Rotor Blades In Forward Flight Research Articles

Numerical investigation on the use of various blade tips for the helicopter's main rotor blade in forward flight regarding aerodynamic performance and HSI noise considerations

The goal of this research is to investigate aerodynamic performance and high-speed impulsive (HSI) noise of various blade tips for the helicopter's main rotor in forward flight. For this purpose, three-dimensional compressible flow field around the rotor blades is numerically simulated by the solution of the unsteady Reynolds averaged Navier-Stokes (URANS) equations with the "SST k-ω" turbulence model. Solution is quantified by the calculation of torque, drag, and thrust coefficients of CQ, CD, CT, and L/D ratio. HSI noise on the rotor plane is assessed by the calculation of sound pressure level at certain points of the flow field using the Ffowcs Williams-Hawking (FW-H) equation. Blade tips are eleven and include anhedral, dihedral, Eagle and combinations of them, and also Blue Edge and rectangular. The combinations are inspired by the tip configurations of the fixed wings. In this regard effects of the tip geometry parameters including curvature and height of its parts have been considered as well. All of the tips are examined for two flight conditions of Caradonna and Tung two-blade rotor with fixed pitch angle and time-varying pitch angles. Based on the present results, it is found that blade tips of dihedral family do not improve the aerodynamic performance of the helicopter blade. Also, addition of anhedral part to the blade tip improves the L/D ratio. According to the present acoustic analysis, it was seen that after the Blue Edge, the Eagle tip stands in order of priority. For the condition of Caradonna and Tung two-blade rotor with tip Mach number of 0.8 and advance ratio of 0.2, HSI noise of the Eagle-tip blade decreases by 2.39 % compared to that of the two-blade rotor with rectangular tip. However, the aerodynamics performance of the Eagle tip is much better than that of the Blue Edge.

Geometrically exact aeroelastic stability analysis of helicopter composite rotor blades in forward flight

In this paper, a sophisticated method for analyzing the geometrically exact aeroelastic stability of composite rotor blades in forward flight has been proposed. The unique advantage of the proposed method is to use the geometrically exact beam model with the updated Variational Asymptotic Beam Sectional analysis (VABS) for structural modeling to not only calculate the blade cross-sectional properties and blade spanwise deformations more accurately, but also take into account the effect of transverse shear deformation and blade initial twist and curvatures sophisticatedly. The Peters finite state airloads theory and the Peters-He finite state dynamic inflow model are coupled to calculate the three-dimensional unsteady airloads applied on the blade. The auto-pilot trim scheme is adopted to determine the blade pitch controls. The finite element spatial discretization method, the Newmark numerical integration method, the Newton-Raphson method and the moving-block analysis method are combined to solve the established geometrically exact aeroelastic equations and obtain the geometrically exact aeroelastic stability of composite rotor blades. The proposed method is validated by correlation with existing analytical results. The investigations show that the effect of transverse shear deformation on the aeroelastic stability of composite hingeless rotors in forward flight is non-negligible. Blade initial curvatures have significant effect on the aeroelastic stability of composite hingeless rotors in forward flight, which is not only related to the blade elastic couplings, but also to the advance ratio.

Geometrically Exact Beam-based Aeroelastic Modeling and Solution Of Composite Rotor Blades in Forward Flight

In this paper, the geometrically exact beam model and aeroelastic solution methods for composite rotor blades in forward flight by the latest variational asymptotic beam sectional analysis (VABS) have been employed. The geometrically exact beam equations of motion in the mixed variational form and the latest VABS are used to deal with one-dimensional blade analysis and the structural property of blade cross section, respectively. The methods can be used for the aeroelastic solution of composite rotor blades with arbitrary cross-sectional shape and material distribution, large deflections and significant nonclassical effects such as cross-sectional warping, transverse shear deformation, and elastic couplings caused by anisotropic material properties. The Peters–He finite state dynamic inflow model and the Peters finite state airloads theory are used to calculate the induced velocity and blade airloads, respectively. An auto-pilot trim scheme is used for calculating the blade pitch controls to meet the trim requirements. The convergence issue encountered when solving the geometrically exact, mixed variational aeroelastic equations in time domain has been successfully addressed. The values of the empirical parameters in the auto-pilot trim scheme for the presented aeroelastic model have been properly selected. The accuracy of the presented aeroelastic modeling and solution methods has been verified against the SA349/2 flight-test data. The influence of transverse shear deformation on the aeroelastic response of composite rotor blades was also investigated, indicating that this effect has a nonnegligible influence on the aeroelastic response of the five different kinds of elastically coupled hingeless composite rotors investigated in this paper.

