Rotor Tip Research Articles

Power and Flow Analysis of Axial Induction Control in an Array of Model-Scale Wind Turbines

As research on wind energy has progressed, it has broadened from a focus on the wind turbine to include the entire wind farm. In particular, methods to mitigate the negative effects of upstream wakes on downstream turbines have received significant attention. One such mitigation method is axial induction control (AIC) in which upstream turbines are derated to reduce the momentum deficits in their wakes, leaving higher speed flow for downstream turbines. If performed correctly, it is theorized that the power production gains in downstream turbines can compensate for the power sacrificed by derating upstream turbines. Previous work has indicated that the “excess” energy left in the wake of the derated turbine is along the edges of the wake such that a turbine placed directly downstream will see little to no increase in power. To address this hypothesis, we performed a control and treatment experiment with model-scale turbines in a wide flume. Five turbines were arranged in three successive streamwise rows, with the first two rows consisting of two aligned turbines, while three turbines with small transverse spacing were placed in the third row, the central of which was also streamwise-aligned with the upstream two turbines. This arrangement was used to evaluate the difference in power production primarily among the turbines in the third row when the upstream turbines were derated. Particle image velocimetry (PIV) was used to measure the wake in the streamwise-vertical planes along the centerline of the array and along the rotor tips of the centerline turbines between all rows, and high accuracy power measurements were recorded from each turbine. The results show that the total power of the array was decreased while implementing AIC but that individual turbine performance differed from predictions. PIV results show that mean kinetic energy (MKE) is redistributed to the edges of the wakes as has been previously hypothesized. We provide an analysis of the results that connects both the power and flow measurements and that highlights several of the aspects of wind turbine wake flows that make them so complex and challenging to study.

Transient analysis of the pre-whirl pump-jet propulsor with different blade numbers

ABSTRACT In this work, the verified mesh is employed to study the effect of the stator number (Ns ) and rotor number (N r) on the pump-jet propulsor performance. Based on the RNG k–ϵ turbulence model, results show that the increase of Ns has an accelerating effect on flow velocity and improves the thrust and torque. The unsteady thrust characteristics are closely related to the combination of Ns and Nr. As Ns increase, the tip vortex intensity decreases and causes a reduction in pressure fluctuation on the rotor tip regime. While for the results of different Nr with constant area ratio, the steady performance almost unchanged. And the tip vortex moves upstream as Nr increases, this results in a stronger fluctuation on the rotor pressure side. Further study of cavitating flow indicates that the increase of Nr lead to an improve of the relative cavitation range, but the cavity volume decreases.

Unsteady simulations of deposition in a rotor passage of the first-stage turbine for an aero-engine

Atmospheric particulate pollutants are prone to deposit in aero-engine turbines due to high-temperature and high-velocity gas flows. The resulting deposition changes the blade profile, leading to a degradation of aerodynamic performance, increase in surface roughness, and blockage of film cooling holes and internal cooling channels, which further reduces cooling performance of the blade. Therefore, the blades are easily ablated, especially for the rotating parts as they have high rotating speed. In present study, unsteady simulations on the effects of particle deposition were carried out by demonstrating the migration trajectories and deposition distributions of particles in turbine rotor passages of an aero-engine that operates at real engine conditions. The effects of rotating speed, blade tip clearance and its cavity depth on the deposition and migration of contaminant particulates were examined. Results reveal that the deposition on the blade surfaces varies with the rotating speeds and the rotor tip clearances. The deposits are mainly concentrated on the pressure side of the blade where multiple rebounds of the particles are observed under a cruise operating condition. At a larger tip clearance, more particles flow into the tip clearance due to stronger leakage flow, and the squealer tip increases the capture efficiency of the particles on the blade tip.

