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In recent years, high‐altitude/long‐endurance airship platforms have generated great interest as a means to provide communications and surveillance capabilities. The purpose of this paper is to develop a model for airship conceptual design and help provide insight into the viability of high‐altitude/long‐endurance airships.

A configuration analysis model with the consideration of pressure difference, temperature difference, and helium purity, etc. was developed. The influences of the airship payload, size and area required of solar cell with environment and operation parameters, such as operation latitude, pressure difference, temperature difference, helium purity, seasons, latitude, and wind speed, etc. were analyzed.

The results show that the area of solar cell required for stratospheric airship is very large under the condition of low altitude, high latitude, wind, and in winter, etc. which might make the design of high‐altitude/long‐endurance airship an elusive goal. They also show that the solar cell efficiency is the key technology in the control of solar cell area required for airships, and the technology advances in regenerative fuel cells and propeller efficiency have significant effects among on the airship payload, size, and solar cell area required for airship.

The paper analyses the energy balance of the high‐altitude/long‐endurance airship.

The purpose of this study is to present a methodology for parameters’ sensitivity analysis of solar-powered airplanes.

The study focuses on a preliminary design and parameters’ relations of a heavier-than-air, solar-powered, high-altitude long-endurance unmanned aerial vehicle. The methodology is founded on the balance of energy production and requirement. An analytic expression with four generalized parameters is derived to determine the airplane flying on the specific altitude. The four generalized parameters’ sensitivities on altitude are then analyzed. Finally, to demonstrate the methodology, a case study is given on the parameters’ sensitivity analysis of a prototype solar-powered airplane.

When using the presented methodology, the nighttime duration and the energy density of batteries are more sensitive on flight altitude of the prototype airplane.

It is not easy to design a solar-powered airplane to realize high-attitude and long-endurance flight. For the current state-of-art, it is a way to figure out the most critical parameters which need prior consideration and immediate development.

This paper provides an analytical methodology for analyzing the parameters’ sensitivities of solar-powered airplanes, which can benefit the preliminary design of a solar-powered airplane.

The purpose of this paper is to propose a methodology to determine the designing parameters for solar powered high‐altitude, long‐endurance (HALE) unmanned aerial vehicles (UAV).

By depicting solar power distribution on earth, along with the efficiencies analysis of photo‐voltaic cells (P‐cell) and lithium‐sulfur battery (LS‐battery), the influence of energy to concept design parameters is analyzed first. Second, the lift efficiency is determined from ground to 20 km for HALE UAV. Third, the methodology to determine design parameters for HALE UAV is generalized by analyzing the carrying ability of some famous HALE UAVs, such as Zephyr, Helios, and so on.

Energy is the key constraint on design of HALE UAV. The questions about where HALE UAVs are capable of operating and how long they could work can be answered according to power density distribution on earth. The total mass of HALE UAV can be divided into two parts: one is the constant mass, the other is the mass increasing with area of wing. The total mass can be estimated by the former one; the later one plays an important role in estimating wing load in the designing process.

The only way to enhance carrying ability of HALE UAVs is to redistribute their wing load: lighter structure materials and a better method to fix P‐cell with lighter fundus are the key technologies to enhance HALE UAVs’ carrying ability. At current technological levels, it is not easy to design a UAV to achieve the aim of high‐altitude long‐endurance.

This paper presents a very efficient and convenient method to determine the designing parameters of HALE UAV.

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.

High Altitude Long Endurance Unmanned Aerial Vehicle (HALE UAV) driven by a hybrid power between battery and solar panel have attracted many researchers. The HALE UAV which develops at Bandung Institute of Technology has design requirements of a 63 kg MTOW with a cruise velocity of 22.1 m/s at an altitude of 60,000 ft propelled by two propellers. The main problems that arise with the propellers gained from the market are these propellers cannot operate properly at the cruise phase due to inadequate thrust and high drag value. This paper aims to design a propeller that solves those problems.

The Larrabee method is used to design this propeller geometry with an output in the form of a chord and twist distribution. The CFD approach method is used to improve the design resulting from the Larrabee method.

This study shows that the inputted thrust value of the propeller designed using the Larrabee method is always higher than the thrust value resulting from the CFD simulation with a difference of around 20% so a design improvement process using CFD is required.

The analysis of propeller implementation in various mission profiles shows that this propeller can operate fully from climbing at sea level to cruising flight at an altitude of 60,000 ft. The same procedure can be applied in other HALE UAV cases to generate a propeller design with different objectives.

Amongst all classes of unmanned aircraft system (UAS), the rise of the Mini UAS class is the most dominant. Mini UASs are field-deployable systems and hence are not expected to operate from a runway. Therefore, the operating terrain plays an important role in the deployment and employment of the Mini UAS. However, there is limited published work in this area. The impact of terrain is more critical for military applications than civilian applications. The purpose of this paper is to explore the implications of various types of terrain on the employment and deployment of Mini UAS.

