Huu Duc Vo

Criteria for Spike Initiated Rotating Stall

A computational study to define the phenomena that lead to the onset of short length-scale (spike) rotating stall disturbances has been carried out. Based on unsteady simulations, we hypothesize there are two conditions necessary for the formation of spike disturbances, both of which are linked to the tip clearance flow. One is that the interface between the tip clearance and oncoming flows becomes parallel to the leading-edge plane. The second is the initiation of backflow, stemming from the fluid in adjacent passages, at the trailing-edge plane. The two criteria also imply a circumferential length scale for spike disturbances. The hypothesis and scenario developed are consistent with numerical simulations and experimental observations of axial compressor stall inception. A comparison of calculations for multiple blades with those for single passages also allows statements to be made about the utility of single passage computations as a descriptor of compressor stall.

Impact of Pressure and Temperature on the Performance of Plasma Actuators

FB = body force induced on the fluid by the actuator per unit of actuator span, mN=m FNet = net actuator force of (FB F ) per unit of actuator span, mN=m F = shear force at the wall, induced by the plasma actuator flow, per unit of actuator span, mN=m u = axial velocity component, m=s v = normal velocity component, m=s x = axial position with respect to actuator, m y = normal position above surface, m = air density, kg=m

Control of Rotating Stall in Axial Compressors Using Plasma Actuators

This paper proposes the use of SDBD plasma actuators to suppress rotating stall inception and extend the stable operating range of axial compressors. Plasma actuators may provide a practical low-power alternative to effectively increase the surge margin of aircraft engines with minimal or even positive impact on compressor performance. A computational study is carried out on a representative subsonic modern compressor rotor geometry to evaluate the proposed casing plasma actuation for suppression of short (spike) as well as long (modal) length-scale rotating stall inception based on their respective flow physics. The objective is to assess the optimum actuator location and required actuation strength to achieve the desired effects at low and medium subsonic compressor speed. Results show that plasma actuation near the rotor leading edge and concentrated in the tip clearance gap region most effectively suppresses both of the criteria for spike stall inception and delays the predicted stall point to a lower flow coefficient with relatively low power input. In addition, the observed increase in rotor pressure rise characteristic from the proposed actuation means that the concept, with a new suggested actuator modification, can also be used to suppress modal stall inception. The simulations indicate that actuation effectiveness decreases with increasing compressor speed and that stronger actuation strength than that of conventional SDBD plasma actuators may be needed. Some implications for the practical implementation of this concept on real compressors are also discussed.

Suppression of Short Length-Scale Rotating Stall Inception With Glow Discharge Actuation

This paper proposes and investigates the pioneering use of glow discharge (plasma) actuation to suppress short length-scale (spike) rotating stall inception. A single dielectric barrier discharge plasma actuator basically consists of two parallel offset thin electrodes separated by a dielectric material. The application of a high frequency AC voltage across the electrodes results in an induced body force on the flow adjacent to the surface. This simple, robust actuator may provide a practical low-power mean to positively alter the tip clearance flow dynamics responsible for spike stall inception. A computational study is carried out on a low-speed compressor rotor with the implementation of a published plasma actuation model in an established turbomachinery CFD code. The objective is to provide a preliminary assessment of the effectiveness of a casing circumferential plasma actuator, with varying actuator location, input voltage and frequency, in suppressing the two flow criteria associated with the formation of spike disturbances leading to stall. Results show that plasma actuation most effectively suppresses both of these flow criteria when placed near the rotor leading edge and delays the predicted stall point to a lower flow coefficient with minimal power input. The simulations also indicate that the effectiveness of the actuation decreases non-linearly with input voltage and frequency. In addition, results indicate that this technology could perhaps be used for suppression of both short and long-length scale stall inception in axial compressors.Copyright

Performance Improvement of Axial Compressors and Fans with Plasma Actuation

This paper proposes the use of plasma actuator to suppress boundary layer separation on a compressor blade suction side to increase axial compressor performance. Plasma actuators are a new type of electrical flow control device that imparts momentum to the air when submitted to a high AC voltage at high frequency. The concept presented in this paper consists in the positioning of a plasma actuator near the separation point on a compressor rotor suction side to increase flow turning. In this computational study, three parameters have been studied to evaluate the effectiveness of plasma actuator: actuator strength, position and actuation method (steady versus unsteady). Results show that plasma actuator operated in steady mode can increase the pressure ratio, efficiency, and power imparted by the rotor to the air and that the pressure ratio, efficiency and rotor power increase almost linearly with actuator strength. On the other hand, the actuators position has limited effect on the performance increase. Finally, the results from unsteady simulations show a limited performance increase but are not fully conclusive, due possibly to the chosen pulsing frequencies of the actuator and/or to limitations of the CFD code.