Speculation of local aerodynamic loads on helicopter rotor blade in forward flight

Design of rotor blade is mainly the most important consideration in helicopter industry. A sufficient input data is needed for various flight conditions that should be accounted for in rotor design process. In this paper, blade element theory is applied to helicopter rotor blade in forward flight. The main objective is to assure the fidelity of using blade element theory in helicopter rotor blade design. Blade element theory is used for calculating the average lift, thrust, power for rotary wings and the local aerodynamic loads. The local non-dimensional normal force coefficients obtained from blade element theory were compared with experimental results of the 7A rotor at five non dimensional radial positions for the high speed test point 312. The results showed the same pattern of variation along the azimuth with a considered convergence along the blade radius.

Parametric analyses on dynamic stall control of rotor airfoil via synthetic jet

The effects of synthetic jet control on unsteady dynamic stall over rotor airfoil are investigated numerically. A moving-embedded grid method and an Unsteady Reynolds Averaged Navier-Stokes (URANS) solver coupled with k-ω Shear Stress Transport (SST) turbulence model are established for predicting the complex flowfields of oscillatory airfoil under jet control. Additionally, a velocity boundary condition modeled by sinusoidal function has been developed to fulfill the perturbation effect of periodic jet. The validity of present CFD method is evaluated by comparisons of the calculated results of baseline dynamic stall case for rotor airfoil and jet control case for VR-7B airfoil with experimental data. Then, parametric analyses are conducted emphatically for an OA212 rotor airfoil to investigate the effects of jet control parameters (jet location, dimensionless frequency, momentum coefficient, jet angle, jet type and dual-jet) on dynamic stall characteristics of rotor airfoil. It is demonstrated by the calculated results that efficiency of jet control could be improved with specific momentum coefficient and jet angle when the jet is located near separation point of rotor airfoil. Furthermore, the dual-jet could improve control efficiency more obviously on dynamic stall of rotor airfoil with respect to the unique jet, and the influence laws of dual-jet’s angles and momentum coefficients on control effects are similar to those of the unique jet. Finally, unsteady aerodynamic characteristics of rotor via synthetic jet which is located on the upper surface of rotor blade in forward flight are calculated, and asa result, the aerodynamic characteristics of rotor are improved compared with the baseline. The results indicate that synthetic jet has the capability in improving aerodynamic characteristics of rotor.

Experimental Investigation of Unsteady Transition on a Pitching Rotor Blade Airfoil

The unsteady flow around a pitching helicopter main rotor blade airfoil, the EDI-M109, was experimentally investigated at conditions similar to those existing on a retreating rotor blade in forward flight. High-speed pressure measurements and hot film anemometry were used to investigate the unsteady transition characteristics of the airfoil. Results are presented for dynamic test points with attached flow, light dynamic stall, and deep dynamic stall at M = 0.3, Re = 1.8 × 106, and f = 6.6 Hz. The results include the discussion of the periodicity of the hot film signals for different flow states. The transition process of the pitching airfoil is analyzed, and the significance of the intermittent region is described. A time delay between the transition and the model motion is discussed, and a linear relationship between the transition position and time is observed. The influences of the pitching amplitude on the transition characteristics are discussed, and the flow separation initiating dynamic stall is analyzed.

