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A longitudinal flight dynamic study of a low mass moment of inertia vehicle is presented. Aerodynamic and stability derivatives of a flying-wing microaerial vehicle (FWMAV) were obtained through detailed subsonic wind tunnel tests at a Reynolds Number of 1.87 × 105. Rate and acceleration derivatives were obtained using the potential flow solver, Tornado®. A novel methodology for the estimation of dimensional derivatives is proposed, and results are compared with the conventional linear time-invariant systems (LTI) approach. Free response for natural frequency, damping coefficient, and time constant as well as forced response upon a unit step and a unit impulse elevon input has been calculated and analyzed. The proposed methodology predicted two pairs of complex conjugates for the longitudinal flight up to a pitch angle of 89° whereas the conventional methodology predicted the same up to 57°. Longitudinal modes sensitivity in terms of stability with the variation of mass, velocity, and pitch angle has also been analyzed. The flying-wing microaerial vehicle was able to sustain straight and level flight during flight trials; however, higher frequencies of phugoid and short period modes were observed. These high frequencies were the consequence of large magnitude of [Formula: see text] (ratio of Z-force derivative with the angle of attack and cruise velocity) and [Formula: see text] (ratio of Z-force derivative with the axial velocity and cruise velocity). It is concluded that the proposed methodology presented a more realistic representation of longitudinal flight modes since classical flight modes are captured till 89° which conventional LTI methodology failed to do so.

Longitudinal damping derivatives, [Formula: see text], of an aerial vehicle is important from an aerodynamic stability point of view. Experimental calculation of longitudinal damping derivatives using wind tunnel is not a cost-effective method; therefore, researchers have developed numerical solutions as an alternative. In this research, the longitudinal damping derivatives of a flying wing micro aerial vehicle (FWMAV) were calculated using numerical simulations by adopting pull-up maneuver and forced harmonic motion in pitch axis. Pull-up maneuver with four steady rotational rates was simulated to obtain pitch rate derivative, C mq. Combined derivative, [Formula: see text], was obtained by simulating forced harmonic motion of FWMAV around a mean angle of attack of 0° with amplitude of oscillation of ± 3° using four reduced frequencies (0.02, 0.03, 0.04, and 0.05). Unstructured surface and volume mesh was used in a spherical domain engulfed inside a large cuboid domain for moving reference frame strategy. Reynolds number taking mean aerodynamic chord as a reference length was 2.33 × 105. Spalart–Allmaras turbulence model was used. Pitch rate derivative, combined derivative, and acceleration derivative were found as − 0.03/rad, − 7.39/rad, and − 7.36/rad, respectively, by the use of a phase method at a reduced frequency of 0.03. During flight dynamic analysis, it was found that [Formula: see text] has a significant contribution on damping in short period mode with no effect on Phugoid mode. The research concluded that for tailless configurations, acceleration derivative [Formula: see text] can exist and can provide necessary damping in the longitudinal flight mode.

Reliability of high precision linear motion system is one of the main concerns in industrial and military systems. The performance and repeatability of these systems are influenced by their respective linear drives and load bearings. A fault in these members severely affects the safe working of overall system. This paper gives a reliable intelligent approach to detect and classify faults for linear motion systems based on deep learning methods. Accuracy in faults identification is highly dependent on improved features extraction. For this purpose, a novel Residual Twin CNN (ResT-CNN) is proposed that uses combination of 1-D and 2-D CNN in parallel learning which improves features extraction performance; followed by knowledge base-Remnant-PCA (Kb-Rem-PCA) architecture in combination with multi-class support vector machine (Mc-SVM). This novel hybrid combination proved very effective in accurate faults identification. The performance of proposed methodology was also validated by IMS-UC (Intelligent Maintenance Systems – University of Cincinnati) public bearing dataset. The results confirm the effectiveness of proposed scheme in comparison to existing state of the art techniques.

A comprehensive design cycle of a nose landing gear strut having an oleo-pneumatic shock absorber for a lightweight aircraft is proposed. Design and analysis of a retractable nose landing gear according to Airworthiness Standards FAA FAR Part 23 have been carried out. This research is focused on mathematical modeling of an oleo-pneumatic strut with an analytical solution of design variables at static and dynamic loading conditions. The variation of spring and damping characteristics of an oleo-pneumatic shock absorber with the stroke length is also presented. Feasibility of equivalent mechanical spring and damper along with comparison of pneumatic as air spring and oleo as hydraulic damper is studied. Numerical integration technique was used to solve the dynamic model of an oleo-pneumatic strut with forcing function of an impact force during touch down scenario. Energy conservation principle was used to determine height required for drop tests. Parametric study of anteversion angle within the constraints of ground clearance and volume in the fuselage determined an optimized angle of nose landing gear strut. Based on the maximum pressure and impact force encountered during landing, the hydraulic cylinder and piston design was finalized. In order to validate the proposed design cycle for preliminary phase, a structural integrity of cylinder and piston assembly was carried out using finite element analysis. Deformation, maximum stresses and factor of safety validated the proposed design cycle of a nose landing gear strut specific to a general aviation aircraft having all up mass of 1600 Kg.

