Search results

The purpose of this paper is to present the development and application of a numerical formulation for the structural dynamics and aeroelastic analysis of new generation helicopter and tiltrotor rotor blades. These are characterized by a curvilinear elastic axis, typically with the presence of tip sweep and anhedral angles.

The structural dynamics model implemented is based on nonlinear, flap‐lag‐torsion, rotating beam equations that are valid for slender, homogeneous, isotropic, non‐uniform, twisted blades undergoing moderate displacements. A second‐order approximation scheme for strain‐displacement is adopted. Aerodynamic contributions for aeroelastic applications are derived from sectional theories, with inclusion of wake inflow models to take into account three‐dimensional effects. The numerical integration is obtained through implementation within the COMSOL Multiphysics Finite‐Element‐Method (FEM) software code, considering the elastic axis of arbitrary curvilinear shape.

The computational tool developed is validated by comparisons with results available in the literature. These demonstrate the capability of the tool to accurately predict structural dynamics and aeroelastic behavior of curved‐axis rotor blades. In particular, the influence of sweep and anhedral angles at the blade tip is successfully captured.

The numerical tool developed is limited to the analysis of isotropic blades, with a simple sectional aerodynamic modeling for aeroelastic applications. However, the flexibility of the process through which the proposed tool has been developed is such that a moderate effort is required for its extension to composite blades and more accurate aerodynamic loads predictions.

The proposed computational solver is a reliable tool for preliminary design and optimal design processes of helicopter and tiltrotor rotor blades.

Computational tools for rotors with advanced‐geometry blades are not commonly available. Therefore, the presentation of a successful way to implement structural dynamics/aeroelastic mathematical formulations for rotor blades with curvilinear elastic axis in highly flexible, multiphysics, FEM‐based, commercial software may be of interest for designers and researchers.

This paper aims to compare the comprehensive rotorcraft analyses using the two different blade section property data sets for the blade natural frequencies, airloads, elastic deformations, the trimmed rotor pitch control angles and the blade structural loads of a small-scale model rotor in a blade vortex interaction (BVI) phenomenon.

The two different blade section property data sets for the first Higher-harmonic control Aeroacoustic Rotor Test (HART-I) are considered for the present rotor aeromechanics analyses. One is the blade property data set using the predicted values which is one of the estimated data sets used for the previous validation works. The other data set uses the measured values for an uninstrumented blade. A comprehensive rotorcraft analysis code, CAMRAD II (comprehensive analytical model of rotorcraft aerodynamics and dynamics II), is used to predict the rotor aeromechanics such as the blade natural frequencies, airloads, elastic deformations, the trimmed rotor pitch control angles and the blade structural loads for the three test cases with and without higher-harmonic control pitch inputs. In CAMRAD II modelling with the two different blade property data sets, the blade is represented as a geometrically nonlinear elastic beam, and the multiple-trailer wake with consolidation model is used to consider more elaborately the BVI effect in low-speed descending flight. The aeromechanics analysis result sets using the two different blade section property data sets are compared with each other as well as are correlated with the wind-tunnel test data.

The predicted blade natural frequencies using the two different blade section property data sets at non-rotating condition are quite similar to each other except for the natural frequency in the fourth flap mode. However, the natural frequencies using the predicted blade properties at nominal rotating condition are lower than those with the measured blade properties except for the second lead-lag frequency. The trimmed collective pitch control angle with the predicted blade properties is higher than both the wind-tunnel test data and the result using the measured blade properties in all the three test cases. The two different blade property data sets both give reasonable predictions on the blade section normal forces with BVI in the three test cases, and the two analysis results are reasonably similar to each other. The blade elastic deformations at the tip using the measured blade properties are correlated more closely with the wind-tunnel test data than those using the predicted blade properties in most correlation examples. In addition, the predictions of blade structural loads can be slightly or moderately improved by using the measured blade properties particularly for the oscillatory flap bending moments. Finally, the movement of the sectional centre of gravity location of the uninstrumented blade has a moderate influence on the blade elastic twist at the tip in the baseline case and the oscillatory flap bending moment in the minimum noise case.

The present comparison study on rotor aeromechanics analyses using the two different blade property data sets will show the influence of blade section properties on rotor aeromechanics analysis.

This paper is the first attempt to compare the aeromechanics analysis results using the two different blade section property data sets for all three test cases (baseline, minimum noise and minimum vibration) of HART-I in low-speed descending flight.

This paper aims to develop a new approach for aeroelastic analysis of hingeless rotor blades.

The aeroelastic approach developed here is based on the geometrically exact fully intrinsic beam equations and three-dimensional unsteady aerodynamics.

The developed approach is accurate, fast and very useful in rotorcraft aeroelastic analysis.

This 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.

This study aims to carry out flutter stability and vibration analysis of a uniform hingeless rotor blade in hovering flight conditions.

The perturbation equations are obtained using the governing differential equations of motion derived in Part I (Aircraft Engineering and Aerospace Technology, Vol. 79 No. 2, 2007). The differential transform method (DTM) is applied to the perturbation equations of motion and the transformed equations are coded in the computer package Mathematica.

The effects of the built‐in pretwist angle and the rotational speed ratio on the natural frequencies are investigated and the results are compared with the results in literature.

This study, in carrying out an analysis of flutter stability and vibration of a uniform hingless rotor blade, is in good agreement with previous studies in the literature.

This study aims to derive the kinetic and the potential energy expressions of a rotating uniform hingeless rotor blade and the aerodynamic loads that act on the blade element in hovering flight conditions.

The blade is modeled as an Euler‐Bernoulli beam. The governing partial differential equations of motion and the associated boundary conditions are derived using the Hamilton's principle.

