10.3390/s21227541.
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There is a growing interest in using unmanned aerial vehicles (UAVs) in the most diverse application areas from agriculture to remote sensing, that determine the need to project and define mission profiles of the UAVs. In addition, solar photovoltaic energy increases the flight autonomy of this type of aircraft, forming the term Solar UAV. This study proposes an extended methodology for sizing Solar UAVs that take off from a runway. This methodology considers mission parameters such as operating location, altitude, flight speed, flight endurance, and payload to sizing the aircraft parameters, such as wingspan, area of embedded solar cells panels, runway length required for takeoff and landing, battery weight, and the total weight of the aircraft. Using the Python language, we developed a framework to apply the proposed methodology and assist in designing a Solar UAV. With this framework, it was possible to perform a sensitivity analysis of design parameters and constraints. Finally, we performed a simulation of a mission, checking the output parameters.
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Sridharan, Ananth, Bharath Govindarajan, and Inderjit Chopra. "A Scalability Study of the Multirotor Biplane Tailsitter Using Conceptual Sizing." Journal of the American Helicopter Society 65, no. 1 (January 1, 2020): 1–18. http://dx.doi.org/10.4050/jahs.65.012009.
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This paper presents a methodology for preliminary sizing of unconventional rotorcraft using a physics-based approach to estimate the weight of primary load-carrying members and rotor efficiencies. The methodology is demonstrated for a quadrotor biplane tailsitter, a tilt-body configuration that can operate in both helicopter and airplane mode. A beam lattice framework for the airframe structure is iteratively adjusted in the sizing loop to accommodate the limit loads. A similar semianalytical approach is followed to size and estimate weight of the rotor blades. Using this analysis, a consistent combination of vehicle macrodimensions (rotor radius, wing span) and tip speed as well as detailed design parameters (spar height, skin thickness, and cross-section weight) are obtained simultaneously. To compare the effectiveness of various power plants within a weight class, the sizing methodology was modified to identify the payload for three different vehicle takeoff weights: 20, 50, and 1000 lb. To enable operation within constrained urban canyons, the effect of restricting maximum vehicle dimensions to 10 ftfor the 1000-lb designs is also examined. An electric transmission model is used in these designs owing to its relative insensitivity of transmission efficiency to the operating RPM. A variable-pitch and variable-RPM rotor design allows for control redundancy within each rotor.
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Rajendran, Parvathy, and Howard Smith. "Development of Design Methodology for a Small Solar-Powered Unmanned Aerial Vehicle." International Journal of Aerospace Engineering 2018 (2018): 1–10. http://dx.doi.org/10.1155/2018/2820717.
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Existing mathematical design models for small solar-powered electric unmanned aerial vehicles (UAVs) only focus on mass, performance, and aerodynamic analyses. Presently, UAV designs have low endurance. The current study aims to improve the shortcomings of existing UAV design models. Three new design aspects (i.e., electric propulsion, sensitivity, and trend analysis), three improved design properties (i.e., mass, aerodynamics, and mission profile), and a design feature (i.e., solar irradiance) are incorporated to enhance the existing small solar UAV design model. A design validation experiment established that the use of the proposed mathematical design model may at least improve power consumption-to-take-off mass ratio by 25% than that of previously designed UAVs. UAVs powered by solar (solar and battery) and nonsolar (battery-only) energy were also compared, showing that nonsolar UAVs can generally carry more payloads at a particular time and place than solar UAVs with sufficient endurance requirement. The investigation also identified that the payload results in the highest effect on the maximum take-off weight, followed by the battery, structure, and propulsion weight with the three new design aspects (i.e., electric propulsion, sensitivity, and trend analysis) for sizing consideration to optimize UAV designs.
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Leuveano, Ab Rahman, Mahmood, and Saleh. "Integrated Vendor–Buyer Lot-Sizing Model with Transportation and Quality Improvement Consideration under Just-in-Time Problem." Mathematics 7, no. 10 (October 11, 2019): 944. http://dx.doi.org/10.3390/math7100944.
