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The purpose of this paper is to create a new fuel flow rate model for the descent phase of the flight using particle swarm optimization (PSO).

A new fuel flow rate model was developed for the descent phase of the B737-800 aircraft, which is frequently used in commercial air transport using PSO method. For the analysis, the actual flight data records (FDRs) data containing the fuel flow rate, speed, altitude, engine speed, time and many more data were used. In this regard, an empirical formula has been created that gives real fuel flow rate values as a function of altitude and true airspeed. In addition, in the fuel flow rate predictions made for the descent phase of the specified aircraft, a different model has been created that can be used without any optimization process when FDR data are not available for a specific aircraft take-off weight condition.

The error analysis applied to the models showed that both models predict real fuel flow rate values with high precision.

Because of the high accuracy of the PSO model, it is thought to be useful in air traffic management, decision support systems, models used for trajectory prediction, aircraft performance models, strategies used to reduce fuel consumption and emissions because of fuel consumption.

This study is the first fuel flow rate model for descent flight using PSO algorithm. The use of real FDR data in the analysis shows the originality of this study.

The purpose of this paper is to create a new fuel flow rate model using cuckoo search algorithm (CSA) for the descending stage of the flight.

Using the actual flight data record data of the B737-800 aircraft, a new fuel flow rate model has been developed for this aircraft type. The created model is to predict the fuel flow rate with high accuracy depending on the altitude and true airspeed. In addition, the CSA fuel flow rate model was used to calculate the fuel consumption for the point merge system, which is used for combining the initial approach to the final approach at Istanbul Airport, the largest airport of Turkey.

As a result of the analysis, the correlation coefficient value is found as 0.996858 for Flight 1, 0.998548 for Flight 2, 0.995363 and 0.997351 for Flight 3 and Flight 4, respectively. The values that are so close to 1 indicate that the model predicts the real fuel flow rate data with high accuracy.

This model is considered to be useful in air traffic management decision support systems, aircraft performance models, models used for trajectory prediction and strategies used by the aviation community to reduce fuel consumption and related emissions.

The importance of this study lies in the fact that to the best of the authors’ knowledge, it is the first fuel flow rate model developed using CSA for the descent stage in the existing literature; the data set used is real values.

This paper aims to investigate the effects of descent time spent with flaps extended on fuel burn (FB) and specific range for five different flight path angles (FPAs) ranging between 2.0° and 4.0° for a commercial aircraft.

A large data set of actual flight data (n = 475) of the same type of a frequently used commercial aircraft were investigated by using statistical methods.

The result of the comparison of the highest and the lowest FBs of flight profiles for each FPAs present that the fuel saving was achieved by keeping at as a high airspeed as possible and deploying flaps as late as possible, which is in line with the objective of delayed deceleration approaches. From analyzing the flight profiles, it was proven that delaying deceleration and also descending without flaps or with flap over a shorter time resulted in less FB of 101.1, 70.9 and 94.9 kg for FPA 2.5°, FPA 3.0° and FPA 3.5°, respectively.

This study differs from prior studies because it focused on the effects of the different vertical profiles on FB. Also, the use of real flight data recorder data in the analysis presents the originality of this study.

Continuous descent approach (CDA) is a method, which allows the aircraft flying its individual optimal vertical profile down to runway threshold with engines operating at low‐thrust power. The main objective of this paper is to provide less‐fuel consumption, less noise and less emission with using CDA procedures instead of conventional procedures.

Conventional and CDA procedures were modelled in the Istanbul terminal area (TMA), which has five entry points. The real speed and the real altitude limitations were maintained on these entry points. System for Assessing Aviation's Global Emissions research results were also used to determine the emission savings.

With CDA procedures, more than 40 kg fuel and 2 min time savings per flight are obtained; furthermore, regarding CO₂ and H₂O, significant emission savings are also noted.

Some of the benefits of CDA procedures are reported for Istanbul TMA by using true flight data.

The purpose of this paper is to present an approach to a polar graph measurement by a flight testing technique and to propose a baseline research method for future tests of UAV polar graphs. The method presented can be used to demonstrate a conceptual and preliminary design process using a scaled, unmanned configuration. This shows how results of experimental flight tests using a scaled flying airframe may be described and analysed before manufacturing the full scale aircraft.

During the research, the flight tests were conducted for two aerodynamic configurations of a small UAV. This allowed the investigation of the influence of winglets and classic vertical stabilizers on the platform stability, performance and therefore polar graphs of a small unmanned aircraft.

A methodology of flight tests for the assessment of a small UAV’s polar graph has been proposed, performed and assessed. Two aerodynamic configurations were tested, and it was found that directional stability had a large influence on the UAV’s performance. A correlation between the speed and inclination of the altitude graph was found – i.e. the higher the flight speed, the steeper the altitude graph (higher descent speed, steeper flight path angle). This could be considered as a basic verification that the recorded data have a physical sense.

