Vortical flow phenomena on helicopter rotors

Chapter 5 illustrates the genetic fuzzy system for health monitoring of a composite helicopter rotor in forward flight. The rotor is the most important component of the helicopter, and its health is critical for helicopter performance and control. Progressive damage accumulation is considered in the composite material. This damage model considers matrix cracking as the first damage type, followed by debonding/delamination, and finally fiber breakage. The damaged helicopter rotor is modeled using a finite element simulation which solves the rotor blade equations and vehicle trim equations. This aeroelastic simulation provides the blade response, blade and hub loads, strains, etc., for a damaged composite helicopter rotor in forward flight. The genetic fuzzy system is developed and tested for this helicopter rotor health monitoring problem. Different combinations of measurements are considered, and their advantages and shortcomings are evaluated. Finally, a life prediction approach is developed based on phenomenological damage growth models, and the genetic fuzzy system is illustrated for damage detection as well as life prediction for a helicopter rotor.

Numerical simulations of the unsteady flowfield of helicopter rotors on moving embedded grids

The wake has a strong influence on the aerodynamics of a helicopter rotor operating in various flight modes. An accurate prediction of its geometry is critical for an improved understanding of aerodynamic performance, vibration and acoustics characteristics. In this paper, a higher-order lifting-surface method is discussed that uses singularity elements having distributed vorticity to model the lifting surfaces and shed wakes, and applies it to the problem of computing the free-wake geometry of a helicopter rotor in forward flight. Such distributed vorticity elements allow the representation of a force-free wake-vortex sheet that does not suffer from numerical singularities and is numerically robust with regard to the wake roll-up behavior. In contrast with other free-wake methods that use discrete vortex filaments, the new method does not require the implementation of vortex core models etc. Using this new method, the distorted wake geometry of a rotor in forward flight are computed, along with the capability of accounting for the effect of inboard active devices on wake development, airloads and downwash distribution.

An overset grid thin-layer Navier-Stokes code has been extended to include dynamic motion of helicopter rotor blades through relative grid motion. The unsteady flowfield and airloads on an AH-IG rotor in forward flight were computed to verify the methodology and to demonstrate the method's potential usefulness towards comprehensive helicopter codes. In addition, the method uses the blade's first harmonics measured in the flight test to prescribe the blade motion. The solution was impulsively started and became periodic in less than three rotor revolutions. Detailed unsteady numerical flow visualization techniques were applied to the entire unsteady data set of five rotor revolutions and exhibited flowfield features such as blade vortex interaction and wake roll-up. The unsteady blade loads and surface pressures compare well against those from flight measurements. Details of the method, a discussion of the resulting predicted flowfield, and requirements for future work are presented. Overall, given the proper blade dynamics, this method can compute the unsteady flowfield of a general helicopter rotor in forward flight.

The simulation capabilities of the URANS solvers TAU and FLOWer of DLR have been extended in order to predict the laminar-turbulent transition onset on helicopter rotors. Therefore an approximate method for rotor blade transition prediction (RBT code) has been developed and coupled to the URANS solvers. The RBT code computes laminar boundary layer quantities based on a combined integral/Pohlhausen method. Transition is empirically predicted respecting Tollmien-Schlichting waves, laminar separation bubbles, crossflow, bypass instabilities and attachment line transition. To study the effects of blade-wake/vortex-interactions on the laminar boundary layer stability, the turbulence level of the rotor wake is taken into account for transition prediction. Test computations for two laminar flow airfoils reproduced the transition onset due to longitudinal and crossflow instabilities in close agreement to the experiment. For a helicopter rotor in forward flight, laminar-turbulent simulations were successfully demonstrated for both URANS flow solvers. A broad spectrum of different boundary layer instabilities could be identified at the blades. The predicted laminar flow at the rotor blades reduced the required rotor power by -4.5% at high speed forward flight compared to a fully turbulent simulation. The rotor thrust remained practically unaffected. The results were in good agreement between both URANS flow simulations and the experimental data.

