Rotor Tip Vortex Research Articles

The Influence of an Upstream Pylon on Open Rotor Aerodynamics at Angle of Attack

The aerodynamic impact of installing a horizontal pylon in front of a contra-rotating open rotor engine, at take-off, was studied. The unsteady interactions of the pylon's wake and potential field with the rotor blades were predicted by full-annulus URANS CFD calculations at 0 deg and 12 deg angle of attack (AoA). Two pylon configurations were studied: one where the front rotor blades move down behind the pylon (DBP), and one where they move up behind the pylon (UBP). When operating at 12 deg AoA, the UBP orientation was shown to reduce the rear rotor tip vortex sizes and separated flow regions, whereas the front rotor wake and vortex sizes were increased. In contrast, the DBP orientation was found to reduce the incidence variations onto the front rotor, leading to smaller wakes and vortices. The engine flow was also time-averaged, and the variation in work done on average midspan streamlines was shown to depend strongly on variation in incidence, along with a smaller contribution related to change of radius.

Rotor–duct aerodynamic and acoustic interactions at low Reynolds number

Ducted single and multi-rotor UAVs (unmanned aerial vehicle) offer better crash worthiness and safety compared to open rotor UAVs. The interactions that affect performance and acoustics are of interest. High speed stereo particle image velocimetry (SPIV), performance, and acoustic measurements were done on a rotor with removable protective duct to study rotor–duct aerodynamic interactions over a range of hover conditions at low Reynolds number. Instantaneous and averaged SPIV velocity field and performance results, as well as acoustic signatures, are compared to baseline un-ducted rotor cases. Increase in rotor figure of merit is strongly correlated with the level of duct interaction with rotor tip vortex cores. Efficiency change is higher for the cases where thrust induced losses take a larger share of the overall aerodynamic losses. While the overall acoustic sound pressure level was not affected significantly, cases with higher vortex core/duct interactions showed higher tonal content. Graphical abstract

The Inner Workings of Axial Casing Grooves in a One and a Half Stage Axial Compressor With a Large Rotor Tip Gap: Changes in Stall Margin and Efficiency

Effects of axial casing grooves (ACGs) on the stall margin and efficiency of a one and a half stage low-speed axial compressor with a large rotor tip gap are investigated in detail. The primary focus of the current paper is to identify the flow mechanisms behind the changes in stall margin and on the efficiency of the compressor stage with a large rotor tip gap. Semicircular axial grooves installed in the rotor's leading edge area are investigated. A large eddy simulation (LES) is applied to calculate the unsteady flow field in a compressor stage with ACGs. The calculated flow fields are first validated with previously reported flow visualizations and stereo particle image velocimetry (SPIV) measurements. An in-depth examination of the calculated flow field indicates that the primary mechanism of the ACG is the prevention of full tip leakage vortex (TLV) formation when the rotor blade passes under the axial grooves periodically. The TLV is formed when the incoming main flow boundary layer collides with the tip clearance flow boundary layer coming from the opposite direction near the casing and rolls up around the rotor tip vortex. When the rotor passes directly under the axial groove, the tip clearance flow boundary layer on the casing moves into the ACGs and no roll-up of the incoming main flow boundary layer can occur. Consequently, the full TLV is not formed periodically as the rotor passes under the open casing of the axial grooves. Axial grooves prevent the formation of the full TLV. This periodic prevention of the full TLV generation is the main mechanism explaining how the ACGs extend the compressor stall margin by reducing the total blockage near the rotor tip area. Flows coming out from the front of the grooves affect the overall performance as it increases the flow incidence near the leading edge and the blade loading with the current ACGs. The primary flow mechanism of the ACGs is periodic prevention of the full TLV formation. Lower efficiency and reduced pressure rise at higher flow rates for the current casing groove configuration are due to additional mixing between the main passage flow and the flow from the grooves. At higher flow rates, blockage generation due to this additional mixing is larger than any removal of the flow blockage by the grooves. Furthermore, stronger double-leakage tip clearance flow is generated with this additional mixing with the ACGs at a higher flow rate than that of the smooth wall.

