scholarly journals A Numerical Analysis of the Dynamic Stall Mechanisms on a Helicopter Rotor from Light to Deep Stall

2020 ◽  
Vol 65 (3) ◽  
pp. 1-17
Author(s):  
Camille Castells ◽  
François Richez ◽  
Michel Costes
Keyword(s):  
Fluid Dynamics ◽  
Separated Flow ◽  
Comprehensive Analysis ◽  
Dynamic Stall ◽  
Helicopter Rotor ◽  
Loose Coupling ◽  
Vortex Interaction ◽  
Thrust Coefficient ◽  
Blade Vortex Interaction ◽  
Rotor Disk

A loose coupling methodology between computational fluid dynamics and Comprehensive Analysis codes (elsA/HOST) is used to simulate a helicopter rotor in dynamic stall condition. Three stalled forward flight conditions have been selected in the wind tunnel 7A rotor test data to investigate the evolution of the stall mechanisms from a light stall to a deep stall condition. A decrease in the RPM is used to increase the rotor load. The lower the RPM the more severe the stall is. A double stall is observed in the lowest RPM case. The simulations are in satisfactory agreement with the experiment and are used to identify the mechanisms leading to the different stall events, notably the blade–vortex interaction. Rotormaps of the flow-separation regions are computed from numerical results, and similar regions of separated flow are observed in all the cases. These flow-separations originate from different aeroelastic mechanisms depending on their position on the rotor disk. As the rotor thrust coefficient is increased, some of these flow separations grow and lead to stall events.

A modified near wake dynamic model for rotor analysis

1999 ◽  
Vol 103 (1021) ◽  
pp. 143-146 ◽  
Author(s):  
T. Wang ◽  
F. N. Coton
Keyword(s):  
High Resolution ◽  
Dynamic Model ◽  
Numerical Evaluation ◽  
Helicopter Rotor ◽  
Vortex Interaction ◽  
Near Wake ◽  
Blade Vortex Interaction ◽  
Wake Model ◽  
Induced Velocity

Abstract The Beddoes near wake model, developed for high resolution blade vortex interaction computations, enables efficient numerical evaluation of the downwash due to trailed vorticity in the near wake of a helicopter rotor. The model is, however, limited by the assumption that the near wake lies in the plane of the rotor and, in some cases, by its inability to accurately evaluate the induced velocity contribution from vorticity trailed from inboard blade sections. In this paper, modifications to the method are proposed which address these issues and allow it to be used with confidence over a wider range of rotor flows.

Development and validation of a hybrid method for predicting helicopter rotor impulsive noise

2018 ◽  
Vol 233 (4) ◽  
pp. 1323-1339 ◽  
Author(s):  
Liangquan Wang ◽  
Guohua Xu ◽  
Yongjie Shi
Keyword(s):  
Hybrid Method ◽  
High Speed ◽  
Stokes Equations ◽  
Impulsive Noise ◽  
Computational Cost ◽  
Helicopter Rotor ◽  
Aerodynamic Load ◽  
Vortex Interaction ◽  
Blade Vortex Interaction ◽  
Hybrid Solver

Prediction of helicopter rotor impulsive noise is practically a very challenging task. This paper describes a hybrid method to predict rotor impulsive noise for both high-speed impulsive noise and blade–vortex interaction noise. The hybrid solver has been developed by combining the advantages of three different methods: (1) a computational fluid dynamics method based on Reynolds-averaged Navier–Stokes equations to account for the viscous and compressible effects near the blade; (2) a vorticity transport model to predict rotor wake system without artificial dissipation; and (3) an acoustic calculation method, based on Ffowcs-Williams Hawkings equation implemented to a permeable data surface. The developed hybrid solver is validated through available test data, for the cases of UH-1H model rotor, AH-1 Operational Loads Survey rotor, and Helishape 7A rotor. Peak sound pressure level of high-speed impulsive noise is accurately predicted with relative errors less than 7%. Additionally, acoustic waveform of blade–vortex interaction noise is well captured though it is sensitive to small changes in aerodynamic load. It is suggested that present hybrid method is versatile for the prediction of rotor impulsive noise with moderate computational cost.

Prediction of Helicopter Rotor Loads Using Time-Spectral Computational Fluid Dynamics and an Exact Fluid–Structure Interface

2011 ◽  
Vol 56 (4) ◽  
pp. 1-15 ◽  
Author(s):  
Seongim Choi ◽  
Anubhav Datta ◽  
Juan J. Alonso
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Dynamic Stall ◽  
Helicopter Rotor ◽  
Fluid Structure ◽  
Blade Shape ◽  
Coupling Procedure ◽  
Time Marching ◽  
Computational Structural Dynamics ◽  
Computational Fluid Dynamics Cfd

The objectives of this paper are to introduce time-spectral computational fluid dynamics (CFD) for the analysis of helicopter rotor flows in level flight and to introduce an exact fluid–structure interface for coupled CFD/computational structural dynamics (CSD) analysis. The accuracy and efficiency of time-spectral CFD are compared with conventional time-marching computations. The exact interface is equipped with an exact delta coupling procedure that bypasses the requirement for sectional airloads. Predicted loads are compared between time-spectral and time-marching CFD using both interfaces and validated using UH-60A flight data for high-vibration and dynamic stall conditions. It is concluded that time-spectral CFD can indeed predict rotor performance and peak-to-peak structural loads efficiently, and hence, open opportunity for blade shape optimization. The vibratory and dynamic stall loads, however, require a large number of time instances, which reduces its efficiency. The exact interface and delta procedure allow coupling to be implemented for arbitrary grids and advanced structural models exactly, without the requirement for two-dimensional sectional airloads.

