Rotor Blade In Hover Research Articles

Geometrically Exact Aeroelastic Stability Analysis of Composite Helicopter Rotor Blades in Hover by Updated VABS

Geometrically Exact Aeroelastic Stability Analysis of Composite Helicopter Rotor Blades in Hover by Updated VABS

Numerical Investigation of Flow Separation Control over Rotor Blades Using Plasma Actuator

The flow separation control over rotating rotor blades in hover using an alternating current dielectric barrier discharge (AC-DBD) plasma actuator is investigated through numerical simulation, possibly for the first time. Three-dimensional unsteady Reynolds-averaged Navier–Stokes calculation with a Reynolds stress turbulence model is adopted to resolve the airflow. The impact of the AC-DBD plasma discharge on the airflow is modeled by an empirical body force model. The objective of this work is to characterize the separated flow over the rotating blades, resolve the flow control procedure, reveal the flow control mechanism, and evaluate the control authority. Compared with the experiment, the numerical simulation can provide satisfactory predictions of the baseline flow over the rotor for the qualitative flow pattern and quantitative aerodynamic features. The good agreement between the simulation and experiment lends credibility to the observations made from the numerical result. It is found that in the baseline flow, the flow separation occurs in the vicinity of the blade tips and exhibits strong three-dimensional characteristics under the influence of centrifugal force. The plasma excitation can significantly inhibit the flow separation over the blades at large angles of attack (AOAs). The flow control is seen to be dominated by the spanwise vortices induced by the plasma discharges, for which the strong entrainment effect makes the separated flow fully or partially reattach to the blade surfaces. The resulting forced flow pattern is influenced by the AOA of the blades. At each high AOA under consideration, the controlled flow is a significant improvement over the baseline case, and the aerodynamic characteristics of the rotor blades are appreciably enhanced.

CFD simulation of helicopter rotor flow based on unsteady actuator disk model

Actuator Disks (AD) can provide characterizations of rotor wakes while reducing computational expense associated with modeling the fully resolved blades. This work presents an unsteady actuator disk method based on surface circulation distribution combined with empirical data, blade element theory and rotor momentum theory. The nonuniform circulation distribution accounts for 3D blade load effects, and in particular, tip loses. Numerical simulations were conducted for the isolated pressure sensitive paint model rotor blade in hover and forward flight using the HMB3 CFD solver of Glasgow University. Validation of CFD results in comparison with published numerical data was performed in hover, for a range of blade pitch angles using fully turbulent flow and the k-ω SST model. In forward flight, the vortex structures predicted using the unsteady actuator disk model showed configurations similar to the ones obtained using fully resolved rotor blades. Despite the reduced grid cells number, the CFD results for AD models captured well the main vortical structures around the rotor disk in comparison to the fully resolved cases.

Aeroelastic stability analysis of hingeless rotor blades in hover using fully intrinsic equations and dynamic wake model

PurposeThis paper aims to develop a new approach for aeroelastic analysis of hingeless rotor blades.Design/methodology/approachThe aeroelastic approach developed here is based on the geometrically exact fully intrinsic beam equations and three-dimensional unsteady aerodynamics.FindingsThe developed approach is accurate, fast and very useful in rotorcraft aeroelastic analysis.Originality/valueThis beam formulation has been never combined with three-dimensional aerodynamic model to be used for aeroelastic analysis of blades. In addition, it is possible to handle the composite blades, as well as blades with initial curvatures and twist with this proposed formulation.