New Approach for Aerodynamic and Aeroacoustic Analysis of Actively Controlled Flaps Rotor

A new approach for aerodynamic and aeroacoustic analysis of actively controlled flaps (ACF) rotor is established to predict airloads and noise reduction. The approach is divided into three parts. The first is an unsteady aerodynamic model of the ACF described through a panel method, and the unsteady pressure under interaction of rotor tip vortex and ACF trailing-edge vortex, which are depicted through a viscous vortex particle method, are taken into account. The second is the influence of the ACF on trajectory of the wake and ACF trailing-edge vortex, which is carried out by a new imaged vortex model to capture blade-vortex interaction (BVI). The third is an aeroacoustic model coupled into the aerodynamic model by transmitting the unsteady pressure to Ffowcs Williams and Hawkings (FW-H) equation. The present approach is applied to AH-1G rotor, and the results show that the predicted unsteady airloads, high-speed impulsive (HSI), and BVI noise are compared favorably with experiment and Reynolds-averaged Navier–Stokes (RANS)/FW-H results. Finally, the variation of airload and noise reduction of ONERA Active Blade Concept rotor (ABC) is predicted and compared with experiment. It is shown that the present approach predicts the airload and noise reduction of the ACF with different phases and flap locations with good accuracy.

Investigation of Slotted Tip Parameters on the Aerodynamic Characteristic of Helicopter Rotor

Numerical simulations are conducted to investigate the effect of helicopter rotor tip vortex alleviation using a slotted tip, and effects of parameters involving location, number, aperture of the slot on tip vortex structure and aerodynamic performance are analyzed in detail. The flow field of hovering helicopter rotor is simulated by solving Navier-Stokes equations in the rotating coordinate system. Finite volume method combined with overset unstructured grids algorithm are adopted. The calculation of inviscid fluxes uses second-order central scheme with artificial dissipation, and turbulence model utilizes Spalart-Allmaras model. Simulation results show that the location of the slot has little effect on tip vortex alleviation. The larger the number and aperture of the slot, the better the effect of slotted tips weakening tip vortex. Lift coefficient of helicopter rotor with a slotted tip increases slightly compared with that without slots, while power coefficient increases drastically.

Wake developments behind different configurations of passive disks and active rotors

The present paper takes a broad view on our previous experimental studies of flows behind different single and dual configurations from passive disks or active rotors to establish new aspects of the wake development [1-4]. The aim of the present examination is to obtain a better understanding of the wake formations and interactions between wind turbines in wind farms. A correlation between independent investigations of the near [1] and far wakes behind single [2] and dual [3-4] systems will be established to the same operating regimes and flow conditions. New examinations of the old data need because two main differences in the wake behaviour for the disk-disk and the rotor-rotor systems were found: the wake intensity grows for the dual disks in comparison with the single one, but in contrast to this, wake intensity behind the dual rotor system is smaller than the one behind a single rotor. These differences may be explained by an influence of the rotor tip vortices which are absent in the disk-disk model. The present retesting of the near and far wake data should provide an evidence of this conclusion.

Effect of Aft Rotor on the Inter-Rotor Flow of an Open Rotor Propulsion System

The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig (ORPR) were examined. A particle image velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image-processing algorithm. The effects of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in reasonably good agreement with computational fluid dynamics (CFD) simulations. The aft rotor position was shown to have a significant upstream effect, with implications for front rotor interaction noise. It was found that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

