Helicopter Rotor Research Articles

Effect of spanwise groove on the dynamic stall characteristics of an airfoil

Numerical simulations are conducted to investigate the effect of triangular groove on the dynamic stall characteristics of a NACA 0012 airfoil at a Reynolds number of 135,000. The right-angled triangular grooves are placed at either 10%, 25%, or 50% chord locations on the suction and have depths of 0.025c and 0.05c, measured normal to the surface of the airfoil. The solutions that are second order accurate in time and space are obtained using pressure-based finite volume solver and the 4-equation transition SST turbulence model viz. γ- Re θt is used to predict transition and viscous stresses accurately. The airfoil is in harmonic pitch motion about its quarter-chord with a maximum circular frequency of 18.67 rad/s. The results suggest that the presence of a groove, except for the deeper grove at 0.5c, quickens the dynamic stall, but with smaller rise in C l,max and a less severe fall in lift at the stall. The mean C l value during the downstroke is improved by up to 8% for the deeper groove at 0.25c, reducing the hysteresis in lift significantly. The grooves at 0.1c, 0.25c, and 0.5c also reduce the drag by 4%, 7%, and 9% during a complete cycle, with subsequent improvements of 54%, 69%, and 63% in the l/d ratio. The current finding can be thus used to enhance the performance of flapping wing MAVs, helicopter rotors, and wind turbine blades as these applications encounter the dynamic stall phenomena frequently.

Analytical plane-stress recovery of non-prismatic beams under partial cross-sectional loads and surface forces

High levels of strength- and stiffness-to-mass ratio can be achieved in slender structures by lengthwise tailoring of their cross-sectional areas. During use, a non-prismatic beam element can be subject to surface forces or loads acting on only a part of their cross-section. Practical examples involve tapered aircraft wings, wind turbine and helicopter rotor blades under fluid pressure and shear forces; arched beams in bridges subject to vehicular traction forces and tensile stresses in tendons of prestressed concrete. Presently, beam theories generalise the external loads to the entire cross-sectional area. However, this technique does not accurately describe surface-load boundary conditions and beams under partial cross-sectional loads. Hence, an efficient analytical plane-stress recovery methodology is introduced in the present study that generalises the external load to a specific sub-area of the cross-section of homogeneous non-prismatic beams with one plane of symmetry. As a result, the transverse stress components become piecewise distributions, i.e. non-smooth but continuous in the thickness direction. Additional novelties include the boundary equilibrium recovered considering applied surface loads and terms up to second-order derivatives of the internal forces to define the transverse stress field. Closed-form solutions for the specific case of non-prismatic beams with a rectangular cross-section loaded both on top and bottom surfaces are presented. For validation purposes, different numerical examples are modelled with results compared to solid-like finite element analyses as well with relevant analytical theories. The results show that the developed formulation predicts the stress field in non-prismatic beams under surface forces and non-uniform loads applied to a part of the cross-sectional area with goods levels of accuracy. The error associated with the proposed method escalates with the taper angle, such that a 10° taper angle could result in a 6% error at the surfaces and reduced values for interior zones, while the analytical state-of-the-art models were not able to predict the transverse stresses correctly.

Development of secondary vortex structures in rotor wakes

The blade tip vortex system is a crucial feature in the wake of helicopter rotors, and its correct prediction represents a major challenge in the numerical simulation of rotor flows. A common phenomenon in modern high-fidelity CFD simulations is the breakdown of the primary vortex system in hover due to secondary vortex braids. Since they are strongly influenced by the numerical settings, the degree to which these secondary vortex structures actually physically occur is still discussed and needs experimental validation. In the current work, the development of secondary vortex structures in the wake of a two-bladed rotor in hover conditions was investigated by combining stereoscopic particle image velocimetry measurements in different measurement planes and high-fidelity simulations. Secondary vortex structures were detected and quantified at different axial locations in the wake by applying an identical scheme to the measured and simulated velocity data. In agreement, it was found that the number of secondary vortices is maximum at a distance of 0.8,R below the rotor. The more intense secondary vortex structures were quantitatively well captured in the simulation, whereas in the experiment a larger number of weaker vortices were detected. No distinct preferential direction of rotation was found for the secondary vortices, but they tended to develop in vortex pairs with alternating sense of rotation. A clustered occurrence of secondary vortices was observed close to the primary tip vortices, where the rolled-up blade shear layer breaks down into coherent vortex structures.Graphical abstract

1D modelling and PID control of helicopter Diesel engine rotational speed in torque changes

