Blade-vortex Interaction Research Articles

Effects of gust on the aerodynamics of multi-rotors

Understanding the aerodynamics of multi-rotor systems under temporally varying discrete gusts is the focus of this work. Axial and edgewise 1-minus-cosine profiles of varying amplitude (0.1ΩR–0.3ΩR) and gust durations (0.3–15 s) were studied for the rigid, unarticulated Harrington single and coaxial rotors, and the Advanced Precision Composites (APC) 12x6E (a commercial off-the-shelf propeller) quad-rotor in hover. An in-house free vortex method (FVM) was utilized to study the time and frequency domain of thrust, span-wise angle of attack distribution, and the rotor wake structure across the gust event. It was observed that edgewise gusts produced higher frequency components in the thrust signal compared to axial gusts for all rotor systems. Strong wake–wake and blade–vortex interactions (BVI) were observed particularly for edgewise gusts, leading to strong variations in the sectional angle of attack and the thrust values. The wake interactions are more pronounced in the APC quad-rotor system, which operates at a low reduced frequency, essentially resulting in a quasi-steady response of the wake to the gust, resulting in a considerably higher number of BVI events, increased peak rotor loads, and higher frequency content. This study can aid in the understanding of the gust response of small- and medium-scale rotors.

BOS and Hot-Film Analysis of a CH-53G Helicopter Wake in Ground Effect

Constant temperature anemometry and background-oriented schlieren (BOS) measurements on a CH-53G helicopter (takeoff mass ≈14t) are used to investigate the rotor wake including the vortex breakdown to statistical turbulence. Hovering flights at a range of hub heights 0.50 < h < 0.83 were investigated. The position of the rotorcraft was optically tracked using markers affixed to the fuselage. A complete breakdown of the tip vortices was noted within 0.3R of the rotor plane, in stark qualitative contrast to small-scale experiments which show the tip vortices persisting to the ground plane. The large (37 m2) BOS field of view showed a small effect of vortex pairing and blade vortex interaction, which was overshadowed by the rapid wake breakdown. Vortex pairing was found to result from varying vortex convection velocities rather than from initial differences in the vortex spacing due to the blade track.

An Experimental Study of Stacked Rotor Tip Vortex Interaction in Hover

Measurements were made using flow visualization to track the blade tip vortices of a 1.108−m radius stacked rotor in hover. Vortex measurements of an isolated, two-bladed rotor were also made and correlated with the Landgrebe wake model. When a second rotor was introduced to the system, the interaction between the upper and lower rotor vortices resulted in deviations from the Landgrebe wake model. Changing the azimuthal spacing revealed large changes in the trajectory of the upper and lower rotor vortices, particularly at small, negative azimuthal spacings. The miss distance, or distance between a rotor blade and vortex, was measured. Positive miss distances where the lower rotor blade passed below the upper rotor tip vortex were associated with greater noise between 1 and 5 kHz, possibly from turbulent boundary layer noise or blade vortex interaction noise. The greatest performance and lowest broadband noise corresponded with large miss distance and minimal vortex???vortex interaction, which occurred at an azimuthal spacing of −45°. The tip vortex trajectory, and thus miss distance, could be modified by utilizing differential collective. This resulted in a 4.5−dB reduction in broadband noise at negative differential collectives by allowing the lower rotor blade to pass above the upper rotor vortex and achieving negative miss distances. At azimuthal spacings where the miss distance was always negative, the airfoil self-noise dictated the broadband noise level.

Numerical investigation on the blade vortex interaction noise reduction using higher harmonic control

Accurately predicting the formation, evolution, and breakdown of helicopter blade tip vortex is crucial for simulating the Blade-Vortex Interaction (BVI) phenomenon. Firstly, the high-order Perturbed polynomial reconstructed Targeted Essentially Non-Oscillatory (TENO-P) scheme proposed by our research group is employed to improve the resolution of the helicopter rotor flowfield solver. The TENO-P scheme, building on the fifth-order TENO5 scheme, achieves one-order of accuracy improvement by adaptively adjusting the values of the free-parameter introduced by perturbed polynomial reconstruction. Subsequently, the AH-1 helicopter model rotor undergoing blade-vortex interaction is analyzed using the improved rotor flowfield solver and the Farassat-1A formula. The implementation of the TENO-P scheme notably enhances the resolution of the rotor flowfield solver in resolving the blade tip vortex structures, and the predicted noise results are in good agreement with the experimental data. Finally, the alteration of the BVI noise and its reduction mechanism of the AH-1 model rotor under different Higher Harmonic Controls (HHC) are analyzed. The findings show that the phase modulation margin increases with a decrease in harmonic order, and the noise reduction effect significantly improves as well. The negative component of the higher harmonic control is beneficial for reducing BVI noise while the positive component of the higher harmonic control exacerbates the BVI noise. The HHC reduces the collective pitch angle in the region where the tip vortex is about to be disturbed, decreasing the strength of the blade tip vortex in this region. This reduction weakens the strength of the parallel and near-parallel interaction on the advancing side, leads to the reduction of the positive peak impulsive of the BVI noise, and consequently lowers the intensity of the BVI noise.

