Main Rotor Blade Research Articles

Numerical investigation of compressibility effect on dynamic stall

A rotorcraft main-rotor blade experiences a broad range of the Mach number in high-speed forward flights or rapid maneuvers. Near the boundary of the flight envelope, the rotor blade often encounters severe dynamic stall which limits the overall aerodynamic performance. The effects of the flow compressibility on dynamic stall due to the variation of the freestream Mach number are studied here. A rotor airfoil, VR-12, is computationally investigated for both static and dynamic stall conditions in the relevant range of the Mach number, 0.2 to 0.4. For the small enough Mach number such as 0.2, both static and dynamic stalls are significantly influenced by flow separation due to adverse pressure gradient with little compressibility effect. For the highest Mach number 0.4, compressibility near the leading edge is no longer negligible, forming a local supersonic pocket there. The shock-induced separation occurs near the leading edge, which dominates both the static and dynamic stall at the high Mach number. At the intermediate Mach number 0.3, it is observed that both adverse pressure gradient and weak shock affect the stall. Unsteady compressible Reynolds-averaged Navier-Stokes equations are numerically solved with the Spalart-Allmaras turbulence model in the current study. Numerical results are compared to relevant wind-tunnel test data.

Computational Investigation of Effects of Expanded Metal Foils on the Lightning Protection Performance of a Composite Rotor Blade

Lightning is one of the most important threats to the safe operation of an aircraft. Rotorcraft can experience lightning strikes on the main rotor blades and tail rotor blades which can seriously affect the rotorcraft and its components in several ways. When lightning strikes a rotor blade made of composite material, electrical–thermal multi-physical phenomena such as the Joule heating effect occur. In this study, an electrical-thermal computational simulation of lightning strikes on composite rotor blades was performed. The simulation analyzed various situations including the presence of a lightning protection system on the composite rotor blade. Emphasis was placed on the effects of geometric parameters of expanded metal foil on the lightning protection performance. In addition, the effects of curvature on thermal and electrical response were investigated. The thermally damaged area was found to substantially increase in inner layers compared to the flat plate case.

An inverse damage location problem applied to AS-350 rotor blades using bat optimization algorithm and multiaxial vibration data

In this study, a damage identification method is proposed using both the finite element method and the bat optimization algorithm applied to the AS-350 helicopter main rotor blade. First, the structure is numerically modeled and evaluated with and without the presence of induced damages. In a second approach, an inverse problem of optimization is constructed in order to identify certain damages in terms of its position and severity level. Three different objective functions are evaluated according to the modal parameters of the rotor blade (vibrations in x, y and z directions). Numerical results, through analysis of variance, showed that local damage significantly modifies the modal response into a non-linear aspect. The modal response used was able to identify, with great efficiency, the actual (noise simulated) damages induced in terms of location and severity. Accordingly, a damage identification method is developed in order to better handle any measurement data (to find/regarding) structural changes (or damages) in complex aerospace structures. The obtained results from these numerical examples indicate that the proposed approach can detect true damage locations and estimate damage magnitudes with satisfactory accuracy, even under high measurement noise.

Performance of Quiet Helicopter

Abstract Noise generated by helicopters is one of the main problems associated with the operation of rotorcrafts. Requirements for reduction of helicopter noise were reflected in the regulations introducing lower limits of acceptable rotorcraft noise. A significant source of noise generated by helicopters are the main rotor and tail rotor blades. Radical noise reduction can be obtained by slowing down the blade tips speed of main and tail rotors. Reducing the rotational speed of the blades may decrease rotor thrust and diminish helicopter performance. The problem can be solved by attaching more blades to main rotor. The paper presents results of calculation regarding improvement of the helicopter performance which can be achieved for reduced rotor speed but with increased number of rotor blades. The calculations were performed for data of hypothetical light helicopter. Results of simulation include rotor loads and blade deformations in chosen flight conditions. Equations of motion of flexible rotor blades were solved using the Galerkin method which takes into account selected eigen modes of the blades. The simulation analyzes can help to determine the performance and loads of a quiet helicopter with reduced rotor speed within the operational envelope of helicopter flight states.

Statistical learning for turboshaft helicopter accidents using logistic regression.