The stability and dynamic positioning of any floating structure is essential during offshore or on shore activities. It is therefore necessary to carry out offshore mooring system analysis in order for it to withstand extreme environmental forces of wind, wave, and current that will act on the structures. This project is focused on designing a fit-for-purpose buoyant mooring system to achieve the stability and dynamic positioning of a floating system. This system must provide a discreet non-linear force distribution within a given range. The sea-water depth range at the shore was also provided. For a certain stretch in the mooring line, due to rise in the level of water, the decrement in the required buoyant force was to decrease non-linearly. In order to meet that requirement a mechanism with only one piston-cylinder connected to mooring line was devised. This was the simplest solution that could be achieved for the desired outcome. Any other mechanism, according to observations, increases the complexity of the system, which was to be avoided (as demanded in the problem).Once the string is stretched the piston is pulled downwards. This increases the total volume occupied by the mechanism thus increasing the buoyant force (Archimedes Principle). But there is a linear relationship between the stretch of string and increase of buoyant force. In order to make this relationship non-linear a multi-groove step pulley was used. The mooring line passing over this pulley produced a non-linear pull in the piston (since the radius of the each step changes non-linearly). As a result buoyant force also changes non-linearly for the same stretch of the mooring line. CATIA was used for 3D modeling of the system, and better understanding of the mechanism. The above mentioned mechanism was the simplest which could be achieved, because no complex tools were required for its construction and it also provided the exact values as demanded.

Micro aerial vehicles using flapping wings are under investigation, as an alternative to fixed-wing and rotary-wing micro aerial vehicles. Such flapping-wing vehicles promise key potential advantages of high thrust, agility, and maneuverability, and have a wide range of applications. These applications include both military and commercial domains such as communication relay, search and rescue, visual reconnaissance, and field search. With the advancement in the computational sciences, developments in flapping-wing micro aerial vehicles have progressed exponentially. Such developments require a careful aerodynamic and aeroelastic design of the flapping wing. Therefore, aerodynamic tools are required to study such designs and configurations. In this paper, the role of several parameters is investigated, including the types of flapping wings, the effect of the kinematics and wing geometry (shape, configuration, and structural flexibility) on performance variables such as lift, drag, thrust, and efficiency in various modes of flight. Kinematic variables have a significant effect on the performance of the flapping wing. For instance, a high flap amplitude and pitch rotation, which supports the generation of the strong leading-edge vortex, generates higher thrust. Likewise, wing shape, configuration, and structural flexibility are shown to have a large impact on the performance of the flapping wing. The wing with optimum flexibility maximizes thrust where highly flexible wings lead to performance degradation due to change in the effective angle of attack. This study shows that the development of the flexible flapping wing with performance capabilities similar to those of natural fliers has not yet been achieved. Finally, opportunities for additional research in this field are recommended.

In this research effect of propeller induced flow on aerodynamic characteristics of low aspect ratio flying wing micro aerial vehicle has been investigated experimentally in subsonic wind tunnel. Left turning tendencies of right-handed propellers have been discussed in literature, but not much work has been done to quantify them. In this research, we have quantified these tendencies as a change in aerodynamic coefficient with a change in advance ratio at a longitudinal trim angle of attack using subsonic wind tunnel. For experimental testing, three fixed pitch propeller diameters (5 inch, 6 inch and 7 inch), three propeller rotational speeds (7800, 10800 and 12300 RPMs) and three wind tunnel speeds (10, 15 and 20 m/s) have been considered to form up 27 advance ratios. Additionally, wind tunnel tests of 9 wind mill cases were conducted and considered as baseline. Experimental uncertainty assessment for measurement of forces and moments was carried out before conduct of wind tunnel tests. Large variation in lift, drag, yawing moment and rolling moment was captured at low advance ratios, which indicated their significance at high propeller rotational speeds and large propeller diameters. Side force and pitching moment did not reflect any significant change. $\frac {L}{D}$ at trim point was found a nonlinear function of propeller diameter to wingspan ratio $\frac {D}{b}$ , and propeller rotational speed. Rate and control derivatives were obtained using unsteady vortex lattice method with propeller induced flow effect modeled by Helical Vortex Modeling approach. In this research, we have proposed improved 6-DOF equations of motion, with a contribution of advance ratio $J$ . It is concluded, that propeller induced flow effects have a significant contribution in flight dynamic modeling for vehicles with large propeller diameter to wingspan ratio, $\frac {D}{b}$ of 22% or more.

