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The purpose of this paper is to devise an analytical approach to calculate conductor winding losses, considering multiple contributing aspects simultaneously. These include the geometric configuration of coil windings, frequency of the electric current and the dependency on the coil temperature, derived studying a coupled fluid–solid model considering the cooling system characteristics. The obtained results allow identifying power loss trends according to such system variables as coolant inlet temperature or overall flow rate of the motor.

An easy-to-use coupled analytical approach is applied, which is suitable for rapid estimations of the impact of parameter variation on the resulting conductor winding power losses that facilitates decision-making in the design process of electric aircraft engines.

In the considered cooling parameters, the overall conductor winding power losses vary approximately between 6 kW and 7.2 kW. More than 95 per cent of this loss is because of direct current losses. These losses cause the variation in maximal coil temperature ranging between 115°C and 170°C.

The SP260D motor is set and was currently tested in Extra 330. It recently broke two world records.

One of the current trends in aircraft engineering is electric aircraft. Advantages of electric aircraft include improved manoeuvrability because of greater torque from electric motors, increased safety because of decreased chance of mechanical failure, less risk of explosion or fire in the event of a collision and less noise. There will be environmental and cost benefits associated with the elimination of dependency on fossil fuels and resultant emissions.

The use of a novel fluid–solid interaction model for predicting conductor winding power loss of the SP260D electric aircraft motor has not been done earlier. A novel alternative derivation of the widely applied Dowell’s formula (Dowell, 1966) is presented for the estimation of proximity losses in square winding conductors.

This paper aims to present airworthiness considerations regarding a shaft of an electric motor. A fatigue lifetime prediction analysis based on one-step load spectrum is performed during high-cycle fatigue. Time-dependent normal and shear stress components are estimated using a high-fidelity digital twin built in Siemens PLM Nx Nastran as a finite element model (FEM). Linear and centrifugal acceleration as well as gyroscopic moment, motor torque, propeller thrust and thermal loads are considered. The equivalent cyclic degree of utilisation and a safety margin against the slip of a press-fitted shaft to rotor hub connection is estimated.

A load analysis using FEM is presented. The numerically obtained results are verified on an analytical and a semi-empirical basis.

The shaft of the electric motor can sustain 74 h of operation if burdened with aerobatic loads. Its load capacity equals 48% for the overall safety factor of 2.25.

The paper presents a specific, easily identifiable advance in knowledge that can be applicable in safety flight analysis issues.

The work presents a rotor of a novel lightweight electric motor for aircraft applications, which is a successor of the electric motor set recently in Extra 330E. The work delivers a computational estimation of the shaft life.

The purpose of this study was to analyze the possibility of using combat aircraft including decommissioned as a platform for launching and carrying space rockets with satellites (nano and microsatellites). Thus, an airborne-launcher-to-space-system may be attractive to countries without ground-based space rocket launch sites.

For considered launch-to-orbit system configurations, simulations of space rocket effects on aerodynamic characteristics were performed using computational fluid dynamics (CFD ANSYS Fluent) methods. In addition, experimental studies were performed in a wind tunnel to verify the numerical simulations. Discrete models of the aircraft structure were developed for analysis using finite element method (FEM). The analysis of simulated structural properties of the models was carried out to test its stiffness and mass characteristics important for solving the static and dynamic problems of the structure. The validation analyses of aircraft models were based on mass distribution estimation and matching the stiffness properties of the individual airframe structural assemblies.

The results of numerical analyses and tunnel tests indicate that the influence of carrier rockets on the change of aerodynamic and strength characteristics of the airframe is rather negligible. The aircraft can be used as launching platforms for space rockets. Simulations have indicated that the aircraft will successfully perform a mission of taking away and launching a rocket of at least about 1,000 kg total weight with a 10 kg space payload included.

The combat aircraft can be used as launch platforms for space rockets, and the air/rocket set can become the equivalent of responsive space assets for countries with small space budgets.

The work presents original results obtained by the authors during a preliminary design of a low-cost satellite launch system consisting of a carrier aircraft and a space rocket orbiter. The possibility of using decommissioned combat aircraft as air-launch-to-orbit platforms was taken into consideration. In the absence of aircraft design documentation, reverse engineering methods and techniques were used to develop aircraft geometry and airframe strength structure. Use of CFD, FEM and simulation methods to evaluate system capabilities was demonstrated. Numerical results from CFD simulations were finally verified in experimental tests.

Tandem wing aircrafts belong to an unconventional configurations group, and this type of design is characterised by a strong aerodynamic coupling, which results in lower induced drag. The purpose of this paper is to determine whether a certain trend in the wingspan impact on aircraft dynamic stability can be identified. The secondary goal was to compare the response to control of flaps placed on a front and rear wing.

The aerodynamic data and control derivatives were obtained from the computational fluid dynamics computations performed by the MGAERO software. The equations of aircraft longitudinal motion in a state space form were used. The equations were built based on the aerodynamic coefficients, stability and control derivatives. The analysis of the dynamic stability was done in the MATLAB by solving the eigenvalue problem. The response to control was computed by the step response method using MATLAB.

The results of this study showed that because of a strong aerodynamic coupling, a nonlinear relation between the wing size and aircraft dynamic stability proprieties was observed. In the case of the flap deflection, stronger oscillation was observed for the front flap.

Results of dynamic stability of aircraft in the tandem wing configuration can be found in the literature, but those studies show outcomes of a single configuration, while this paper presents a comprehensive investigation into the impact of wingspan on aircraft dynamic stability. The results reveal that because of a strong aerodynamic coupling, the relation between the span factor and dynamic stability is nonlinear. Also, it has been demonstrated that the configuration of two wings with the same span is not the optimal one from the aerodynamic point of view.