This paper presents two methods for predicting the noise from helicopter rotors in forward flight. Aerodynamic and acoustic solutions in the near field are computed with a finite-difference solver for the Euler equations. Two different Kirchhoff acoustics methods are then used to propagate the acoustic signals to the far field in a computationally-efficient manner. One of the methods uses a Kirchhoff surface that rotates with the rotor blades. The other uses a nonrotating Kirchhoff surface. Results from both methods are compared to experimental data for both high-speed impulsive noise and blade-vortex interaction noise. Agreement between experimental data and computational results is excellent for both cases. The rotating and nonrotating Kirchhoff methods are also compared for accuracy and efficiency. Both offer high accuracy with reasonable computer resource requirements. The Kirchhoff integrations efficiently extend the near-field finite-difference results to predict the far field helicopter noise.

: It is assumed that a helicopter rotor in forward flight can be represented by a flat plate. Expressions are developed relating the pressure field to the steady aerodynamic thrust and moment and the time-dependent flapping moment. This pressure field satisfies the zero pressure condition at the center and edge of the disc, the first and second harmonic variation in lift, and Laplace's equation. A computer program was written to calculate the velocity field and streamlines. Two sample cases corresponding to high- and low- speed conditions for the UH-1 rotor were chosen for computation. Both cases showed the tip-vortex phenomenon. In order to examine the sensitivity of blade load calculations to wake geometry, an existing airloads program was modified to accept the computed wake geometry. A comparison of results of this modified program with experimental data indicated that the effect of distorted wake is greater at low advance ratios than at high advance ratios. The reduced sensitivity at high advance ratios may be due to the wake being blown farther behind the rotor and/or the predominance of other aeroelastic parameters.

Computationally fast harmonic balance methods for unsteady aerodynamic predictions of helicopter rotors

A new integral equation, and the corresponding boundary-element method, is presented for the aerodynamic analysis of compressible, subsonic, potential flows around bodies in arbitrary motion, with applications to helicopter rotors in hover. The formulation is closely related to that introduced by Morino for the aerodynamic analysis of aircraft: in that paper, the surface of the aircraft is assumed to be in arbitrary motion with respect to a frame of reference that moves in uniform translation. Such a formulation is very useful for the analysis of aircraft undergoing small oscillations with respect to a reference configuration. However, for the analysis of more complex motions (e.g. helicopter rotors in forward flight), it is more convenient to use a frame of reference that moves in arbitrary motion (in particular, for the case of a body moving in rigid-body motion, a frame of reference rigidly connected with the body).

Aeroelastic stability of complete rotors with application to a teetering rotor in forward flight

ACCURATE simulation of the viscous flow around a helicopter rotor is still a challenging problem. To predict the blade loading accurately, the numerical method must have a capability to capture the vortical wake generated from the rotor blade tip precisely, which significantly affects the aerodynamic performance of the rotor. Therefore, both the numerical scheme with low numerical dissipation and sufficient dense mesh far away from the blade surface are required for rotor computations to simulate the tip vortex with minimal distortion. Furthermore, for a hovering case, a large numerical integration time is needed for the capture of the fully developed vortical wake [1]. In addition, the existence of incompressible flow near the root will also make the convergence slow. Compared with fixed-wing computations, hovering rotor computations not only require more accurate numerical scheme but also will take much more run time to converge. During the past two decades many numerical methods [2–7] have been developed to simulate the viscous flow around a hovering rotor. The methods [2,3] with the single-block structured grid suffer from the grid-quality problem, because the mesh must be stretched and skewed to cover the complete computational domain and to match the periodic boundary condition. To avoid this problem, the chimera technique [4–7] suggested by Benek et al. [8] was introduced into hovering rotor computations. To capture the tip vortex without unphysical distortion and predict the blade loading accurately, some high-order schemes and vortex-adapted chimera methods [6,7] were presented. Although significant advances in the accurate simulation of the tip vortex have been made with these advanced methods, the large computation cost and the low robustness, as well as the complexity make them far away from engineering applications. As mentioned above, hovering rotor computations, especially the computations with high-order methods, require more run time to converge than their fixed-wing counterparts. A few attempts [1,9,10] have been made to use multigrid methods to accelerate the convergence of hovering rotor calculations. Up to now, there is few works concerning the multigrid method for the viscous flow computation about a hovering rotor based on overset grids. The main objective of the present study is to develop an efficient multigrid algorithm coupled with chimera grids for Navier–Stokes computations about a hovering rotor. It should be mentioned that the experience of this work can be extended to accelerate the computation about a rotor in forward flight using moving overset grids,which is themainmotivation of this research. Furthermore, this experience can also be the basis of the future multigrid computation using high-order methods and chimera grids for rotor flows.