Research on the aerodynamic characteristics of electrically controlled rotor under Parallel Blade Vortex Interaction using Lattice Boltzmann Method

An electrically controlled rotor (ECR), also known as a swashplateless rotor, employs a trailing edge flap (TEF) system for primary rotor control instead of a swashplate, demonstrating the significant potential in rotor vibration and noise reduction. To investigate the aerodynamic characteristics of the blade flap segment of the ECR under parallel blade vortex interaction, an aerodynamic analysis model based on the lattice Boltzmann method (LBM) is established using the D3Q27 lattice model. The model is validated against experimental data of both the airfoil with trailing edge flap and conventional airfoil under vortex interaction, showing that the LBM can effectively predict variations in aerodynamic loads under both conditions. Based on this model, the effects of different flap deflection angles and miss distances on the aerodynamic characteristics of the ECR under parallel BVI are analyzed. The results indicate that under strong vortex interaction, a region of flow separation forms due to the entrainment effect of the vortex and adverse pressure gradient. The small-scale vortex structure upstream of the flap can also be observed and is believed to contribute to the unsteady flow phenomena such as vortex structure splitting, development, and separation on the upper surface of the flap. The different flap deflection angles mainly affect the scale and type of vortex structures developed on the flap upper surface. As the miss distance increases, the interaction effect is significantly weakened compared to strong vortex interaction. However, as the vortex moves downstream along the airfoil lower surface, it entrains vorticity from the lower surface, ultimately forming a negative pressure region on the lower surface of the flap. The different flap deflection angles will influence the structural characteristics during the downstream motion of the vortex, which changes the size of the negative pressure region, causing differences in the magnitude of the variations in aerodynamic parameters.

A novel efficient calculation method for rotor aeroacoustic characteristics based on multiple aerodynamic surfaces source model

Based on the simplified acoustic source model of multiple aerodynamic surfaces, an efficient calculation method for rotor aeroacoustic characteristics has been established. Firstly, the aerodynamic characteristics database of the airfoil is obtained based on the Computational Fluid Dynamics (CFD) method. The inflow environment of each blade element is calculated using the free wake method, and the distribution of chord-wise loading is obtained by combining the established database. Subsequently, based on the F1A equation, a simplified formula for loading noise with pressure difference as a parameter is derived. The thickness noise is obtained from the blade shape and operating conditions, and the blade surface mesh is constructed by using the adaptive curvature distribution strategy of airfoil points. The validation is carried out by comparing with experimental values in the hover and forward flight states. Then, A hybrid weighted interpolation scheme is proposed based on inverse distance weighted interpolation (IDW) and bidirectional linear interpolation method, which is suitable for capturing blade/vortex interaction (BVI) noise. A set of parameters suitable for engineering applications is obtained through the parametric sensitivity analysis. The computational time and memory usage are reduced to 1/47 and 1/4 of the original amount, respectively. Finally, Using the above method, the influence of rotational speeds (ω) on rotor aeroacoustic characteristics are studied. The results show that ω and sound pressure level (SPL) are linearly related and the SPL decreases by approximately 5 dB for every 0.1 Ω decrease in ω. The rotational speed ω≈90 %Ω can effectively suppress rotor aeroacoustic levels, benefiting the stealth design of helicopters.

