Main Rotor Hub Research Articles

Helicopter performance enhancement by alleviating retreating blade stall using active flow control

In this study, a model was created numerically to study the effect of active flow control on helicopter blade to improve the retreating blade stall. Active flow control is applied to helicopter blades to improve retreating blade stall along with the overall helicopter performance. The novel aspect of this study is using active control in a helicopter with a complex aerodynamic environment and using it in all significant flight modes. This has the advantage of practical application since active control can be used constantly without having to turn it off. The flow control is established by using blowing jets at the leading edge of the airfoil that is used at the tip part of the main rotor blade. 2D and 3D computational fluid dynamics (CFD) Turbulent-unsteady models are developed to investigate the effect of applying blowing jets on different helicopter modes of flight (hover and forward flight). Different excitation parameters and turbulence models are compared to get the maximum possible enhancement of the aerodynamic characteristic of the flow both in 2D and 3D. The results are validated against a NASA helicopter main rotor hub with known geometry and available performance parameters. Both models illustrated a good agreement with the published benchmark problem. The results revealed that the active control blowing jets can enhance the stall characteristics and aerodynamic performance of helicopter blades. The method delays the flow separation and increase the blade lift by about 10% in addition to a 40% increase in the lift to drag ratio. Thus, we can conclude that the present active control technique can enhance the helicopter in forward flight by delaying the retreating blade stall. This research offers insightful information on the usage of blowing jets for rotor flow control.

A method for reducing parasite drag of medium transport helicopter

At present, more and more stringent emission requirements are being imposed on helicopters. Carbon dioxide emissions are a function of fuel consumption, and fuel consumption is directly dependent on the power required for a helicopter cruise flight. The power, required for horizontal flight of a helicopter directly depends on the design of the fuselage and other non-lifting elements, and accordingly, on parasite drag. Therefore, the desire to reduce parasite drag is expected and understandable. Helicopter parasite drag reduction was hailed as one of the key topics of the Seventh Clean Sky Framework Conference. For sixty years of research on this topic, a tremendous amount of experience has been accumulated. In particular, it was found that the main contributors to the total parasite drag of the helicopter were the main rotor hub and swashplate. The subject matter of this article is the study of the parasite drag of the hub and swashplate of a medium passager/transport helicopter Mi-8/17. This study aimed to develop methods to reduce the parasite drag of the Mi-8/17 helicopter. Mi-8/17 is the most mass-produced twin-engine helicopter in the world. Currently, more than 4,5 thousand units are in operation around the world. Therefore, the task of reducing parasite drag for this helicopter is very relevant. Modern methods of computation fluid dynamics (CFD) are used. Modern methods of computational fluid dynamics with an acceptable degree of accuracy make it possible to consider the interference phenomena of the components of the hub, swashplate, and helicopter fuselage, which is not always possible, to obtain correctly during expensive wind tunnel tests. During the work, the following results were obtained. An analysis was made of the contribution of each element of the design of the hub and the main rotor swashplate to the total parasite drag. A method for reducing the parasite drag of a helicopter using a mast fairing is proposed. Several variants of the mast fairing were considered, and the most effective one was chosen. The effect, as well as its degree of influence, of the airflow blown by the fan cooling the air-oil radiators, on the general parasite drag of the helicopter was revealed. The results of the studies, given in this article have been implemented in the production.

Numerical study of the main rotor hub and swashplate parasite drag estimation of a medium transport helicopter

This work is devoted to the determination of the parasite drag of the components of the main rotor hub and the swashplate of a medium transport helicopter Mi-8 type. As well as interference of components between themselves and the fuselage. The calculations were carried out using modern methods of computational fluid dynamics (CFD). The work was carried out taking into account the decision of The European Commission 7th Framework Programme "Clear Sky" Joint Technology Initiative (JTI).

Effectiveness of the Compound Helicopter Configuration in Rotorcraft Performance Increase

Abstract The article presents the results of calculations applied to compare flight envelopes of varying helicopter configurations. Performance of conventional helicopter with the main and tail rotors, in the case of compound helicopter, can be improved by applying wings and pusher propellers which generate an additional lift and horizontal thrust. The simplified model of a helicopter structure, consisting of a stiff fuselage and the main rotor treated as a stiff disk, is applied for evaluation of the rotorcraft performance and the required range of control system deflections. The more detailed model of deformable main rotor blades, applying the Galerkin method, is used to calculate rotor loads and blade deformations in defined flight states. The calculations of simulated flight states are performed considering data of a hypothetical medium class helicopter with the take-off mass of 6,000kg. In the case of both of the helicopter configurations, the articulated main rotor hub is taken under consideration. According to the Galerkin method, the elastic blade model allows to compute blade deformations as a combination of the blade bending and torsional eigen modes. Introduction of additional wing and pusher propellers allows to increase the range of operational speed over 300 km/h. Results of the simulation are presented as time-runs of rotor loads and blade deformations and in a form of disk distribution plots of rotor parameters. The simulation method can be useful in defining requirements for a high speed rotorcraft.