Airborne visualization of helicopter blade tip vortices

Blade tip vortices and their interaction with the helicopter blades play an important role in the generation of noise on rotorcraft. Full-scale vortex visualization is essential for the understanding of these effects and the validation of sub-scale experiments and numerical simulations. In the present work, the reference-free background-oriented schlieren (BOS) method was used to visualize blade tip vortices of a full-scale BO 105 helicopter in hover and under free flight conditions. A ground-based dual-camera BOS system with an artificial background was applied for the visualization of the helicopter’s main and tail rotor tip vortices during take-off. In a second flight test, a similar dual-camera BOS system was used to visualize the main rotor blade tip vortices of the BO 105 during high speed forward, curve, and accelerating forward flight. The camera system was deployed aboard a microlight airplane flying above the helicopter, with vegetation on the ground serving as natural backgrounds. Different natural and artificial backgrounds were analyzed and compared to assess their suitability for the BOS evaluation. The airborne setup enabled the visualization of vortices up to a maximum vortex age of $$\psi _v=630^\circ$$ at distances of up to one rotor diameter behind the rotor plane. Blade-vortex interaction effects and deformations of the vortex system were observed. The visualization results obtained during the inflight measurements show that the reference-free BOS method is highly suitable for the detection of helicopter blade tip vortices during full speed forward and maneuvering flight, therefore removing the restrictions imposed by ground-based measurements.

Reynolds-averaged and wall-modelled large-eddy simulations of impinging jets with embedded azimuthal vortices

We have performed simulations of an impinging jet with embedded azimuthal vortices, a model of the wake of a helicopter hovering near the ground. This problem has considerable practical importance since, when the landing area is covered with sand or snow, the interaction between the helicopter wake, the rotor-tip vortices and the solid particles on the ground can result in the lift up of a cloud of sediment that limits the pilot’s vision, causing accidents and potentially loss of life. We compare the results of well-resolved large-eddy simulations (LES) at laboratory scale, with solutions of the Unsteady Reynolds-Averaged Navier–Stokes (URANS) equations with three turbulence models, and also with LES in which the wall layer is modelled (WMLES), using the Delayed Detached Eddy Simulation (DDES) approach, or employing approximate boundary conditions. The URANS solutions do not yield a reliable prediction of the development of the azimuthal vortices (a well-known shortcoming of most eddy-viscosity turbulence models in this type of configuration), while the WMLES and DDES predict vortex decay in good agreement with the resolved LES data. Simulations at a Reynolds number higher by a factor of 20 (intermediate between the laboratory scale and the real configuration) were also carried out. All the simulation methods predict very little effect of the Reynolds number: the flow appears to be driven by the evolution of the large-scale embedded vortices. The trends are the same observed at laboratory-scale: the RANS turbulence models predict a much faster decay of the embedded vortices compared to the WMLES and DDES. The need for reliable experimental data is highlighted.

Computation of the Aero-Optical Environment of a Helicopter Using Prescribed-Wake Methods

A fast low-order computational approach is presented for estimating the aero-optic effect of the rotor tip vortex system on helicopter-borne optical systems. The approach employs prescribed-wake methods that have been developed by the helicopter design community to rapidly generate realistic approximations to the vortex-wake systems of helicopters in hover and forward flight. With the geometry of the vortex-wake system defined, an experimentally validated computational method that employs the isentropic assumption is then used to compute pressure, temperature, density, and index-of-refraction fields in the selected computational domain. Predictions determined using the computational approach for the spatially and temporally resolved aero-optic aberrations on an optical system mounted on a helicopter in hover and forward flight are presented.

