Numerical Study on Low-Reynolds Compressible Flows around Mars Helicopter Rotor Blade Airfoil

High-speed rotor rotation under the low-density condition creates a special low-Reynolds compressible flow around the rotor blade airfoil where the compressibility effect on the laminar separated shear layer occurs. However, the compressibility effect and shock wave generation associated with the increase in the Mach number (M) and the trend change due to their interference have not been clarified. The purpose is to clear the compressibility effect and its impact of shock wave generation on the flow field and aerodynamics. Therefore, we perform a two-dimensional unsteady calculation by Computational fluid dynamics (CFD) analysis using the CLF5605 airfoil used in the Mars helicopter Ingenuity, which succeeded in its first flight on Mars. The calculation conditions are set to the Reynolds number (Re) at 75% rotor span in hovering (Re = 15,400), and the Mach number was varied from incompressible (M = 0.2) to transonic (M = 1.2). The compressible fluid dynamics solver FaSTAR developed by the Japan aerospace exploration agency (JAXA) is used, and calculations are performed under multiple conditions in which the Mach number and angle of attack (α) are swept. The results show that a flow field is similar to that in the Earth’s atmosphere above M = 1.0, such as bow shock at the leading edge, whereas multiple λ-type shock waves are observed over the separated shear layer above α = 3° at M = 0.80. However, no significant difference is found in the Cp distribution around the airfoil between M = 0.6 and M = 0.8. From the results, it is found that multiple λ-type shock waves have no significant effect on the airfoil surface pressure distribution, the separated shear layer effect is dominant in the surface pressure change and aerodynamic characteristics.

Foreign Object Damage to Fan Rotor Blades of Aeroengine Part II: Numerical Simulation of Bird Impact

Bird impact is one of the most dangerous threats to flight safety. The consequences of bird impact can be severe and, therefore, the aircraft components have to be certified for a proven level of bird impact resistance before being put into service. The fan rotor blades of aeroengine are the components being easily impacted by birds. It is necessary to ensure that the fan rotor blades should have adequate resistance against the bird impact, to reduce the flying accidents caused by bird impacts. Using the contacting-impacting algorithm, the numerical simulation is carded out to simulate bird impact. A three-blade computational model is set up for the fan rotor blade having shrouds. The transient response curves of the points corresponding to measured points in experiments, displacements and equivalent stresses on the blades are obtained during the simulation. From the comparison of the transient response curves obtained from numerical simulation with that obtained from experiments, it can be found that the variations in measured points and the corresponding points of simulation are basically the same. The deforming process, the maximum displacements and the maximum equivalent stresses on blades are analyzed. The numerical simulation verifies and complements the experiment results.

Foreign Object Damage to Fan Rotor Blades of Aeroengine Part I:Experimental Study of Bird Impact

The conditions of experiment for bird impact to blades have been improved. The experiment of bird impact to the fan rotor blades of an aeroengine is carried out. Through analyzing the transient state response of blades impacted by bird and the change of blade profile before and after the impact, the anti-bird impact performance of blades in the first fan rotor is verified. The basis of anti-foreign object damage design for the fan rotor blades of an aeroengine is provided.

Modeling and Dynamic Analysis for the Foundation of a Blade-Rotor System

In order to achieve the model-based fault monitoring and diagnosis,an accurate model for the rotor system is necessary to locate and quantify faults.Since the dynamic characteristics of a blade-rotor system is influenced by foundation flexibility,the modeling and dynamic analyses on the foundation were sequentially investigated.Firstly,the effect of element size on the model convergence was investigated using the forward difference quotient as the slope of the frequency difference,which found that the model converged when the element size refined to 4mm.Secondly,a modal analysis and a harmonic response analysis were performed to obtain the dynamic characteristics of the foundation structure.Finally,an optimization to the foundation utilizing an additional stiffener was conducted to reduce the foundation response and make the critical speed far away from the working frequency band of 20—50Hz.