Optimisation of aspects of rotor blades in forward flight

This work presents a method for the optimisation of aspects of rotor blade shape in forward flight. The proposed technique employs CFD in conjunction with artificial neural networks (ANNs) and genetic algorithms (GAs). The developed method was used to optimise the anhedral and sweep of the UH60-A rotor blade in forward flight. A parameterisation method was defined, a specific objective function was created using the initial CFD data and the metamodel was used for evaluating the objective function during the optimisation. The obtained results suggest optima in agreement with engineering intuition but provide precise information about the shape of the final lifting surface and its performance. The results were checked using different optimisation methods and metamodels and were not sensitive to the employed techniques with substantial overlap between the outputs of the selected methods. The main CPU cost was associated with populating the CFD database necessary for the metamodel.

Chaotic Response of an Airfoil due to Aeroelastic Coupling and Dynamic Stall

Flight-test data of helicopters indicate that vibratory levels in the fuselage exhibit a wide spectrum of frequencies, including a few below the rotor revolutions per minute. It is well known that helicopter blades operate in a complex aerodynamic environment, involving time-varying heave, pitch, and pulsating oncoming flow. During operation, some sections of the rotor blade undergo dynamic stall once in a revolution. This paper attempts to understand the reason for the existence of several frequencies in the response of the fuselage and the possible cause for this observed phenomenon by analyzing the effects of dynamic stall and aeroelastic couplings on the response of 2-D airfoil. The ONERA dynamic stall model developed by Petot is modified by incorporating a higher-order rational approximation of Theodorsen's lift deficiency function. This improved model is shown to provide a better correlation with experimental stall data. The response characteristics of a 2-D airfoil undergoing pitching and plunging motion in a pulsating oncoming flow, simulating the response of a cross section of a helicopter rotor blade in forward flight are analyzed. This study shows significant difference in the response characteristics of the airfoil for unsteady (dynamic stall model) and quasi-steady aerodynamic models. It is observed that the nonlinear aerodynamics (dynamic stall effects) in association with aeroelastic couplings above a certain level lead to a bounded chaotic motion of the airfoil.

On the effect of matrix cracks in composite helicopter rotor blade

This paper investigates the feasibility of an on-line damage detection capability for helicopter main rotor blades made of composite material. Damage modeled in the composite is matrix cracking. A box-beam with stiffness properties similar to a hingeless rotor blade is designed using genetic algorithm for the typical [± θ m /90 n ] s family of composites. The effect of matrix cracks is included in an analytical model of composite box-beam. An aeroelastic analysis of the helicopter rotor based on finite elements in space and time is used to study the effects of matrix cracking in the rotor blade in forward flight. For global fault detection, rotating frequencies, tip bending and torsion response, and blade root loads are studied. It is observed that the effect of matrix cracking on lag bending and elastic twist deflection at the blade tip and blade root yawing moment is significant and these parameters can be monitored for online health monitoring. For implementation of local fault detection technique, the effect on axial and shear strain, for matrix cracks in the whole blade as well as matrix cracks occurring locally is studied. It is observed that using strain measurement along the blade it is possible to locate the matrix cracks as well as to predict density of matrix cracks.

A new experimental method for determining local airloads on rotor blades in forward flight

This paper presents a new approach for determining local airloads on helicopter rotor blade sections in forward flight. The method is based on the momentum equation in which all the terms are expressed by means of the velocity field measured by a laser Doppler velocimeter. The relative magnitude of the different terms involved in the momentum and Bernoulli equations is estimated and the results are encouraging.