The article discusses the results of simulation tests concerning the operation of a diesel engine for a light helicopter. The tests were carried out in the AVL Boost software which is used to analyze dynamic phenomena in internal combustion engines. The research object was a newly designed diesel engine of a V8 structure and a power of 330 kW. This engine was designed to be used in the construction of a light class helicopter. The created one-dimensional simulation model included all the main engine components as well as the connection to the helicopter main transmission and the helicopter rotor. The tests consisted in selecting the P value in the PID controller used to control the amount of fuel injected into the engine. The change in the P value indirectly influenced the reaction of the engine to a change in power and torque during horizontal flight of a helicopter. These changes were introduced by changing thrust torque in the helicopter rotor. The fuel injection regulator was designed to maintain a constant engine rotational speed. The maximum speed deviations from the nominal speed of the engine operation due to both increasing and decreasing speed were analyzed. Additionally, the sum of the deviation values was analyzed until the rotational speed of the tested object stabilized. The results showed that the change of the P parameter affects all the analyzed parameters of the engine operation; however, the minimum deviation values for each parameter occur at non-equal PID settings, which makes it difficult to clearly indicate the appropriate value of the P element.

A novel high-order scheme for numerical simulation of wake flow over helicopter rotors in hover

Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor. A new high-order scheme (we call it WENO-K) proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors. WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction, and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities. Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid. To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid, a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid. Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation. Numerical results for Caradonna-Tung (C-T) rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%. Compared with WENO-JS scheme, WENO-K scheme achieves about 40% improvement on accuracy of predicting rotor thrust with only 4.1% extra computational cost. More importantly, WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270° compared to WENO-JS scheme.

Computational investigation of the aerodynamic performance of reversible airfoils for a bidirectional tidal turbine

A reversible airfoil is an airfoil that has equal performance when the flow is reversed. Such airfoils are relevant for many different applications, including use in ventilation fans, helicopter rotors, wind turbines and tidal turbines. Compared to traditional airfoils, reversible airfoils have different performance characteristics and have been less explored in the scientific literature. This work investigates the aerodynamic performance of some selected reversible airfoils using computational fluid dynamics. The selected airfoils are based on existing NACA 6 profiles and a profile using B-spline parameterization. The results show reduced performance for the reversible airfoils compared to a unidirectional airfoil. Of the investigated airfoils, the B-spline airfoil has the highest performance, with a maximum aerodynamic efficiency which is 87 % of the unidirectional design.

High-speed PIV investigation of the flow created by the model rotor in hover mode

A study of the instantaneous and average velocity and vorticity fields in the flow created by the model helicopter rotor in the hover mode was carried out. The velocity fields of the flow generated by the model rotor were obtained by a two-dimensional TR PIV system, which provided two components of the velocity vector in the diagnostic light plane. The processing of the obtained raw images was carried out using a two-frame algorithm with adaptive interrogation windows. The experiments carried out have shown the possibility of using the PIV technique to visualize the tip vortex structure descending from the rotor blade. This possibility seems to be especially interesting as one of the means of validation of the numerical methods for calculating rotor aerodynamics and acoustics.

Validation of actuator disc circulation distribution for unsteady virtual blades model

The actuator disc method is an engineering approach to reduce computer resources in computational fluid dynamics (CFD) simulations of helicopter rotors or aeroplane propellers. Implementation of an actuator disc based on rotor circulation distribution allows for approximations to be made while reproducing the blade tip vortices. Radial circulation distributions can be formulated according to the nonuniform Heyson-Katzoff “typical load” in hover. In forward flight, the nonuniform disk models include “azimuthal” sin and cos terms to reproduce the blade cyclic motion. The azimuthal circulation distribution for a forward flight mode corresponds to trimmed conditions for the disk rolling and pitching moments. The amplitude of the cos harmonic is analysed and compared here with presented in references data and CFD simulations results.

Structurally coupled characteristics of rotor blade using new rigid-flexible dynamic model based on geometrically exact formulation

In rotor dynamics, blades are normally modelled as a slender beam, in which elastic deformations are coupled with each other. To identify these coupling effects, new rigid-flexible structural model for helicopter rotor system is proposed in this paper. Finite rotations of the whole blade (on flapwise, lagwise, and torsional) are described as three global rigid degrees of freedom. The nonlinear deformation geometrics of the beam is built on geometrically exact beam theory. New expressions for blade strain energy, kinetic energy, and virtual work of various kinds of external forces are derived as functions of finite rotations and elastic deformations. To quantify the coupling characteristics, following the definition of coupling factor in electromagnetics, a new coupling factor between two modal components on each mode is introduced in modal analysis. Simulations show that the new structural model is highly capable of solving static and dynamic problems in rotor system and the maximum deformation that moderate deformation beam theory can predict might be 15% of beam length. After the new coupling factor is applied to study structurally coupled characteristics of rotor blade, it can be concluded that closeness of natural frequencies likely indicates considerable coupling between corresponding DOFs in structure.