Detection of Blade Mistuning in a Low Pressure Turbine Rotor Resulting from Manufacturing Tolerances and Differences in Blade Mounting

For a serious prediction of vibration characteristics of any structure, a detailed knowledge of the modal characteristic is essential. This is especially important for bladed turbine rotors. Mistuning of the blading of a turbine rotor can appear due to manufacturing tolerances or because of the blading process itself due to unequal mounting of the blades into the disk. This paper investigates the mistuning of the individual blades of a low pressure turbine with respect to the effects mentioned above. Two different rotors with different aerodynamic design of the blades were investigated. The blades were mounted to the disk with a so-called hammer head root which is especially prone to mounting irregularities. For detailed investigations, the rotor was excited with a shaker system to detect the forced response behavior of the individual blades. The measurements were done with a laser vibrometer system. As the excitation of rotor structure was held constant during measurement, it was possible to detect the line of nodes and mode shapes as well. It could be shown that the assembly process has an influence on the mistuning. The data were analyzed and compared with numerical results. For this, different contact models and boundary conditions were used. The above described characterization of modal behavior of the rotor is the basis for the upcoming aeroelastic investigations and especially for the blade vibration measurements of the rotor, turning with design and off-design speeds.

ON INFLUENCE OF KINEMATICS TO EQUIVALENT LINEAR DAMPING OF HELICOPTER BLADE HYDRAULIC DAMPER

An analytical model of hydraulic damper was presented in forward flight accounting for pitch/flap/lag kinematic coupling and its nonlinear force-velocity curve. The fourth order Runge-Kutta was applied to calculate the damper axial velocity in time domain. Fourier series based moving block analysis was applied to calculate equivalent linear damping in terms of transient responses of damper axial velocity. Results indicate that equivalent linear damping will be significantly reduced if pitch/flap/lag kinematic coupling introduced for notional model and flight conditions.

Transverse vibration of the blade for unmanned micro helicopter using rayleigh-ritz method

A rotor manipulation mechanism for micro unmanned helicopter utilizing the inertia and the elasticity of the rotor is introduced. The lagging motion equation of the rotor blades is established, and then the natural frequencies and mode shapes of the blade for the helicopter are studied by using beam characteristic orthogonal polynomials by the Rayleigh-Ritz method. The variation of natural frequencies with the speed of rotation and the mode shapes at different rotational speeds are plotted. The using of orthogonal polynomials for the bending shapes enables the computation of higher natural frequencies of any order to be accomplished without facing any difficulties.

Static aeroelasticity analysis of a rotor blade using a Gauss-Seidel fluid-structure interaction method

The present study introduces a Gauss-Seidel fluid-structure interaction(FSI)method including the flow solver,structural statics solver and a fast data transfer technique,for the research of structural deformation and flow field variation of rotor blades under the combined influence of steady aerodynamic and centrifugal forces.The FSI method is illustrated and validated by the static aeroelasticity analysis of a transonic compressor rotor blade,NASA Rotor 37.An improved local interpolation with data reduction(LIWDR)algorithm is introduced for fast data transfer on the fluid-solid interface of blade.The results of FSI calculation of NASA Rotor 37 show that when compared with the radial basis function(RBF)based interpolation algorithm,LIWDR meets the interpolation accuracy requirements,while the calculation cost can be greatly improved.The data transmission time is only about 1%of that of RBF.Moreover,the iteration step of steady flow computation within one single FSI has little impact on the converged aerodynamic and structural results.The aerodynamic load-caused deformation accounts for nearly 50%of the total.The effects of blade deformation on the variations of aerodynamic performance are given,demonstrating that when static aeroelasticity is taken into account,the choke mass flow rate increases and the peak adiabatic efficiency slightly decreases.The impact mechanisms on performance variations are presented in detail.