Reduction of Fan and Compressor Wake Defect Using Plasma Actuation for Tonal Noise Reduction

This paper proposes a new technique to reduce the noise generated by rotor-stator interaction (tonal noise) in fans and compressors. The method involves the use of single dielectric barrier discharge (plasma) actuators near the blade trailing edge to reduce blade wakes. Plasma actuators are a new and simple type of active flow control device consisting of two parallel and offset electrodes separated by a layer of dielectric material. The application of a high ac voltage at high frequency to the electrodes generates a body force on the flow in the vicinity of the electrodes to inject momentum without mass addition. A preliminary assessment of the proposed concept is performed with a computational study on modern low-speed compressor rotor geometry. A plasma actuator model is implemented in an established turbomachinery CFD code. Simulations are carried out to evaluate the effect of the actuator strength, location, and actuation method (continuous versus pulsed) on the rotor wake. Results show that plasma actuators operated in continuous mode near the trailing edge can significantly influence the wake of the rotor with relatively little power consumption. The effectiveness of the actuation is proportional to actuator strength (induced body force). The exact position of the actuator in the trailing edge region has little effect on the effectiveness of the actuation. The results from simulations with pulsed actuation show very low time-averaged influence on the wake and are not fully conclusive, due possibly to the frequencies simulated and the limitations of the RANS CFD tool.

Blade Tip Clearance Flow and Compressor Nonsynchronous Vibrations: The Jet Core Feedback Theory as the Coupling Mechanism

This paper investigates the role of tip clearance flow in the occurrence of nonsynchronous vibrations (NSVs) observed in the first axial rotor of a high-speed high-pressure compressor in an aeroengine. NSV is an aeroelastic phenomenon where the rotor blades vibrate at nonintegral multiples of the shaft rotational frequencies in operating regimes where classical flutter is not known to occur. A physical mechanism to explain the NSV phenomenon is proposed based on the blade tip trailing edge impinging jetlike flow, and a novel theory based on the acoustic feedback in the jet potential core. The theory suggests that the critical jet velocity, which brings a jet impinging on a rigid structure to resonance, is reduced to the velocities observed in the blade tip secondary flow when the jet impinges on a flexible structure. The feedback mechanism is then an acoustic wave traveling backward in the jet potential core, and this is experimentally demonstrated. A model is proposed to predict the critical tip speed at which NSV can occur. The model also addresses several unexplained phenomena, or missing links, which are essential to connect tip clearance flow unsteadiness to NSV. These are the pressure level, the pitch-based reduced frequency, and the observed step changes in blade vibration and mode shape. The model is verified using two different rotors that exhibited NSV.

Role of Tip Clearance Flow in Rotating Instabilities and Nonsynchronous Vibrations

Single- and multiple-blade-passage simulations of an isolated subsonic axial compressor rotor show that flow oscillations in the tip region, known as rotating instabilities and a driver for nonsynchronous vibrations, occur when only one of the two criteria for short-length-scale rotating stall inception is satisfied. This criterion is tip clearance backflow below the trailing-edge blade tip. The flow oscillations associated with rotating instabilities most likely result from impingement of this tip clearance backflow on the rear pressure side of the blade. This phenomenon could plausibly be modeled with an impinging jet subject to a lateral pressure gradient and lateral shear flow. The findings have important practical implications on the prediction and suppression of nonsynchronous vibrations.

Flight Control Using Wing-Tip Plasma Actuation

A concept for lift modification on a conventional aircraft wing for roll control at low angle of attack with dielectric barrier discharge plasma actuators is proposed and assessed through computational fluid dynamics simulations and preliminary wind-tunnel experiments. The concept consists of placing plasma actuators around the wing tip to add momentum in the direction opposite to that of the flow forming the tip vortex. Because of the limited strength of existing plasma actuators, the assessment is carried out for a relatively small two-dimensional wing (NACA 4418) with a rounded tip at zero angle of attack and 15 m/s for a Reynolds number in the range of 1.5 x 10 5 . Computational fluid dynamics simulations show a significant alteration of the vorticity field downstream of the trailing edge characterized by a more diffused vortex surrounded by zones of negative vorticity induced by the actuators and, but not necessarily, outboard displacement of the tip vortex. This leads to a reduced downwash that results in a change in lift of up to almost 20% for actuator strength levels that should be achievable in the short term with a new generation of dielectric barrier discharge actuators. The actuator placed on the suction side contributes the most to the lift increase, with its induced jet blocking the flow around the wind tip at the origin of the formation of the tip vortex. Wind-tunnel experimental results support the computational fluid dynamics predictions in both magnitude and trend. Furthermore, preliminary computational fluid dynamics simulations are carried out for a symmetric nonlifting wing (NACA 0018), representative of aircraft tail surfaces at zero angle of attack to generate lift for pitch and yaw control. Results indicate lift generation that increases and becomes larger than drag at higher actuator strengths. These promising results show a potential for the proposed concept to replace movable flight control surfaces on future aircraft wings and empennages.

Rotating Stall Suppression in Axial Compressors with Casing Plasma Actuation

optimumactuatorlocationandrequiredactuationstrengthtoachievethedesiredeffectsatlowandmediumsubsonic compressor speeds. Results show that plasma actuation near the rotor leading edge and concentrated in the tip clearancegapregionmosteffectivelysuppressesbothofthecriteriaforspikestallinceptionanddelaysthepredicted stall point to a lower flow coefficient with relatively low power input. In addition, the observed increase in rotor pressure-rise characteristic from the proposed actuation means that the concept, with a new suggested actuator modification,canalsobeusedtosuppressmodalstallinception.Thesimulationsindicatethatactuationeffectiveness decreaseswithincreasingrotortipspeed,thattherequiredactuatorstrengthscaleswiththisspeed,andthatstronger actuation strength than that of conventional single dielectric barrier discharge plasma actuators may be needed. Some implications for the practical implementation of this concept on real compressors are also discussed.