Helicopter Rotor In Forward Flight Research Articles

Experimental investigation on the performance degradation of helicopter rotor due to ice accretion

Helicopter rotor performance can be significantly altered by ice accretion which directly degrades the handling qualities and safety during flight. To obtain understanding into the effect of ice accretion on a helicopter rotor in forward flight under different icing environments, an experimental campaign was conducted in an icing wind tunnel using a rotor model with diameter at 2 m. Research was emphasized on revealing effect of liquid water content and static temperature on the rotor performance degradation. The dynamic thrust and torque of the rotor were measured by a six-component balance. Ice shape of the airfoils at specific rotor blade spanwise locations was obtained through the ice cutting process; meanwhile, the entire ice topology on the blade was scanned by means of the three-dimensional scanning technique. Results showed that ice was mainly accreted on the leading edge and lower surface of the blade. The static temperature of a icing cloud has a noticeable effect on the regime of accreted ice. The ice formation was transparent glaze at −8°C and transformed to milky rime when changed the static temperature to −20°C. Comparing to rime ice, the glaze ice on the blade is more influential to the rotor performance degradation. Particular attention is given to the fact that under extreme icing conditions, over 30% degradation of transient rotor thrust was observed.

Some improvements of hybrid trim method for a helicopter rotor in forward flight

The main difficulty of the trimming for helicopter rotor CFD (Computational Fluid Dynamics) lies in the high computational cost and low computational efficiency. The BET/CFD hybrid trim method, which uses the low-fidelity BET (Blade Element Theory) to calculate the Jacobian matrix and high-fidelity CFD method to get the aerodynamic force and moment for comparison with the target values, has been widely used in industry due to its high efficiency and low computational cost compared with the pure CFD method. However, this method also has two disadvantages: a complex C81 table needs to be established before the BET works; because of the low accuracy in computing the Jacobian matrix, the number of trim iterations is too large and the trimming may not even converge. In this paper, two improvements of the hybrid trim method have been proposed. Firstly, an Artificial Neural Networks (ANN) model has been developed to predict the aerodynamic coefficients in the BET. The ANN model greatly improves the efficiency and works for all kinds of airfoils without any extra operation. The second improvement is to divide the trimming process into two stages: the preliminary stage in which the Jacobian matrix is calculated using the BET with an ANN model, and the accurate stage in which the Jacobian matrix is obtained by the CFD method. The two-stage method avoids the second shortcoming mentioned above. The AH-1G rotor and ONERA 7A rotor in forward flight have been investigated to validate the present method. The trim results agree well with the flight test results and other computational results.

Numerical investigation of dynamic stall reduction on helicopter blade section in forward flight by an airfoil deformation method

In the present work, airflow over blade section of helicopter rotor in forward flight is investigated numerically to offer a proposal for reduction or alleviation of dynamic stall on retreating blade. Unsteady 2-D compressible Reynolds-averaged Navier–Stokes equations are solved at flow conditions experienced by retreating blade of UH-60A helicopter rotor. For validation and verification purposes, flows around static and pitching airfoils in constant and variable flow Mach numbers are simulated and compared with other numerical results and experimental data. Having defined two deformation parameters, nine different permanently-deformed airfoils are produced from SSC-A09 airfoil. Flow field around these nine airfoils are simulated in variable flow Mach number with defined pitching angles. Coefficients of $$C_{{\text{l}}}$$ , $$C_{{\text{d}}}$$ and $$C_{{\text{m}}}$$ of these deformed SSC-A09 airfoils are compared with those of original airfoil to demonstrate that airfoil deformation proposed here is able to reduce and alleviate dynamic stall. Quantitatively, $$C_{{{\text{d}},\max }}$$ and absolute value of $$C_{{{\text{m}},\min }}$$ have been reduced up to 49.2% and 25%, respectively. In addition it is worth to mention that in some flow conditions dynamic stall is even eliminated with the proposed airfoil deformation. These are all achieved by permanent airfoil deformation which does not impose design complexities and expenses needed by dynamic airfoil deformation as proposed by others.

Computational fluid dynamics trimming of helicopter rotor in forward flight

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.

Numerical Research on the Unsteady Evolution Characteristics of Blade Tip Vortex for Helicopter Rotor in Forward Flight

A high-resolution numerical method is developed to simulate the generation and evolution progress of the blade tip vortex for helicopter rotor in forward flight. In the current method, the Navier–Stokes equation is adopted as the governing equation. Along with fifth order WENO scheme, the upwind Roe scheme is employed for the calculation of convective flux. A dual time-stepping scheme is utilized for time discretization, and the LU-SGS scheme is used for every pseudo time step. The Spalart–Allmaras model is selected as the turbulence model. To trim the rotor in forward flight, a high-efficiency trim method is introduced to the current solver. Results for cases in different forward flight conditions show that trimming process has a great influence on the evolution characteristics of the rotor tip vortex. Compared with the cases with different Mach numbers, the evolution characteristics of rotor tip vortex have larger difference for cases with different advance ratios.