Numerical investigation of plasma actuator induced forcing direction on the performance of a low-speed isolated axial compressor rotor

This paper investigates the effects of changing the dielectric barrier discharge (DBD) plasma actuator forcing direction on stall margin improvement and aerodynamic performance of a low-speed axial compressor rotor. The effects of plasma actuator induced force at yaw angles of 0, +30, −30° relative to the rotational axis of the compressor were evaluated through the incorporation of body force spatial distribution into Navier-Stokes equation. The evaluations focused on compressor performance comprising rotor efficiency, total pressure rise coefficient, span-wise loss coefficient, and diffusion factor on one hand and the tip leakage flow structure on the other hand in each configuration. The results indicated that reduction of mass flow rate could be associated with rotor tip spillage and trailing edge backflow, leading to the occurrence of the stall phenomena and flow blockage in the rotor tip region. However, using plasma actuators led to the energized tip region flow structure and improved aerodynamic performance by reducing the size of the tip leakage vortex and flow blockage. Specifically, among proposed configurations, the forcing direction at a yaw angle of −30° improved the compressor stall margin by about 8%. This is mainly due to the higher plasma jet velocity in the relative frame of reference versus to other configurations resulting in higher injected momentum to the tip leakage flow. The results propose counter-flow induced forcing direction as a favorable configuration for stall margin improvement.

Research on the Effect of Tip Surface Coatings on High-Speed Spindles’ Noise

The contact interface between the stator and the rotor tip of the spindle could be destructed when the spindle is rotating continually at high speed, which will cause strong noise and severe vibration. In order to reduce the sound pressure level of the noise generated by the rotating spindle, three different coating materials, that is, Al-Ti-Cr-C, Ti-C and DLC, were applied to the rotor tip surface of the spindle. The effects of the coating materials on the sound pressure levels of the rotating spindles were studied by using the treated spindles and the untreated spindles. Results showed that compared with Coating Al-Ti-Cr-C, the Coating Ti-C containing only the two main elements of Ti and C produces the smallest sound pressure level in the experiment speed range; the surface roughness of Coating DLC is smaller, but the sound pressure level of the entire spindle becomes larger than Coating Ti-C; the sound pressure level of the spindles with surface coating treatment is obviously lower than that of the spindles without coating treatment. The research results can provide basic data for the design and production of noiseless spinning spindles.

Spike type of stall inception in the tandem rotor

There are two known mechanisms of inception of rotating stall in the compressor, modal wave inception and spike inception. However, the stall inception mechanisms for a tandem-rotor have not been explored so far in detail. The paper explores the unsteady behaviour of the tandem rotor. Unsteady casing pressure measurements are carried out with the help of seven equispaced unsteady pressure sensors. The casing pressure traces are analysed using wavelet transforms, fast Fourier transform, discrete spatial Fourier series, travelling wave energy method and correlation techniques to identify the nature of the stall cells. The changes observed in the stall cells during the transition from the stable mode of operation to the fully stalled mode and back from the stall regime to the stable operating condition are investigated in detail. It is observed that the stall incepts in the form of a low-intensity spike, which later evolves as a long-length scale disturbance in the fully developed stall region. The initial stall disturbance is found to be rotating with 75% of the rotor speed. The speed of the stall cells is reduced to 61% of the rotor speed in a fully developed region. The number and size of stall cells increase as the mass flow is throttled. The number and size of the stall cells are strongly coupled with the leading-edge incidence pattern. During the transition from the stall regime to the stable operating condition, the mass flow of the compressor increases, which reduces the flow incidence angle of the tandem rotor. Thus, the separation region shrinks during this transition leading to a reduction in the size of the stall cells. The stall in the tandem-rotor is primarily associated with the tip leakage flow. The rotating disturbances in the tandem-rotor originate due to the leading-edge flow separation of the forward rotor blade. Leading-edge separation of the aft rotor, which arises due to interaction of the forward rotor tip leakage vortex, is effectively suppressed with the help of the gap-nozzle flow. For a comprehensive unsteady analysis, the different techniques need to be used in combination.

Three-dimensional flow investigation of a high-pressure turbine with rotor tip desensitization based on Winglet geometry

Three-dimensional flow investigation of a high-pressure turbine with rotor tip desensitization based on Winglet geometry