The objective of this work is to advance the understanding of helicopter accidents by examining and quantifying the association between helicopter-specific configurations (number of main rotor blades, number of engines, rotor diameter, and takeoff weight) and the likelihood of accidents. We leverage a dataset of 8,338 turboshaft helicopters in the U.S. civil fleet and 825 accidents from 2005 to 2015. We use the dataset to develop a logistic regression model using the method of purposeful selection, which we exploit for inferential purposes and highlight the novel insights it reveals. For example, one important question for the design and acquisition of helicopters is whether twin-engine turboshaft helicopters exhibit a smaller likelihood of accidents than their single-engine counterparts, all else being equal. The evidence-based result we derive indicates that the answer is contingent on other covariates, and that a tipping point exists in terms of the rotor diameter beyond which the likelihood of accidents of twin-engines is higher (worse) than that of their single-engine counterparts. Another important result derived here is the association between the number of main rotor blades and likelihood of accidents. We found that for single-engine turboshaft helicopters, the four-bladed ones are associated with the lowest likelihood of accidents. We also identified a clear coupling between the number of engines and the rotor diameter in terms of likelihood of accidents. In summary, we establish important relationships between the different helicopter configurations here considered and the likelihood of accident, but these are associations, not causal in nature. The causal pathway, if it exists, may be confounded or mediated by other variables not accounted for here. The results provided here lend themselves to a rich set of interpretive possibilities, and because of their significant safety implications they deserve careful attention from the rotorcraft community.

Processing of videos of the fatigue cracks propagation in the rotor blades of the helicopter

The article describes the processing of videos of growing fatigue cracks in the main rotor blades of a helicopter, which are formed during their fatigue tests on the stand. A specially created computer program allows to determine the geometrical parameters of a crack based on video recording (length, width, extreme points, center of gravity). Each frame of the video is binarized, after which a binary matrix is created on its basis, which is subject to element-by-element processing. The obtained data are presented in the form of a graph showing the dependence of the crack growth rate on the time of its growth. The graph is traced in the Mathcad program, as a result, two data vectors are created — the time and crack growth rate, which undergo an interpolation procedure with the construction of a third-degree polynomial. According to the results of the integration of the polynomial, the critical length of the crack is determined, after reaching which its rapid growth begins.

헬리콥터 주로터 블레이드의 동적밸런싱 시험에 대한 데이터베이스 설계 및 적용

헬리콥터 주로터 블레이드의 동적 밸런싱 시험은 헬리콥터의 비행시험 전 각 블레이드의 트랙과 피치로드 하중을 동일한 수준으로 조절하기 위해 수행되는 지상회전시험이다. 이러한 동적밸런싱 시험의 목적은 블레이드 간 특성차이로 인해 나타나는 주로터 시스템 진동을 감소시키기 위함이다. 고흥 항공센터의 훨타워 시험설비에서는 오랜 기간에 걸쳐 수리온 주로터 블레이드에 대한 동적밸런싱 시험을 수행하고 데이터를 축적해 왔다. 이에 데이터양이 점차 방대해짐에 따라 시험결과를 효율적으로 관리하기 위한 데이터베이스 프로그램의 개발의 필요성이 제기되어 왔다. 본 연구에서는 훨타워 시험설비에서 수행한 동적밸런싱 시험결과를 바탕으로 구축한 데이터베이스 시스템에 대해 기술하였다. 본 연구에서 구현된 데이터베이스 시스템은 동적밸런싱 시험 결과를 각 정보별로 논리적으로 모델링하고 각각의 데이터가 상호 관계를 이룰 수 있도록 연결하여 관계형 구조를 이루도록 설계되었다. 아울러 본 논문을 통해 기 설계된 데이터베이스 시스템이 동적밸런싱 시험 결과의 축적과 시험 결과 분석에 효과적으로 활용될 수 있음을 보였다.

Влияние несущего винта и других элементов конструкции вертолета на качество функционирования размещаемой на нем аппаратуры потребителей спутниковых радионавигационных систем

We analysed how signal diffraction caused by the main rotor, signal reflection from helicopter components, and receiver antenna shielding by the main rotor blades affect the signal-to-noise ratio in the correlator band of satellite radionavigation system user equipment for GLONASS and GPS, its navigation precision and error probability when transmitting overhead information from navigation spacecraft to the helicopter. The investigation showed that the effect of helicopter design on the performance of user equipment mounted on it manifests as antenna shielding and diffraction waves appearing at the edges. The presence of diffraction waves and reduction in the original signal amplitude as a result of shielding the user equipment antenna leads to the Bit Error Rate in the overhead information transmission channel increasing by one or two orders of magnitude. Effect on navigation accuracy is negligible.