Helicopter Rotor Boundary Layer Transition Measurement in Forward Flight Using an Infrared Camera

Boundary layer transition measurement was demonstrated using differential infrared thermography (DIT) on the top side of a helicopter rotor in forward flight, which detects the difference in the convective heat transfer at the boundary layer transition position. The tests used a FLIR X8500xc SLS long wave infrared camera to observe the DLR EC135 test helicopter rotor. The boundary layer transition was detected for hover out of ground effect (150 ft) and for forward flight at 80 kt (1700 ft). The measured boundary layer transition positions are consistent with previous measurements of the EC135 hovering in ground effect, and with predicted boundary layer transition positions. A method for the analysis of DIT images for a rotor in forward flight is shown, based on computational analysis of a pitching airfoil with varying inflow.

Optimization of differential infrared thermography for unsteady boundary layer transition measurement

Differential infrared thermography (DIT) is a method of analyzing infrared images to measure the unsteady motion of the laminar–turbulent transition of a boundary layer. It uses the subtraction of two infrared images taken with a short-time delay. DIT is a new technique which already demonstrated its validity in applications related to the unsteady aerodynamics of helicopter rotors in forward flight. The current study investigates a pitch-oscillating airfoil and proposes several optimizations of the original concept. These include the extension of DIT to steady test cases, a temperature compensation for long-term measurements, and a discussion of the proper infrared image separation distance. The current results also provide a deeper insight into the working principles of the technique. The results compare well to reference data acquired by unsteady pressure transducers, but at least for the current setup DIT results in an additional measurement-related lag for relevant pitching frequencies.

Investigations of synthetic jet control effects on helicopter rotor in forward flight based on the CFD method

ABSTRACTTo investigate the control effect of the synthetic jet on the aerodynamic characteristic of rotors, a numerical simulation procedure for the rotor flowfield is established. First, a moving-embedded grid method and an unsteady Reynolds Averaged Navier–Stokes (URANS) solver are established for predicting the complex flowfield of rotors. A velocity jet boundary condition over the jet actuator orifice is constructed, and a numerical method for simulating the active flow control on rotors is developed. Then, the effectiveness of the simulation method is validated by comparing the numerical results of jet control on NACA 0015 aerofoil with the experimental data. At last, the aerodynamic characteristic of rotors with synthetic jet actuators located on the suction surface of the blade in forward flight is calculated. The results indicate that the synthetic jet has the capability of improving the aerodynamic characteristic of rotors, especially in inhibiting the flow separation over the surface. In addition, the increase of the jet momentum coefficient and the jet angle can both enhance the lift coefficient in the retreating side. Compared with a single jet, jet arrays have better control effects on improving the aerodynamic characteristic of rotors in forward flight.

Aerodynamic prediction of helicopter rotor in forward flight using blade element theory

In this study, the helicopter blade in forward-flight condition was investigated. The blade element theory (BET) was used throughout this analysis to investigate the angle of attack variations at the blade cross sections, lift distribution along the blade and effects of increasing helicopter speed. Prouty's helicopter data was used to validate the analysis results. In this analysis, the helicopter blade was divided into 50 equally spaced elements and the azimuth ψ was set at 7.2° for each movement of the blade. The helicopter speed of 80 m/s was considered. The analysis revealed that the computation results were in good agreement with Prouty's diagram. Furthermore, it was also evident that in the case of a helicopter in forward-flight condition, the blade at retreating side was generally at low angle of attack and experienced low lift, in contrast to the blade at advancing side. The increment of the helicopter speed affected the lift distribution along the blade. The reverse flow area was widened two times from that given by the original Prouty's diagram. In addition, it was proven that each helicopter has its own speed limit called velocity never exceed (VNE). It was also shown that BET is important in conducting the analysis to modify the helicopter blade design for the aerodynamic characteristics' improvement as well as stability and general performance enhancement for the helicopter.