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This paper deals with the problem of transportation and quality within a Just-in-Time (JIT) inventory replenishment system. Formerly, transportation and quality problem are often modelled separately in most integrated inventory lot-sizing models. Hence, this paper develops an integrated vendor-buyer lot-sizing model by considering transportation and quality improvements into a JIT environment. The model is developed for minimising a total vendor–buyer system cost by optimising decisions such as delivery quantity, production batch, number of shipments, and process quality. Numerical examples and sensitivity analysis are provided to illustrate the proposed model. The developed model was also compared with an enumeration method to analyse the effectiveness of the proposed model to find the optimum solution. The results emphasise that the proposed model contributes to a new approach and obtains a near optimum solution for inventory replenishment decisions. The results are also beneficial to JIT practices as the model can improve the transport payload and reduce the chance of defective products and improving quality-related costs.
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Chudoba, B., G. Coleman, L. Gonzalez, E. Haney, A. Oza, and V. Ricketts. "Orbital transfer vehicle (OTV) system sizing study for manned GEO satellite servicing." Aeronautical Journal 120, no. 1226 (April 2016): 573–99. http://dx.doi.org/10.1017/aer.2016.3.
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ABSTRACTIn an effort to quantify the feasibility of candidate space architectures for astronauts servicing Geosynchronous Earth Orbit (GEO) satellites, a conceptual assessment of architecture-concept and operations-technology combinations has been performed. The focus has been the development of a system with the capability to transfer payload to and from geostationary orbit. Two primary concepts of operations have been selected: (a) Direct insertion/re-entry (Concept of Operations 1 – CONOP 1); (b) Launch to low-earth orbit at Kennedy Space Center inclination angle with an orbital transfer to/from geostationary orbit (Concept of Operations 2 – CONOP 2). The study concludes that a capsule and de-orbit propulsion module system sized for the geostationary satellite servicing mission is feasible for a direct insertion/re-entry concept of operation CONOP 1. Vehicles sized for CONOP 2 show overall total mass savings when utilising the aero-assisted orbital transfer vehicle de-orbit propulsion module options compared to the pure propulsive baseline cases. Overall, the consideration of technical, operational and cost factors determine if either the aero-assisted orbital transfer vehicle concepts or the re-usable/expendable ascent/de-orbit propulsion modules is the preferred option.
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Zong, Jianan, Bingjie Zhu, Zhongxi Hou, Xixiang Yang, and Jiaqi Zhai. "Sizing and Mission Profile Analysis of the Hybrid-Electric Propulsion System for Retrofitting a Fixed Wing VTOL Aircraft." International Journal of Aerospace Engineering 2022 (February 7, 2022): 1–10. http://dx.doi.org/10.1155/2022/9384931.
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Hybrid-electric technology can be expected to improve the performance of fixed wing vertical takeoff and landing (VTOL) aircraft. In this paper, we demonstrated a method of retrofitting a single-energy propulsion system prototype with a hybrid-electric propulsion system. Since the hybrid-electric system has several working modes, the optimal design results have strong coupling with mission performance. Therefore, we propose an analysis method of the mission profile to determine the design point. Finally, the payload-range sensitivity is studied. The results show that the hybrid-electric propulsion system can greatly increase the mission profile of aircraft. The analysis method of the mission profile also provides perspective for the hybrid-electric propulsion system design.
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DİNÇ, Ali. "SIZING OF A TURBOPROP ENGINE POWERED HIGH ALTITUDE UNMANNED AERIAL VEHICLE AND IT`S PROPULSION SYSTEM FOR AN ASSUMED MISSION PROFILE IN TURKEY." First Issue of 2019, no. 2019.01 (December 18, 2019): 19–23. http://dx.doi.org/10.23890/ijast.2019.0103.
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ABSTRACT In this study, preliminary sizing of a turboprop engine powered high altitude unmanned aerial vehicle and it`s propulsion system for an assumed mission profile in Turkey was performed. Aircraft mission profile is one of the most important design inputs in aircraft design. While the aircraft is dimensioned according to the requirements in the specification (useful payload, range, target cost, etc.), parameters such as cruise altitude and speed within the mission profile affect the engine type, power level, fuel quantity, and therefore the overall dimensions and total weight of the aircraft. The unmanned aerial vehicle with turboprop engine investigated in this study, can stay in the air for at least 24 hours at high altitude (40000 ft) and can be used for border surveillance, coast control, forest fires and land exploration.