Anomaly of Rotor Dynamics in Ultra-Light Helicopter – Robinson R22

Abstract The article presents the analysis of anomaly of rotor dynamics in ultra-light helicopter - Robinson R22. Robinson R22 is two-seat, two-blade main rotor and single-engine helicopter, well known as simple and common used aircraft because of low price and high availability. At the same time, large number of accidents and strictly defined rules (recommended piloting technique) and weather condition for safe flight show disadvantages of Robinson R22. The reason for considering this topic is the analysis of different flight properties and helicopter behaviour as well as easy entering into dangerous flight manoeuvre. In the article different flight properties and loss of control during the pull-up, manoeuvre or vertical gusts of wind and mast bumping accidents were analysed. Analysis shows that problem may be caused by construction of three-hinged rotor hub, designed and patented by F. Robinson. Article presents model of rotor hub and review of main rotor and rotor hub construction in light helicopters. Because of number of accidents, caused by the unusual behaviour of Robinson R22, restrictive pilotage rules were introduced: prohibition of flight in certain weather conditions, the necessity of attending additional training conducted by trained instructors. To reduce the probability of an accident a special instruction for specific Robinson R22 properties was created. Moreover, the statistics of accidents resulting from loss of control and review of legal changes caused by Robinson R22 accidents are presented.

Investigating the impact of wakes on single turbine availability

The current trend in wind energy leads to closely packed wind farms. These short distances result in a higher turbulence and consequently higher loads. In this paper an approach is proposed to evaluate single turbine availability losses due to the wake effect by modeling material fatigue behavior of selected turbine components. First, a characteristic wake scenario is defined: While one turbine is exposed to a wind field conforming to wind class IIB, another turbine experiences higher turbulence intensities according to Frandsen’s wake model. Using a dynamic multibody simulation, turbine loads are calculated which then are analyzed using the rainflow counting algorithm (giving load range and cycle counts) before being extrapolated to design lifetime of 20 years. Subsequently, artificial material curves are estimated representing the minimum load cycles to withstand during design lifetime. The slope of these curves is determined using the slope of a representative S-N-curve for each component. Position is then determined applying Miner’s rule with unity damage at the end of design lifetime. Lastly, failure behavior is approached by using a lognormal probability distribution. As a result increased failure probabilities for the rotor hub unit (main bearing and rotor hub) are found in the wake case. First results for an annual time in wake fraction of 20 % show an increase in total failure probability from 2.30 % to up to 41.85 % at a distance between the two turbines of 3D. Applying a mean down time per failure from literature (based on modern EU wind farms) finally leads to an overall technical availability decrease due to hub unit failure of 0. 035 % in the wake case compared to the free stream case.

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.

Optimal Control Approach to Helicopter Noise Abatement Trajectories in Nonstandard Atmospheric Conditions

This paper discusses the development of a software suite that aims to optimize helicopter trajectories with respect to the noise impact on the ground. The software suite has been developed around an advanced gradient-based optimization algorithm based on optimal control theory. The helicopter trajectories are modeled using an eight-degree-of-freedom flight mechanics model. To determine the noise impact on the ground, a helicopter noise model developed in this work consists of three modules, namely, a helicopter source noise model, a propagation model, and a noise impact model. To determine the noise levels on the ground, the source noise levels are determined from a database of aeroacoustically determined noise levels for varying flight conditions and projected on a hemisphere centered at the main rotor hub. The noise propagation model included in the suite is capable of determining the propagation loss between the source and the receiver in nonstandard atmospheric conditions, and it yields the total noise level in individual receiver locations. Finally, these noise levels can be quantified into a single noise impact criterion, which can be used as an optimization criterion in the optimal control formulation. To exemplify the capabilities of the suite, a hypothetical city center approach procedure is optimized for the noise impact in communities surrounding the helispot.

Single rotor helicopter parasite drag estimation in the preliminary design stage

This paper presents the methodology for the rapid estimation of the single rotor helicopter parasite drag in the forward flight in the preliminary design phase. The aerodynamic drag of fuselage, main rotor hub, landing gear and the other, non-rotating defective surfaces contribute with more than 50% of the total power required for the forward flight at higher cruising speeds and therefore requires a more approximate determination in the design stages before the prototype development. The parasite drag determination is based on the evaluation of the basic value of The Equivalent Parasite Drag Flat Plate Area and its changes depending on the fuselage angle of attack. Plenty of the statistical indicators, aerotunnel tests results and the recent CFD investigations were taken into consideration during the process of developing and testing this methodology.