Main landing gears are a critical structural part of an aircraft that support all up weight in static conditions and absorb kinetic energy in a form of impact loads during touchdown. High strength and low weight are critical design considerations for any landing gear strut. During this research, carbon reinforced composite strut was optimally designed such that it undertakes static loads and impact loads of a UAV having GTOW of 1600 kg. CAD models of the designs were developed in CATIA V5R20® and the analysis was carried out in ANSYS Workbench®. Rectangular, circular and elliptical cross-sectional shapes were modeled in ANSYS ACP® with different ply orientation sequences and varying weight distributions from central axis. Least total deformation along with least equivalent stress and aerodynamic drag was considered for the selection of landing gear strut configurational shape. Further design changes in the selected CAD model were incorporated keeping in view the safety factor of the struts. We concluded that 60 preimpregnated carbon reinforced fiber with thermoset polymer matrix having 0.27 mm of thickness for each ply at specified concentration and orientation gave the best results. The orientation of plies was found to be 60 %/ 0°, 30 %/ 45° and 10%/ 90°. Maximum deformation for the final model was 42.154 mm at the joints where the axle was connected with the strut whereas the maximum stress was 712.92 MPa at the joints representing the connection of the landing gears with the fuselage. Drag force for each cross-sectional shape was computed at the takeoff speed of 30 m/s. Drag force for the rectangular cross section was 50.23 N whereas for the circular and elliptical cross section was observed to be 36.837 N and 16.744 N respectively. Hence research concluded that elliptical cross-sectional shape presented least stress distribution and deformation values.

Among different kinds of actuators, pressure-driven ones are significant and utilized because of their high capacity to weight proportion. The most significant utilization of pressuredriven actuators is related to flywheels. Utilizing actuators in blend with flywheels empowers us to control rotational speed (RPMs) and force of any instrument in a successful way. Such constructions are frequently observed in mechanical arms, boats, and marines. In this study, the design and analysis of control actuators with the flywheel are presented. One such device is a dynamometer which gives quick input estimation about the extent and power applied. Demonstrating such a framework needs a brought together way to deal with speak to the nonlinear conduct qualities. The bond diagram model is ideal for this undertaking. Moreover, traditional PID control methodology is utilized to accomplish the necessary exactness and accuracy, which is straightforward, programmed, less tedious, and proficient. Reproductions and investigations on the actuator are completed to assess the adequacy of the proposed control technique. The dynamometer framework is utilized for the testing and assessment of engine and powertrain segments. So We can utilize can dynamometer sensors to accomplish a close loop control framework as an actuator in the flywheel. flywheels accessible for when the motor driveline needs to incorporate a latency incentive to coordinate with the airframe transmission. Most flywheels are provided mounted in an independent base casing. Be that as it may, for specific applications their systems can be used as half couplings Dynamometer has likewise been utilized as a fundamental device. The standard of the dynamometer will be able to retain the power of the car and gauge force and speed and decide the motor yield power or the powertrain power to adjust the force and speed.