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Dinç, Ali. "Sizing of a Turboprop Engine Powered High Altitude Unmanned Aerial Vehicle and It`s Propulsion System for an Assumed Mission Profile in Turkey." International Journal of Aviation Science and Technology vm01, is01 (September 10, 2020): 5–8. http://dx.doi.org/10.23890/ijast.vm01is01.0101.
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In this study, preliminary sizing of a turboprop engine powered high altitude unmanned aerial vehicle and it`s propulsion system for an assumed mission profile in Turkey was performed. Aircraft mission profile is one of the most important design inputs in aircraft design. While the aircraft is dimensioned according to the requirements in the specification (useful payload, range, target cost, etc.), parameters such as cruise altitude and speed within the mission profile affect the engine type, power level, fuel quantity, and therefore the overall dimensions and total weight of the aircraft. The unmanned aerial vehicle with turboprop engine investigated in this study, can stay in the air for at least 24 hours at high altitude (40000 ft) and can be used for border surveillance, coast control, forest fires and land exploration.
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Schwinn, Dominik B., P. Weiand, and M. Buchwald. "Structural Sizing of a Rotorcraft Fuselage Using an Integrated Design Approach." Journal of the American Helicopter Society 65, no. 4 (October 1, 2020): 1–12. http://dx.doi.org/10.4050/jahs.65.042008.
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For many years, the primary design objective of new helicopters was the design of the main rotor(s). Within the last couple of years, this approach has changed into an assessment of all helicopter components as an overall system, thus turning rotorcraft design into a highly interdisciplinary process. For instance, aerodynamics, flight mechanics, and the structural evaluation strongly affect each other, and these mutual influences are taken into account from the early phase of the conceptual design. Weight prediction in early design stages represents an essential part of the design process as it determines the basic properties of the rotorcraft. Owing to its function to carry crew and payload but also to serve as the central mounting for all components, the fuselage represents a major part of the rotorcraft. Therefore, the structural design of the fuselage airframe constitutes a significant factor of the rotorcraft design at the preliminary level.<br/> In this paper, an approach to include a higher fidelity method using finite elements for the structural analysis of rotorcraft fuselages within an integrated design environment is presented. Model generation and static analysis are conducted automatically. The helicopter is described using a common parametric data model during the complete design process, therefore providing a fast analysis of model changes. The generic finite element model presented in this paper was generated and structurally sized in about 2.5 min using a standard office computer, thus offering the integration of higher fidelity methods into early design sizing loops.
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Ng, Wanyi, Mrinalgouda Patil, and Anubhav Datta. "Hydrogen Fuel Cell and Battery Hybrid Architecture for Range Extension of Electric VTOL (eVTOL) Aircraft." Journal of the American Helicopter Society 66, no. 1 (January 1, 2021): 1–13. http://dx.doi.org/10.4050/jahs.66.012009.
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The objective of this paper is to study the impact of combining hydrogen fuel cells with lithium-ion batteries through an ideal power-sharing architecture to mitigate the poor range and endurance of battery powered electric vertical takeoff and landing (eVTOL) aircraft. The benefits of combining the two sources is first illustrated by a conceptual sizing of an electric tiltrotor for an urban air taxi mission of 75 mi cruise and 5 min hover. It is shown that an aircraft of 5000–6000 lb gross weight can carry a practical payload of 500 lb (two to three seats) with present levels of battery specific energy (150 Wh/kg) if only a battery–fuel cell hybrid power plant is used, combined in an ideal power-sharing manner, as long as high burst C-rate batteries are available (4–10 C). A power plant using batteries alone can carry less than half the payload; use of fuel cells alone cannot lift off the ground. Next, the operation of such a system is demonstrated using systematic hardware testing. The concepts of unregulated and regulated power-sharing architectures are described. A regulated architecture that can implement ideal power sharing is built up in a step-by-step manner. It is found only two switches and three DC-to-DC converters are necessary, and if placed appropriately, are sufficient to achieve the desired power flow. Finally, a simple power system model is developed, validated with test data and used to gain fundamental understanding of power sharing.