Interference mechanism of trailing edge flap shedding vortices with rotor wake and aerodynamic characteristics

A robust Reynolds-Averaged Navier-Stokes (RANS) based solver is established to predict the complex unsteady aerodynamic characteristics of the Active Flap Control (AFC) rotor. The complex motion with multiple degrees of freedom of the Trailing Edge Flap (TEF) is analyzed by employing an inverse nested overset grid method. Simulation of non-rotational and rotational modes of blade motion are carried out to investigate the formation and development of TEF shedding vortex with high-frequency deflection of TEF. Moreover, the mechanism of TEF deflection interference with blade tip vortex and overall rotor aerodynamics is also explored. In non-rotational mode, two bundles of vortices form at the gap ends of TEF and the main blade and merge into a single TEF vortex. Dynamic deflection of the TEF significantly interferes with the blade tip vortex. The position of the blade tip vortex consistently changes, and its frequency is directly related to the frequency of TEF deflection. In rotational mode, the tip vortex forms a helical structure. The end vortices at the gap sides co-swirl and subsequently merge into the concentrated beam of tip vortices, causing fluctuations in the vorticity and axial position of the tip vortex under the rotor. This research concludes with the investigation on suppression of Blade Vortex Interaction (BVI), showing an increase in miss distance and reduction in the vorticity of tip vortex through TEF phase control at a particular control frequency. Through this mechanism, a designed TEF deflection law increases the miss distance by 34.7% and reduces vorticity by 11.9% at the target position, demonstrating the effectiveness of AFC in mitigating BVI.

Wind-tunnel experimental investigation on rotor-rotor aerodynamic interaction in compound helicopter configuration

In recent years, the compound helicopter configuration, featuring the addition of lateral propellers to a helicopter main rotor, has gained renewed interest for its ability to achieve higher flight speed. The aerodynamic behaviour of compound rotorcraft is dominated by mutual interactions between the rotors and the wakes they generate, which can affect their performance and the handling qualities of the aircraft. In order to study the effects of the rotor-wake and wake-wake interactions, a test rig was developed and an extensive wind-tunnel test campaign was carried out measuring the performance of rotor and propellers under different flight conditions. Hovering flight and forward flight at different advance ratios were considered, as well as crosswind flight under various wind directions. In general, an increase in thrust for both main rotor and propellers is found due to the interaction. In particular, the significant increase in the propeller thrust is attributed to the influence of the main rotor downwash, which impacts edgewise on the propellers. In some specific conditions, a decrease in the propeller's thrust is observed, which might be related to blade-vortex interaction (BVI) effects. To aid in the interpretation of the experimental results, numerical simulations with a mid-fidelity aerodynamic code were also performed.

Numerical Study on Rotor–Building Coupled Flow Field and Its Influence on Rotor Aerodynamic Performance under an Atmospheric Boundary Layer

In urban settings, buildings create complex turbulent conditions, affecting helicopter flight performance during missions and increasing safety risks during takeoff and landing. A numerical study on rotor–building coupled flow field is carried out to address rotor aerodynamic performance under building interferences in natural atmospheric conditions. A high-fidelity atmospheric boundary layer (ABL) model described by an exponential law is established herein. The solution of the coupled flow field is based on the Reynolds-averaged Navier–Stokes (RANS) equations, with the rotor’s rotation achieved through the overset grid method. Based on the dominant wind features, the building flow field is distributed into four regions, where the updraft along the headwind side impacts the rotor, bringing about a 76% increase in pitching moment. On the lateral side of the building, distorted rotor wake squeezed upward into the rotor disk, leading to severe blade–vortex interaction (BVI). During low-altitude hovering over rooftops, the mixing of building shed vortices with forward flow wakes causes the formation of a circulation region on the rotor’s windward side, resulting in a thrust loss of approximately 7.8%. Meanwhile, the flow environment on the leeward side of the buildings is more stable. Therefore, it is recommended that helicopters adopt a headwind approach during rooftop operations. However, an 11.4% loss in the average hover figure of merit is observed due to consistent thrust losses caused by the recirculation region.

A high-resolution numerical investigation of unsteady wake vortices for coaxial rotors in hover

