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The purpose of this paper is to define minimum cost technique of turbo fan transport aircraft in the presence of dynamic change of aircraft performance. Results can be practical applicable in airlines for achieving minimal operation costs.

Logarithmic differential is applied for defining conditions in order to achieve optimal Mach number for minimal climb cost. This condition is solved numerically by using Newton‐Ramphson method, to obtain optimal Mach number distribution with altitude. Conclusion about optimal top of climb (TOC) is defined after analyses for different aircraft mass and cost indexes.

Proposed method of minimum cost climb resulting in potential savings up to 5 per cent compared to Aircraft Flight Manual climb law. Proposed method also made correction of climb law and optimal TOC under existence of aircraft performance degradation.

Use of defined climb law and optimal TOC will minimize cost of en route flight profile.

At present, there is no definition of climb technique for minimum cost of en route flight profile, under dynamic degradation of aircraft performance. Final results are standardized to become applicable and easy to use with modern and old type of flight management system.

The purpose of this paper is to derive and solve equations for the fuel required by internal combustion engine airplanes on trajectories with a constant rate of climb or descent. Three modes of flight are considered: constant angle of attack, constant speed, and constant Mach number.

Newton's second law of motion is used to derive the equations. These are Riccati equations or reduced to such equations after neglecting a small term. A change of variable transforms them into second order linear differential equations that are solved exactly.

Formulas are found for the weight of fuel, speed, altitude, horizontal distance, time to climb, and power required.

The formulas obtained have direct applications for the analysis of aircraft performances and mission planning.

The formulas obtained are new and fundamental for the analysis of aircraft dynamics.

WE first derive for jet aeroplanes the formulae for the least thrust sufficient for a specified angle of climb. In the mathematical derivation of formulae we make use of the method represented by H. Freeman in J. of Ae. Sci. No. 3/1947, ‘Simple Analytic Equations for the Velocity of an Aeroplane in Unaccelerated Level, Climbing and Diving Flight’. We then see that through a simple deformation the equations can be brought into such a form as to make it possible to represent the steepest climbing angle, the corresponding flight velocity and the rate of climb obtainable at a specified thrust by means of a combined, very simple, dimensionless vector diagram.

What follows considers only steady climbing conditions and ignores the possible use of kinetic energy, as in zooming.

The current air traffic management (ATM) operational approach is changing; “time” is now integrated as an additional fourth dimension on trajectories. This notion will impose on aircraft the compliance of accurate arrival times over designated checkpoints (CPs), called time windows (TWs). This paper aims to clarify the basic requirements and foundations for the practical implementation of this functional framework.

This paper reviews the operational deployment of 4D trajectories, by defining its relationship with other concepts and systems of the future ATM and communications, navigation and surveillance (CNS) context. This allows to establish the main tools that should be considered to ease the application of the 4D-trajectories approach. This paper appraises how 4D trajectories must be managed and planned (negotiation, synchronization, modification and verification processes). Then, based on the evolution of a simulated 4D trajectory, the necessary corrective measures by evaluating the degradation tolerances and conditions are described and introduced.

The proposed TWs model can control the time tolerance within less than 100 s along the passing CPs of a generic trajectory, which is in line with the expected future ATM time-performance requirements.

The main contribution of this work is the provision of a holistic vision of the systems and concepts that will be necessary to implement the new 4D-trajectory concept efficiently, thus enhancing performance. It also proposes tolerance windows for trajectory degradation, to understand both when an update is necessary and what are the conditions required for pilots and air traffic controllers to provide this update.

The aim of this article is to determine the simplest and clearest relations existing between the principal characteristics of aeroplanes and their performances.

Aims to present a methodology for analysing a solar‐electric, high‐altitude, long‐endurance, unmanned aircraft.

The study focuses on the aerodynamics, flight performance and power requirements of a heavier‐than‐air, solar‐electric, HALE UAV. The methodology is founded on using an analytical approach to determine the power required to undertake various flight manoeuvres. An analytical approach is also undertaken in determining the intensity of the solar radiation available to the aircraft. Finally to demonstrate the methodology, a HALE concept was generated and evaluated.

When using estimates of current solar‐electric propulsion and energy conversion efficiencies, the HALE concept was only able to sustain year round, level flight up to latitudes of 10°N.

Further analysis needs to be undertaken into the effect of altitude on the intensity of solar radiation, which could be as much as 25 per cent higher at an altitude of 21.3 km (70,000 ft). Further study into this subject area may provide proof that sustained flight is possible at more northerly latitudes.

This paper provides a simple methodology for persons wishing to undertake an initial feasibility study of a solar‐electric HALE concept.

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.

A hybrid-electric unmanned aerial vehicle (HE-UAV) model has been developed to address the problem of low endurance of a small electric UAV. Electric-powered UAVs are not capable of achieving a high range and endurance due to the low energy density of its batteries. Alternatively, conventional UAVs (cUAVs) using fuel with an internal combustion engine (ICE) produces more noise and thermal signatures which is undesirable, especially if the air vehicle is required to patrol at low altitudes and remain undetected by ground patrols. This paper aims to investigate the impact of implementing hybrid propulsion technology to improve on the endurance of the UAV (based on a 13.6 kg UAV).

A HE-UAV model is developed to analyze the fuel consumption of the UAV for given mission profiles which were then compared to a cUAV. Although, this UAV size was used as reference case study, it can potentially be used to analyze the fuel consumption of any fixed wing UAV of similar take-off weight. The model was developed in a Matlab-Simulink environment using Simulink built-in functionalities, including all the subsystem of the hybrid powertrain. That is, the ICE, electric motor, battery, DC-DC converter, fuel system and propeller system as well as the aerodynamic system of the UAV. In addition, a ruled-based supervisory controlled strategy was implemented to characterize the split between the two propulsive components (ICE and electric motor) during the UAV mission. Finally, an electrification scheme was implemented to account for the hybridization of the UAV during certain stages of flight. The electrification scheme was then varied by changing the time duration of the UAV during certain stages of flight.

Based on simulation, it was observed a HE-UAV could achieve a fuel saving of 33% compared to the cUAV. A validation study showed a predicted improved fuel consumption of 9.5% for the Aerosonde UAV.

The novelty of this work comes with the implementation of a rule-based supervisory controller to characterize the split between the two propulsive components during the UAV mission. Also, the model was created by considering steady flight during cruise, but not during the climb and descend segment of the mission.

THE trend of design in the modern aeroplane has been toward improved performance realised through external cleanness. It is apparent that the number of essential units comprising a modern aeroplane is nearly a minimum at the present stage of the art, and it appears also that the possibilities of further striking reductions in the drag of these units, due to change in form or shape either individually or in combination, are not great.