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Khalil, Mostafa, Anwer Hashish, and Hamed M. Abdalla. "A preliminary multidisciplinary design procedure for tactical missiles." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 9 (September 4, 2018): 3445–58. http://dx.doi.org/10.1177/0954410018797882.
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During missile system development, multidisciplinary design procedure is iteratively implemented based on the missile objective and target nature including internal ballistic, warhead function, and airframe configuration. By applying missile preliminary design, a good estimation for different design parameters can be obtained which will be useful through further detail design process. The aim of this paper is to build a preliminary design procedure for an unguided tactical missile that uses single-stage solid propellant motor to deliver a defined payload mass to a desired ground range. Based on data of available similar mature missile systems, two empirical formulas are developed to serve in the initial sizing of the missile with consideration of slenderness ratio, warhead mass, and desired ground range. Two different design concepts are implemented for tubular and star grains with different propellant compositions and chamber filling coefficients while the body-alone airframe configuration is adopted. The results demonstrate the capability of the proposed design procedure in defining the detailed design parameters. The impact of changing the propellant compositions and chamber filling coefficients on the obtained ground range is also explored.
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Billingsley, Ethan, Mehdi Ghommem, Rui Vasconcellos, and Abdessattar Abdelkefi. "On the Aerodynamic Analysis and Conceptual Design of Bioinspired Multi-Flapping-Wing Drones." Drones 5, no. 3 (July 18, 2021): 64. http://dx.doi.org/10.3390/drones5030064.
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Many research studies have investigated the characteristics of bird flights as a source of bioinspiration for the design of flapping-wing micro air vehicles. However, to the best of the authors’ knowledge, no drone design targeted the exploitation of the aerodynamic benefits associated with avian group formation flight. Therefore, in this work, a conceptual design of a novel multi-flapping-wing drone that incorporates multiple pairs of wings arranged in a V-shape is proposed in order to simultaneously increase the propulsive efficiency and achieve superior performance. First, a mission plan is established, and a weight estimation is conducted for both 3-member and 5-member configurations of the proposed air vehicle. Several wing shapes and airfoils are considered, and aerodynamic simulations are conducted, to determine the optimal planform, airfoil, formation angle, and angle of attack. The simulation results reveal that the proposed bioinspired design can achieve a propulsive efficiency of 73.8%. A stability analysis and tail sizing procedure are performed for both 3-member and 5-member configurations. In addition, multiple flapping mechanisms are inspected for implementation in the proposed designs. Finally, the completed prototypes’ models of the proposed multi-flapping-wing air vehicles are presented, and their features are discussed. The aim of this research is to provide a framework for the conceptual design of bioinspired multi-flapping-wing drones and to demonstrate the sizing, weight estimation, and design procedures for this new type of air vehicles. This work establishes the first multi-flapping-wing drone design which exploits the aerodynamic features of the V-formation flight observed in birds to achieve superior performance in terms of payload and endurance.
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Centracchio, Francesco, Monica Rossetti, and Umberto Iemma. "Approach to the Weight Estimation in the Conceptual Design of Hybrid-Electric-Powered Unconventional Regional Aircraft." Journal of Advanced Transportation 2018 (October 17, 2018): 1–15. http://dx.doi.org/10.1155/2018/6320197.