Non-synchronous blade vibration analysis of a transonic fan

This study numerically investigates the aeromechanic behavior of a transonic fan model with a flat tip-leading-edge on the NASA rotor 67 test case. Single-passage unsteady calculations at a near stall operating point of 82% design speed show that the dominant frequencies of mass flow were not the harmonics of the rotor rotational frequency. A full-annulus fluid–structure interaction analysis was subsequently carried out to examine the unsteady flows and their interactions with blade vibrations. The results show that the modal displacement of the backward traveling seventh nodal diameter of the second torsion mode grew exponentially, which reveals that the blade vibration was non-synchronous. The vibration pattern indicates that the aerodynamic mode was resonant with the structural vibration mode. Around the rotor tip, the circumferential vortical propagation induced by interactions among the main flow, tip leakage flow, and tip clearance vortex was the source of aerodynamic excitation. To clarify the mechanism of the non-synchronous vibration, the coupling between aerodynamic disturbance and structural response, i.e., aliasing, was summarized. The frequency spectra of the fluctuating pressure show that an aerodynamic Backward Traveling Wave (BTW) was co-aliased to a structural BTW due to the propagation of the circumferential vortex. The correlation between the frequency and free convective speed of the aerodynamic disturbance determined the directions of aliasing.

Method of Improving Stability of a Highly-Loaded Axial Compressor Stage by Coupling Different Casing Treatments

Casing treatment is a powerful method for improving the stability of aircraft compressors. An optimized slot-type casing treatment was tested on the first rotor of a highly-loaded two-stage compressor, and the results showed that the casing treatment could not increase the compressor stability at the design and off-design speeds. A coupled casing treatment (CCT), which is built with an injector, a bridge, a plenum chamber, and several slots for a recirculating loop, is proposed and optimized to enhance the stability of the compressor in the present study. The optimized CCT improves the compressor stability and efficiency under the design condition by 75.8% and 0.71%, respectively. The coupling effect, which is established with an inner circulation in the slots and an outer circulation from the slots to the injectors, accounts for the excellent stability enhancement. The coupling effect reduces the amount of tip leakage flow, depresses the development of the tip leakage vortex (TLV), and greatly decreases the blockage in the rotor tip which is primarily induced by the interaction of the shock-wave and boundary-layer at the blade suction surface. The parametric study shows that improving the coupling effect has a positive effect on reducing the rotor tip blockage, but a negative effect on the stability of the compressor stage. This is because the inflow condition of the stator is tremendously distorted while the coupling effect is excessively strong, which can cause a stall in the stator rather than in the rotor. The compressor stability can be maximally enhanced by adjusting the strength of the coupling effect to make a compromise of the improved rotor tip flow and the deteriorated stator flow.

Loss evaluation and aerodynamics investigation of an aggressive intermediate turbine duct under off-design conditions

To maximize the performance of the intermediate turbine duct (ITD) under off-design conditions, the loss generation in a one-half stage turbine was quantified using entropy generation and the global entropy generation rate. The numerical results solved by the unsteady Reynolds-averaged Navier–Strokes equations were first verified. Then, the aerodynamic losses within the high-pressure turbine stage were evaluated by efficiency loss under nine operating conditions composed of three rotor speeds and three rotor tip gaps. Finally, the disturbance modes caused by the upstream wake were captured by the dynamic mode decomposition method. Different from the influence of tip gaps, losses of the high-pressure turbine and the ITD are due to the swirl angle display an opposite trend. Under the influence of the interaction between the tip leakage flow and the shroud flow of the ITD, the viscous dissipation and turbulent dissipation increase with a larger tip gap owing to the dominant counter-rotating vortices and secondary flow occurring near the upstream of the ITD shroud. In addition, a large gap seems to enlarge the swirl component of the inflow angle, especially over 80% passage height, leading to greater dissipation losses in these areas. At the ITD inlet, two pairs of counter-rotating vortices at the shroud and the hub are, respectively, captured by the axial velocity mode. Large tip gaps enhance endwall vortices near the shroud and make the up vortex pairs merge into one pair.