Experimental and Numerical Investigation of Non-synchronous Blade Vibration Excitation in a Transonic Axial Compressor

The complex flow phenomenon of rotating instability(RI) and its induced non-synchronous vibration(NSV) have become a significant challenge as they continuously increase aerodynamic load.This study aims to provide an understanding of the non-synchronous blade vibration phenomenon caused by the rotating instability of a transonic axial compressor rotor.In this case,blade vibrations and non-synchronous excitation are captured by strain gauges and unsteady wall pressure transducer sensors.Unsteady numerical simulations for a full-annulus configuration are used to obtain the non-synchronous flow excitation.The results show that the first-stage rotor blade exhibits an NSV close to the first bending mode;NSV is accompanied by a sharp increase in pressure pulsation;amplitude can reach 20%,and unsteady aerodynamic frequency will lock in a structural mode frequency when the blade vibrates in a large-amplitude motion.The predicted NSV frequency aligns well with the experimental results.The dominant mode of circumferential instability flow structure is approximately 47% of the number blades,and the cell size occupies 2-3 pitches in the circumferential direction.The full-annulus unsteady simulations demonstrate that the streamwise oscillation of the shedding and reattachment vortex structure is the main cause of NSV owing to the strong interaction between the tip leakage and separation vortices near the suction surface.

Heat Transfer and Flow Structure of a Turbine Blade's Air-cooled Leading Edge Considering Different Hole Shapes and Additional Flow Angles

A numerical study is conducted to elucidate the impact of hole shapes and additional flow angles on the flow structure of the coolant and temperature field in the leading edge area of the gas turbine rotor.Four typical hole shapes are considered for the GE-E3 blade.The impact of the additional flow angle(E)within each hole shape on the temperature field is investigated.The results indicate that for the leading edge area and suction surface,the fan-shaped hole case performs best in decreasing temperatures,with a decrease of about 43 K.This is mainly due to the fact that the fan-shaped hole has the maximum expansion in hole spanwise direction.For the pressure surface,the console hole case performs best in decreasing temperatures,with a maximum reduction of about 47.2 K.The influence of E on the surface temperature at leading edge area varied between the different hole shapes.For the cylinder hole and console hole,the E=-20°case has the lowest area-averaged temperature.Because both the fan-shaped hole and the 7-7-7 shaped hole are expansion holes,the pattern of variation of the leading edge area temperature with increasing E is similar for the fan-shaped hole case and 7-7-7 shaped hole case.The E=20°case shows the lowest spanwise-averaged temperature near the hole outlet,and the E=-20°case shows the lowest spanwise-averaged temperature further downstream.

叶顶刷子汽封在汽轮机上的应用

叙述了叶顶刷子汽封的结构、原理、作用、安全性、寿命,介绍了其在200MW汽轮机动叶叶顶上实际使用后的良好效果。刷子汽封可用于大、中、小型汽轮机和其它涡轮机上。

Aerodynamic design optimization of helicopter rotor blades including airfoil shape for hover performance

This study proposes a process to obtain an optimal helicopter rotor blade shape for aerodynamic performance in hover flight. A new geometry representation algorithm which uses the class function/shape function transformation （CST） is employed to generate airfoil coordinates. With this approach, airfoil shape is considered in terms of design variables. The optimization process is constructed by integrating several programs developed by author. The design variables include twist, taper ratio, point of taper initiation, blade root chord, and coefficients of the airfoil distribution function. Aerodynamic constraints consist of limits on power available in hover and forward flight. The trim condition must be attainable. This paper considers rotor blade configuration for the hover flight condition only, so that the required power in hover is chosen as the objective function of the optimization problem. Sensitivity analysis of each design variable shows that airfoil shape has an important role in rotor performance. The optimum rotor blade reduces the required hover power by 7.4% and increases the figure of merit by 6.5%, which is a good improvement for rotor blade design.

Experimental Study of the Vibration Phenomenon of Compressor Rotor Blade Induced by Inlet Probe Support

An experimental investigation was conducted in order to understand the installation effects of inlet measurement probes on the vibration characteristics of the rotor blades in two axial compressors.The vibration signal of the rotor blades was analyzed for different layouts of the inlet measurement probes.For the three-stage axial compressor,the first-order resonance occurs on the first stage of the rotor blades at an engine order excitation condition,which is induced by the cylindrical probe support with a diameter of 10 mm.When the size of the probe support is decreased,the vibration level reduces evidently.In contrast,for the six-stage axial compressor,the first-order resonance occurs on the first stage of the rotor blades at an excitation source of 6th order,which is triggered by the inlet measurement probes and the upstream struts.When the number of the inlet measurement probes is changed,the resonance of the rotor blades vanishes.