Investigation of Engine–Airframe Vibration Due to Main Rotor Hub Loads Using a Substructuring Framework

This paper presents a methodology to analyze the coupled structural dynamic response of an elastic airframe and engines of a helicopter in response to main rotor hub loads. Transfer functions of individual components (airframe, engine, mount struts, and torque tube) are coupled together using a substructuring approach to obtain consistent coupled solutions of the entire system. Using this approach, a twin-engine, four-bladed helicopter is analyzed using NASTRAN-based models of the airframe and engines. This efficient substructuring approach is validated against the fully coupled NASTRAN model using forced response studies. Characteristics of the mount properties, i.e., the torque tube stiffness, and aft mount stiffness and damping are systematically varied to study their effect on the engine vibration response. The fore and aft mount element properties for minimizing the 8P engine response are identified without increasing 4P response. A compromise between 4P and 8P response is also identified from parametric studies of rear mount properties, using just three parameters to represent the design space. Using the substructuring approach presented here, future studies can be performed to rapidly match airframe characteristics with available engines at approximately 1000 times the speed of running a detailed finite element model (millions of degrees of freedom), without any reduction in accuracy.

Development of wireless strain-gauge measurement system for structural tests

A wireless sensor system for data acquisition and processing during structural and fatigue tests is considered. A prototype of wireless strain-gauge measurement system is used in fatigue tests of a utility helicopter’s main rotor hub. The setup of the test rig and software are described.

Design of automatic rotor blades folding system using NiTi shape memory alloy actuator

This present paper will study the requirements for development of a new Automatic Rotor Blades Folding (ARBF) system that could possibly solve the availability, compatibility and complexity issue of upgrading a manual to a fully automatic rotor blades folding system of a helicopter. As a subject matter, the Royal Malaysian Navy Super Lynx Mk 100 was chosen as the baseline model. The aim of the study was to propose a design of SMART ARBF's Shape Memory Alloy (SMA) actuator and proof of operating concept using a developed scale down prototype model. The performance target for the full folding sequence is less than ten minutes. Further analysis on design requirements was carried out, which consisted of three main phases. Phase 1 was studying the SMA behavior on the Nickel Titanium (NiTi) SMA wire and spring (extension type). Technical values like activation requirement, contraction length, and stroke- power and stroke-temperature relationship were gathered. Phase 2 was the development of the prototype where the proposed design of stepped-retractable SMA actuator was introduced. A complete model of the SMART ARBF system that consisted of a base, a main rotor hub, four main rotor blades, four SMA actuators and also electrical wiring connections was fabricated and assembled. Phase 3 was test and analysis whereby a PINENG-PN968s-10000mAh Power Bank's 5 volts, which was reduced to 2.5 volts using LM2596 Step-Down Converter, powered and activated the NiTi spring inside each actuator. The bias spring (compression type), which functions to protract and push the blades to spread position, will compress together with the retraction of actuators and pull the blades to the folding position. Once the power was removed and SMA spring deactivated, the bias spring stiffness will extend the SMA spring and casing and push the blades back to spread position. The timing for the whole revolution was recorded. Based on the experimental analysis, the recorded timing for folding sequence is 2.5 minutes in average and therefore met the required criteria.

A Case Study on the Air Flow Characteristics of the Hirobo-FALCON 505 Controllable Helicopter's Main Rotor Blade

The aerodynamics of the helicopter rotor is one of the most elating and exigent tribulations faced by the aerodynamicists today. Study through flow visualization process plays a key role in understanding the airflow distinctiveness and vortex interaction of the helicopter main rotor blade. Inspecting and scrutinizing the effects of wake vortices during operation is a great challenge and imperative in designing effective rotor system. This study aimed to visualize the main rotor airflow pattern of the Hirobo-FALCON 505 controllable subscale helicopter and seek for the vortex flow at the blade tip. The experimental qualitative data is correlated with quantitative data to perform scrupulous study on the airflow behavior and characteristics along with its distinctiveness spawned by the main rotor blade. Simulation using design software is performed in analogous stipulations to endow with comparability between the flow visualization results. Throughout the blade span several dissimilar flow patterns have been identified. The main rotor hub has turbulent flow at its center due to low energy of air amassed in this region whereas in the middle portion of the rotor blade, the air encompasses high kinetic energy with a clear straight streamline pattern.