Sweep Effects on Fan–Intake Aerodynamics at High Angle of Attack

Abstract Sweep in a transonic fan is conventionally used to reduce design point losses by inclining the passage shock relative to the incoming flow. However, future low pressure ratio fans operate to lower Mach numbers meaning the role of sweep at cruise is diminished. Instead, sweep might be repurposed to improve the performance of critical high Mach number off-design conditions such as high angle of attack (AOA). In this article, we use unsteady computational fluid dynamics to compare two transonic low pressure ratio fans, one radially stacked and one highly swept, coupled to a short intake design, at the high AOA flight condition. The AOA considered is 35 deg, which is sufficient to separate the intake bottom lip. The midspan of the swept fan was shifted upstream to add positive sweep to the outer span. Based on previous design experience, it was hypothesized that the swept fan would reduce transonic losses when operating at high AOA. However, it was found that the swept fan increased the rotor loss by 24% relative to the radial fan. Loss was increased through two key mechanisms. (i) Rotor choking: flow is redistributed around the intake separation and enters the rotor midspan with high Mach numbers. Sweeping the fan upstream reduced the effective intake length, which increased the inlet relative Mach number and amplified choking losses. (ii): Rotor-separation interaction (RSI): the rotor tip experiences low mass flow inside the separation, which increases the pressure rise across the casing to a point where the casing boundary layer separates. The swept fan diffused the casing streamtube, causing the casing separation to increase in size and persist in the passage for longer. High RSI loss indicated that the swept fan was operating closer to the rotating stall point.

High durability variable geometry turbine for commercial vehicle turbochargers

Abstract Transportation sector constitutes a major portion in the Green House Gas (GHG) emissions of the world. Reduction of GHG emission requires the development of highly efficient Internal Combustion Engines (ICE) which are downsized using turbochargers that provide compressed air for cleaner combustion. Variable Geometry (VG) turbochargers are widely accepted for this application because of the ability of VG turbine to function at wider flow range. This ensures the supply of compressed air at low engine speeds also. However, the interaction between stator vanes and turbine rotor has been a major issue in terms of High Cycle Fatigue (HCF) of such a turbocharger. Till now, a conservative design approach constrained the maximum rotation speed to below resonance speed for vibration mode 2. But, to achieve further downsizing it would be necessary to challenge higher rotation speeds. This paper will firstly introduce the tip-timing experiment and CFD-FEA simulations used to evaluate the HCF of VG turbine, and secondly it discusses the results of new VG turbine system which achieved safe operability even at vibration mode 3, both on simulation and experiment. The tip-timing experiment was conducted at 680°C on hot gas stand test bench with optical sensors installed circumferentially around the rotor tip. These measurements have been verified and validated with the safety factor calculated from blade surface pressure data obtained from transient rotor CFD simulations.

Multi-Objective Optimization of Circumferential Groove Casing Treatment in an Ultra-Highly Loaded Low-Reaction Transonic Compressor Rotor

This study concerns a multi-objective optimization of circumferential single grooved casing treatment for a low-reaction transonic rotor with ultra-high loading. The axial location, width and depth of the groove are investigated as design variables. The optimization problem seeks to fully extend the operation range of the rotor while minimizing efficiency degradation. Artificial neural network of radial basis function is applied to construct the surrogate model. The optimal groove configuration was determined using non-dominated sorting genetic algorithm II (NSGA II) in conjunction with technique for order preference by similarity to ideal solution (TOPSIS). Detailed analysis of flow field reveals that two flow features involving stability enhancement for the low-reaction rotor are the inhibition of shock/vortex interaction in the rotor tip region and the reduction or elimination of double-leakage tip gap flow. The blocking region located right downstream of the interface between the tip leakage flow and the main flow is decreased due to the tip unloading effect and recirculation flow induced by the groove. Additionally, the efficiency improvement can be observed as the intensity of tip leakage vortex decreases. Based on the single groove optimization, the prospect of a particular multiple groove casing configuration consisting of component grooves with varied geometrical dimensions is also discussed in the paper. The simulation results indicate that the new-type multiple groove configuration is more advantageous to the rotor’s performance.

A study on the flow physics of a supersonic partial admission impulse turbine for different tip clearances