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The present work deals with the development of an innovative approach to the weight estimation in the conceptual design of a Hybrid-Electric-Powered (HEP) Blended Wing Body (BWB) commercial aircraft. In the last few decades, the improvement of the environmental impact of civil aviation has been the major concern of the aeronautical engineering community, in order to guarantee the sustainable development of the system in presence of a constantly growing market demand. The sustained effort in the improvement of the overall efficiency of conventional aircraft has produced a new generation of vehicles with an extremely low level of emissions and noise, capable of covering the community requirements in the short term. Unfortunately, the remarkable improvements achieved represent the asymptotic limit reachable through the incremental enhancement of existing concepts. Any further improvement to conform to the strict future environmental target will be possible only through the introduction of breakthrough concepts. The aeronautical engineering community is thus concentrating the research on unconventional airframes, innovative low-noise technologies, and alternative propulsion systems. The BWB is one of the most promising layouts in terms of noise emissions and chemical pollution. The further reduction of fuel consumption that can be achieved with gas/electric hybridisation of the power-plant is herein addressed in the context of multidisciplinary analyses. In particular, the payload and range limits are assessed in relation to the technological development of the electric components of the propulsion system. The present work explores the potentialities of an energy-based approach for the initial sizing of a HEP unconventional aircraft in the early conceptual phase of the design. A detailed parametric analysis has been carried out to emphasise how payload, range, and degree of hybridisation are strictly connected in terms of feasible mission requirements and related to the reasonable expectations of development of electric components suitable for aeronautical applications.
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Sahu, Dr Anil. "Design and Development of Multipurpose Delta Robot." International Journal for Research in Applied Science and Engineering Technology 9, no. 8 (August 31, 2021): 1471–75. http://dx.doi.org/10.22214/ijraset.2021.37608.
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Abstract: This report represents an designing and simulating ideal pick and place robot which should carry out the operations in minimum time and should also be cost efficient. It is four degrees of freedom parallel configuration used for very high speed pick and place operations. The objectives of this report are designing a Delta robot capable of carrying 1kg payload, achieving a cycle rate of 120 cycles per minute covering a work volume of 400x300x200 mm3. The project involves Kinematic & Dynamic modeling of the robot for the above specifications. The kinematic parameters, involving the lengths of the bicep and forearm, are calculated based on the work volume requirements and the dynamic parameters, involving the motor torque and speed, are calculated based on the maximum acceleration requirements and the inertia of the system. The project further involves the structural analysis of the robot which deals with the proper sizing of the mechanical structure which should be capable of withstanding the high torque and acceleration required for smooth and fast motion. The future work involves integrating the mechanical system with the control system and programming the system for a particular application
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Akash, Arumugam, Vijayaraj Stephen Joseph Raj, Ramesh Sushmitha, Boga Prateek, Sankarasubramanian Aditya, and Veloorillom Madhavan Sreehari. "Design and Analysis of VTOL Operated Intercity Electrical Vehicle for Urban Air Mobility." Electronics 11, no. 1 (December 22, 2021): 20. http://dx.doi.org/10.3390/electronics11010020.
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This paper discusses the conceptual design of an intercity electrical vertical take-off-and-landing aircraft. A literature survey of existing eVTOL aircrafts, configuration selection, initial sizing, weight estimation, modelling and analysis was conducted. The present intercity eVTOL aircraft has the capability to carry four passengers along with one pilot for a distance of 500 km. Two specific aircraft modes, such as air-taxi and air-cargo mode, are considered in the present design. Market entry is predicted before 2031. Subsequently, innovative technologies are incorporated into the design. The present design features an aerodynamically shaped fuselage, tapered wing and a V-tail design. It can carry a nominal payload of 500 kg to a maximum range of 500 km at a cruise speed of Mach 0.168. The present eVTOL is comprised of a 5 m-long fuselage and an 11 m wingspan. It utilizes six tilt-rotor propeller engines. The maximum take-off weight and empty weight are 1755 kg and 1255 kg, respectively. The unit price is expected to be between USD 14.83 and 17.36 million. This aircraft has an aesthetically pleasing, intelligent and feasible design.
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Chen, Shenyan, Zihan Yang, Minxiao Ying, Yanwu Zheng, Yanjie Liu, and Zhongwen Pan. "Parallel Load-Bearing and Damping System Design and Test for Satellite Vibration Suppression." Applied Sciences 10, no. 4 (February 24, 2020): 1548. http://dx.doi.org/10.3390/app10041548.