The polar graph and therefore glide ratio of the aircraft is a major factor for determining its performance and power required for flight. Using the right flight test procedure can speed-up the process of measuring glide ratio, making it easier, faster, robust, more effective and accurate in future research of novel, especially unorthodox configurations. This paper also can be useful for the proper selection of requirements and preliminary design parameters for making the design process more economically effective.

This paper presents a very efficient method of assessing the design parameters of UAVs, especially the polar graph, in an early stage of the design process. Aircraft designers and producers have been widely performing flight testing for years. However, these procedures and practical customs are usually not wide spread and very often are treated as the company’s “know how”. Results presented in this paper are original, relatively easily be repeated and checked. They may be used either by professionals, highly motivated individuals and representatives of small companies or also by ambitious amateurs.

The purpose of this paper is to present the development of an optimal design framework for high altitude long endurance solar unmanned aerial vehicle. The proposed solar aircraft design framework provides a simple method to design solar aircraft for users of all levels of experience.

This design framework consists of algorithms and user interfaces for the design of experiments, optimization and mission analysis that includes aerodynamics, performance, solar energy, weight and flight distances.

The proposed sizing method produces the optimal solar aircraft that yields the minimum weight and satisfies the constraints such as the power balance, the night time energy balance and the lift coefficient limit.

The design conditions for the sizing process are given in terms of mission altitudes, flight dates, flight latitudes/longitudes and design factors for the aircraft configuration.

The framework environment is light and easily accessible as it is implemented using open programs without the use of any expensive commercial tools or in-house programs. In addition, this study presents a sizing method for solar aircraft as traditional sizing methods fail to reflect their unique features.

Solar aircraft can be used in place of a satellite and introduce many advantages. The solar aircraft is much cheaper than the conventional satellite, which costs approximately $200-300m. It operates at a closer altitude to the ground and allows for a better visual inspection. It also provides greater flexibility of missions and covers a wider range of applications.

This study presents the implementation of a function that yields optimized flight performance under the given mission conditions, such as climb, cruise and descent for a solar aircraft.

During the time that the Lockheed Flight Management System (FMS) has been in service in TriStar aircraft since certification by the FAA in 1977, sophistication and capability have been enhanced and the system has been simplified procedurally. Fuel conservation and cost savings are major considerations and continuing efforts in the regime of flight management are made to these ends.

The author developed a theory of optimal trajectories for air vehicles with variable wing area and conventional wings. He applied a new theory of singular optimal solutions and obtained the optimal flight in many cases. At first glance, the results may seem strange however, this is correct and this paper will show how this new theory may be used. The main idea of the research is in using the vehicle's kinetic energy for increasing the range of missiles and projectiles. The author shows that the range of a ballistic warhead can be increased 3‐4 times if an optimal wing is added to the ballistic warhead, especially a wing with variable area. If increased range is not needed, the warhead mass can be increased. The range of big gun shells can also be increased 3‐9 times. The range of aircraft may be improved 3‐15 percent and more. The results can be used for the design of aircraft, missiles, flying bombs and shells of big guns.

The purpose of this paper is to develop a real‐time simulation environment for the validation of controller for an autonomous small‐scale helicopter.

The real‐time simulation platform is developed based on the nonlinear model of a series of small‐scale helicopters. Dynamics of small‐scale helicopter is analyzed through simulation. The controller is designed based on the extracted linear model.

The model‐based linear controller can be effectively designed and tested using real‐time simulation platform. The hover controller is demonstrated to be robust against wind disturbance.

To use the real‐time simulation environment to test and validate controllers for small‐scale helicopters, basic helicopter parameters need to be measured, calculated or estimated.

The real‐time simulation environment can be used generically to test and validate controllers for small‐scale helicopters.

The paper presents the design and development of a low‐cost hardware in the loop simulation environment using xPC target critical for validating controllers for small‐scale helicopters.

The purpose of this paper is to generate optimised trajectories for an unmanned aerial vehicle (UAV) during a forest fire detection mission. It is assumed that the UAV flies 3D curvature-constrained Dubins manoeuvres and has a limited amount of battery energy that prevents it from covering the entire search area in a single trip.

In this paper, the search area is discretised into a grid of multiple targets, and each target assigned with a score that is proportional to the time elapsed since the last UAV visit. This problem, known as Dubins Airplane Orienteering Problem, consists of finding the number and order of targets to visit and the UAV heading at each target that maximises the total trip score without exceeding the UAV battery energy. The solution is found using the Randomised Variable Neighbourhood Search metaheuristic. All target scores are updated after each trajectory generation according to the elapsed time since the last UAV visit.

The proposed approach produced feasible results when generating optimised trajectories for a fire detection mission context where energy battery constraints are important.

The authors carry out the planning of UAV missions with limited amounts of onboard energy such as a real fire detection mission using a single electric propulsion and fixed-wing UAV.

This paper introduces an energy-based approach to the Dubins Airplane Orienteering Problem, which takes into account the UAV performance and energy budget when generating optimised trajectories.