scholarly journals Impact of Tip-Vortex Modeling Uncertainty on Helicopter Rotor Blade–Vortex Interaction Noise Prediction

2021 ◽  
Vol 66 (1) ◽  
pp. 1-13
Author(s):  
Stavros Vouros ◽  
Ioannis Goulos ◽  
Calum Scullion ◽  
Devaiah Nalianda ◽  
Vassilios Pachidis
Keyword(s):  
Vortex Core ◽  
Numerical Procedure ◽  
Tip Vortex ◽  
Analytical Models ◽  
Aerodynamic Noise ◽  
Helicopter Rotor ◽  
Vortex Interaction ◽  
Blade Vortex Interaction ◽  
Free Wake ◽  
The Impact

Free-wake models are routinely used in aeroacoustic analysis of helicopter rotors; however, their semiempiricism is accompanied with uncertainty related to the modeling of physical wake parameters. In some cases, analysts have to resort to empirical adaption of these parameters based on previous experimental evidence. This paper investigates the impact of inherent uncertainty in wake aerodynamic modeling on the robustness of helicopter rotor aeroacoustic analysis. A free-wake aeroelastic rotor model is employed to predict high-resolution unsteady airloads, including blade–vortex interactions. A rotor aeroacoustics model, based on integral solutions of the Ffowcs Williams–Hawkings equation, is utilized to calculate aerodynamic noise in the time domain. The individual analytical models are incorporated into an uncertainty analysis numerical procedure, implemented through nonintrusive Polynomial Chaos expansion. The potential sources of uncertainty in wake tip-vortex core growth modeling are identified and their impact on noise predictions is systematically quantified. When experimental data to adjust the tip-vortex core model are not available the uncertainty in acoustic pressure and noise impact at observers dominated by blade–vortex interaction noise can reach up to 25% and 3.50 dB, respectively. A set of generalized uncertainty maps is derived, for use as modeling guidelines for aeroacoustic analysis in the absence of the robust evidence necessary for calibration of semiempirical vortex core models.

A Numerical Investigation of the Influence of the Blade–Vortex Interaction on the Dynamic Stall Onset

2021 ◽  
Author(s):  
Camille Castells ◽  
François Richez ◽  
Michel Costes
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Dynamic Stall ◽  
Test Case ◽  
Vortex Interaction ◽  
Fluid Structure ◽  
Speed Test ◽  
Vortex Model ◽  
Flow Parameters ◽  
Blade Vortex Interaction

Recently, fluid–structure coupling simulations of helicopter rotors in high-thrust forward flight suggested that dynamic stall might be triggered by the blade–vortex interaction. However, no clear evidence of a correlation between dynamic stall and blade–vortex interaction has yet been given. We propose in this paper a simplified two-dimensional numerical model that can be used to indicate the role that the blade–vortex interaction plays in dynamic stall onset for different flight conditions. In this model, the rotor blade element is considered in pitching oscillation motion with a nonuniform translation, and a simplified vortex model can be introduced or not in the simulation to highlight the effect of blade–vortex interaction. All flow parameters of this simplified model are deduced from data provided by previous three-dimensional high-fidelity fluid–structure simulations. The method is used for validation and analysis of three flight conditions. The results show that, for the two cases with moderate advance ratio, the dynamic stall event is only triggered when a blade–vortex interaction occurs in the stall region. For the high-speed test case, the dynamic stall event seems to be only triggered by the very high angle of attack due to the motion of the blade.

Parallel Blade-Vortex Interaction Modelling for Helicopter Rotor Noise Control Synthesis

2014 ◽  
Vol 13 (7-8) ◽  
pp. 587-606 ◽  
Author(s):  
Sara Modini ◽  
Giorgio Graziani ◽  
Giovanni Bernardini ◽  
Massimo Gennaretti
Keyword(s):  
Noise Control ◽  
Helicopter Rotor ◽  
Control Synthesis ◽  
Vortex Interaction ◽  
Rotor Noise ◽  
Blade Vortex Interaction ◽  
Interaction Modelling

scholarly journals Computational fluid dynamics trimming of helicopter rotor in forward flight

2020 ◽  
Vol 12 (5) ◽  
pp. 168781402092525 ◽  
Author(s):  
Chenglong Zhou ◽  
Ming Chen
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Test Data ◽  
Pitch Angle ◽  
Flight Test ◽  
Helicopter Rotor ◽  
Main Rotor ◽  
Forward Flight ◽  
Thrust Coefficient ◽  
Moment Coefficient

A computational fluid dynamics (CFD) trimming method based on wind tunnel and flight test data is proposed. Aerodynamic coefficients obtained for a helicopter rotor using this method were compared with both experimental data from a test report and CFD results based on the control parameters that were reported in the same document. The method applies small disturbances to the collective pitch angle, the lateral cyclic pitch angle and the longitudinal cyclic pitch angle of the helicopter’s main rotor during forward flight to analyze the effects of each disturbance on the thrust coefficient, the pitching moment coefficient and the rolling moment coefficient of the rotor. Then, by solving a system of linear equations, the collective pitch angle, the lateral cyclic pitch angle and the longitudinal cyclic pitch angle of the main rotor in the CFD trim state are obtained. The AH-1G rotor is used in this paper. NASA has conducted a comprehensive flight test program on this model and has published detailed test reports. Using this method, the pitch moment and the roll moment can be corrected to almost zero and the calculated thrust coefficient is more consistent with the test data when compared with results from direct CFD simulations.

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