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The traditional series-type satellite vibration suppression scheme significantly decreases satellite frequency, which leads to difficulty in controlling the amplitude. In the present work, a new parallel viscous damping scheme is adopted on the Payload Adaptor Fitting (PAF), which aims to integrate a load-bearing design and vibration reduction. The vibration amplitude and weight are the most important design requirements of the damping system. The Finite Element (FE) model of PAF was established. Through a series of analyses, the appropriate number and coefficient of dampers were determined. The damping force was calculated according to the damping coefficient and the relative velocity between the two ends of the damper. Based on the damping force and the installation dimensions, the damping rod was designed. The force–velocity test was carried out on the damping rod prototype, which showed its performance satisfies the requirements. With the topology optimization and sizing optimization technology, the light-weight supports were designed and manufactured. One damping rod and two supports were assembled as one set of dampers. Eight sets of dampers were installed on the PAF. Vibration tests were conducted on the damping state PAF. The results showed that the proposed system is effective at suppressing vibration and maintaining stiffness simultaneously.
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Rompokos, Pavlos, Sajal Kissoon, Ioannis Roumeliotis, Devaiah Nalianda, Theoklis Nikolaidis, and Andrew Rolt. "Liquefied Natural Gas for Civil Aviation." Energies 13, no. 22 (November 13, 2020): 5925. http://dx.doi.org/10.3390/en13225925.
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The growth in air transport and the ambitious targets in emission reductions set by advisory agencies are some of the driving factors behind research towards new fuels for aviation. Liquefied Natural Gas (LNG) could be both environmentally and economically beneficial. However, its implementation in aviation has technical challenges that needs to be quantified. This paper assesses the application of LNG in civil aviation using an integrated simulation and design framework, including Cranfield University’s aircraft performance tool, Orion, and engine performance simulation tool Turbomatch, integrated with an LNG tank sizing module and an aircraft weight estimation module. Changes in tank design, natural gas composition, airframe changes, and propulsion system performance are assessed. The performance benefits are quantified against a Boeing 737–800 aircraft. Overall, LNG conversion leads to a slightly heavier aircraft in terms of the operating weight empty (OWE) and maximum take-off weight (MTOW). The converted aircraft has a slightly reduced range compared to the conventional aircraft when the maximum payload is considered. Compared to a conventional aircraft, the results indicate that although the energy consumption is increased in the case of LNG, the mission fuel mass is decreased and CO2 emissions are reduced by more than 15%. These benefits come with a significant reduction in fuel cost per passenger, highlighting the potential benefits of adopting LNG for aviation.
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Riboldi, Carlo E. D., Alberto Rolando, and Gregory Regazzoni. "On the Feasibility of a Launcher-Deployable High-Altitude Airship: Effects of Design Constraints in an Optimal Sizing Framework." Aerospace 9, no. 4 (April 11, 2022): 210. http://dx.doi.org/10.3390/aerospace9040210.
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When ground observation or signal relaying in the vicinity of an unfriendly operative scenario are of interest, such as for military actions or disaster relief, high-altitude airships (HAA) offer some technical benefits. Featuring a milder cost and higher deployment flexibility with respect to lower-Earth orbit satellites, these platforms, often baptized as high-altitude pseudo-satellites (HAPS), operate sufficiently far from the ground to provide better imaging coverage and farther-reaching signal relaying than standard low-flying systems, such as aircraft or helicopters. Despite the atmospheric conditions in the higher atmosphere, they offer stable airstreams and highly-predictable solar energy density, thus ideally giving the chance of smooth operation for a prolonged period of time. The design of airships for the task is often conditioned by the need to go through the lower layers of the atmosphere, featuring less predictable and often unstable aerodynamics, during the climb to the target altitude. With the aim of simultaneously largely increasing the ease and quickness of platform deployment, removing most of the design constraints for the HAPS induced by the crossing of the lower atmosphere, and thus allowing for the design of a machine best suited to matching optimal performance at altitude, the deployment of the HAA by means of a missile is an interesting concept. However, since the HAA platform should take the role of a launcher payload, the feasibility of the mission is subject to a careful negotiation of specification, such that the ensuing overall weight of the airship is as low as possible. A preliminary design technique for high-altitude airships is therefore introduced initially, customized to some features typical to missile-assisted deployment, but with the potential for broader applications. The proposed procedure bends itself to the inclusion in an optimal framework, with the aim of seeking a design solution automatically. A validation of the adopted models and assumptions on existing HAPS is proposed first. The design of the airship is then carried out in a parameterized fashion, highlighting the impact of operative and technological constraints on the resulting sizing solutions. This allows for the marking of the boundaries of the space of design solutions for a launcher-deployable airship.