Full-field optical deformation measurement and operational modal analysis of a flexible rotor blade

Abstract This study describes the deformation measurement and operational modal analysis (OMA) of an extremely flexible rotor blade in hover. Blade deformation of a 0.45 m-diameter two-bladed rotor was measured at two different root pitch angles and two rotational speeds up to 1500 RPM by a time-resolved digital image correlation (DIC) technique. The DIC technique successfully measured the time history of three dimensional displacements over the entire rotor blade at approximately 900 measurement locations for the flap, lead-lag, and torsional degrees of freedom. Measured blade deformation data were then processed with the Complexity Pursuit (CP) algorithm, to determine the dynamic characteristics of the rotor blade without input excitation information. The identified modal frequencies were within 5% of those obtained by another OMA technique called NExT-ERA as well as a numerical model. There was good agreement between the mode shapes estimated by the CP algorithm and the numerical model; however there was some discrepancy with the mode shapes identified by NExT-ERA. Although modal damping values were extracted by the CP algorithm, they were less than 3 % for all test cases, which is too low to account for aerodynamic damping. Overall, the modal identification process with the CP algorithm was found to be easier to implement and more accurate than the conventional NExT-ERA approach.

Systematic evaluation of the helicopter rotor blades: design variables and interactions

PurposeThe effects of rotor blade design variables and their mutual interactions on aerodynamic efficiency of helicopters are investigated. The aerodynamic efficiency is defined based on figure of merit (FM) and lift-to-drag responses developed for hover and forward flight, respectively.Design/methodology/approachThe approach is to couple a general flight dynamic simulation code, previously validated in the time domain, with design of experiment (DOE) required for the response surface development. DOE includes I-optimality criteria to preselect the data and improve data acquisition process. Desirability approach is also implemented for a better understanding of the optimum rotor blade planform in both hover and forward flight.FindingsThe resulting system provides a systematic manner to examine the rotor blade design variables and their interactions, thus reducing the time and cost of designing rotor blades. The obtained results show that the blade taper ratio of 0.3, the point of taper initiation of about 0.64 R within a SC1095R8 airfoil satisfy the maximum FM of 0.73 and the maximum lift-to-drag ratio of about 5.5 in hover and forward flight.Practical implicationsThe work shows the practical possibility to implement the proposed optimization process that can be used for the advanced rotor blade design.Originality/valueThe work presents the rapid and reliable optimization process efficiently used for designing advanced rotor blades in hover and forward flight.

Aeroelastic Response of a Hingeless Rotor Blade in Hover

AbstractIn this paper, the aeroelastic response of a hingeless rotor blade in hover is investigated. The hingeless rotor blade was modeled by using the geometrically exact fully intrinsic beam equa...

The effect of a movable mass on the aeroelastic stability of composite hingeless rotor blades in hover

In this paper, the aeroelastic stability of a composite hingeless rotor blade with a chordwise movable mass is investigated. The point mass is located near the tip of the blade and its chordwise location is variable with respect to the elastic axis and can be moved during the flight. This movable mass is added to the blade to actuate the blade twist during flight. By actuating the mass in the chord direction of the blade during the flight, a bending moment which is the result of the centrifugal force of the mass and its offset is induced on the blade. This bending moment induces twist in the blade, due to bend-twist coupling in the composite lamination. The blade is modelled by using the geometrically exact fully intrinsic beam equations along with the variational asymptotic beam sectional analysis. The aerodynamic loads are simulated by using the two-dimensional strip theory combined with a uniform inflow. The nonlinear partial differential aeroelastic equations are discretized by a time-space scheme, and the converged results are compared with those reported in the literature and a very good match is observed. The results show that by positioning the mass near the tip of the blade, and also by using the ply angle of about 30 degree in this configuration, the highest possible twist change is achieved when the mass moves from the leading edge to the trailing edge of the blade. Moreover, the spanwise location of the mass slightly changes the stability boundaries, while the chordwise movement significantly affects the aeroelastic instability.

Prediction of helicopter rotor noise in hover using FW-H analogy

The integral formulation of Ffowcs Williams and Hawkings (FW-H) analogy developed by Farassat is implemented to understand the sound generation and propagation of a rotating slender body like the helicopter rotor blade in hover. Using the linear acoustic theory, the thickness and loading noise terms are implemented in a flexible, new, in-house aero-acoustic tool. The pressure distribution on the blade surface is obtained from CFD simulation which sufficiently captures the correct flow characteristics of the rotor blade. Although, there is an improved estimation of the acoustic pressure signal in comparison to existing numerical noise signal results, there is still an under prediction of the peak amplitude compared to experimental values. The code is validated against the experiment for the UH-1H model helicopter rotor equipped with two rectangular blades having NACA0012 section.