A fully automated framework for helicopter rotor blades design and analysis including aerodynamics,structure, and manufacturing

This study describes an integrated framework in which basic aerospace engineering aspects(performance, aerodynamics, and structure) and practical aspects(configuration visualization and manufacturing) are coupled and considered in one fully automated design optimization of rotor blades. A number of codes are developed to robustly perform estimation of helicopter configuration from sizing, performance analysis, trim analysis, to rotor blades configuration representation. These codes are then integrated with a two-dimensional airfoil analysis tool to fully design rotor blades configuration including rotor planform and airfoil shape for optimal aerodynamics in both hover and forward flights. A modular structure design methodology is developed for realistic composite rotor blades with a sophisticated cross-sectional geometry. A D-spar cross-sectional structure is chosen as a baseline. The framework is able to analyze all realistic inner configurations including thicknesses of D-spar, skin, web, number and ply angles of layers of each composite part,and materials. A number of codes and commercial software(ANSYS, Gridgen, VABS, Pre VABS,etc.) are implemented to automate the structural analysis from aerodynamic data processing to sectional properties and stress analysis. An integrated model for manufacturing cost estimation ofcomposite rotor blades developed at the Aerodynamic Analysis and Design Laboratory(AADL),Aerospace Information Engineering Department, Konkuk University is integrated into the framework to provide a rapid and dynamic feedback to configuration design. The integration of three modules has constructed a framework where the size of a helicopter, aerodynamic performance analysis, structure analysis, and manufacturing cost estimation could be quickly investigated. All aspects of a rotor blade including planform, airfoil shape, and inner structure are considered in a multidisciplinary design optimization without an exception of critical configuration.

Numerical Simulation of Ice Accretion Phenomena on Rotor Blade of Axial Blower

Ice accretion is the phenomenon that super-cooled water droplets impinge and accrete on a body. It is well known that ice accretion on blades and airfoils leads to performance degradation and severe accidents. For this reason, experimental investigations have been carried out using flight tests or icing tunnels. However, it is too expensive, dangerous, and difficult to set actual icing conditions. Hence, computational fluid dynamics is useful to predict ice accretion. A rotor blade is one of jet engine components where ice accretes. Therefore, the authors focus on the ice accretion on a rotor blade in this study. Three-dimensional icing phenomena on the rotor blade of a commercial axial blower are computed here, and ice accretion on the rotor blade is numerically investigated.

Dynamic behavior of aero-engine rotor with fusing design suffering blade off

Fan blade off（FBO） from a running turbofan rotor will introduce sudden unbalance into the dynamical system,which will lead to the rub-impact,the asymmetry of rotor and a series of interesting dynamic behavior.The paper first presents a theoretical study on the response excited by sudden unbalance.The results reveal that the reaction force of the bearing located near the fan could always reach a very high value which may lead to the crush of ball,journal sticking,high stress on the other components and some other failures to endanger the safety of engine in FBO event.Therefore,the dynamic influence of a safety design named ‘‘fusing＂ is investigated by mechanism analysis.Meantime,an explicit FBO model is established to simulate the FBO event,and evaluate the effectiveness and potential dynamic influence of fusing design.The results show that the fusing design could reduce the vibration amplitude of rotor,the reaction force on most bearings and loads on mounts,but the sudden change of support stiffness induced by fusing could produce an impact effect which will couple with the influence of sudden unbalance.Therefore,the implementation of the design should be considered carefully with optimized parameters in actual aero-engine.