Time-Dependent Aeroelastic Adjoint-Based Aerodynamic Shape Optimization of Helicopter Rotors in Forward Flight

A formulation for sensitivity analysis of fully coupled time-dependent aeroelastic rotors in forward flight with trim constraints is given in this paper. The formulation includes the effect of blade shape parameters as well as blade cyclic pitch control parameters to enable analysis and optimization of rotors in a trimmed condition. Both forward sensitivity and adjoint sensitivity formulations are derived that correspond to analogs of the nonlinear aeroelastic analysis problem. Both sensitivity analysis formulations make use of the same iterative disciplinary solution techniques used for analysis, and they make use of an analogous coupling strategy. The fully coupled aeroelastic analysis formulation is first verified to effectively perform performance predictions of a four-bladed rotor in forward flight started impulsively from rest. Upon successful verification of the fully coupled adjoint formulation, it is used to perform trim and aerodynamic shape design optimization for helicopter rotors in forward-flight conditions.

Optimal Design and Acoustic Assessment of Low-Vibration Rotor Blades

An optimal procedure for the design of rotor blade that generates low vibratory hub loads in nonaxial flow conditions is presented and applied to a helicopter rotor in forward flight, a condition where vibrations and noise become severe. Blade shape and structural properties are the design parameters to be identified within a binary genetic optimization algorithm under aeroelastic stability constraint. The process exploits an aeroelastic solver that is based on a nonlinear, beam-like model, suited for the analysis of arbitrary curved-elastic-axis blades, with the introduction of a surrogate wake inflow model for the analysis of sectional aerodynamic loads. Numerical results are presented to demonstrate the capability of the proposed approach to identify low vibratory hub loads rotor blades as well as to assess the robustness of solution at off-design operating conditions. Further, the aeroacoustic assessment of the rotor configurations determined is carried out in order to examine the impact of low-vibration blade design on the emitted noise field.

A 3D mesh deformation technique for irregular in‐flight ice accretion

SummaryAircraft holding around busy airports may be requested to sustain as much as 45 min of icing before landing or being diverted to another airport. In this paper, a three‐dimensional mesh deformation scheme, based on a structural frame analogy, is proposed for the numerical simulation of ice accretion during extended exposure to adverse weather conditions. The goal is to provide an approach that is robust and efficient enough to delay or altogether avoid re‐meshing while preserving (enforcing) nearly orthogonal elements at the highly distorted ice surface. Robustness is achieved by suitably modifying the axial and torsional stiffness components of the frame elements in order to handle large and irregular grid displacements typical of in‐flight icing. Computational efficiency is obtained by applying the mesh displacement to an automatically selected small subset of the entire computational domain. The methodology is validated first in the case of deformations typical of fluid‐structure interaction problems, including wing bending, a helicopter rotor in forward flight, and the twisting of a high‐lift wing configuration. The approach is then assessed for aero‐icing on two swept wings and compared against experimental measurements where available. Copyright © 2015 John Wiley & Sons, Ltd.

Efficient Numerical Techniques for Simulating a Rotorcraft Flow Field with Overlapping Grids

We propose efficient and accurate techniques for simulating a time-periodic, unsteady rotorcraft flow. The Reynolds-averaged Navier–Stokes equations are solved on overlapping grids using a multigrid method and time spectral method. A new implementation of the multigrid algorithm for overlapping grids is proposed to accelerate hovering and forward-flight computations. Using the time-lagged periodic boundary condition, the time spectral method is combined with the multigrid scheme to significantly improve the simulation efficiency of a helicopter rotor in forward flight. The resulting equations are solved by the implicit lower-upper symmetric-Gauss–Seidel scheme. Some hovering and forward-flight cases are simulated, and the results are compared with experimental data, demonstrating the high efficiency and accuracy of the proposed method.

Aeroelastic control of non-rotating and rotating wings using the dynamic stiffness modulation principle via piezoelectric actuators

Carleton University’s Rotorcraft Research Group is working on the development of an active rotor control system that incorporates a mechanism for helicopter blade pitch dynamic stiffness modulation at the root, the Active Pitch Link. This system overcomes stroke limitations of smart material and attains superior performance for helicopter rotor-induced vibration reduction. The system was tested at the whirl tower facility and this article reports the achievements obtained with a dynamically similar hinged rotor blade model. Up to 100% reduction in the transmitted loads occurred at the target 2/rev frequency when the blade was excited by a transversal jet to mimic the asymmetric flow of the helicopter rotor in forward flight. An open-loop control algorithm optimized to a target higher-harmonic frequency of the rotor also minimized the impact on the rotor fundamental cyclic control frequency at 1/rev. In another experiment at University of São Paulo, semi-passive control techniques using shunted piezoelectric materials were investigated for the aeroelastic control of fixed wings. Flutter oscillations of a typical section were controlled out over a range of airflow speeds. Finally, the similarity between both control techniques is discussed and recognized that they are based on a dynamic stiffness modulation control principle.

