In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-...In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-degree-of-freedom pitch-plunge aeroelastic oscillations was coupled with the unsteady Reynolds-averaged Navier-Stokes equations to perform flutter simulation.Meanwhile,the influence mechanism of SDR on flutter boundary is explained through aerodynamic work and the correlated shock wave location.The results show that the SDR delays the shock wave shifting downstream,and the Mach number corresponding to reaching freeze region increases as the split angle increases.Therefore,the peak value of aerodynamic moment coefficient amplitude and the sharp ascent process of phase occurs at higher Mach number,which leads to the delay in the occurrence of the transonic dip.Besides,before the transonic dip of airfoil without SDR occurs,the aerodynamic moment phase of airfoil with the SDR decreases slowly due to the decrease in the speed of shock wave moving downstream.This results in an increased flutter speed when employing the SDR before the transonic dip of airfoil without SDR occurs.Meanwhile,the effects of asymmetric split angles on the transonic flutter characteristics are also investigated.Before the transonic dip of airfoil without SDR occurs,the flutter characteristic is dominated by the smaller split angle.展开更多
Limit Cycle Oscillation(LCO)quenching of a supercritical airfoil(NLR 7301)considering freeplay is investigated in transonic viscous flow.Computational Fluid Dynamics(CFD)based on Navier-Stokes equations is implemented...Limit Cycle Oscillation(LCO)quenching of a supercritical airfoil(NLR 7301)considering freeplay is investigated in transonic viscous flow.Computational Fluid Dynamics(CFD)based on Navier-Stokes equations is implemented to calculate transonic aerodynamic forces.A loosely coupled scheme with steady CFD and an efficient graphic method are developed to obtain the aerodynamic preload.LCO quenching phenomenon is observed from the nonlinear dynamic aeroelastic response obtained by using time marching approach.As the airspeed increases,LCO appears then quenches,forming the first LCO branch.Following the quenching region,LCO occurs again and sustains until the divergence of the response,forming the second LCO branch.The quenching of LCOs was addressed physically based on the aerodynamic preload and the linear flutter characteristic.An“island”of stable region is observed in the flutter boundary,i.e.the flutter speed versus the mean Angle of Attack(AoA).The LCO quenches when the aerodynamic preload crosses this stable region with the increasing of airspeed.The LCO quenching of this model in transonic flow is essentially induced by destabilizing effect from aerodynamic preload,since the flutter speed is sensitive to AoA due to aerodynamic nonlinearity.展开更多
The Unsteady Adaptive Stochastic Finite Elements(UASFE)approach is a robust and efficient uncertainty quantification method for resolving the effect of random parameters in unsteady simulations.In this paper,it is sho...The Unsteady Adaptive Stochastic Finite Elements(UASFE)approach is a robust and efficient uncertainty quantification method for resolving the effect of random parameters in unsteady simulations.In this paper,it is shown that the underlying Adaptive Stochastic Finite Elements(ASFE)method for steady problems based on Newton-Cotes quadrature in simplex elements is extrema diminishing(ED).It is also shown that the method is total variation diminishing(TVD)for one random parameter and for multiple random parameters for first degree Newton-Cotes quadrature.It is proven that the interpolation of oscillatory samples at constant phase in the UASFE method for unsteady problems results in a bounded error as function of the phase for periodic responses and under certain conditions also in a bounded error in time.The two methods are applied to a steady transonic airfoil flow and a transonic airfoil flutter problem.展开更多
The trend of increasing the power-to-weight ratios of aircraft turbofan engines and efficiency of steam turbines leads to designs with long and slender blades often operating at transonic flow conditions.Such blades a...The trend of increasing the power-to-weight ratios of aircraft turbofan engines and efficiency of steam turbines leads to designs with long and slender blades often operating at transonic flow conditions.Such blades are prone to undesirable and possibly destructive vibra-tions caused by engine-order excitation or induced by flow itself.To shed more light on this problem and to extend the existing knowledge,this paper presents experimental and numerical study on torsional mode vibration of one blade in a linear blade cascade of flat profiles.In this study,dynamic loading and pressure distributions were investigated at subsonic,supercritical and transonic flow regimes while the blade was kinematically excited by a motor and shaft mechanism at reduced frequencies up to k Z 0.47.Dynamic flow structure development was documented and analyzed based on numerical simulations.Furthermore,dependence of energy transfer over an oscillation cycle on frequency and exit Mach number was investigated.Results revealed significant hysteresis in the flow field configuration particularly at supercrit-ical and transonic cases.Hysteresis is manifested namely by different development of supersonic regions when the oscillating blade passes through the zero deflection during upstroke and downstroke.Resulting aerodynamic moment is non-harmonic and there is an increasing phase lag with respect to the blade deflection when oscillation frequency increases.In majority of investigated regimes,hysteresis resulted in aerodynamic damping of the blade oscillation.展开更多
文摘In this paper,a series of flutter simulations are carried out to investigate the effects of split drag rudder(SDR)on the transonic flutter characteristic of rigid NACA 64A010.A structural dynamic model addressing two-degree-of-freedom pitch-plunge aeroelastic oscillations was coupled with the unsteady Reynolds-averaged Navier-Stokes equations to perform flutter simulation.Meanwhile,the influence mechanism of SDR on flutter boundary is explained through aerodynamic work and the correlated shock wave location.The results show that the SDR delays the shock wave shifting downstream,and the Mach number corresponding to reaching freeze region increases as the split angle increases.Therefore,the peak value of aerodynamic moment coefficient amplitude and the sharp ascent process of phase occurs at higher Mach number,which leads to the delay in the occurrence of the transonic dip.Besides,before the transonic dip of airfoil without SDR occurs,the aerodynamic moment phase of airfoil with the SDR decreases slowly due to the decrease in the speed of shock wave moving downstream.This results in an increased flutter speed when employing the SDR before the transonic dip of airfoil without SDR occurs.Meanwhile,the effects of asymmetric split angles on the transonic flutter characteristics are also investigated.Before the transonic dip of airfoil without SDR occurs,the flutter characteristic is dominated by the smaller split angle.