This paper explores the implications of various types of terrain on the employment and deployment of Mini UAS.

Mini UAS with field deployable requirements is often launched within the tactical battle area in case of military applications or in close proximity to the intended target area for civilian applications. Due to the size and weight of the Mini UAS, launch and recovery becomes an important factor to be considered. Rotary wing or fixed-wing vertical take-off and landing configuration UAS overcomes the limitations of Mini UAS and hence it is the preferred option. Impact of the terrain is significantly higher for military applications as compared to civil applications. Mountain terrain is the most challenging for Mini UAS operations.

This paper will help the designers configure the UAS as per the operating terrain.

Terrain affects the deployment and employment of Mini UAS and the capabilities of the system with respect to terrain in which it is expected to operate must be considered during the design of a Mini UAS. The paper will help the designers configure the UAS as per the operating terrain.

The high probability of the occurrence of separation bubbles or shocks and early transition to turbulence on surfaces of airfoil makes it very difficult to design high-lift and high-speed Natural-Laminar-Flow (NLF) airfoil for high-altitude long-endurance unmanned air vehicles. To resolve this issue, a framework of uncertainty-based design optimization (UBDO) is developed based on an adjusted polynomial chaos expansion (PCE) method.

The γ ̄Re-θt transition model combined with the shear stress transport k-ω turbulence model is used to predict the laminar-turbulent transition. The particle swarm optimization algorithm and PCE are integrated to search for the optimal NLF airfoil. Using proposed UBDO framework, the aforementioned problem has been regularized to achieve the optimal airfoil with a tradeoff of aerodynamic performances under fully turbulent and free transition conditions. The tradeoff is to make sure its good performance when early transition to turbulence on surfaces of NLF airfoil happens.

The results indicate that UBDO of NLF airfoil considering Mach number and lift coefficient uncertainty under free transition condition shows a significant deterioration when complicated flight conditions lead to early transition to turbulence. Meanwhile, UBDO of NLF airfoil with a tradeoff of performances under both fully turbulent and free transition conditions holds robust and reliable aerodynamic performance under complicated flight conditions.

In this work, the authors build an effective uncertainty-based design framework based on an adjusted PCE method and apply the framework to design two high-performance NLF airfoils. One of the two NLF airfoils considers Mach number and lift coefficient uncertainty under free transition condition, and the other considers uncertainties both under fully turbulent and free transition conditions. The results show that robust design of NLF airfoil should simultaneously consider Mach number, lift coefficient (angle of attack) and transition location uncertainty.

To provide an overview of design activity undertaken within the CAPECON Project supported by European Commission and devoted to development of HALE UAV being proposed for long endurance flights.

Selected research methods devoted mainly to the improvement of dynamic stability of unmanned aerial vehicles have been described and their application into design optimisation are shown. The main goal of this research was to improve an economic effectiveness, safety and a modular arrangement of on‐board systems, especially with respect to sensors being easy replaceable for different missions.

The research and design process included an aerodynamic optimisation of swept wing, stability analysis, weight balance, some on‐board redundant systems, reliability and maintability analysis, safety improvement, cost and performance optimisation. A number of design iterations were performed to achieve the required aircraft performances and characteristics. This iteration number was relatively moderate (four cycles only) due to employing a modern software and the essential role of theoretical analysis performed parallel to the design and redesign process.

Analysis and design methodology is limited to surveillance, high altitude long endurance platforms, where the design objectives are high reliability, safety and low cost of production and operation.

A very useful source of design information and patterns to follow, especially for engineering students and engineers dealing with unmanned aviation.

This paper offers practical help for designers being involved with an unmanned platform to be well optimised for high altitude long endurance mission, giving a lot of practical information about many aspects of longitudinal and lateral stability of Blended Wing Body configuration, on‐board sensors and systems integrated with loading structure.

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.

This paper aims to present the main results achieved in the frame of the TIVANO national-funded project which may anticipate, in a stepped approach, the evolution and the design of the enabling technologies needed for a hybrid/electric medium altitude long endurance (MALE) unmanned aerial vehicle (UAV) to perform persistent intelligence surveillance reconnaissance (ISR) military operations.

Different architectures of hybrid-propulsion system are analyzed pointing out their operating modes to select the more suitable architecture for the reference aircraft. The selected architecture is further analyzed together with its electric power plant branch focusing on electric system architecture and the selected electric machine. A final comparison between the hybrid and standard propulsion is given at aircraft level.

The use of hybrid propulsion may lead to a reduction of the total aircraft mass and an increase in safety level. However, this result comes together with a reduced performance in climb phase.

This study can be used as a reference for similar studies and it provides a detailed description of propulsion operating modes, power management, electric system and machine architecture.

This study presents a novel application of hybrid propulsion focusing on a three tons class MALE UAV for ISR missions. It provides new operating modes of the propulsion system and a detailed electric architecture of its powertrain branch and machine. Some considerations on noise emissions and infra-red traceability of this propulsion, at aircraft level.