High-resolution numerical simulations for wake vortical flows have long been a challenge in rotor aerodynamics. A novel spectrum-optimized sixth-order Weighted Essentially Non-Oscillatory (WENO) scheme is proposed to discretize inviscid fluxes on moving overset grids, and the Improved Delayed Detached Eddy Simulation (IDDES) is employed to resolve turbulent vortices. The integration of these methods facilitates a comprehensive numerical investigation into the unsteady vortical flows over coaxial rotors in hover. The results highlight the substantial improvement in numerical resolution, in terms of both spatial structure and temporal evolution of unsteady multiscale wake vortices. Coaxial rotors in hover manifest three primary scales of wake vortex structures: (A) the helical evolution of primary blade tip vortices and the periodic occurrence of strong Blade-Vortex Interactions (BVI); (B) the continuous shedding of small-scale horseshoe-shaped vortices from the trailing edges of rotor blades, forming the vortex sheets; (C) the emergence of small-scale secondary vortex braids induced by interactions between rotor tip vortices and the vortex sheets. These vortex structures and their interactions cause high-frequency oscillations in rotor disk loads and induce unsteady perturbations in the local flow field. Interactions among these primary vortices, coupled with the generation of secondary vortices, result in the dissipation, distortion, and breakup of the rotor tip vortices, ultimately forming a vortex soup. Notably, a substantial quantity of seemingly weak small-scale secondary vortex braids significantly contribute to energy dissipation during the evolution of wake vortices for coaxial rotors in hover.

Wind Tunnel Experiments on Parallel Blade–Vortex Interaction with Static and Oscillating Airfoil

This study aims to experimentally investigate the effects of parallel blade–vortex interaction (BVI) on the aerodynamic performances of an airfoil, in particular as a possible cause of blade stall, since similar effects have been observed in literature in the case of perpendicular BVI. A wind tunnel test campaign was conducted reproducing parallel BVI on a NACA 23012 blade model at a Reynolds number of 300,000. The vortex was generated by impulsively pitching a second airfoil model, placed upstream. Measurements of the aerodynamic loads acting on the blade were performed by means of unsteady Kulite pressure transducers, while particle image velocimetry (PIV) techniques were employed to study the flow field over the blade model. After a first phase of vortex characterisation, different test cases were investigated with the blade model both kept fixed at different incidences and oscillating sinusoidally in pitch, with the latter case, a novelty in available research on parallel BVI, representing the pitching motion of a helicopter main rotor blade. The results show that parallel BVI produces a thickening of the boundary layer and can induce local flow separation at incidences close to the stall condition of the airfoil. The aerodynamic loads, both lift and drag, suffer important impulsive variations, in agreement with literature on BVI, the effects of which are extended in time. In the case of the oscillating airfoil, BVI introduces hysteresis cycles in the loads, which are generally reduced. In conclusion, parallel BVI can have a detrimental impact on the aerodynamic performances of the blade and even cause flow separation, which, while not being as catastrophic as in the case of dynamic stall, has relatively long-lasting effects.

Aerodynamic interference and unsteady loads for a trimming intermeshing rotor compound propeller helicopter at level forward flight

Traditional trim methods for isolated rotor helicopters are not suitable for intermeshing rotor compound propeller helicopters (IRcPHs) due to inherent structural differences. This study aims to establish a novel trim method for an intermeshing rotor with an auxiliary propeller. First, the relationship between the rotor cyclic pitch and the flapping response in the fully articulated rotor system with hinged extension is derived. The study then formulates the equilibrium state equations and the corresponding flight conditions for trimming the IRcPH in level forward flight. In addition, the unsteady vortex lattice method is employed to calculate the periodic averaging states on the rotor disk and correct for induced losses in flight dynamics. Finally, the paper focuses on the aerodynamic interference, blade–vortex interaction, and harmonic load patterns of the IRcPH, serving as a guide for rotor active control systems. Results show that the positions of the mutual aerodynamic interference occur constantly at 160° and 340° azimuths on the intermeshing rotor disk, serving as the phase basis for suppressing vibration. Moreover, the propeller enters the comfort zone as the forward flight speed increases, and the loads only have frequency orders of (2/nw)/rev with respect to the main frequency of the intermeshing rotor.

Influence of Rotor Inflow, Tip Loss, and Aerodynamics Modeling on the Maximum Thrust Computation in Hover

Comprehensive rotorcraft simulation codes are the workhorses for designing and simulating helicopters and their rotors under steady and unsteady operating conditions. These codes are also used to predict helicopters’ limits as they approach rotor stall conditions. This paper focuses on the prediction of maximum rotor thrust when hovering (due to stall limits) and the thrust and power characteristics when the collective control angle is further increased. The aerodynamic factors that may significantly affect the results are as follows: steady vs. unsteady aerodynamics, steady vs. dynamic stall, blade tip losses, curvature flow, yaw angle, inflow model, and blade-vortex interaction. The inflow model and tip losses are found to be the most important factors. For real-world applications vortex-based inflow models are considered the best choice, as they reflect the blade circulation distribution within the inflow distribution. Because the focus is on the impact of aerodynamic modeling on rotor stall, the blade design and its flexibility are intentionally not considered.