Multi-probe linear fitting and time of arrival linear correction method to analyze blade vibration based on blade tip timing without once-per-revolution

Blade vibration monitoring can ensure the safe operation of aeroengine rotor blades.Among the methods of blade vibration monitoring,Blade Tip Timing(BTT)method has attracted more and more attention because of its advantages of non-contact measurement.However,it is difficult to install the Once-Per-Revolution(OPR)probe in the confined space of aeroengine,and the failure and instability of the OPR signal will reduce the reliability of the blade vibration analysis results,which directly affects the accuracy of the blade vibration parameters identification.The Multi-Probe linear fitting and Time of Arrival(ToA)Linear Correction method based on the BTT(MP-LC-BTT)without OPR is proposed to reduce the errors of single probe linear fitting method for blade vibration displacement analysis.The proposed method can also correct the calculation error of blade vibration displacement due to the nonlinear change of rotation speed,which can improve the analysis accuracy of the blade vibration displacement.A new blade vibration model conforming to the actual vibration characteristics is established,and the effectiveness of the proposed method is verified by numerical simulation.Finally,the reliability and accuracy of the MP-LC-BTT method have been verified by the experiments which include two high-speed blade test-benches and an industrial axial fan.This method can be used in the actual aero-engine monitoring instead of the BTT method with OPR.

Kill probability modeling of multi-purpose guided missile against gunship and its application

Aimed at current deficiencies of multi-purpose guided missile kill probability model against gunship, the concept of the important coefficient of vulnerability blade unit is proposed in this paper. Laser fuze actuation model and warhead condition kill probability model of rotor blades are established by Monte Carlo method and kinetics theory with new ideas. Based on limited data, armor thickness of gunship is estimated, and a complete multi-purpose guided missile kill probability mathematical model is established, which provides necessary mathematical tool for the accurate and objective analysis of multi-purpose guided missile kill probability against gunship. Based on the establishment of the model, sensitivity analysis and optimal design of the main factors of multi-purpose guided missile kill probability are conducted, and the results show that the single multi-purpose guided missile lethality performance can be improved significantly by sensitivity analysis and optimization.

Numerical Investigation on Super-cooled Large Droplet Icing of Fan Rotor Blade in Jet Engine

Icing(or ice accretion) is a phenomenon in which super-cooled water droplets impinge and accrete on a body.It is well known that ice accretion on blades and vanes leads to performance degradation and has caused severe accidents.Although various anti-icing and deicing systems have been developed,such accidents still occur.Therefore,it is important to clarify the phenomenon of ice accretion on an aircraft and in a jet engine.However,flight tests for ice accretion are very expensive,and in the wind tunnel it is difficult to reproduce all climate conditions where ice accretion can occur.Therefore,it is expected that computational fluid dynamics(CFD),which can estimate ice accretion in various climate conditions,will be a useful way to predict and understand the ice accretion phenomenon.On the other hand,although the icing caused by super-cooled large droplets(SLD) is very dangerous,the numerical method has not been established yet.This is why SLD icing is characterized by splash and bounce phenomena of droplets and they are very complex in nature.In the present study,we develop an ice accretion code considering the splash and bounce phenomena to predict SLD icing,and the code is applied to a fan rotor blade.The numerical results with and without the SLD icing model are compared.Through this study,the influence of the SLD icing model is numerically clarified.

Criterion of aerodynamic performance of large-scale offshore horizontal axis wind turbines

With the background of offshore wind energy projects, this paper studies aerodynamic performance and geometric characteristics of large capacity wind turbine rotors （1 to 10 MW）, and the main characteristic parameters such as the rated wind speed, blade tip speed, and rotor solidity. We show that the essential criterion of a high- performance wind turbine is a highest possible annual usable energy pattern factor and a smallest possible dimension, capturing the maximum wind energy and producing the maximum annual power. The influence of the above-mentioned three parameters on the pattern factor and rotor geometry of wind turbine operated in China＇s offshore meteoro- logical environment is investigated. The variation patterns of aerodynamic and geometric parameters are obtained, analyzed, and compared with each other. The present method for aerodynamic analysis and its results can form a basis for evaluating aerodynamic performance of large-scale offshore wind turbine rotors.