A test was conducted to examine the feasibility of using the wide-field shadowgraph method for visualizing the wake of a small-scale helicopter rotor in forward flight. A wide range of test conditions were examined, including thrust and forward speed variations. Shadowgraphs from the side and top of the rotor were obtained. The visibility of the tip vortices using the shadowgraph method was documented for advance ratios up to 0.175. It was demonstrated that shadowgraphs of the rotor wake in forward flight can be used to obtain qualitative and quantitative information about wake geometry, wake/body interactions, and blade/vortex interactions. Video cameras were used for the first time to record the shadowgraphs. These provided several advantages compared to still cameras and greatly enhanced the capabilities of the shadowgraph method.

An individual blade controller designed to attenuate the aeroelastic response of helicopter rotors in forward flight by tailoring the blade root attachment conditions is developed. A feasibility analysis indicates that the open-loop controller which incorporates both passive and active design techniques is energy efficient and may be realized by available adaptive structures.

Based on CFD method and FW-Hpds equations, a numerical method for predicting aeroacoustic characteristics of coaxial rotor is established. In order to accurately simulate the flowfield of coaxial rotor, several techniques including moving-embedded grids, implicit LU-SGS scheme, ROE-MUSCL scheme and Spalart-Allmaras turbulence model are adopted to solve the RANS equations. The aeroacoustic characteristics of coaxial rotor in forward flight are calculated and analyzed by using FW-Hpds equations. The calculated acoustic characteristics of the AH-1/OLS model in forward flight for validation are in good agreements with the experimental data. The acoustic characteristics of coaxial-rotor in forward flight including the sound pressure time history and directivity patterns are investigated in detail. Some meaningful conclusions are obtained from the calculated results. The counter-rotation of upper and lower rotors has significant effects on the sound pressure time history and leads to different phases of peak value at different observation points. The noise radiation intensity of coaxial rotor in forward flight exhibits an approximately symmetrical distribution on both sides of the forward flight direction, which is one of the important differences between the coaxial rotor and the single rotor.

This paper examines the flow physics and principles of force production on a cycloidal rotor (cyclorotor) in forward flight. The cyclorotor considered here consists of two blades rotating about a horizontal axis, with cyclic pitch angle variation about the blade quarter-chord. The flow field at the rotor mid-span is analyzed using smoke flow visualization and particle image velocimeV are compared with flow fields predicted using 2D CFD and time-averaged force measurements acquired in an open-jet wind tunnel at three advance ratios. It is shown that the experimental flow field is nearly two dimensional at μ = 0.73 allowing for qualitative comparisons to be made with CFD. The incoming flow velocity decreases in magnitude as the flow passes through the retreating (upper) half of the rotor and is attributed to power extraction by the blades. A significant increase in flow velocity is observed across the advancing (lower) half of the rotor. The aerodynamic analysis demonstrates that the blades accelerate the flow through the lower aft region of the rotor, where they operate in a high dynamic pressure environment. This is consistent with CFD-predicted values of instantaneous aerodynamic forces which reveal that the aft section of the rotor is the primary region of force production. Phase-averaged flow field measurements showed two blade wakes in the flow, formed by each of the two blades. Analysis of the blades at several azimuthal positions revealed two significant blade-wake interactions. The locations of these blade-wake interactions are correlated with force peaks in the CFD-predicted instantaneous blade forces and highlight their importance to the generation of lift and propulsive force of the cyclorotor.