前進飛行時のヘリコプタ・ロータのワール不安定現象

前進飛行時のヘリコプタ・ロータのワール不安定現象

High order accurate and low dissipation method for unsteady compressible viscous flow computation on helicopter rotor in forward flight

An improved fifth-order weighted essentially non-oscillatory (WENO-Z) scheme combined with the moving overset grid technique has been developed to compute unsteady compressible viscous flows on the...

High order accurate and low dissipation method for unsteady compressible viscous flow computation on helicopter rotor in forward flight

An improved fifth-order weighted essentially non-oscillatory (WENO-Z) scheme combined with the moving overset grid technique has been developed to compute unsteady compressible viscous flows on the helicopter rotor in forward flight. In order to enforce periodic rotation and pitching of the rotor and relative motion between rotor blades, the moving overset grid technique is extended, where a special judgement standard is presented near the odd surface of the blade grid during search donor cells by using the Inverse Map method. The WENO-Z scheme is adopted for reconstructing left and right state values with the Roe Riemann solver updating the inviscid fluxes and compared with the monotone upwind scheme for scalar conservation laws (MUSCL) and the classical WENO scheme. Since the WENO schemes require a six point stencil to build the fifth-order flux, the method of three layers of fringes for hole boundaries and artificial external boundaries is proposed to carry out flow information exchange between chimera grids. The time advance on the unsteady solution is performed by the full implicit dual time stepping method with Newton type LU-SGS subiteration, where the solutions of pseudo steady computation are as the initial fields of the unsteady flow computation. Numerical results on non-variable pitch rotor and periodic variable pitch rotor in forward flight reveal that the approach can effectively capture vortex wake with low dissipation and reach periodic solutions very soon.

Advancement of aerofoil section dynamic stall synthesis methods for rotor design

AbstractDynamic stall is a complex process encountered by an aerofoil in the unsteady flow environment such as a helicopter rotor in forward flight as well as a fixed wing aircraft in manoeuvres and in other unsteady situations. The onset of dynamic stall effectively determines the flight envelope of the helicopter. Significant effort are being made to develop CFD to capture the dynamic stall behaviour, however traditional engineering models based on lifting line theory still offer fast turn-round and broad understanding required for the rotor design process. This paper describes a new engineering model for dynamic stall, developed originally for wind turbine application at a typical Mach number of 0·12. The new dynamic stall model, with a better definition of stall onset, is based on improvements made to Beddoes’ original trailing edge stall model. This paper will describe and demonstrate the improvements in identifying both the stall-onset and the pitching moment break at high pitch rates, when being applied to a generic rotor aerofoil RAE 9651 at M = 0·3. Further validation against oscillatory tests and other Mach numbers are still required. However the study has provided sufficient confidence for it to be employed in a rotor analysis code.

Multiblade Reduced-Order Aerodynamics for State-Space Aeroelastic Modeling of Rotors

Reduced-order aerodynamic models are tools that may be conveniently applied in a wide range of research and design applications in the aeronautical and mechanical fields. This paper presents a methodology for the identification of a reduced-order model (ROM) describing the linearized unsteady aerodynamics of helicopter rotors in arbitrary steady flight, which is particularly suited for the derivation of the state-space perturbation aeroelastic operators and is hence useful for stability analysis and aeroservoelastic applications. It is defined in terms of multiblade coordinates and yields a (finite state) constant-coefficient, linear, differential form relating them to the corresponding multiblade aerodynamic loads. This approach requires the prediction of a set of harmonic perturbation responses by an aerodynamic solver. The accuracy of the identified ROM in describing unsteady aerodynamics phenomena is strictly connected to that of the aerodynamic solver. Complex aerodynamic effects (like wake roll-up and wake―blade interactions) are included in the ROM if they are taken into account in evaluating the harmonic responses. Numerical results concerning a flap-lag helicopter rotor in forward flight are presented. These examine the accuracy the aerodynamic ROM introduced both in terms of aerodynamic loads predictions and in terms of aeroelastic stability analysis. An aeroservoelastic application is also included in order to demonstrate the suitability of the ROM proposed for the design of controllers.

Cramer–Rao Bound Development for Linear Time Periodic Systems

System identification techniques are often used to determine the parameters required to define a model of a linear time invariant (LTI) system. The Cramer–Rao bound can be used to validate those parameters in order to ensure that the system model is an accurate representation of the system. Unfortunately, the Cramer–Rao bound is only valid for LTI systems and is not valid for linear time periodic (LTP) systems such as a helicopter rotor in forward flight. This paper describes an extension of the Cramer–Rao bound to LTP systems and demonstrates the methodology for a simple LTP system.