Hovering Helicopter Rotors Modeling Using the Actuator Line Method

An implementation of the actuator line method (ALM) is applied to a hovering helicopter rotor. This method, which is widely used for wind turbine simulations, replaces the rotor blades by momentum source terms in the unsteady Reynolds-averaged Navier–Stokes equations. The removal of the blade mesh significantly reduces the computational mesh size, thus lowering the computational cost. The ALM is presented along with some improvements, notably the choice and treatment of the projection kernel. A parameter sweep is performed showcasing the importance of proper selection of the Gaussian smearing coefficient for accurate rotor performance predictions with a value of scaled around a quarter chord in size. With this value, a new set of simulations on a refined mesh is performed and analyzed covering global rotor performance coefficients, sectional blade loading, tip vortex characteristics in terms of positions, circulation, and core radius. The ALM is benchmarked against an equivalent blade resolved case on the well-known S-76 rotor. Results confirm the appropriateness of the ALM model for a hovering rotor for main flow features and performance metrics, although there was a small loss of accuracy on the tip blade loading in the presence of a blade–vortex interaction. Finally, computational performances indicate an elapsed-time speed-up between 3 and 4× in addition to a greater parallel efficiency in favor of the ALM.

Revising of the Near Ground Helicopter Hover: The Effect of Ground Boundary Layer Development

The interaction of a helicopter rotor with the ground in hover flight is addressed numerically using a hybrid Eulerian–Lagrangian CFD model. When a helicopter takes off or lands, its wake interferes with the ground. This interaction, depending on the height-to-rotor diameter ratio, causes the altering of the rotor loading and performance as compared to the unconstrained case and gives rise to the development of a complex outwash flow field in the surrounding of the helicopter. The present study aims to characterize the interactional phenomena occurring in the early stages of the rotor wake development and in particular the interference of the starting vortex with the ground boundary layer and the effect of this interaction in the motion of the vortex in the rotor outwash flow. The hybrid CFD method employed combines a standard URANS compressible finite volume solver, the use of which is restricted to confined grids around solid bodies, and a Lagrangian approximation of the entire flow field in which conservation equations are solved in their material form, disctretized using particle representation of the flow quantities. The two methods are strongly coupled to each other through an appropriate iterative scheme. The main advantage of the proposed methodology is that it can conveniently handle complex configurations with several bodies that move independently from one another, with affordable computational cost. In this paper, thrust coefficient predictions of the hybrid model are compared to predictions of a free wake code and to experimental data indicating that consistent prediction of the rotor load requires the inclusion of the ground boundary layer in the analysis. Moreover, detailed comparisons of the rotor wake evolution predicted by the hybrid model are presented.

Hysteresis Compensation for a Piezoelectric Actuator of Active Helicopter Rotor Using Compound Control.

Active rotor with trailing-edge flaps is a promising method to alleviate vibrations and noise level of helicopters. Hysteresis of the piezoelectric actuators used to drive the flaps can degrade the performance of an active rotor. In this study, bench-top tests are conducted to measure the nonlinear hysteresis of a double-acting piezoelectric actuator. Based on the experimental data, a rate-dependent hysteresis model is established by combining a Bouc–Wen model and a transfer function of a second order system. Good agreement is exhibited between the model outputs and the measured results for different frequencies. A compound control regime composed of a feedforward compensator and PID (Proportional–Integral–Derivative) feedback control is developed to suppress the hysteresis of this actuator. Bench-top test results demonstrate that this compound control regime is capable to suppress hysteresis at different frequencies from 10 Hz to 60 Hz, and errors between the desired actuator outputs and the measured outputs are reduced dramatically at different frequencies, revealing that this compound control regime has the potential to be implemented in an active helicopter rotor to suppress actuator hysteresis.