A unique approach of analyzing jet exhaust nozzle integrated to aircraft and propulsion system is presented in this paper. Engine exhaust nozzle is usually omitted in Wind Tunnel Testing and numerical analysis of aircraft due to complexities involved in integration of nozzle and presence of high pressure / temperature inside exhaust nozzle. Also, the flow properties are non-uniform and highly turbulent in the vicinity of nozzle. Therefore, exhaust nozzle is usually analyzed in isolation and these results often lead to inaccuracies from actual scenario where nozzle is integrated with aircraft and its propulsion system. This research aims to integrate engine exhaust nozzle on a supersonic fighter aircraft and analyze its flow characteristics and variation in performance parameters due to its integration. Engine propulsion characteristics and parameters such as nozzle inlet temperature and total pressure have been analyzed through an in-house validated engine analytical model developed by some of the authors of this study. In the first part of paper, exhaust plume structure has been analyzed to study the flow behaviour (flow turbulence and flow distortion etc) at nozzle exit. Later, nozzle performance parameters such as Exit Velocity, Nozzle Pressure Ratio (NPR), Engine Pressure Ratio (EPR), and Engine Temperature Ratio (ETR) have been calculated when exhaust nozzle is integrated with the aircraft. Finally, the results are compared and validated with analytical calculations to compare the performance of nozzle when it is in isolation and when it is integrated on aircraft. It is observed that nozzle flow has no significant effect on aircraft major surfaces such as fuselage, wing upper and lower surfaces, and nose section. However, there is a prominent effect of exhaust nozzle flow on horizontal stabilizers, vertical tail and rear fuselage area of the aircraft. An average difference of 18% in NPR, 12% in EPR, and 9% in ETR is observed between integrated nozzle and isolated nozzle which further signifies the importance of integrating exhaust nozzle in aircraft analysis. This proposed methodology will allow more accurate analysis of the effects of exhaust nozzle on the overall performance of aircraft. The methodology can further be used for proposing design changes in existing nozzle configurations.

Multirotor unmanned aerial vehicles (UAVs) have gone from a single rotor to multiple rotors. Multirotor unmanned aerial vehicles (UAVs) are now used for surveillance and rescue operations. These are used for surveillance over disastrous areas and to perform necessary rescue action. The capacity of multirotor UAVs to carry heavy payload is less with small and limited propellers. Therefore, the number of propellers have been increased, and ultimately the size of multirotor UAV has also been increased. The rotor arrangement is fundamental in the thrust of multirotor UAVs. The opensource software has been developed to integrate with the flight controller hardware. The current focus of research is to reduce the size of multirotor UAVs and increase the payload capacity of multirotor UAVs. The thrust has been reduced in case of overlapping rotor configuration, such as in coaxial and half propeller overlap configuration. These configurations are adopted to reduce the size of multirotor UAV. In this study, different multirotor UAV configurations have been studied for selection of best configuration for small but heavy-lift multirotor. The overlapping configurations have been found to increase the payload and reduce the size of the multirotor. The thrust has been increased to about 90% in the case of some propeller overlap configuration with a rotor plane angle. For coaxial configuration, the rotor thrust was about 76% of the thrust of the independent rotor. The thrust generated by the flat configuration is more than the coaxial and less propeller overlap configuration. The flight controllers are responsible for the stability and control of UAVs. Many flight controllers are working on the latest technology and can be programmed for various features. The flight controllers also contribute to increasing stability. Mostly open-source software is available to integrate with the flight controller hardware. The ARM-based flight controllers like Pixhawk flight controllers are the best to control the multirotor. These flight controllers can be programmed quickly and can be installed on the multirotor.

Pakistan has been frequently hit by natural disasters in the recent past, and the frequency of such disasters is on a rise. National agencies work hard to cope up with the situations to save precious human lives; however, there is still a dire need to develop capable and potent platforms which can fly remotely over difficult terrains and can provide comprehensive picture of the affected areas. For this purpose, in this research, a long range video telemetry system has been developed for use on an aerial platform, an RC helicopter, to make it fly worthy beyond visual range through video signal and other flight sensors data. The proposed vision system consists of a camera, GPS, altimeter, stabilization sensor, air speed sensor, G-sensor and temperature sensors. In order to save the payload capacity of our helicopter, a box has been designed and manufactured from glass reinforced plasticto make it lighter. This box contains separate compartments for the transmitter and other flight data sensors. The video signal is obtained from on-board camera, and other data is overlaid on this video stream and is then transmitted at 2.4 GHz frequency from a long distance. After receiving and amplifying the video signal, the telemetry data is extracted and used as a real time visual feedback to control and maneuver the RC helicopter through communication system when it goes beyond the visual site. The developed vision system is successfully tested from 48 km line-of-sight distance. It is a low cost, small size, light weight and power efficient vision system that can be installed on any ready-to-fly RC helicopter with as low pay load capacity as 2.0 kg.