Compared to traditional turbines, supersonic partial admission impulse (SPAI) turbines have significantly different flow physics owing to partial admission configuration and high supersonic flow at the rotor entrance. The objective of the present study is to reveal the intrinsic flow physics of an SPAI turbine for different rotor tip clearance dimensions. Numerical simulations were carried out based on the unsteady Reynolds-averaged Navier-Stokes (URANS) method in an SPAI turbine whose time-averaged relative Mach number at the rotor inlet was up to 2.1. The partial admission configuration led to a severer blockage in the hub region of the rotor entrance than that in the mid-span and tip regions. As a consequence, corner separation occurred near the leading edge of the rotor hub. The tip leakage flow was divided into two vortices. Due to the strong shock/tip leakage vortex interaction, the first vortex broke down and disappeared immediately. The second vortex originated from the leading edge shock foot, migrated downstream toward the mid-span (near the rotor trailing edge). It subsequently interacted with the trailing edge shock and exhibited an expansion process without experiencing a breakdown. Under the unsteady conditions, slight variations were observed in the shock train at the nozzle as the rotor passed. In contrast, due to the partial admission, the blade loading and the rotor leading edge shock experienced significant periodic variations. The rotor blade exhibited the maximum loading when the bulk flow originated from the nozzle flowed toward the rotor blade, whereas the minimum loading was achieved when the nozzle induced wake streamed toward the rotor blade. The leading edge shock/tip leakage vortex interaction occurred periodically, resulting in a periodic vortex breakdown. For the lowest efficiency condition, the highest total to static pressure ratio was achieved and the maximum intensity was exhibited for the leading edge shock of the rotor blade. This generated a significant loss due to the shock and the shock-induced flow separation.

Effect of the Axial Casing Groove Geometry on the Production and Distribution of Reynolds Stresses in the Tip Region of an Axial Compressor Rotor

AbstractStereo-PIV data are used for investigating the effect of axial casing groove (ACG) geometry on the distribution, evolution, and production rates of turbulent kinetic energy (TKE) and Reynolds stresses near a rotor tip. The ACGs delay the onset of stall by entraining the tip leakage vortex (TLV) and cause periodic changes to incidence angle. These effects are decoupled using semicircular, U-shaped, and S-shaped grooves that have similar inlets, but different outflow directions. Most TKE distribution trends can be explained by the local turbulence production rates, elucidating the different mechanisms involved and providing a unique database for turbulence modeling. Interaction of the tip flow with the ACGs modifies the highly anisotropic and inhomogeneous passage turbulence. In all cases, the TKE is high in the TLV center and shear layer connecting the TLV to the rotor tip. At prestall flowrate, TLV entrainment reduces the passage turbulence level, but introduces elevated turbulence in the corner vortex formed at the downstream corner of grooves, and in shear layers developing at the exit from grooves. The location of peaks and the dominant components vary among grooves. Near the best efficiency point, interactions of the TLV with the circumferentially negative outflow from the U and semicircular ACGs generate high turbulence levels, which extend deep into the passage. In contrast, interactions with S grooves are limited, resulting in a substantially lower turbulence level. Accordingly, the S groove maintains the untreated endwall efficiency, while the U and semicircular grooves reduce the peak efficiency.

Effect of Differential Tip Clearance on the Performance of a Tandem Rotor

Abstract Tandem blade is an interesting concept that promises a higher total pressure rise per stage. Owing to two separate tip leakage vortices and their interaction, losses are likely to increase particularly near the tip region. Although rotors are designed with optimum tip clearance, the clearance changes during engine operation as well as during its service life. In the case of tandem rotors, the forward and the aft rotors can have different tip clearances. This will also impact the performance of the stage. Six different tip clearances have been investigated. ANSYS CFX is used for steady RANS computational analysis. The results suggest that the performance of the tandem rotor is highly sensitive to the forward rotor tip clearance. Higher tip clearance adversely affects the total pressure rise and operation stability of the tandem rotor. At design mass flowrate, the performance degradation for tandem configuration with the higher tip clearance (Case2, Case 3, Case 5, and Case 6) is attributed to the vortex breakdown of TLV1, which leads to the sudden expansion of the blockage region near the rotor tip. Vortex breakdown primarily depends upon the swirling strength of TLV1 and TLV2 as well as on the adverse pressure gradient. Near the stall point, the role of the adverse pressure gradient becomes more dominant in the vortex breakdown.