MATHEMATICAL MODELS OF THE RADAR SIGNAL REFLECTED FROM A HELICOPTER MAIN ROTOR IN APPLICATION TO INVERSE SYNTHESIS OF ANTENNA APERTURE

Introduction. The basis for solving the problem of aircraft recognition is the formation of radar portraits, reflecting the constructive features of aerial vehicles. Portraits, which are radar images of the propellers of aerial vehicles, have high informativeness. These images allow us to distinguish the number and relative position of the propeller blades, as well as the direction of its rotation. The basis for obtaining such images are mathematical models of reflected signals. Objective. The aim of this paper is to develop mathematical models of the radar signal reflected from the helicopter main rotor applied to inverse synthetic aperture radar (ISAR). Methods and materials. ISAR processing is used to produce a radar image of a propeller in a radar with a monochromatic probing signal. The propeller blades in the models are approximated by different geometric shapes. The models used to describe the reflection from the propellers of helicopters and fixed-wing aircraft have significant differences. In the process of moving each blade of the helicopter main rotor makes characteristic movements (flapping, dragging, feathering), as well as bends in a vertical plane. Such movements and bendings of the blades are influence the phase of the signal reflected from the main rotor. It is necessary to take the phase change of the reflected signal into account as accurately as possible when developing an ISAR algorithm for imaging the main rotor. Results. We found that in the centimeter wavelength range the mathematical model of the signal reflected from the helicopter main rotor as a system of blades is most accurately described by representing each blade with a set of isotropic reflectors located on the main rotor’s blade leading and trailing edges. Taking into account the flapping movements and curved shapes of the blades in the model allows you to get as close as possible to the features of the real signal. Conclusion. The developed model which takes into account the flapping movements and bends of the helicopter main rotor blades can be used to improve the ISAR algorithms providing the radar imaging of aerial vehicles.

КОНСТРУКТИВНО-ТЕХНОЛОГІЧНІ ОСОБЛИВОСТІ НЕСУЧИХ ПОВЕРХОНЬ ВЕРТОЛЬОТА З ПОЛІМЕРНИХ КОМПОЗИЦІЙНИХ МА-ТЕРІАЛІВ

The analysis of both foreign and domestic experience in the use of polymer composite materials in the structures of the lifting surfaces of a helicopter (blades, wing, stabilizer, fin), as well as the technological features of the manufacture of such structures is performed. The design and technological and operational advantages and disadvantages of composite blades compared to all-metal blades are noted. The load-carrying structures of the composite main rotor blades of a helicopter, as well as the schemes of special joints of the MR blade and the polymer composite materials package are considered. The route scheme for manufacturing of the composite blade of a helicopter is presented. The technological features of manufacturing the composite spar of a blade are analyzed by the methods of lay-up and winding-on. The scheme and route technology for manufacturing the tail section of a blade, including the scheme of lay-up and molding of composite covers and ribs, are presented. The efficiency of the equipment for automatic cutting of the roll prepreg and manual lay-up of the prepreg layers with the help of a laser projector is noted. The technological tasks of special software for lay-up CNC equipment are given. The technological transitions of the general assembly-bonding of a helicopter composite blade in the assembly jig are presented. Recommendations are given on the method of manufacturing a monoblock lifting surfaces of type of a wing, tail pylon and helicopter stabilizer from polymer composite materials using automated winding-on. A scheme of surface reinforcement is proposed. It allows to obtain a variable wall thickness from the root rib to the end rib. The trajectories of movement of the working bodies of a three-coordinate CNC winding machine for their implementation are determined. A variant of a computer simulation of the stages of lay-up of a reinforcing tape on a technological mandrel is shown. The technology of manufacturing a low aspect wing made of polymer composite materials by the method of automated winding is presented.