Effect of Sweep on a Pitching Finite Wing

An investigation was performed into the effect of positive and negative sweep angle on the boundary layer transition and dynamic stall behavior of a finite wing. The finite wing had a 6:1 aspect ratio, modern (SPP8) tip shape, and positive twist, moving the maximum load on the wing away from the wind tunnel wall. Experiments were performed with sweep Λ = ±30° and Λ = 0° for static polars and sinusoidal pitching. The positively twisted wing shows a S-shaped boundary layer transition on the pressure side similar to that previously seen for helicopter rotor blades in hover. The transition positions on the suction side of the wing are comparable for the same local angle of attack at all values of the sweep at each of the three pressure sections, and for dynamic pitching motions a hysteresis around the static transition positions is seen. Sweeping the wing led to later stall and higher maximum lift for both static polars and dynamic stall, except for a single case. The negative aerodynamic damping is worse for the swept wing than for the unswept wing, except where the delay of stall led to the flow remaining attached.

Exploring the Detached-Eddy Simulation for Main Rotor Flows

This paper applies the Detached-Eddy Simulation (DES) method to resolve a larger part of the flow spectrum around rotor blades in hover and forward flight. A comparison between DES and Unsteady Reynolds–Averaged Navier–Stokes simulation was carried out for the case of a forward flying rotor suggesting that DES has great potential for rotor applications.

Aeroelastic Stability of Hingeless Rotor Blades in Hover Using Fully Intrinsic Equations

In this paper, the aeroelastic stability of a hingeless rotor blade in hover is investigated. The blade is modeled using the geometrically exact, fully intrinsic beam equations, whereas the aerodynamic loads applied on the blade are simulated based on the quasi-steady Greenberg aerodynamic theory. Using the generalized differential quadrature method, the resultant coupled aeroelastic equations are discretized and solved, and the eigenvalues of the linearized system are determined. By inspecting the eigenvalues, the stability of the system is examined for variations in different parameters. The obtained results are validated through comparison against those reported in the literature. Furthermore, the effects of offset of the aerodynamic center from the reference axis, the precone angle, the lead–lag frequency, static stall, and the drag coefficient on the stability boundaries of the system are evaluated. It is found that, by using the geometrically exact, fully intrinsic beam equations along with the generalized differential quadrature method, the aeroelastic instabilities of the hingeless rotor blades in hover can be determined accurately.

Numerical simulation of the tip aerodynamics and acoustics test

The application of an efficient flow control system on helicopter rotor blades may lead to improved aerodynamic performance. Recently, our invention of Rod Vortex Generators (RVGs) has been analyzed for helicopter rotor blades in hover with success. As a step forward, the study has been extended to forward flight conditions. For this reason, a validation of the numerical modelling for a reference helicopter rotor (without flow control) is needed. The article presents a study of the flow-field of the AH-1G helicopter rotor in low-, medium- and high-speed forward flight. The CFD code FLOWer from DLR has proven to be a suitable tool for the aerodynamic analysis of the two-bladed rotor without any artificial wake modelling. It solves the URANS equations with LEA (Linear Explicit Algebraic stress) k-ω model using the chimera overlapping grids technique. Validation of the numerical model uses comparison with the detailed flight test data gathered by Cross J. L. and Watts M. E. during the Tip Aerodynamics and Acoustics Test (TAAT) conducted at NASA in 1981. Satisfactory agreements for all speed regimes and a presence of significant flow separation in high-speed forward flight suggest a possible benefit from the future implementation of RVGs. The numerical results based on the URANS approach are presented not only for a popular, low-speed case commonly used in rotorcraft community for CFD codes validation but preferably for medium- and high-speed test conditions that have not been published to date.