Numerical investigation of aerodynamic interactions for the coaxial rotor system in low-speed forward flight

This study is conducted to investigate the aerodynamic characteristics of a coaxial rotor in low-speed forward flight using a high-fidelity computational fluid dynamics (CFD) solver. Numerical simulations are performed on the coaxial rotor and its isolated upper and lower rotors at two small advance ratios, 0.12 and 0.24. Interaction events in the flowfield, especially blade-vortex-interaction (BVI), are studied in detail. As the forward speed increases, the airload distributions become concentrated on the front part of the disk. Fluctuations in the loading indicate self-BVIs within each rotor, and blade-crossover events and rotor-to-rotor BVIs between the coaxial upper and lower rotors. Principal BVI events are identified in the well-captured wake structures by the 8th-order TENO reconstruction scheme. Self-BVIs on all rotors reduce in strength as the forward speed increases, while rotor-to-rotor BVIs on the coaxial lower rotor increase in strength with the increasing forward speed. In the coaxial configuration, the wake of the upper rotor is induced to move backwards and downwards due to the exertion of the lower rotor wake. The wake of the coaxial lower rotor features the most complex flow nature among all rotors. At both advance ratios, the coaxial lower rotor experiences a parallel rotor-to-rotor BVI, causing highly impulsive loading fluctuations along the blade span.

Aerodynamic model comparison for an X-shaped vertical-axis wind turbine

Abstract. This article presents a comparison study of different aerodynamic models for an X-shaped vertical-axis wind turbine and offers insight into the 3D aerodynamics of this rotor at fixed pitch offsets. The study compares six different numerical models: a double-multiple streamtube (DMS) model, a 2D actuator cylinder (2DAC) model, an inviscid free vortex wake model (from CACTUS), a free vortex wake model with turbulent vorticity (from QBlade), a blade-resolved unsteady Reynolds-averaged Navier–Stokes (URANS) model, and a lattice Boltzmann method (from PowerFLOW). All models, except URANS and PowerFLOW use the same blade element characteristics other than the number of blade elements. This comparison covers the present rotor configuration for several tip-speed ratios and fixed blade pitch offsets without unsteady corrections, except for the URANS and PowerFLOW which cover a single case. The results show that DMS and 2DAC models are inaccurate – especially at highly loaded conditions, are unable to predict the downwind blade vortex interaction, and do not capture the vertical/axial induction this rotor exhibits. The vortex models are consistent with each other, and the differences when compared against the URANS and PowerFLOW mostly arise due to the unsteady and flow curvature effects. Furthermore, the influence of vertical induction is very prominent for this rotor, and this effect becomes more significant with fixed pitch offsets where the flow at the blade root is considerably altered.

Air and Structural Loads Analysis of a 5-Ton Class Rotorcraft in a Pull-Up Maneuver Using CFD/CSD Coupled Approach

The air and structural loads of a 5-ton class light helicopter (LH) rotor in a 2.24 g pull-up maneuver are investigated using a coupling between the computational structural dynamics (CSD) and computational fluid dynamics (CFD) methods. The LH rotor is characterized by a five-bladed system with elastomeric bearings and inter-bladed dampers. The periodic trim solution along with the converged CFD/CSD delta airloads obtained in steady-level flight (advance ratio of 0.287) are used to perform the transient CSD maneuver analysis. The resulting vehicle attitude angles and velocity profiles of the aircraft are then prescribed in the quasi-static (QS) CFD maneuver analysis. It is demonstrated that the present QS approach provides an effective means for the maneuver loads’ analysis. The important flow behaviors such as BVI (blade–vortex interaction)-induced oscillations and the negative pitching moment peaks met in maneuver flight are captured nicely with the proposed method. Either the vortex trajectories or the surface pressure distributions are examined to identify the sources of the oscillations. A loose CFD/CSD coupling (LC) is used to predict the blade elastic motions, structural moments, and pitch link loads at the specified maneuver revolution of the rotor and also to correlate these with the transient CSD-based predictions. A reasonable correlation is obtained. The LC results show more pronounced 5P (five per revolution) oscillations on the structural response than those of the CSD-based methods.