Design of Composite Gyrocopter Main Rotor Blade Involving Rib and Spar Elements

Abstract Gyrocopter or gyroplane is a type of rotorcraft that uses an unpowered main rotor in free autorotation to develop lift. Gyrocopter rotor blades have smaller cord length and longer span compared to helicopters blades. National Advisory Committee for Aeronautics (NACA) 8-H-12 gyrocopter rotor blade profile, unsymmetrical airfoil sections were used for this research. An attempt has been made in this work to investigate the effect of ribs and spar elements in response to applied load. Three possible modeling alternatives were studied to predict the actual induced stress and deformation of the blade: Model I is by considering the blade shell part only, Model II is blade shell with 25 numbers of ribs and without the spar element and Model III is blade shell with 25 numbers of ribs and with spar element. The rotor blade was sized based on single seat open frame and high-wind-start gyrocopter. Structural static analysis has been carried out to evaluate the strength of composite rotor blade using ANSYS Workbench 15. The results show that among these three proposed models; Model III had registered minimum Von Mises stress and deformation. Also the result reveals that by considering ribs and spar element during analysis of gyrocopter blade is crucial because, it will help to know the actual induced stress and deformation. The predicted value of induced stress and deformation is closer to the actual values will help the designer not to overdesign the parts. Consequently, the main drawbacks related to overdesign increase in weight and cost will be minimized; thereby the product operational efficiency will be improved.

Recent Experiences of Helicopter Main Rotor Blade Damage

Results of a survey investigating commonly occurring minor rotor blade damage incidents are presented in this paper. Over 100 participants worldwide ranging from test pilots to commercial pilots and licensed engineers answered the survey. The focus of this work was to provide a user-oriented context that can inform the decision-making process for integrating state-of-the-art instrumentation systems for rotor blade health monitoring onboard operational helicopters. This paper highlights the dichotomy faced by designers who have a choice to follow either a reactive strategy based on operational experience or a preventative approach based on technological trends.

A Vision-Based Technique for in-Flight Measurement of Helicopter Blade Motion

The measurement of helicopter main rotor blade angles during flight is a key capability to implement advanced applications, such as strategies for the reduction of emitted noise and to develop innovative flight control laws. The approach proposed in this work for the real-time estimation of blade angles is based on a stereoscopic system mounted on the top of the main rotor and pointing to an optical target placed on the blade root. An advanced image-processing algorithm was developed to match the target features in the left and right camera images, which was required for the 3D reconstruction of the target based on a triangulation method. This algorithm was customized for the target used in this specific application, in order to implement a procedure that is both reliable in blob matching and characterized by a very low computation effort. This allowed the system to speed up the triangulation procedure aimed at obtaining the 3-D coordinates of the features, in view of real-time applications, even with very compact processing units that can be accommodated on the main rotor head. An inverse problem for the 3-D rotation of the target was solved using the Singular Value Decomposition technique, thus improving the robustness of the measurement. The stereoscopic system was developed in order to be integrated on board an AW139 helicopter main rotor hub, equipped with a synchronous lighting device and a pre-processing unit. The latter enabled the system to automatically extract the minimum set of information to be transferred, by means of a slipring, to the processing unit hosted in the fuselage. For the full assessment of the reliability and accuracy of the integrated system, harsh dynamic and accuracy trials were conducted on laboratory test benches. The harsh dynamic tests demonstrated that the system can work continuously in realistic conditions without any structural or data acquisition problems. The accuracy tests, based on a robot test rig simulating the motion of the blade, demonstrated the capability of the system and the accuracy of the measurement technique developed. The discrepancy between the reference blade angles and the estimated ones was found to be less than 0.360 for all the realistic blade angle combinations tested.

Simulation of Boundary States of Helicopter Flight

Abstract Results of simulation of main rotor blade loads and deformations, which can be generated during boundary states of helicopter flight, are presented. Concerned cases of flight envelope include hover at maximum height, level flight at high velocity, pull-up manoeuvres applying cyclic pitch and mixed collective and cyclic control. The simulation calculations were executed for data of light helicopter with three-bladed articulated rotor. For analysis, the real blades are treated as elastic axes with distributed masses of blade segments. The model of deformable blade allows for out-of-plane bending, in plane bending, and torsion. For assumed flight state of helicopter, the equations of rotor blades motion are solved applying Runge-Kutta method. According to Galerkin method, for each concerned azimuthal position of blade the parameters of its motions are assumed as a combination of considered bending and torsion eigen modes of the blade. The loads of rotor blades generated during flight depend due to velocity of flight, helicopter mass, position of rotor axis in air and deflections of swashplate that correspond to collective and cyclic pitch angle applied to rotor blades. The results of simulations presenting rotor loads and blade deformations are shown in form of time-runs and as plots of rotor-disk distributions. The simulations of helicopter flight states may be useful for prediction the conditions of flight-tests without exceeding safety boundaries or may help to define limitations for manoeuvre and control of helicopter.