基金the financial support by the National Natural Science Foundation of China(No.12102317).
文摘Limit Cycle Oscillation(LCO)quenching of a supercritical airfoil(NLR 7301)considering freeplay is investigated in transonic viscous flow.Computational Fluid Dynamics(CFD)based on Navier-Stokes equations is implemented to calculate transonic aerodynamic forces.A loosely coupled scheme with steady CFD and an efficient graphic method are developed to obtain the aerodynamic preload.LCO quenching phenomenon is observed from the nonlinear dynamic aeroelastic response obtained by using time marching approach.As the airspeed increases,LCO appears then quenches,forming the first LCO branch.Following the quenching region,LCO occurs again and sustains until the divergence of the response,forming the second LCO branch.The quenching of LCOs was addressed physically based on the aerodynamic preload and the linear flutter characteristic.An“island”of stable region is observed in the flutter boundary,i.e.the flutter speed versus the mean Angle of Attack(AoA).The LCO quenches when the aerodynamic preload crosses this stable region with the increasing of airspeed.The LCO quenching of this model in transonic flow is essentially induced by destabilizing effect from aerodynamic preload,since the flutter speed is sensitive to AoA due to aerodynamic nonlinearity.
基金This research was supported by the Technology Foundation STW,applied science division of NWO and the technology programme of the Ministry of Economic Affairs.
文摘The Unsteady Adaptive Stochastic Finite Elements(UASFE)approach is a robust and efficient uncertainty quantification method for resolving the effect of random parameters in unsteady simulations.In this paper,it is shown that the underlying Adaptive Stochastic Finite Elements(ASFE)method for steady problems based on Newton-Cotes quadrature in simplex elements is extrema diminishing(ED).It is also shown that the method is total variation diminishing(TVD)for one random parameter and for multiple random parameters for first degree Newton-Cotes quadrature.It is proven that the interpolation of oscillatory samples at constant phase in the UASFE method for unsteady problems results in a bounded error as function of the phase for periodic responses and under certain conditions also in a bounded error in time.The two methods are applied to a steady transonic airfoil flow and a transonic airfoil flutter problem.
基金supported by the Ministry of Educa-tion Youth and Sports of the Czech Republic under the grant LUAUS23231 Origins and mechanisms of flutter and non-synchronous vibration in modern turbomachines oper-ating at wide range of regimes.
文摘The trend of increasing the power-to-weight ratios of aircraft turbofan engines and efficiency of steam turbines leads to designs with long and slender blades often operating at transonic flow conditions.Such blades are prone to undesirable and possibly destructive vibra-tions caused by engine-order excitation or induced by flow itself.To shed more light on this problem and to extend the existing knowledge,this paper presents experimental and numerical study on torsional mode vibration of one blade in a linear blade cascade of flat profiles.In this study,dynamic loading and pressure distributions were investigated at subsonic,supercritical and transonic flow regimes while the blade was kinematically excited by a motor and shaft mechanism at reduced frequencies up to k Z 0.47.Dynamic flow structure development was documented and analyzed based on numerical simulations.Furthermore,dependence of energy transfer over an oscillation cycle on frequency and exit Mach number was investigated.Results revealed significant hysteresis in the flow field configuration particularly at supercrit-ical and transonic cases.Hysteresis is manifested namely by different development of supersonic regions when the oscillating blade passes through the zero deflection during upstroke and downstroke.Resulting aerodynamic moment is non-harmonic and there is an increasing phase lag with respect to the blade deflection when oscillation frequency increases.In majority of investigated regimes,hysteresis resulted in aerodynamic damping of the blade oscillation.