Tip Clearance Investigation of a Ducted Fan Used in VTOL Unmanned Aerial Vehicles—Part II: Novel Treatments Via Computational Design and Their Experimental Verification

Ducted fan based vertical lift systems are excellent candidates to be in the group of the next generation vertical lift vehicles, with many potential applications in general aviation and military missions. Although ducted fans provide high performance in many “vertical take-off and landing” (VTOL) applications, there are still unresolved problems associated with these systems. Fan rotor tip leakage flow adversely affects the general aerodynamic performance of ducted fan VTOL unmanned aerial vehicles (UAVs). The current study utilized a three-dimensional Reynolds-averaged Navier–Stokes (RANS) based computational fluid dynamics (CFD) model of ducted fan for the development and design analysis of novel tip treatments. Various tip leakage mitigation schemes were introduced by varying the chordwise location and the width of the extension in the circumferential direction. Reduced tip clearance related flow interactions were essential in improving the energy efficiency and range of ducted fan based vehicles. Full and inclined pressure side tip squealers were also designed. Squealer tips were effective in changing the overall trajectory of the tip vortex to a higher path in radial direction. The interaction of rotor blades and tip vortex was effectively reduced and the aerodynamic performance of the rotor blades was improved. The overall aerodynamic gain was a measurable reduction in leakage mass flow rate. This leakage reduction increased thrust significantly. Experimental measurements indicated that full and inclined pressure side tip squealers increased thrust obtained in hover condition by 9.1% and 9.6%, respectively. A reduction or elimination of the momentum deficit in tip vortices is essential to reduce the adverse performance effects originating from the unsteady and highly turbulent tip leakage flows rotating against a stationary casing. The novel tip treatments developed throughout this study are highly effective in reducing the adverse performance effects of ducted fan systems developed for VTOL UAVs.

Tip Flow Unsteadiness and Blade Row Interactions for a Low-Speed Compressor Rotor

A numerical investigation has been conducted to determine the effects of self-excited tip-leakage flow unsteadiness, upstream stator wakes, and downstream blade row interactions on the performance prediction of a low-speed research compressor rotor. Calculations included a single blade row rotor configuration as well as two multiblade row configurations: one in which the rotor was modeled with an upstream stator and a second in which the rotor was modeled with a downstream stator. Steady-state and time accurate calculations were performed using a Reynolds averaged Navier–Stokes solver, and the results were compared with detailed experimental data at operating conditions corresponding to the compressor design point, near stall, and a midpoint between the two. Differences in the performance predictions between the three configurations were then used to determine the effect of the upstream stator wakes and the downstream blade row interactions. Results obtained show that for this compressor with rather large tip clearance time accurate calculations are a vast improvement over steady-state calculations, but the upstream stator wakes and the downstream blade row interactions have only a small effect on the numerical solution. In this paper, changes in the rotor flowfield due to the upstream and downstream stators are investigated and discussed to explain the limited impact of the stator wakes and downstream blade row interactions on this tip dominated flow. In addition, the unsteady nature of the rotor tip vortex at near stall conditions is discussed to explain the improvement of time accurate analysis over steady-state computations.

Unsteady Effects in Axial Compressors: A Multistage Simulation

This work presents a numerical simulation of a four-stage axial compressor with cantilevered stators and repeating stages. Two simulations have been accomplished, namely, unsteady with sliding plane and steady with mixing plane, over a wide range of operating conditions. The solver is an unsteady Reynolds-averaged Navier–Stokes code with the Spalart–Allmaras turbulent model. It has been found that the greatest effect of unsteadiness is near the end walls. The stall point for the unsteady simulation is 40% closer to the experimental data when compared with the steady one with mixing planes. This is due to the recovery and segregation of the rotor tip vortex near the casing, which has a stabilizing effect in terms of stall limit. The unsteady simulation is able to reproduce total pressure loss experimentally observed near the hub at the stator exit but not found in the steady simulation. A previous experimental campaign suggested this loss was generated by the wakes of upstream inlet guide vanes propagating through the machine. However, this work found that it is generated by the migration of hub leakage flow induced by the incoming rotor wakes. The unsteady simulation shows higher performance at midspan. This has been found to be related to the rotor wake rectification with a 70% inviscid recovery of wake.