Multi-criteria decision making (MCDM) is the method of making accurate decision among many available criteria. It is a scientific method of making decisions. The decisions made by using MCDM are very accurate. MCDM steps have been divided into decision structuring, criteria specification, alternative performance measurement, alternative scoring, and criteria weight. There are many methods available for MCDM. In this study many direct and indirect methods of MCDM have been studied. Analytic hierarchy process (AHP) and technique for order of preference by similarity to ideal solution (TOPSIS) have been found very powerful methods for MCDM. AHP and TOPSIS methods have been compared in this study after using their application formulas. COPRAS method application has also been discussed in this study. After comparing these methods, it was found that the TOPSIS method was ideal to be applied because of its vast applications. TOPSIS method has been found a better and easy method to apply for decision making. TOPSIS method steps including formulas have been presented. The TOPSIS method has been applied at the drones available in the market at desired cost. The formulas of TOPSIS method have been applied at the selected drones. The non-beneficial and beneficial weights have been assigned to the drones. The normalized matrices have been calculated for each drone. The Euclidean distance for ideal best and ideal worst values has been calculated for each drone and assigned performance score. The drones have been allotted ranks starting from highest to the lowest values. The highest rank drone has been selected.

Electrical linear actuators reduce the weight of landing gear retraction / extension mechanism which helps in keeping the gross takeoff weight of aircraft weight low. Reduction in weight of the aircraft has manifold advantages in which noticeable are increased endurance, range, climb performance, acceleration performance and fuel efficiency. Use an electrical linear actuator in landing gear system is a simplification in the kinematics of landing gear retraction / extension sequence. In this research, a retraction / extension mechanism for nose landing gear (NLG) of a MALE UAV has been designed and tested according to FAR CFR 25.729. The design process of a Retractable Landing Gear System Mechanism (RLGS) started with selection of pivot point on NLG and axis for the rotation in free space available which resulted in least power requirement. Pivot point was selected such that no part of gear posed hindrance to other part of an aircraft. Next, the path of rotation about pivoting axis and position kinematics were calculated using vector loop closure equation. Essential parameters regarding retraction / extension sequence were calculated from trigonometry. Forces and power required for complete retraction / extraction sequence were calculated. During the research, it was found that retraction force on NLG depended upon the aerodynamic drag, inertial loads and gravitational force. In this research analysis of these forces is presented with much detail. Effect of velocity variation of linear actuator on the retraction force is commented upon. Computational analysis is carried out using multibody dynamics MSC ADAMS® software. The research concluded with experimental validation of a planner triangular retraction mechanism of NLG actuated by electrical linear actuator.

Owing to non-availability of adequate data for situational analysis and requisite rapid surveillance equipment, initial response for relief efforts remains delayed during national calamities in Pakistan. As every second counts in humanitarian logistics, many precious lives are lost until initiation of relief and rescue operations. It necessitates development of a potent system able to spearhead relief and rescue operations by providing realtime situational analysis right at the onset. The research objective is to develop a disaster relief aerial platform indigenously, capable of providing quick and cost effective solution for initial disaster survey and relief efforts. Firstly, Rotary Unmanned Aerial Vehicle was scrutinized as basic platform, which was assembled after classification of each component. To achieve enhanced range operations, Flight Control Communication System with enhanced range and customized Beyond Visual Line Of Sight Telemetry system were integrated on the RUAV. Evaluations to verify the intended results and reliability of the developed system conceded a number of ground tests alongwith successful flight tests. The developed RUAV with integrated Enhanced Range Flight Control System and BVLOS Telemetry System is a novel platform able to provide feedback for situational analysis as well as for relief efforts during calamities.

This paper presents a review of recent progress made towards the applications of the meshfree particle methods (MPMs) for solving coupled fluid-structure interaction (FSI) problems. Meshfree methods are categorized based on their mathematical formulation and treatment of computational data points. The advantages and limitations of these methods, particularly related to FSI applications, have been identified. A detailed account of salient work related to the FSI problems involving complex geometries, viscous flows, and large structural deformations has been presented and the benchmark solutions are identified for future research. Compared to their mesh-based counterparts, MPMs are found better suited in negotiating moving boundaries and complex geometries, features that are the hallmark of FSI problems. However, the biggest challenge to their wider acceptability is their implementation and programming complexity, higher computational cost, and lack of commercial software packages. So far, meshfree methods have mostly been limited to applications, where conventional methods show limited performance. Owing to its promising growth potential, partitioned FSI is the prime emphasis of this paper. Various aspects of partitioned FSI have been identified and classified for meshfree FSI problems, which include problem formulation strategies, domains discretization approaches, solver coupling methodology, interface treatment, benchmark problems, computational load, and availability of commercial software. Furthermore, various challenges involved in employing MPMs for FSI have also been identified and discussed along with the state-of-the-art techniques used in meshfree methods and FSI applications, and a future way forward has been proposed. In essence, this paper is an effort to identify and classify key aspects of MPM applications for FSI and suggest potential avenues to explore the full potential of MPM capabilities for the solution of coupled problems.