Technical Note: Aerodynamic Design of Helicopter Rotor Blade in Forward Flight Using Response Surface Methodology

Technical Note: Aerodynamic Design of Helicopter Rotor Blade in Forward Flight Using Response Surface Methodology

Laplace-domain approximation to the transfer functions of a rotor blade in forward flight

AbstractA continuous frequency domain method with roots on the classic Hill’s determinant analysis is presented to approximate the time-varying characteristics of a linear periodic system. The method is particularly useful to derive a time-invariant equivalent form of the time-varying aeroelastic problem of a rotor blade in forward flight. The proposed technique allows methodology usually employed in fixed wing aircraft to obtain closed-loop control laws be extended to rotary wings. The method is first validated solving Mathieu’s equation. Next, the two-degree-of-freedom (flap bending and torsion) problem of rotating beam subject to unsteady and incompressible aerodynamics in forward flight is solved in the laplace domain. As a demonstration of the proposed method, the transfer functions in the ‘s’ plane between a sudden and uniformly distributed input pressure perturbation applied along the beam and the output response of the two elastic degrees of freedom considered are obtained at a set of local sections.

Aeroelastic response and stability analysis of composite rotor blades in forward flight

Abstract The aeroelastic phenomena of a composite rotor blade are investigated using a finite element method. The aeroelastic equations of motion of the rotor blade are formulated using a large deflection-type beam model that has no artificial restrictions on the magnitudes of displacements and rotations due to the degree of nonlinearity. The sectional elastic constants of a composite box beam including warping deformations are determined from the refined cross-sectional finite element method. The nonlinear, periodic blade steady response is obtained using a time finite element method on full finite element equation. The aeroelastic stability of a rotor blade is investigated by linearizing the dynamics of the rotor blade about the nonlinear equilibrium position. Numerical results of the steady equilibrium deflections and the aeroelastic modal damping are presented for various configurations of composite rotor blades and are compared with those obtained from a previously published modal analysis using a moderate deflection-type beam theory.

Aeroelastic Analysis of a Hingeless Rotor Blade in Forward Flight

The aeroelastic response and stability of isotropic and composite rotor blades are investigated using a large deflection-type beam theory. The finite element equations of motion for beams undergoing arbitrary large displacements and rotations, but small strains, are obtained from Hamilton's principle. The sectional elastic constants of a composite box beam including warping deformations are determined from the refined cross-sectional finite element method. The analysis is performed for a soft-in-plane hingeless rotor in free flight propulsive trim. The nonlinear periodic blade steady response is obtained by integrating the full finite element equation in time through a coupled trim procedure with a vehicle trim. After the coupled trim response is computed, the aeroelastic response is calculated through a time-marching solution procedure under small perturbations assumption, and then the stability analysis is performed by using a moving block analysis. Numerical results of rotating natural frequencies, blade response, and aeroelastic stability are presented. The results of the full finite element analysis using the large deflection-type beam theory are quite different from those of a previously published modal analysis using the moderate deflection-type beam theory.

Aeroelastic analysis of a hingeless rotor blade in forward flight

Aeroelastic analysis of a hingeless rotor blade in forward flight

Periodic Flap Control of a Helicopter Blade in Forward Flight

The primary objective of this paper is to present new periodic control strategies for the control of flapping motion of an individual helicopter rotor blade in forward flight, which is represented by a differential equation with periodic coefficients. First, an algebraic procedure based on Chebyshev polynomial expansion is employed to control the periodic flapping motion. In this approach, the state vector and the elements of the periodic system matrix have been expanded in terms of shifted Chebyshev polynomials of the first kind over the principal period. Later, optimal control theory in conjunction with Floquet theory has been used to design full-state and observer-state feedback controllers. In the second method, the feedback controllers have been designed in the time-invariant domain through the application of a linear, invertible, periodic transformation known as Liapunov-Floquet (L-F) transformation to the periodic system model. It is shown that by using both methods the periodic control gains can be obtained as explicit functions of time and, therefore, a control scheme more suitable for real-time implementation can be achieved. The design procedures have been found to be much simpler in character when compared to those techniques that have appeared in the literature.