Failure Analysis of a Helicopter's Main Rotor Bearing

Presented results report some of the findings of a detailed failure analysis carried out on a main rotor hub assembly, which had symptoms of burning and mechanical damage. The analysis suggests environmental degradation of the grease which causes pitting on bearing-balls. The consequent inefficient lubrication raises the temperature which leads to the smearing of cage material (brass) on the bearing-balls and ultimately causes the failure. The analysis has been supported by the microstructural studies, thermal analysis and micro-hardness testing performed on the affected main rotor bearing parts.

무베어링 헬리콥터 로터의 지상공진 불안정성 특성 해석

The ground resonance instability of a helicopter with bearingless main rotor hub were investigated. The ground resonance instability is caused by an interaction between the blade lag motion and hub inplane motion. This instability occurs when the helicopter is on the ground and is important for soft-inplane rotors where the rotating lag mode frequency is less than the rotor rotational speed. For the analysis, the bearingless rotor was composed of blades, flexbeam, torque tube, damper, shear restrainer, and pitch links. The fuselage was modeled as a mass-damper-spring system having natural frequencies in roll and pitch motions. The rotor-fuselage coupling equations are derived in non-rotating frame to consider the rotor and fuselage equations in the same frame. The ground resonance instabilities for three cases where are without lead-lag damper and fuselage damping, with lead-lag damper and without fuselage damping, and finally with lead-lag damper and fuselage damping. There is no ground resonance instability in the only rotor-fuselage configuration with lead-lag damper and fuselage damping.

Numerical simulation of the slant fracture of a helicopter's rotor hub with ductile damage failure criteria

ABSTRACTThe main rotor hub is one of the most critical components in helicopter structures. Thus, to assess the behaviour of the component, a fatigue test until failure has been carried out on a real Ti–6Al–4V hub. A fracture mechanics analysis is used to estimate the critical crack dimensions. Subsequently, the slant fracture due to a propagating fatigue crack has also been simulated numerically with an explicit finite element model, using the Bao–Wierzbicki ductile damage criterion. Fracture toughness and failure calibration of Ti alloy have been carried out to obtain reliable results. The numerical fracture surface obtained has been compared with the experimental one.

Failure analysis of a composite main rotor helicopter hub

A full scale fatigue test of a helicopter main rotor hub MR/H has been performed by spectrum loading. The test has been supported by a detailed finite element model of the component providing relevant information about its behaviour under loading. The MR/H has been made of Ti–6Al–4V wrapped up by a circular carbon band. Stress and strain from FE model have been validated by strain survey addressing the most critical locations. Fatigue life evaluation has been carried out by analysis in good agreement with the test results. The combined approach by test and validated numerical models has proved to be a reliable tool for fatigue substantiation of critical parts.

회전익 항공기 형상의 기체공력 특성에 관한 실험적 연구

This paper describes the test carried out on an experimental study of fuselage drag and stability characteristics of a helicopter configuration and the test techniques developed for the testing and the lessons learned in the Agency for Defense Development Low Speed Wind Tunnel(ADD-LSWT). The main objective of this test is to determine the drag and stability characteristics of helicopter configurations according to the various configuration changes. The fuselage model with a highly modular structure is a representation of 1:8 scale of the external contour of the conceptual design helicopter configuration with rotating main rotor hub including blade stubs capable of rotating up to 500 rpm. The test results are compared with the available similar data and fair to good agreement is obtained.

Detection and Avoidance of Main Rotor Hub Moment Limits on Rotorcraft

There is interest in developing carefree maneuvering capability on future rotorcraft. This has driven the need for advanced algorithms that predict the onset of structural limits and cue the pilot using tactile feedback. A new method is presented for the detection and avoidance of static load limits on the main rotor hub. An algorithm was developed that uses linear models to estimate constraints on longitudinal and lateral cyclic stick positions that ensure the transient response of hub moments remain bounded within prescribed limits. The system was tested using a high-fidelity nonlinear simulation of a UH-60A helicopter (GENHEL). The system was shown to be successful in simultaneously constraining stick travel in both lateral and longitudinal axes to prevent hub moment limit violations. The most critical conditions occurred during control reversals, at which point the system effectively imposed rate limits on the stick motion. The algorithm was shown to be robust to changes in aircraft weight and c.g. location.

49892 Measurements of acoustic properties of the composite materials constituting the main rotor hub of the Agusta-Westland helicopter EH-101 (civil version)

49892 Measurements of acoustic properties of the composite materials constituting the main rotor hub of the Agusta-Westland helicopter EH-101 (civil version)