Effect of Film Position on the Flow and Heat Transfer Characteristics of Full-Ribbed Rotor Tip

This article numerically studied the flow and heat transfer characteristics of the full-ribbed fluted tip of a rotor with four types of distribution of cooling films, namely, the mid-arc uniform film at the bottom of the groove, the uniform film on full perimeter ribs, the uniform film on the suction-side rib, and the uniform film on the pressure-side rib. The evolution of secondary flow in the tip clearance under these four conditions was discussed, and the effect of cooling film position on the tip leakage flow and tip heat transfer was also studied. The results showed that films in the ribs were more conducive to lower leakage flow rates and can reduce total energy losses. Adding cooling films on the single-side rib had a better effect on the control of leakage flow rate, with the most significant improvement due to adding cooling films on the suction-side rib. The mid-arc jet at the bottom of the groove can greatly contain the swirl structure in the groove, but it weakened the air sealing effect, which led to a rise in leak flow. The main reason for the swirl structure in the groove was that the leakage flow from the tip of the rotor was blocked by the cold air jet from the cooling films to form a backflow. The vortex inside the groove was the most violent when the air film hole was set on the pressure side, so the loss coefficient was higher, while the vortex inside the groove was the simplest when the cooling film hole was set on the suction side, and the energy loss caused by the vortex inside the groove was the least, and this structure was more favorable to the cooling of the leaf tip. It was also found that the uniformity of the outlet airflow angle distribution improved in the presence of cold air injection on the tip of the rotor.

Influence of wet compression on aerodynamic performance and stability boundary of transonic compressor

The gas–liquid two-phase flow in a transonic compressor stage, Stage 35, under wet compression was numerically simulated by Computational Fluid Dynamics (CFD) technique along with Euler–Lagrange approach. The evaporation and cooling effect of water droplets on compressor performance and stable flow range was investigated by changing the inlet spraying conditions (droplet sizes and injected water flow rates). Further, the influencing mechanism was explained by the changes of rotor velocity triangle, rotor specific work, blade loading, and tip leakage flow before and after spraying. The results indicate that wet compression can improve the pressure ratio and efficiency of the compressor, but, under different comparison conditions before and after spraying, the influence on compressor power consumption has different trends. Wet compression improves the compressor specific power consumption under constant flow rate conditions, while it reduces compressor specific power consumption under equal pressure ratio conditions. Additionally, the stalling flow rate of compressor increases by wet compression effect, and the increasing degree of stalling flow rate is positively correlated with the increasing level of pressure ratio. The evaporative and cooling effect of droplets reduces the airflow temperature at rotor outlet, leading to the increase of airflow density, the decrease of axial velocity, which results in the promotion of rotor specific work—this is the main reason for the improvement of pressure ratio under constant flow rate conditions. Under wet compression condition, the enhanced blade loading near tip strengthens the tip leakage flow, making the compressor achieve stall criteria at larger flow rate, which is the cause for wet compression influencing the compressor stall boundary.

Transient Aeroelastic Response of a Rotor During Rotor Speed Transition in Forward Flight

Compound helicopters often have to reduce their rotor tip speeds to achieve higher forward speeds. The rotors, therefore, have to undergo a transient process of rotor speed changes. To investigate the transient loads and the related transformation from the rotating frame to the nonrotating frame, a rotor model is used to predict the transient aeroelastic responses. The rotor model is validated by using test data of a teetering model rotor during engagement and disengagement operations. The investigations indicate that a small transient lagwise root bending moment can result in a large transient contribution to rotor torque. This transient component is triggered by the force related with the sudden change of the angular acceleration of the rotor speed. Increasing the blade damping in the lagwise direction can significantly decrease the transient components of the lagwise root bending moment and the corresponding rotor torque. Making the angular acceleration continuous will not trigger the transient components, suggesting that this is a better way for the reduction of the transient rotor loads during speed changes.

Control Allocation Reconfiguration for Actuator Failure on a Coaxial-Pusher Helicopter

An elastic blade flight dynamics model for a coaxial helicopter platform based on the Sikorsky X2 Technology™ Demonstrator is presented and validated with steady trim and frequency response flight-test data. A full authority explicit model following control architecture along with pseudoinverse control allocation is implemented for the model in hover and cruise at 180 kt using CONDUIT® in order to stabilize the vehicle and meet a set of stability, handling qualities, and performance requirements. Different fault scenarios are considered including failure of rotor swashplate actuators and tail surface actuators in hover and forward flight, which are compensated for by recalculating the pseudoinverse control mixing accordingly. The approach is shown to maintain aircraft stability through the fault transient and into a new steady trim state for the vehicle. Though the implemented controller is successful in maintaining the aircraft state, different fault cases lead to violations in rotor tip clearance limits, which will require additional effort to account for in flight.