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The main purpose of this paper is to investigate the effects of icing on unmanned aerial vehicles (UAVs) at low Reynolds numbers and to highlight the differences to icing on manned aircraft at high Reynolds numbers. This paper follows existing research on low Reynolds number effects on ice accretion. This study extends the focus to how variations of airspeed and chord length affect the ice accretions, and aerodynamic performance degradation is investigated.

A parametric study with independent variations of airspeed and chord lengths was conducted on a typical UAV airfoil (RG-15) using icing computational fluid dynamic methods. FENSAP-ICE was used to simulate ice shapes and aerodynamic performance penalties. Validation was performed with two experimental ice shapes obtained from a low-speed icing wind tunnel. Three meteorological conditions were chosen to represent the icing typologies of rime, glaze and mixed ice. A parameter study with different chord lengths and airspeeds was then conducted for rime, glaze and mixed icing conditions.

The simulation results showed that the effect of airspeed variation depended on the ice accretion regime. For rime, it led to a minor increase in ice accretion. For mixed and glaze, the impact on ice geometry and penalties was substantially larger. The variation of chord length had a substantial impact on relative ice thicknesses, ice area, ice limits and performance degradation, independent from the icing regime.

The implications of this manuscript are relevant for highlighting the differences between icing on manned and unmanned aircraft. Unmanned aircraft are typically smaller and fly slower than manned aircraft. Although previous research has documented the influence of this on the ice accretions, this paper investigates the effect on aerodynamic performance degradation. The findings in this work show that UAVs are more sensitive to icing conditions compared to larger and faster manned aircraft. By consequence, icing conditions are more severe for UAVs.

Atmospheric in-flight icing is a severe risk for fixed-wing UAVs and significantly limits their operational envelope. As UAVs are typically smaller and operate at lower airspeeds compared to manned aircraft, it is important to understand how the differences in airspeed and size affect ice accretion and aerodynamic performance penalties.

Earlier work has described the effect of Reynolds number variations on the ice accretion characteristics for UAVs. This work is expanding on those findings by investigating the effect of airspeed and chord length on ice accretion shapes separately. In addition, this study also investigates how these parameters affect aerodynamic performance penalties (lift, drag and stall).

This paper aims to present an alternative airspeed computation method based on artificial neural networks (ANN) without requiring pitot-static system measurements.

The data set used to train proposed neural model is obtained from the Digital Flight Data Acquisition Unit records of a Boeing 737 type commercial aircraft for real flight routes. The proposed method uses the flight parameters as inputs of the ANN. The Levenberg–Marquardt training algorithm was used to train the neural model.

The predicted airspeed values obtained with ANN are in good agreement with the measured airspeed values. The proposed neural model can be used as an alternative airspeed computation method.

The proposed alternative airspeed computation method can be used when the air data computer or pitot-static system has failed.

The proposed method uses flight parameters as inputs for the ANN. As such, airspeed is calculated using flight parameters instead of the pitot-static system measurements.

The purpose of this study is to detect and reconstruct a fault in pitot probe and static ports, which are components of the air data system in commercial aircrafts, without false alarm and no need for pitot-static measurements. In this way, flight crew will be prevented from flying according to incorrect data and aircraft accidents that may occur will be prevented.

Real flight data collected from a local airline was used to design the relevant system. Correlation analysis was performed to select the data related to the airspeed and altitude. Fault detection and reconstruction were carried out by using adaptive neural fuzzy inference system and artificial neural networks, which are machine learning methods. MATLAB software was used for all the calculations.

No false alarm was detected when the fault test following the fault modeling was carried out at 0–2 s range by filtering the residual signal. When the fault was detected, fault reconstruction process was initiated so that system output could be achieved according to estimated sensor data.

The presented alternative analytical redundant airspeed and altitude calculation scheme could be used when the pitot-static system contains any fault condition.

Instead of using the methods based on hardware redundancy, the authors designed a new system within the scope of this study. Fault situations that may occur in pitot probes and static ports are modeled and different fault scenarios that can be encountered in all flight phases have been examined.

According to the study of the space flight market, there is a demand for space suborbital flights including commercial tourist flights. However, one of the challenges is to design a mission and a vehicle that could offer flights with relatively low G-loads. The project of the rocket-plane in a strake-wing configuration was undertaken to check if such a design could meet the FAA recommendation for this kind of flight. The project concept assumes that the rocket plane is released from a slowly flying carrier plane, then climbs above 100 kilometers above sea level and returns in a glide flight using a vortex lift generated by the strake-wing configuration. Such a mission has to include a flight transition during the release and return phases which might not be comfortable for passengers. Verification if FAA recommendation is fulfilled during these transition maneuvers was the purpose of this study.