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Frediani, A., Vittorio Cipolla, K. Abu Salem, V. Binante, and M. Picchi Scardaoni. "Conceptual design of PrandtlPlane civil transport aircraft." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 234, no. 10 (February 1, 2019): 1675–87. http://dx.doi.org/10.1177/0954410019826435.
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According to aircraft manufacturers and several air transportation players, the main challenge the civil aviation will have to deal with in the future is to provide a sustainable growth strategy, in order to face the growing demand of air traffic all over the world. The sustainability requirements are related to air pollution, noise impact, airport congestion, competitiveness of the air transportation systems in terms of travel time and passengers' comfort. Among the possible ways to allow a sustainable growth of the air transportation systems, disruptive aircraft configurations have been object of study for several years, in order to demonstrate that the improvement of aircraft performance can enable the envisaged growth. This paper presents the study of a possible novel configuration called “PrandtlPlane,” having a box-wing layout derived from Prandtl's “Best Wing System” concept. The paper deals with the definition of top level requirements and faces the conceptual study of the overall configuration, focusing on fuselage sizing as well as on the aerodynamic design of the box-wing system. This latter is designed through an optimization-driven strategy, carried out by means of a low-fidelity aerodynamic tool, which simulates the flow condition in the subsonic range and introduces corrections to take the transonic effects into account. Design procedures and tools are presented, showing preliminary results related to a PrandtlPlane compliant with ICAO Aerodrome Reference Code “C” standard, such as Airbus A320 and Boeing 737, whose wingspan is limited to 36 m. Activities and results here shown are part of the first phase of the research project “PARSIFAL” (Prandtlplane ARchitecture for the Sustainable Improvement of Future AirpLanes), funded by the European Commission under the Horizon 2020 Program, which aims to demonstrate that the PrandtlPlane configuration can improve aircraft payload capability, keeping their dimensions compatible with present airport infrastructures.
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Ng, T. T. H., and G. S. B. Leng. "Design Optimization of Rotary-Wing Micro Air Vehicles." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 220, no. 6 (June 1, 2006): 865–73. http://dx.doi.org/10.1243/09544062jmes104.
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In this paper, a new design methodology is introduced to automate the configuration layout design and geometric sizing of rotary-wing micro air vehicles (MAV). The objective of this design-optimization problem is to organize a given set of components and payloads such that the resulting flight vehicle has the most compact overall size and still fulfils the given physical and control constraints. Genetic algorithm (GA) is chosen as the optimization engine because of its proven robust performance. A detailed discussion is presented to explain how the rotary-wing MAV design problem can be formulated as a GA optimization problem. From the case study performed, it is demonstrated that the proposed methodology is able to achieve the design goal.
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"Design and Analysis of UAV for High Payload." International Journal of Innovative Technology and Exploring Engineering 9, no. 2 (December 10, 2019): 3733–36. http://dx.doi.org/10.35940/ijitee.b6694.129219.
Full textAbstract:
This project presents design process for a medium range unmanned aircraft. The objective is to design an aircraft which carries high payload with restricted wing span. The project deals with study of various innovative techniques to improve the aerodynamic performance of the aircraft. It emphasizes on usage of box wing technology after considering various aspect of design and fabrication. The design process involved is a conventional design cycle with initial sizing and trade studies followed by analysis for the validation of the design. The study of the downwash and effect of the stagger angle due to two wings on their aerodynamic performance was also carried out. Flow analysis for the required conditions of cruise is performed and had shown that the lift is sufficient and drag is slightly high. For the applied loads the structure designed has stresses within the limits of the structural integrity. The propulsive force required was comparatively high when compared to single wing of same area.