Experimental Investigation of Lift Enhancement and Drag Reduction of Discrete Co-Flow Jet Rotor Airfoil

Rotor airfoil design involves multi-point and multi-objective complex constraints. How to significantly improve the maximum lift coefficient and lift-to-drag ratio of rotor airfoil is a fundamental problem, which should be solved urgently in the development of high-performance helicopter rotor blades. To address this, discrete co-flow jet (DCFJ) technology is one methods with the most potential that can be harnessed to improve the performance of the rotor airfoil. In this study, wind tunnel experiments are conducted to study the effect of DCFJ technology on lift enhancement and drag reduction of OA312 airfoil. Furthermore, the performance improvement effects of the open co-flow jet (CFJ) and DCFJ technologies are studied. In addition, the influence of fundamental parameters, such as the obstruction factor and relative unit length, are analyzed. Results demonstrate that DCFJ technology is better than CFJ technology on the performance enhancement of the OA312 airfoil. Moreover, the DCFJ rotor airfoil can significantly reduce the drag coefficient and increase the maximum lift coefficient and the stall angle of attack. The maximum lift coefficient can be increased by nearly 67.3%, and the stall angle of attack can be delayed by about 12°. The DCFJ rotor airfoil can achieve the optimal performance when the obstruction factor is 1/2 and the relative unit length is 0.025.

Mission-based optimal morphing parameters for rotors with combined chord and twist morphing

Background: The fixed geometry rotor blades in today’s helicopters do not give the best performance throughout the duration of any mission. However, low-speed and high-speed flights have different geometrical requirements for the shape of the most efficient rotor blades. With advancements in morphing technologies, these can be applied to change the shape of the blades in different flight regimes. Methods: Two different helicopter rotor morphing concepts – namely, the linearly variable chord extension and the torque-tube based twist - under the framework of the European project SABRE were investigated for their optimal geometric parameters using a Particle Swarm Optimization (PSO) algorithm. Since the morphing parameters were dependent on the mission profile, three different missions representing typical helicopter applications were chosen. The optimization problem was posed both as single objective (power) and as multi-objective (power, tip elastic torsion and vibratory hub load). Based on the insights drawn from these investigations, a rotor was set up including both morphing concepts in a single blade. Results: The rotor with combined chord and twist morphing was shown to have better performance than the baseline blade, while keeping the penalty on the elastic torsion and vibration of the rotor to a minimum. Conclusions: Chord and twist are both important parameters determining the efficiency of a rotor blade. Since they have non-overlapping requirements, combining the two morphing concepts into a single blade can yield higher performance than the individual ones.

Design procedures and experimental verification of an electro-thermal deicing system for wind turbines

Abstract. There has been a substantial growth in the wind energy power capacity worldwide, and icing difficulties have been encountered in cold climate locations. Rotor blade icing has been recognized as an issue, and solutions to mitigate accretion effects have been identified. Wind turbines are adapting helicopter rotor and propeller ice protection approaches to reduce aerodynamic performance degradation related to ice formation. Electro-thermal heating is one of the main technologies used to protect rotors from ice accretion, and it is one of the main technologies being considered to protect wind turbines. In this research, the design process required to develop an ice protection system for wind turbines is discussed. The design approach relies on modeling and experimental testing. Electro-thermal heater system testing was conducted at the Adverse Environment Rotor Test Stand at Penn State, where wind turbine representative airfoils protected with electro-thermal deicing were tested at representative centrifugal loads and flow speeds. The wind turbine sections tested were half-scale models of the 80 % span region of a generic 1.5 MW wind turbine blade. The icing cloud impact velocity was matched to that of a 1.5 MW wind turbine at full power production. Ice accretion modeling was performed to provide an initial estimate of the power density required to de-bond accreted ice at a set of icing conditions. Varying icing conditions were considered at −8 ∘C with liquid water contents of the cloud varying from 0.2 to 0.9 g/m3 and water droplets from 20 µm median volumetric diameter to 35 µm. Then, ice accretion thickness gradients along the span of the rotor blade for the icing conditions were collected experimentally. Given a pre-determined maximum power allocated for the deicing system, heating the entire blade was not possible. Heating zones were introduced along the span and the chord of the blade to provide the required power density needed to remove the accreted ice. The heating sequence for the zones started at the tip of the blade, to allow de-bonded ice to shed off along the span of the rotor blade. The continuity of the accreted ice along the blade span means that when using a portioned heating zone, ice could de-bond over that specific zone, but the ice formation could remain attached cohesively as it is connected to the ice on the adjacent inboard zone. To prevent such cohesive retention of de-bonded ice sections, the research determined the minimum ice thickness required to shed the accreted ice mass with the given amount of power availability. The experimentally determined minimum ice thickness for the varying types of ice accreted creates sufficient tensile forces due to centrifugal loads to break the cohesive ice forces between two adjacent heating zones. The experimental data were critical in the design of a time sequence controller that allows consecutive deicing of heating zones along the span of the wind turbine blade. Based on the experimental and modeling efforts, deicing a representative 1.5 MW wind turbine with a 100 kW power allocation required four sections along the blade span, with each heater section covering 17.8 % span and delivering a 2.48 W/in.2 (0.385 W/cm2) power density.