Response of a nonrotating rotor blade to lateral turbulence Part II: Experiment

In Part I of this work, a theoretical simulation study of rotor blade response to turbulence in forward flight is presented. For verification of this theoretical computational method, a new experimental method based on a special gust field generated by a rotating slotted cylinder with an airfoil (RSC/airfoil) in a wind tunnel is developed to simulate the aerodynamic environment of a rotating rotor blade in forward flight. This gust generator can produce a single harmonic gust wave and also turbulence with uniform power spectral density over a certain frequency band in the lateral and longitudinal directions. In this article, quantitative comparisons are also made with theoretical results for both random and nonrandom parametric excitation. The quantitative agreement between theory and experiment indicates that this experimental method is useful.

Response of a Nonrotating Rotor Blade to Lateral Turbulence Part I: Theory

Theoretical simulation of a rotor blade in forward flight by a nonrotating rotor blade in a longitudinal sinusoidal pulsating flow, and the flapping and torsional response of a flexible nonrotating rotor blade model to lateral turbulence have been investigated. A direct time domain computational method using a modified linear ONERA aerodynamic model and a time-frequency approach using the classical aerodynamic model have been proposed. A theoretical lift comparison between the classical aerodynamic theory and the modified ONERA model is made. The numerical calculations indicate that the statistically quantitative agreement for both flap and torsional variance responses between the linear ONERA and classical aerodynamic models is reasonably good. The effects of random parametric excitation (when the longitudinal flow includes a turbulence component) and parameter variations are discussed. The numerical results are used to confirm the validity of a new experimental method presented as a companion paper, Part II. a.c b, b C, Cm dFM, dMv dM0 E e

Euler Calculations of Unsteady Interaction of Advancing Rotor with a Line Vortex

The unsteady, three-dimensional flowfield of a helicopter rotor blade in forward flight encountering a concentrated line vortex is calculated using an implicit, finite difference numerical procedure for the solution of Euler equations. A prescribed vortex method is adopted to preserve the structure of the interacting vortex. The test cases considered for computation correspond to the two-bladed model rotor experimental conditions of Caradonna et al. and consist of parallel and oblique interactions

Optimal output feedback for linear time-periodic systems

An approach is developed for applying optimal output feedback control theory to the design of fixed gain controllers for time-periodic systems. Constant feedback gains based on plant outputs are calculated by minimizing a linear quadratic performance cost functional. A solution algorithm is developed using Floquet- Lyapunov theory and a generalized harmonic expansion technique. The theory is applied to the control of a helicopter rotor blade in forward flight through individual blade control. It is shown that constant gain output feedback can be used to augment the stability of the resulting time-periodic system, with the potential for greatly reducing the implementation complexity. HE primary purpose of this paper is to present a general framework for applying output feedback techniques to linear time-periodic systems. The motivation for such an ap- proach is, in part, to reduce the control implementation re- quirements. The use of constant gain optimal output feedback in designing active feedback controllers provides many advan- tages over existing linear quadratic Gaussian (LQG) design methods. In particular, an on-line observer is not required, and the method can be extended to include compensator de- sign wherein the order of the compensator may be prespecified during the design process. Unneeded feedback channels may also be eliminated. 1 The number of sensors may be reduced while eliminating the need to estimate all of the unmeasurable states in real time. Recently, efficient compensator design formulations have been developed for linear time-invariant systems including robustness considerations for both unstruc- tured uncertainty and parametric uncertainty.2'3 The potential advantages of designing a fixed-order linear time-invariant controller are especially appealing for systems whose equations of motion are characterized by linear peri- odic coefficient dynamics. Aerospace applications include he- licopter forward flight stability and control and satellite or- bital stability problems. For helicopter aeromechanical and aeroelastic stability and control applications, the number of states required to model the system accurately continues to increase as research in structural elasticity, dynamic inflow, and aeroelasticity progresses. However, the feasibility and cost effectiveness of measuring/est imating all of the higher frequency structural elastic and dynamic inflow positions and rates for helicopter rotors in real time in an operational envi- ronment is questionable. This also drives the need for poten- tial alternatives to the full state feedback solution. An output feedback approach provides additional flexibility in that higher fidelity state space physical models may be used for the