Dynamic Stability of Hingeless Rotor Blade in Hover Using Padé Approximations

The dynamic stability of a hingeless rotor blade in hover is investigated using a finite-state, time-domain representation of the aerodynamic loads. The linearized equations of motion are obtained by perturbing the motion variables about nonlinear equilibrium deflections. The flutter analysis is performed in the time domain using three different approximations based on Loewy’s deficiency function. The most accurate approximation is obtained using a general nonlinear least-square optimization for the deficiency response function. This problem is subjected to zero and infinite frequency limits and nonnegative poles constraints. Numerical results for both divergence and flutter boundaries are depicted for a hingeless rotor blade configuration and compared to previous work. Two different scenarios for the flutter boundary are observed, depending on the range of lead–lag frequency, collective pitch, and the type of the approximations used.

Prediction of helicopter rotor noise in hover

Two mathematical models are used in this work to estimate the acoustics of a hovering main rotor. The first model is based on the Ffowcs Williams-Howkings equations using the formulation of Farassat. An analytical approach is followed for this model, to determine the thickness and load noise contributions of the rotor blade in hover. The second approach allows using URANS and RANS CFD solutions and based on numerical solution of the Ffowcs Williams-Howkings equations. The employed test cases correspond to a model rotor available at the KNRTU- KAI aerodynamics laboratory. The laboratory is equipped with a system of acoustic measurements, and comparisons between predictions and measurements are to be attempted as part of this work.

Aerodynamic analysis of potential use of flow control devices on helicopter rotor blades

The interest in the application of flow control devices has been rising in the last years. Recently, several passive streamwise vortex generators have been analysed in a configuration of a curved wall nozzle within the framework of the UFAST project (Unsteady Effects of Shock Wave Induced Separation, 2005 – 2009). Experimental and numerical results proved that the technology is effective in delaying flow separation. The numerical investigation has been extended to helicopter rotor blades in hover and forward flight applying the FLOWer solver (RANS approach) implementing the chimera overlapping grids technique and high performance computing. CFD results for hover conditions confirm that the proposed passive control method reduces the flow separation increasing the thrust over power consumption. The paper presents the numerical validation for both states of flight and the possible implementation of RVGs on helicopter rotor blades.

Computational aeroelastic analysis of a hovering W3 Sokol blade with gurney flap

This paper demonstrates the potential effect of a gurney flap on the performance of the W3-Sokol rotor blade in hover. A rigid blade was first considered and the calculations were conducted at several thrust settings. The gurney flap was extended from 46%R to 66%R and it was located at the trailing edge of the main rotor blade. Four different sizes of gurney flaps were studied, 2%, 1%, 0.5% and 0.3% of the chord. The biggest flap proved to be the most effective. A second study considered elastic blades with and without the gurney flap. The results were trimmed at the same thrust values as the rigid blade and indicate an increase of aerodynamic performance when the gurney flap is used, especially for high thrust cases.

Optimizing Rotor Blades with Approximate British Experimental Rotor Programme Tips

This work presents a framework for the optimization of certain aspects of a British Experimental Rotor Programme-like rotor blade in hover and forward flight so that maximum performance can be obtained from the blade. The proposed method employs a high-fidelity, efficient computational fluid dynamics technique that uses the harmonic balance method in conjunction with artificial neural networks as metamodels, and genetic algorithms for optimization. The approach has been previously demonstrated for the optimization of blade twist in hover and the optimization of rotor sections in forward flight, transonic aerofoils design, wing and rotor tip planforms. In this paper, a parameterization technique was devised for the British Experimental Rotor Programme-like rotor tip and its parameters were optimized for a forward flight case. A specific objective function was created using the initial computational fluid dynamics data and the metamodel was used for evaluating the objective function during the optimization using the genetic algorithms. The objective function was adapted to improve forward flight performance in terms of pitching moment and torque. 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 main computational cost was associated with the population of the genetic algorithms database necessary for the metamodel, especially because a full factorial method was used. The computational time of the optimization process itself, after the database has been obtained, is relatively insignificant. Therefore, the computational time was reduced with the use of the harmonic balance method as opposed to the time marching method. The novelty in this paper is two-fold. Optimization methods so far have used simple aerodynamic models employing direct “calls” to the aerodynamic models within the optimization loops. Here, the optimization has been decoupled from the computational fluid dynamics data allowing the use of higher-fidelity computational fluid dynamics methods based on Navier–Stokes computational fluid dynamics. This allows a more realistic approach for more complex geometries such as the British Experimental Rotor Programme tip. In addition, the harmonic balance method has been used in the optimization process.

