Experimental studies of a helicopter rotor in ground effect aim to quantify the effect of the ground proximity on their performance. As a result of the performed experimental work, dependencies between the produced thrust (T) and torque (Q), varying the rotor’s distance to the ground (H) and blade pitching angle (8) are drawn from one perspective, and the correlations between the generated thrust (T) and torque (Q), and their corresponding coefficients C_T and C_Q, varying the distance to the ground (H) and blade pitching angle (8) at several fixed rotational frequencies (n). Additionally, a dependency between the generated relative thrust coefficients (C_TIGE/C_TOGE) against the relative distance to the ground (H/D) at a fixed rotational frequency of 750 RPM, varying the blade inclination angle (8) is also derived and presented in this study. In summary, the research successfully quantified the effect of the rotor’s static characteristics in the ground effect zone and posed topics for further research.

The general objective of this paper is to present the initial results obtained as an outcome of applying a coupled Experimentally Derived – Blade Element Momentum Theory (BEMT) model for evaluating the airflow-defining parameters of a hovering helicopter rotor close to obstacles and making performance predictions. Several empirical models are described, and proper comparison with the experimentally obtained data is conducted. In detail, the characterization of the rotor inflow ratio (λ), when operating at fixed rotational frequency (n), at different relative distances to the ground (H/R), varying the pitching angle (θ) is discussed. The dependencies show an increase in the rotor inflow ratio parameter (λ), when increasing the collective pitch angle (θ) in hovering regime at fixed constant relative distance to the ground surface (H/R). On the contrary, the inflow ratio (λ) is experiencing a decline once the helicopter rotor operates closer to the ground surface. Moreover, the inflow ratio characterization (λ) along the blade span can be applied to the total generated thrust (T). As a result, the corresponding thrust coefficients (CT) are calculated and graphically represented. The overall characterization of the thrust coefficients (CT) will allow the definition of the ground effect zone.

This article represents a full estimation of helicopter rotor aerodynamic characteristics in ground effect conditions through the application of a coupled empirical blade element–momentum theory algorithm. The main focus of this research includes the evaluation of the required weighted power coefficients $\left({\displaystyle \raisebox{1ex}{$C_P$}\!\left/ \!\raisebox{-1ex}{$\sigma $}\right.}\right)$ for a hovering state in close proximity to obstacles and their relation to the weighted thrust force coefficients’ values $\left({\displaystyle \raisebox{1ex}{$C_T$}\!\left/ \!\raisebox{-1ex}{$\sigma $}\right.}\right),$ varying the relative distance from the helicopter rotational plane to the ground surface $\left({\displaystyle \raisebox{1ex}{$H$}\!\left/ \!\raisebox{-1ex}{$R$}\right.}\right)$ and the rotor’s collective pitch angle (θ). The represented numerical and experimental results show that an increase in the collective pitch angles (θ) leads to a rise in the generated weighted thrust force coefficients $\left({\displaystyle \raisebox{1ex}{$C_T$}\!\left/ \!\raisebox{-1ex}{$\sigma $}\right.}\right)$ and in the weighted power coefficients $\left({\displaystyle \raisebox{1ex}{$C_P$}\!\left/ \!\raisebox{-1ex}{$\sigma $}\right.}\right)$ for every individual fixed normalized distance from the ground surface $\left({\displaystyle \raisebox{1ex}{$H$}\!\left/ \!\raisebox{-1ex}{$R$}\right.}\right)$. Moreover, a decline in the relative distance from the ground $\left({\displaystyle \raisebox{1ex}{$H$}\!\left/ \!\raisebox{-1ex}{$R$}\right.}\right)$ requires less power to keep the rotation going in hover. The dependencies indicate that the ground effect zone covers a distance of up to 2R from the rotational plane to the ground surface.

This study represents the evaluation of the aeroelastic flutter stability characteristics of a helicopter rotor, expressed as correlations between the flutter frequency ratio and the relative center of gravity coordinate, considering the varying distance from the ground surface and the location of the aerodynamic center for a fixed flapping frequency. For every constant aerodynamic center’s coordinate, a decrease in the relative distance from the ground surface produces a reduction in the flutter stability zone for the given helicopter rotor. Conversely, moving the center of gravity coordinates in the backward direction from ${{\xi }_A}_1=0.1$ to ${{\xi }_A}_3=0.8$ leads to a rise in the flutter stability zone and then improves the stability characteristics. A nonlinear representation of the lift force coefficient variations with the angle of attack, when the reduced frequency changes from $k_1=0$ to $k_6=10$, is conducted with the incorporation of Theodorsen’s function. Eventually, an increase in the reduced frequency magnitude leads to a rise in the produced nonlinearities in the lift force coefficients when the angle of attack changes.

This study represents a comprehensive analysis of the implementation of twin counter-rotating common-axis (coaxial) rotor systems in the design process and technical application of rotorcraft and Unmanned Aerial Vehicles (UAVs). In detail, the conducted literature review clearly illustrates the already usable vehicles and the existing experimental models. The application of a system of two coaxial rotors, one above the other, rotating in opposite directions, eliminates the need for a tail rotor for the provision of directional stability and leads to several additional advantages such as the reduction in the rotorcraft’s weight and enhancement of the directional stability qualities in comparison with the single main rotor configurations. However, the inclusion of two rotors, one above the other, affects lift generation, reducing its magnitude on the lower rotor, while the lift and the thrust of the upper rotor remain relatively unchanged. The implementation of two-rotor systems requires complex algorithms with respect to cyclic and collective pitch control. Ultimately, the analyzed cases indicate the existing gap in the research into coaxial systems concerning the influence of the distance between the two rotors and the collective pitch on the generated thrust and the interferences in cross-flow conditions.

This article represents a numerical approach for estimating the helicopter rotor’s blade longitudinal dynamic characteristics considering several operational parameters. A characterization regarding the blade’s longitudinal, flapping, and vertical responses is derived by solving a system of differential equations that fully describe the rotor’s longitudinal dynamics. Dependencies between the lagging (longitudinal) velocity, the flapping velocity, and the vertical velocity over time are illustrated, taking into consideration the varying flapping frequency in a hovering regime. Additionally, the blade’s longitudinal parameters were evaluated in ground effect conditions when hovering at ${\displaystyle \frac{H}{R}}=1$ and ${\displaystyle \frac{H}{R}}=0.5$. The studied time domain variables represent the rotor’s natural reaction capabilities when a disbalancing condition occurs. Eventually, an increase in the blade’s flapping frequency leads to a rise in the required period for reaching a stable condition with regard to the longitudinal and vertical responses. The ground effect zone reduces the blade’s lagging and flapping reactions as well.