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Saias, Chana Anna, Ioannis Goulos, Ioannis Roumeliotis, Vassilios Pachidis, and Marko Bacic. "Preliminary Design of Hybrid-Electric Propulsion Systems for Emerging UAM Rotorcraft Architectures." Journal of Engineering for Gas Turbines and Power, August 9, 2021. http://dx.doi.org/10.1115/1.4052057.
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Abstract The increasing demands for air-taxi operations together with the ambitious targets for reduced environmental impact have driven significant interest in alternative rotorcraft architectures and propulsion systems. The design of Hybrid-Electric Propulsion Systems (HEPSs) for rotorcraft is seen as being able to contribute to those goals. This work aims to conduct a comprehensive design and trade-off analysis of hybrid powerplants for rotorcraft, targeting enhanced payload-range capability and fuel economy. An integrated methodology for the design, performance assessment and optimal implementation of HEPSs for conceptual rotorcraft has been developed. A multi-disciplinary approach is devised comprising models for rotor aerodynamics, flight dynamics, HEPS performance and weight estimation. All models are validated using experimental or flight test data. The methodology is deployed for the assessment of a hybrid-electric tilt-rotor, modelled after the NASA XV-15. This work targets to provide new insight in the preliminary design and sizing of optimally designed HEPSs for novel tilt-rotor aircraft. The paper demonstrates that at present, current battery energy densities (250Wh/kg) severely limit the degree of hybridization if a fixed useful payload and range are to be achieved. However, it is also shown that if advancements in battery energy density to 500Wh/kg are realized, a significant increase in the level of hybridization and hence reduction of fuel burned and carbon output relative to the conventional configuration can be attained. The methodology presented is flexible enough to be applied to alternative rotorcraft configurations and propulsion systems.
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Gkoutzamanis, Vasilis G., Mavroudis D. Kavvalos, Arjun Srinivas, Doukaini Mavroudi, George Korbetis, Konstantinos G. Kyprianidis, and Anestis I. Kalfas. "Conceptual Design and Energy Storage Positioning Aspects for a Hybrid-Electric Light Aircraft." Journal of Engineering for Gas Turbines and Power 143, no. 9 (June 17, 2021). http://dx.doi.org/10.1115/1.4050870.
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Abstract This work is a feasibility study of a 19-passenger hybrid-electric aircraft, to serve the short-haul segment within the 200–600 nautical miles. Its ambition is to answer some dominating research questions, during the evaluation and design of aircraft based on alternative propulsion architectures. The potential entry into service (EIS) is foreseen beyond 2030. A literature review is performed to identify similar concepts under research and development. After the requirements' definition, the first level of conceptual design is employed. The objective of design selections is driven by the need to reduce CO2 emissions and accommodate aircraft electrification with boundary layer ingestion engines. Based on a set of assumptions, a methodology for the sizing of the hybrid-electric aircraft is described to explore the basis of the design space, incorporating a parametric analysis for the consideration of boundary layer ingestion effects. Additionally, a methodology for the energy storage positioning is provided to highlight the multidisciplinary aspects between the sizing of an aircraft, the selected architecture (series/parallel partial hybrid), and the storage characteristics. The results show that it is not possible to fulfill the initial design requirements (600 nmi) with a fully-electric aircraft configuration, due to the far-fetched battery necessities. It is also highlighted that compliance with airworthiness standards is favored by switching to hybrid-electric aircraft configurations and relaxing the design requirements (targeted range, payload, battery technology). Finally, the lower degree of hybridization (40%) is observed to have a higher energy efficiency (−12% energy consumption) compared to the higher degree of hybridization (50%) and greater CO2 reduction, with respect to the conventional configuration.
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Vervaeke, Michael. "Optical interconnects for satellite payloads: sizing up the state of the art." SPIE Newsroom, 2010. http://dx.doi.org/10.1117/2.1201003.002685.
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