Lag-twist coupling sensitivity and design for a composite blade cross-section with D-spar

In this paper, the effect of various parameters of a specific rotor blade cross-section on the effectiveness of a twist morphing concept is investigated. Then, by considering different constraints, a cross-section consistent with this morphing concept with high lag-twist coupling and low extension-twist, is developed. This lag bending-torsion coupling is used to change the twist of the blade during the flight, while the high values of extension-twist coupling is avoided. To this end, a concentrated mass is added to the blade, where its chordwise location varies in flight. When the mass moves in the chordwise direction, a local lag bending is introduced into the blade. This in-plane bending moment then changes the blade twist distribution through lag-twist coupling induced through stiffness tailoring in the blade cross-section. Therefore, this coupling plays an important role in this morphing concept. The one-dimensional dynamics of the blade is modelled by using the geometrically exact fully intrinsic bean equations while the 2D cross-sectional stiffness values are determined by using the VABS software. First, a blade which resembles the BO-105 main rotor blade in the fundamental frequencies is designed. Then, the effect of various parameters of the cross-section on the fundamental frequencies, the lag-twist coupling, and the extension-twist coupling are determined. It is found that the skin of the spar has the highest contribution to both the extension-twist and the lag-twist coupling. Finally, a cross-section compatible with the proposed morphing concept is designed and it is demonstrated that the twist of the blade may be changed significantly.

ALGORITHM OF MESH DEFORMATION FOR ACCOUNTING CYCLIC BLADE CONTROL AND BLADES FLAPPING IN THE PROBLEM OF HELICOPTER MAIN ROTOR MODELING

This paper presents the developed algorithm for numerical grid deformation for solving the problems of modeling the flow around the helicopter main rotor in the horizontal flight mode with allowance for flapping movements and cyclic changes in the angle of the blade installation. In general, this algorithm can be applied to simulate the aerodynamics of solid bodies deviating from its initial position at angles up to 90 degrees in the vertical and horizontal planes relative to the origin point, and also performing a rotational motion at an angle up to 90 degrees around the axis through the center of coordinates and the body mass center. The first part provides a brief overview of the existing methods of the computational grid deformation for solving various problems of numerical simulation. These include methods for rebuilding the grid, moving grids and "Chimera" grids. The second part describes the algorithms for allocating of grid deformation and for finding the final coordinate of the computational grid nodes in the presence of a predetermined blade control law. The equations of the deformation zones shape in numerical grid are given. The influence of variables on zones sizes is shown. The third part presents the results of methodological calculations confirming the performance and limitations when choosing mesh deformation zones. The influence of the size and shape of the deformation zones of the numerical grid on the quality of the mesh elements is also shown. This work is methodical in nature and is a preliminary stage in the numerical modeling of the flow around the helicopter main rotor taking into account the automatic main rotor balancing and blades flapping.

Computational Simulation of Fully Trimmed Flight of a Helicopter in Hover

Abstract In trimmed flight of a helicopter, all the forces and moments, aerodynamic, inertial, and gravitational, are in balance. Keeping the helicopter in trimmed state, needs a precise adjustment of flight controls. The methodology of simulation of a fully trimmed flight of rotorcraft has been developed and applied to simulate hover of a helicopter. The presented approach is based on a solution of Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations. In contrast to typical solutions of such problem, in the newly developed methodology, the flight controls corresponding to the trimmed-flight conditions are also determined based on the solution of URANS equations. The methodology is based on coupling of several computational models of Computational Fluid Dynamics and Flight Dynamic. The URANS equations are solved in a three-dimensional region surrounding the flying helicopter, using the ANSYS FLUENT code. The approach is truly three-dimensional, with truly modelled geometry and kinematics of main and tail rotor blades. This applies to modelling of blade flapping and lead-lag motion, too. The trimming procedure uses six independent parameters (i.e. collective and cyclic pitch of main rotor blades, collective pitch of tail rotor blades, pitch, and bank angles of a helicopter) that should be adjusted so as to balance all forces and moments acting on the helicopter. The detailed description of the developed methodology as well as the results of simulation of trimmed hover of the helicopter was presented.