The project was focused on the numerical investigation of a possibility to perform transition maneuvers mentioned above in a passenger-friendly way. The numerical simulations of a full-scale rocket-plane were performed using the simulation and dynamic stability analyzer (SDSA) software package. The influence of an elevator deflection change on flight parameters was investigated in two cases: a transition from the steep descent at high angles of attack to the level glide just after rocket-plane release from the carrier and an analogous transition after re-entry to the atmosphere. In particular, G-loads and G-rates were analyzed.

As a result, it was found that the values of these parameters satisfied the specific requirements during the separation and transition from a steep descent to gliding. They would be acceptable for an average passenger.

To verify the modeling approach, a flight test campaign was performed. During the experiment, a rocket-plane scaled model was released from the RC model helicopter. The rocket-plane model was geometrically similar only. Froude scales were not applied because they would cause excessive technical complications. Therefore, a separate simulation of the experiment with the application of the scaled model was performed in the SDSA software package. Results of this simulation appeared to be comparable to flight test results so it can be concluded that results for the full-scale rocket-plane simulation are also realistic.

It was proven that the rocket-plane in a strake-wing configuration could meet the FAA recommendation concerning G-loads and G rates during suborbital flight. Moreover, it was proven that the SDSA software package could be applied successfully to simulate flight characteristics of airplanes flying at angles of attack not only lower than stall angles but also greater than stall angles.

The application of rocket-planes in a strake-wing configuration could make suborbital tourist flights more popular, thus facilitating the development of manned space flights and contributing to their cost reduction. That is why it was so important to prove that they could meet the FAA recommendation for this kind of service.

The original design of the rocket plane was analyzed. It is equipped with an optimized strake wing and is controlled with oblique, all moving, wingtip plates. Its post-stall flight characteristics were simulated with the application of the SDSA software package which was previously validated only for angles of attack smaller than stall angle. Therefore, experimental validation was necessary. However, because of excessive technical problems caused by the application of Froude scales it was not possible to perform a conventional test with a dynamically scaled model. Therefore, the geometrically scaled model was built and flight tested. Then a separate simulation of the experiment with the application of this model was performed. Results of this separate simulation were compared with the results of the flight test. This comparison allowed to draw the conclusion on the applicability of the SDSA software for post-stall analyzes and, indirectly, on the applicability of the proposed rocket-plane for tourist suborbital flights. This approach to the experimental verification of numerical simulations is quite unique. Finally, a quite original method of the model launching during flight test experiment was applied.

The intention of this paper is to compare the performance of five Chipmunk aircraft, and to investigate the assumption that the performance of any operational aircraft can be considered to be representative of that type of aircraft. All the aircraft were tested, with the same instrument panel fitted, and were weighed in order to determine the variation in weight and centre of gravity. Flight tests were then conducted to investigate the following aspects of performance: turning, climbing, drag, engine power, stalling and spinning. The results, reduced from the readings taken during flight tests, showed that there are two main causes of variation between the performances of the aircraft. First the age and condition of the engine were seen to affect the power dependent aspects of performance. Secondly there was a wide variation of profile drag which was probably due to distortion of the airframes.

The paper aims to establish a comprehensive optimization analysis model for a helicopter roll on the ground and take off based on optimal control method. The trajectory and control of the entire process are studied, and the key factors affecting the helicopter takeoff distance are analyzed.

First, based on the equivalent stiffness and damping, the landing gear model is established, and a six-degree-of-freedom helicopter model is formed. Then, the simulation of the roll-on takeoff is transformed into a nonlinear optimal control problem (NOCP). Meanwhile, a hybrid single-multiple shooting method-based transcription process is used for discretizing the problem, leading to a finite nonlinear programming model, which is solved by sequential quadratic programming. Finally, the process was calculated and compared with flight test data, which verified the feasibility of the NOCP. The influence of takeoff weight, takeoff power and liftoff airspeed on the takeoff distance of the helicopter was analyzed.

The results show that the takeoff weight can be increased by 17% under the maximum takeoff power, which is roll-on takeoff at an altitude of 0 m. When the helicopter takes off with the maximum weight at an altitude of 5000 m, the liftoff airspeed should be over 49.2 km/h.