고속압력감응페인트를 이용한 로터 블레이드 표면 압력 측정

본 논문은 고속압력감응페인트기법(Fast-Responding PSP)을 소개하고, 이를 이용하여 정지비행상태에 있는 축소형 로터 블레이드의 표면(윗면) 압력을 측정해봄으로서 PSP를 이용한 로터 블레이드 표면 압력 측정의 정확성과 그에 따른 실험기법의 타당성을 검증하기 위하여 수행되었다. 실험을 위한 광원으로는 532 nm 파장을 가지는 Pulsed laser를 사용하였고, PSP 측정 기법으로는 Lifetime 기법을 적용하였다. 또한, 모델 표면에 도포된 압력감응페인트는 반응성이 높은 Porous PSP가 사용되었다. 로터 블레이드는 NACA0012 익형을 가지고 있으며 길이 340 mm, 코드 40 mm의 직사각 형상 1종과 끝단의 후퇴각이 다른 4종의 형상을 사용하였다. 로터 블레이드의 콜렉티브 피치각 변화에 따른 표면 압력 분포를 측정하였으며 측정된 결과를 통해 콜렉티브 피치각이 증가할수록 윗면의 압력이 낮아지는 것을 정성적으로 확인하였고, 정량적인 압력계수는 NASA의 실험 데이터와 비교하여 약 0.4~0.7 정도 높은 경향성을 보였다. The present study was conducted by using fast-responding PSP technique to measure the surface pressure on a small-scale rotor blade in hover. Also, the study was performed to verify the accuracy and investigate its possibility of PSP application for rotor blade pressure measurement. Pulsed laser which has 532 nm wavelength was used as a light source. Lifetime measurement technique was applied. Also, the coated paint on a rotor blade was porous PSP which has faster response time than conventional PSP. The blades had NACA0012 airfoils. The length of rotor blade was 340 mm and chord was 40 mm with rectangular shape 1 set, and 4 sets had several tip sweepback angles. The measured results qualitatively showed that the upper surface pressure decreases with increasing the collective pitch angle. Quantitative pressure coefficients of PSP results were higher approximately 0.4 to 0.7 than the pressure tap data of the NASA experiment.

Panel/full-span free-wake coupled method for unsteady aerodynamics of helicopter rotor blade

A full-span free-wake method is coupled with an unsteady panel method to accurately predict the unsteady aerodynamics of helicopter rotor blades in hover and forward flight. The unsteady potential-based panel method is used to consider aerodynamics of finite thickness multi-bladed rotors, and the full-span free-wake method is applied to simulating dynamics of rotor wake. These methods are tightly coupled through trailing-edge Kutta condition and by converting doublet-wake panels to full-span vortex filaments. A velocity-field integration technique is also adopted to overcome singularity problem during the interaction between the rotor wake and blades. Helicopter rotors including Caradonna–Tung, UH-60A, and AH-1G rotors, are simulated in hover and forward flight to validate the accuracy of this approach. The predicted aerodynamic loads of rotor blades agree well with available measured data and computational fluid dynamics (CFD) results, and the unsteady dynamics of rotor wake is also well simulated. Compared to CFD, the present method obtains accurate results more efficiently and is suitable to rotorcraft aeroelastic analysis.