Detailed Computational Investigation of a Hovering Microscale Rotor in Ground Effect

A compressible, structured, overset Reynolds-averaged Navier–Stokes-based solver is used to simulate a microhovering rotor in ground effect to demonstrate the capability of the code to provide accurate tip-vortex-flowfield predictions, and to provide a good understanding of the ground–wake interactions in conditions prevalent in helicopter brownout. The performance validation at different rotor heights above the ground shows good correlation with the experimental thrust and power measurements. A detailed comparison of the predicted tip-vortex flowfield shows good agreement with the vorticity contours and radial-velocity profiles obtained from the particle-image-velocimetry measurements during experiments. The examination of the in ground effect tip-vortex flowfield suggests high degree of instabilities close to the ground. The induced velocities arising from the proximity of the rotor tip vortices cause flow separation at the ground. An analysis of the eddy-viscosity contours near the ground indicates higher turbulence levels in the flowfield at smaller rotor heights above the ground. With a decreasing rotor height, the increased interactions between the tip vortices and the ground boundary layer result in the intermixing of eddy viscosity from the tip vortex to the ground boundary layer.

Effect of tip vortex aperiodicity on measurement uncertainty

Vortex aperiodicity introduces random uncertainty in the measured vortex center location. Unless corrected, this may lead to systematic uncertainty in the vortex properties derived from the measured velocity field. For example, the vortex core size derived from averaged or mean flow field appears larger because of aperiodicity. Several methodologies for aperiodicity correction have been developed over the past two decades to alleviate this systematic uncertainty. However, these do not always reduce the accompanying random uncertainty. The current work shows that the analysis methods used to derive the vortex properties from the measured velocity field play an important role in the resultant random uncertainty in these properties; perhaps, even more important role than the aperiodicity correction methodology itself. It is hypothesized that a class of methods called global methods, which use a large extent of measured data, yield a smaller measurement uncertainty compared to local methods. This hypothesis is verified using a newly proposed global method based on a planar least-squares fit. The general applicability of the method is demonstrated using previous particle image velocimetry measurements of rotor tip vortices. The results clearly demonstrate a reduced random uncertainty in the vortex core properties, even in the presence of secondary vortical structures. Furthermore, the results are independent of the choice of aperiodicity correction methodology.

Helicopter Rotor Noise in the Merged Tip-Vortex and Blade Interaction Condition

BVI (blade vortex interaction) phenomena cause unsteady loading and noise in the helicopter rotor. It is difficult to accurately predict BVI aerodynamics and noise, because rotor wake and tip-vortices are numerically dissipated. In the present study, CFD and a time-marching free-wake coupling analysis were used to describe inflow and outflow without wake dissipation. Rotor noise was also computed using the Farassat formula, derived from the Ffowcs-Williams Hawkings equation. To validate aerodynamics and the noise prediction solver, a two-blade rotor was simulated and the results were compared with experimental data. To investigate the vortex interaction of a multi-blade rotor, a four-blade rotor was also simulated. In a slow descent flight condition, tip vortices are merged at the advancing side. The merged tip-vortex and blade interaction were simulated and are discussed in this paper.

Helicopter Tail Rotor Thrust and Main Rotor Wake Coupling in Crosswind Flight

The tail rotor of a helicopter with a single main rotor configuration can experience a significant reduction in thrust when the aircraft operates in crosswind flight. Brown’s vorticity transport model has been used to simulate a main rotor and tail rotor system translating at a sideslip angle that causes the tail rotor to interact with the main rotor tip vortices as they propagate downstream at the lateral extremities of the wake. The tail rotor is shown to exhibit a distinct directionally dependent mode during which tail rotors that are configured so that the blades travel forward at the top of the disk develop less thrust than tail rotors with the reverse sense of rotation. The range of flight speeds over which this mode exists is shown to vary considerably with the vertical location of the tail rotor. At low flight speeds, the directionally dependent mode occurs because the tail rotor is immersed within not only the downwash from the main rotor but also the rotational flow associated with clusters of largely disorganized vorticity within the main rotor wake. At higher flight speeds, however, the tail rotor is immersed within a coherent supervortex that strongly influences the velocity field surrounding the tail rotor.