Review of Modern Helicopter Constructions and an Outline of Rotorcraft Design Parameters

This work contains the results of a modern helicopter construction analysis. It includes the comparison of almost seventy rotorcraft constructions in terms of size in line with EASA requirements – large and small helicopters. The helicopters are also divided because of a mission purpose. The proposed division for large aircrafts is: transport, multipurpose, attack and for small aircrafts: observation, training, and utility. The aircraft construction features are described. Average dimension values of airframes and rotors are shown. Helicopter rotor arrangements are presented in terms of an operational purpose. Next, the rotorcraft design inputs are described. The mathematical formulas for design inputs are given. The ratios are calculated and gathered for the compared aircrafts. Correlation between the analysed parameters is presented on charts. Design inputs are also presented in the paper as a function of MTOW. The function trends are determined to provide an evaluation tool for helicopter designers. In addition, the parameters are presented as possible optimisation variables.

Fracture analysis of AA6061-graphite composite for the application of helicopter rotor blade

The main objective of the work is to study the fracture behavior of AA6061-graphite material using both experimental technique and finite element simulation by considering helicopter rotor blade as a case study. From the case study, it has been observed that the helicopter rotor blade, made of AA6061, has been failed at the threaded portion of the hole. Experimental fracture toughness is carried out using the compact tension specimens as per ASTM standard testing procedure. Modeling of compact tension specimens and the threaded portion of the bolt hole was utilized to analyze the fracture toughness using a simulation tool. From the results and the comparison, it is recommended to use AA6061-9wt% graphite material as a replacement of AA6061 in the application of main rotor blades of the helicopter.

Validation of correlations-based transition modeling strategies applied to the Spalart-Allmaras turbulence model for the computation of separation-induced transition

Predicting the laminar-turbulent transition is becoming a necessity in order to reduce margin costs in the design of turbines and helicopter rotors. Most RANS models lack the capacity to create turbulence sufficiently abruptly to accurately predict separation-induced transition. This article aims at evaluating the performances of the γ−Reθ,t strategy applied to the Spalart-Allmaras and k−ω−SST models (ν˜−γ−Reθ,t and k−ω−γ−Reθ,t). The chosen test case features a laminar separation bubble over a flat plate. The results are compared to the flows obtained using classical algebraic criteria and validated against DNS data. The effects of the correction by Dacles-Mariani et al. are detailed as well as those of two sets of correlations. The turbulence production mechanism of the present models is too weak downstream of the first increase of intermittency, and the laminar separation bubbles are longer than in the reference flow. More specifically, the ν˜−γ−Reθ,t model with Medida and Baeder's correlations fails to reach the correct levels of turbulence production in the fully turbulent boundary layer. Switching the correlations corrects this last behavior. Finally, using maps of the activation zones of the Dacles-Mariani correction, a sensor is exhibited that could be used in future separation-induced transition modeling efforts.

Fatigue Crack Propagation Monitoring Using Fibre Bragg Grating Sensors

The paper presents the analysis of the possibility of fatigue crack detection and monitoring its propagation process using fibre Bragg grating (FBG) sensors. The investigations were carried out on an aluminium alloy sample (a part of the Mi-2 helicopter rotor blade). During the fatigue test, the sample was equipped with FBG sensors applied for strain measurement and the vibration-based strain monitoring. It was observed that the strain curves determined by the FBG sensors agreed well with the fatigue force profile. However, the strain curves were almost insensitive to the crack propagation process, except in the last stage of the test, when the crack length was equal to 25 mm. The strain values and the natural frequencies of the sample that were determined experimentally were compared with the values achieved from the finite element method model, with both methods showing good agreement. Additionally, spectrogram-based analyses were performed, focused on the acoustic waves phenomena related to a crack propagation process. It was confirmed that the proposed signal processing method, based on spectrogram analyses, can be applied for the detection of fatigue crack development in metal structures.

Development and dynamics of a 2SPU+UPU+SP parallel rotor of helicopter

A prototype of a 2SPS+UPU+SP type parallel rotor of helicopter is developed. Its kinematics model and dynamics model are established. Firstly, the structural principle and the advanced performances of the parallel rotor of the helicopter are explained, and the degree of freedom of the parallel mechanism is calculated. Secondly, the kinematics formulas are derived for solving the displacement, Jacobian matrix, general velocity, Hessian matrix and acceleration of the moving links in the parallel mechanism. Thirdly, a dynamics model of the parallel mechanism for the helicopter rotor is established based on the derived kinematic formulas. Finally, the theoretical kinematics and dynamics solutions are verified by the simulation mechanism of the parallel rotor of the helicopter.