The novelty of this paper lies in the comprehensive consideration of helicopter taxiing and taking-off phases, and the application of optimal control theory to establish a comprehensive analysis model, which can quickly analyze the maximum takeoff weight, takeoff distance, optimal liftoff speed and so on. Meanwhile, the method is verified based on the flight data.

The purpose of this paper is to design a controller for the unmanned aerial vehicle (UAV).

In this study, the constrained multivariable multiple-input and multiple-output (MIMO) model predictive controller (MPC) has been designed to control all outputs by manipulating inputs. The aim of the autopilot of UAV is to keep the UAV around trim condition and to track airspeed commands.

The purpose of using this control method is to decrease the control effort under the certain constraints and deal with interactions between each output and input while tracking airspeed commands.

By using constraint, multivariable (four inputs and seven outputs) MPC unlike the relevant literature in this field, the UAV tracked airspeed commands with minimum control effort dealing with interactions between each input and output under disturbances such as wind.

The purpose of this study is to provide conflict-free operations in terminal manoeuvre areas (TMA) using the point merge system (PMS), airspeed reduction (ASR) and ground holding (GH) techniques. The objective is to minimize both total aircraft delay (TD) and the total number of the conflict resolution manoeuvres (CRM).

The mixed integer linear programming (MILP) is used for both single and multi-objective optimization approaches to solve aircraft sequencing and scheduling problem (ASSP). Compromise criterion and ε-constraint methods were included in the methodology. The results of the single objective optimization approach results were compared with baseline results, which were obtained using the first come first serve approach, in terms of the total number of the CRM, TD, the number of aircraft using PMS manoeuvres, ASR manoeuvres, GH manoeuvres, departure time updates and on-time performance.

The proposed single-objective optimization approach reduced both the CRM and TD considerably. For the traffic flow rates of 15, 20 and 25 aircraft, the improvement of CRM was 53.08%, 41.12% and 32.6%, the enhancement of TD was 54.2%, 48.8% and 31.06% and the average number of Pareto-optimal solutions were 1.26, 2.22 and 3.87, respectively. The multi-objective optimization approach also exposed the relationship between the TD and the total number of CRM.

The proposed mathematical model can be implemented considering the objectives of air traffic controllers and airlines operators. Also, the mathematical model is able to create conflict-free TMA operations and, therefore, it brings an opportunity for air traffic controllers to reduce frequency occupancy time.

The mathematical model presents the total number of CRM as an objective function in the ASSP using the MILP approach. The mathematical model integrates air traffic controllers’ and airline operators’ perspective together with new objective functions.

The air data system of an aircraft includes, amongst other items, the airspeed indicator, altimeter and machmeter all of which derive their readings from measurement of air pressures. The instruments are designed on the assumption that they will be fed with pressures from the undisturbed free‐stream in which the aircraft is flying; this is not generally the case. In flight the aircraft disturbs the air mass and in doing so causes a pressure field around itself which produces the forces necessary for lift and control. The pressure sensors which detect the airstream pressures may be located within the aircraft pressure field and the pressures transmitted to the instruments may not correspond to the free‐stream pressures; if this is so then the instrument indications will be in error.

The purpose of this paper was to elaborate the performance model of the remotely piloted aircraft systems (RPAS) which was destined for simulations of the construction characteristics, airspeeds and trajectory of flight in the controlled, non-segregated airspace according to the standard instrument departure and arrival procedures (SIDs and STARs).

This study used systems engineering approach: decomposition of RPAS performance model into components, relations and its connection with components of controlled the airspace system. Fast-time simulations (FTS) method, which included investigation of many scenarios of the system work, minimizing the number of input variables and low computing power demand, is also used.

Performance envelope of many fixed-wing RPAS was not published. The representative RPAS geometry configuration was feasible to implement. Power unit model and aerodynamic model needed to be accommodated to RPAS category. The range of aircraft minimum drag coefficient differed in the investigated range of take-off mass and wing loading.

Fixed-wing RPAS of small and medium categories cover take-off mass (25–450 kg), wing loading (40–900 N/m²) and power loading (8–40 W/N).

This is a research on integration of the RPAS in the controlled, non-segregated airspace. The results of the work may be used in broadening the knowledge of the RPAS characteristics from the perspective of operators, designers and air traffic services.

The elaborated performance model of the RPAS used the minimum number of three input variables (take-off mass, wing loading and power loading) in identification of the complete RPAS characteristics, i.e. construction features (aerodynamic, propulsion and loads) and flight parameters (airspeeds and flight trajectory).