Cracking of Erosion Strip of Main Rotor Blade of a Military Helicopter

Erosion strip fitted on main rotor blade (MRB) of a military helicopter (flying in coastal/marine environment) had cracked in service. The erosion strip was made of austenitic stainless steel material of specification MIL-S-5059D Type 301 and was in quarter hard condition. The chemical composition, the hardness and the microstructure (condition) of the erosion strip met all the specified requirements. Two cracks were observed in the erosion strip. Scanning electron microscopy of the fracture surfaces showed the presence of fatigue striations amidst thick corrosion products indicating toward failure by corrosion fatigue. Energy-dispersive X-ray spectroscopy (EDS) showed that the corrosion products on the fracture surfaces were predominantly oxides of iron consisting of corrosive species like chlorine, which is typical of corrosion in marine/coastal environment (i.e., the service environment of the helicopter). Fractographic studies showed that fatigue failure had originated from the inner surface of the erosion strip that is bonded to the MRB. Intergranular and transgranular cracks and corrosion were observed in the erosion strip material from its inner surface at the failure origin. These appeared to be stress corrosion cracks as a result of crevice corrosion that had occurred on the inner surface of the erosion strip at the failure origin. Chlorides present in the environment (viz. marine/coastal) had apparently concentrated inside the crevice (at the bonding between the erosion strip and the MRB), attacked the protective oxide surface film on stainless steel and also prevented the re-forming of the protective oxide layer. Thus, the material in the crevice was attacked rapidly and the corrosion continued. Stress corrosion cracks can develop from the base of crevice where the stress concentration is greatest. Thus, the cracking of the erosion strip is a typical example of environment-induced mechanical failure of a fairly passive metallic material.

A simulation investigation of helicopter ground resonance phenomenon

PurposeThe purpose of this paper is to present a simulation method applied for investigation of helicopter ground resonance phenomenon.Design/methodology/approachThe considered physical model of helicopter standing on ground with rotating rotor consists of fuselage and main transmission gear treated as stiff bodies connected by elastic elements. The fuselage is supported on landing gear modeled by spring-damper units. The main rotor blades are treated as set of elastic axes with lumped masses distributed along blade radius. Due to Galerkin method, parameters of blades motion are assumed as a combination of bending and torsion eigen modes. A Runge–Kutta method is applied to solve equations of motions of rotor blades and helicopter fuselage.FindingsThe presented simulation method may be applied in preliminary stage of helicopter design to avoid ground resonance by proper selection of landing gear units and blade damper characteristics.Practical implicationsGround resonance may occur in form of violently increasing mutual oscillations of helicopter fuselage and lead-lag motion of rotor blades. According to changes of stiffness and damping characteristics, simulations show stable behavior or arising oscillations of helicopter. The effects of different blade balance or defect of blade damper are predicted.Originality/valueThe simulation method may help to determine the envelope of safe operation of helicopter in phase of take-off or landing. The effects of additional disturbances as results of blades pitch control as swashplate deflection are introduced.

State-of-the-Art Finite Element Modeling of Rotorcraft Main Rotor Blade for Bird Strike Damage Analysis

This paper demonstrates the state-of-the-art composite modeling methodology to investigate the bird strike phenomenon using available numerical bird models through experimental tests and simulation tools. The present work is based on the application of nonlinear explicit finite element analysis to simulate the response of rotorcraft main rotor blade root end under high velocity impact load. The damage behavior of blade made of composites under soft body impact depends upon bird size, blade size, blade span-wise location of impact and bird orientation with respect to hitting location, blade rotational speed and rotorcraft cruise speed. Bird model is considered as hydrodynamic with length-to-diameter ratio of 2. A bird strike event is characterized by loads of high intensity and short duration. A transient explicit nonlinear finite element-based impact analysis using Autodyn has been carried out to predict bird strike resistance to withstand 1.0 kg bird at critical flight condition. Numerical analysis indicates that blades do not tear which agrees well with the physical test conducted.