構造・非構造格子カップリングによるヘリコプター全機周りの流体数値計算

The objective of this study is to develop a new CFD solver by combining the structured and unstructured grid methods so that the tip vortices can be predicted accurately for the realistic geometry of a helicopter. Flow around a helicopter has unique features caused by the rotor and a variety of equipment on the helicopter fuselage surface such as hoist crane and landing gear. There are difficulties in CFD analysis for rotorcraft, particularly in accurately capturing of rotor tip vortices and solving flow around a complicated geometry. Generally, structured grid method is suitable for capturing tip vortices, while unstructured grid method can solve flow around complicated geometry. In this study, two validations are conducted to show effectiveness of this coupled CFD solver. The first validation is related to a steady flow around ROBIN fuselage (Mach # 0.062). Computational results are compared with both experimental data and existing CFD results. The results show that the coupled CFD solver can predict surface pressure distributions in good agreement with experimental data and existing CFD results. The second validation is related to an unsteady flow around ROBIN fuselage with a rotor (µ = 0.15, CT = 0.0064). Results show that coupled CFD solver can predict periodical surface pressure also in good agreement with experimental data and existing CFD results. Furthermore, the vortex structure predicted has a good agreement between the coupled and existing structured grid CFD solvers. Finally, the coupled CFD solver is applied to a flow around the Eurocopter 135 model to demonstrate its capability for a realistic complicated helicopter geometry.

High-Resolution Computational and Experimental Study ofRotary-Wing Tip Vortex Formation

The formation and rollup of a tip vortex trailed from a hovering helicopter rotor blade is studied in detail using both computations and measurements. The compressible Reynolds-averaged Navier-Stokes equations are computationally solved on an overset mesh system. The flow measurements are made using stereoscopic particle image velocimetry. The high resolution of both the numerics and the measurements reveal multiple coherent structures in the evolving rotor tip vortex flowfield. Secondary and tertiary vortices that result from crossflow separations near the blade tip are identified. These vortices, along with a part of the trailed wake, are ultimately entrained into the tip vortex that is formed downstream of the blade's trailing edge. The simulations clearly demonstrate the resolution required to accurately represent the complex three-dimensional flowfield. The advantage of particle image velocimetry, which has the ability to make planar measurements at a given instant of time, has been fully used to validate the computational fluid dynamics predictions. Even though linear eddy viscosity models are expected to inadequately represent the details of the turbulent quantities, good agreement is seen to be achieved with the particle image velocimetry measurements of the mean flowfield. The various sources of computational and measurement uncertainties are discussed.

BILU implicit multiblock Euler/Navier–Stokes simulation for rotor tip vortex and wake convection

AbstractIn this paper, a block incomplete lower–upper (BILU) decomposition method is incorporated with a multiblock three‐dimensional Euler/Navier–Stokes solver for simulation of hovering rotor tip vortices and rotor wake convection. Results of both Euler and Navier–Stokes simulations are obtained and compared with experimental observations. The comparisons include surface pressure distributions and tip vortex trajectories. The comparisons suggest that resolution of the boundary layer is important for the accurate evaluation of the blade surface loading, but is less so for the correct prediction of the vortex trajectory. Numerical tests show that, using Courant–Friedrichs–Lewy (CFL) number of 10 or 30 with the developed BILU implicit scheme can be 6–7 times faster than an explicit scheme. The importance of solution acceleration schemes that increase the permitted time‐step is illustrated by comparing the evolving wake structures at different stages of the calculation. In contrast to fixed wing simulations, the extent of the wake structures is shown to require resolution of large physical time. This observation explains the poor performance that is obtained when employing convergence acceleration strategies originally intended for solution of equilibrium problems, such as the multigrid methods. Copyright © 2007 John Wiley & Sons, Ltd.