The derivations of the energy expressions and the aerodynamic loads are made in a detailed way by including several explanatory tables. The resultant equations of motion are in good agreement with the literature. Additionally, in this work the hub radius effect is included in the equations of motion.

Arguably this study achieves a breakthrough in deriving the kinetic and the potential energy expressions of a rotating uniform hingeless rotor blade and the aerodynamic loads that act on the blade element in hovering flight conditions.

THE first prototype WG.13 Lynx made its first flight at Yeovil on 21 March, 1971, piloted by Mr W. R. Gellatly, the Chief Test Pilot of Westland Helicopters Ltd. It was the last of the three helicopters in the Anglo/French package deal to fly and the only one with design leadership on this side of the Channel. It also has the distinction of being the first new Westland design for many years. Although the Lynx looks conventional, it incorporates many technical advances, particularly noteworthy in the fundamental design of the rotor head, blades and gearbox. Notable among the new features are the semi‐rigid or hingeless rotor head, strikingly simple compared with the conventional rotor head, and the ‘conformal’ gear train which can transmit twice the power as could conventional gearing of the same size. The key factor in the design has been the requirement of reliability and this has been achieved wherever possible by simplicity.

Studies of the hingeless rotor helicopter dynamic stability and control laws are conducted. A new method is given for the calculation of stability and controllability of a helicopter in flight condition with lateral velocity. First, the rotary wing dynamic model considered is the one of flap‐pitch (including the elastic deformation of control system) – torsion coupling. The induced velocity non‐uniform distribution derived from vortex theory is taken into account. Then, according to the established motion model of the helicopter, the effects of induced velocity distribution, flap‐pitch‐torsion coupling and lateral velocity on the stability and controllability of the helicopter are analyzed. Based on the analyses of dynamic stability of the helicopter, the unstable mode and the necessity of installation of stability augmentation system (SAS) are recognized. Finally, the control laws of SAS for helicopter pitching, rolling and yawing motions are presented. After establishing helicopter flight control state equations, the performance analyses and step response simulation for helicopter SAS are carried out.

In this paper results of numerical analysis of the ground resonance are shown. This analysis is performed making use of the complete model of the one‐main rotor helicopter. This model is adopted from flight mechanics. For the basic analysis it is assumed that the helicopter fuselage is a rigid body and the main rotor consists of four rigid blades. Each blade performs motions about its horizontal flapping hinge and vertical lagging hinge. The tail rotor is treated as a hingeless and weightless source of thrust, which equilibrates the drag moment and ensures directional control of the helicopter. For analysis of the ground resonance phenomenon, forces and moments produced by the landing gear are taken into account – their rigidity and damping are included into consideration.

This paper aims to validate a comprehensive aeroelastic analysis for a helicopter rotor with the higher harmonic control aeroacoustic rotor test (HART‐II) wind tunnel test data.

Aeroelastic analysis of helicopter rotor with elastic blades based on finite element method in space and time and capable of considering higher harmonic control inputs is carried out. Moderate deflection and coriolis nonlinearities are included in the analysis. The rotor aerodynamics are represented using free wake and unsteady aerodynamic models.

Good correlation between analysis and HART‐II wind tunnel test data is obtained for blade natural frequencies across a range of rotating speeds. The basic physics of the blade mode shapes are also well captured. In particular, the fundamental flap, lag and torsion modes compare very well. The blade response compares well with HART‐II result and other high‐fidelity aeroelastic code predictions for flap and torsion mode. For the lead‐lag response, the present analysis prediction is somewhat better than other aeroelastic analyses.

Predicted blade response trend with higher harmonic pitch control agreed well with the wind tunnel test data, but usually contained a constant offset in the mean values of lead‐lag and elastic torsion response. Improvements in the modeling of the aerodynamic environment around the rotor can help reduce this gap between the experimental and numerical results.

Correlation of predicted aeroelastic response with wind tunnel test data is a vital step towards validating any helicopter aeroelastic analysis. Such efforts lend confidence in using the numerical analysis to understand the actual physical behavior of the helicopter system. Also, validated numerical analyses can take the place of time‐consuming and expensive wind tunnel tests during the initial stage of the design process.

While the basic physics appears to be well captured by the aeroelastic analysis, there is need for improvement in the aerodynamic modeling which appears to be the source of the gap between numerical predictions and HART‐II wind tunnel experiments.

This paper aims to present the design of a stability augmentation system (SAS) in the longitudinal and lateral axes for an unstable helicopter.

The feedback controller is designed using linear quadratic regulator (LQR) control with full state feedback and LQR with output feedback approaches. SAS is designed to meet the handling qualities specification known as Aeronautical Design Standard (ADS‐33E‐PRF). A helicopter having a soft inplane four‐bladed hingeless main rotor and a four‐bladed tail rotor with conventional mechanical controls is used for the simulation studies. In the simulation studies, the helicopter is trimmed at hover, low speeds and forward speeds flight conditions. The performance of the helicopter SAS schemes are assessed with respect to the requirements of ADS‐33E‐PRF.

The SAS in the longitudinal axis meets the requirement of the Level 1 handling quality specifications in hover and low speed as well as for forward speed flight conditions. The SAS in the lateral axis meets the requirement of the Level 2 handling quality specifications in both hover and low speed as well as for forward speed flight conditions. The requirements of the inter axis coupling is also met and shown for the coupled dynamics case. The SAS in lateral axis may require an additional control augmentation system or adaptive control to meet the Level 1 requirements.

The study shows that the design of a SAS using LQR control algorithm with full state and output feedbacks can be used to meet ADS‐33 handling quality specifications.