Reduction of rotorcraft BVI using synthetic jets; computational studies using LES

In rotorcraft, the blade-vortex interaction (BVI) has been recognized as one of the main sources of noise and vibrations, affecting its aerodynamics performance and stability. Active and passive flow control techniques, for the reduction of BVI, have been investigated. The passive flow control techniques have proved to be inefficient due to the complex aerodynamics of rotorcraft. Thus, active flow control techniques has become of particular interest. In the present research we propose a novel approach, for the reduction of rotorcraft’s noise and vibrations due to BVI. The main idea of this approach is to inject air at the leading-edge of the blade to alter the characteristics (strength and core size) of the incoming vortex and thus, minimize the noise and vibrations associated with BVI. Computational studies are carried out using the large-eddy simulation (LES) with the dynamic Smagorisky sub-grid scale model. The computations are carried out for a flow of Reynolds number R e = 1.3 × 10 6 , based on the chord of NACA0012 airfoil and free-stream velocity. The studies show that injecting air at the leading edge of the blade, the influence of blade-vortex interaction on the aerodynamic coefficients is significantly reduced. Higher reduction of BVI effect, on the aerodynamic coefficients, was achieved with the increase of the speed of the synthetic jet. The reduction of the BVI, also, leads to the reduction of the vibrations.

Detached-Eddy Simulation of UH-60A Airloads Using an Immersed Boundary Method

The complete helicopter configuration of UH-60A Black Hawk in forward flight is numerically studied using an immersed boundary method and an actuator surface model. Detached-eddy simulation is conducted where the UH-60A fuselage is modeled by the immersed boundary method, while the main and tail rotors are represented using the actuator surface model. The low-speed case (flight counter 8513) from the UH-60A Airloads Program by NASA and the U.S. Army is considered. This case poses challenges such as significant blade–vortex interaction and rotor–airframe interaction. The validation of the test case is conducted for both the isolated main rotor and the full configuration, and the results are compared with flight-test measurements and computational data from the literature. Both the main rotor blade airload and fuselage drag data agree well with the reference data. The surface pressure data over the centerline of the fuselage, horizontal stabilator, and vertical stabilizer are presented. Qualitative data demonstrate massive flow separation from subcomponents of the fuselage that are often not included in computational fluid dynamics simulations. It also features significant main rotor–tail rotor interactions. The computational effort for three grid levels of each test case is provided.

Effects of Ambient Conditions on Helicopter Harmonic Noise Radiation: Theory and Experiment

The effects of ambient atmospheric conditions, air temperature, and density on rotor harmonic noise radiation are characterized using theoretical models and experimental measurements of helicopter noise collected at three different test sites at elevations ranging from sea level to 7000 ftabove sea level. Significant changes in the thickness, loading, and blade–vortex interaction noise levels and radiation directions are observed across the different test sites for an AS350 helicopter ying at the same indicated airspeed and gross weight. However, the radiated noise is shown to scale with ambient pressure when the flight condition of the helicopter is defined in nondimensional terms. Although the effective tip Mach number is identi ed as the primary governing parameter for thickness noise, the nondimensional weight coefficient also impacts lower harmonic loading noise levels, which contribute strongly to low-frequency harmonic noise radiation both in and out of the plane of the horizon. Strategies for maintaining the same nondimensional rotor operating condition under different ambient conditions are developed using an analytical model of single main rotor helicopter trim and confirmed using a CAMRAD II model of the AS350 helicopter. The ability of the Fundamental Rotorcraft Acoustics Modeling from Experiments (FRAME) technique to generalize noise measurements made under one set of ambient conditions to make accurate noise predictions under other ambient conditions is also validated.

Helicopter Noise Source Separation Using an Order Tracking Filter

Due to the importance of understanding the aeroacoustics of rotorcraft with continually changing noise sources, this paper presents a new technique for source separation from ground-based acoustic measurements. The source separation process is based on combining a time-domain de-Dopplerization method with the Vold–Kalman order tracking lter approach. This process can extract rotor harmonic noise even when the sources are continuously changing with time, including impulsive events such as blade–vortex interaction noise. The advantage of this approach over traditional methods such as harmonic averaging is that the phase and amplitude relationship of acoustic signals is preserved throughout the extraction process. The approach is applied to the measured acoustic data from a Bell 430 helicopter. The measured data were separated into main rotor harmonic, tail rotor harmonic, and broadband residual components. For steady-state conditions, the extracted components could be depropagated to form acoustic hemispheres showing the directivity of the separated main and tail rotor components. The source separation process was also applied to a maneuvering ight condition. Each component has different pulse shapes and directivity trends, consistent with aeroacoustic theory.