Rotor In Hover Research Articles

Study on the effect of transition process on rotor hovering simulation

Hovering is one of important statuses to evaluate the aerodynamic performance of a rotor. With the development of the computer technology and CFD technique, the numerical methods based on the first principle are usually employed to evaluate the hovering performance of the rotor. The transition process will evidently affect the results from the RANS-based numerical simulations in some steady cases for the fixed wing aircrafts, which should be taken into consideration in the design process. But it's not clear whether the transition process would affect the numerical results for the rotor simulation. To provide the reference in designing and evaluating the rotorcraft, the effect of the transition process in the rotor simulation needs to be discussed further. The PSP rotor proposed by NASA is calculated using the in-house solver based on the overset grid in this paper. Simulations are performed with fully turbulent model as well as the transitional model and the results are compared to the experimental data. The results prove the superior ability to simulate the flow around a hovering rotor of the in-house solver. The relative errors of the numerical results are under 5%. The range of the laminar flow on the blade is proportional to the rotor thrust, which causes a higher Figure of Merit in transition simulation than the fully turbulent simulation. The sectional pressure distribution and torque distribution along the blade apparently suffer from the transition process, which doesn't affect the thrust distribution along the blade and the blade vortex wake flow under the rotor disk. An obvious flow separation on the surface of the blade can be observed in the transition simulation compared to the fully turbulent simulation.

Flight performance of helicopter tail rotors with extendable chord

To investigate the potential of Statically and Dynamically Extendable Chord (SEC/DEC) in performance improvement of helicopter tail rotors, a prediction model of helicopter flight performance is developed and validated. It is suitable to deploy the SEC and DEC close to the blade tip for power reduction of tail rotor in hover and high-speed flight. The deployment locations of SEC and DEC for tail rotors are similar to that of main rotors. In hover, the deployment of the SEC and DEC results in little power reduction. From low to medium speed flight, the deployment of the DEC and SEC increases the power. In high-speed flight, the SEC and DEC can significantly decrease the power, especially at a large take-off weight. At 300 km/h, when the chord extension is 20% of blade chord, the SEC and DEC with 1/rev input decrease the power by 14.44% and 20.47% at a take-off weight coefficient of 0.0091. The DEC shows greater potential in power reduction than the SEC. There is an optimal phase input for the maximum power reduction, which does not vary with forward speed. For the DEC with 1/rev input, the optimal phase is [Formula: see text]. The optimal phase input for the maximum power reduction of tail rotors is different from that of main rotors.

Investigation on Aerodynamic Noise Characteristics of Coaxial Rotor in Hover

A numerical method based on Reynolds Averaged Navier–Stokes (RANS) equations and a high-efficiency trim model is developed to simulate the aerodynamics of a coaxial rotor. Farassat 1A equations are used for the prediction of thickness and loading noise. Hover cases of different thrust coefficients with torque balance are conducted. The sound pressure history of different observation points positioned below the rotor disk plane is analyzed. Results indicate that the special noise characteristics of the coaxial rotor are mainly caused by the noise superposition of the twin rotors and the unsteady loads of aerodynamic interaction. A new kind of impulsive loading noise is induced by the blade-meeting interaction. In contract to the single rotor, the loading noise of a coaxial rotor has a much larger sound pressure level in the high-frequency band. The loading noise is obviously enhanced around the blade-meeting azimuths. The maximum noise of the coaxial rotor is located immediately below the rotor disk center, while for the single rotor, it is the minimum location.

An Experimental Apparatus for Icing Tests of Low Altitude Hovering Drones

The icing facilities of the Anti-Icing Materials International Laboratory AMIL have been adapted to reproduce icing conditions on a Bell APT70 drone rotor, typical of small-to-medium UAV models. As part of an extensive icing test campaign, this paper presents the design and preliminary testing of the experimental setup and representative icing conditions calibration in the laboratory’s cold chamber. The drone rotor used has four blades with a diameter of 0.66 m and a maximum tip speed of 208 m/s. For the icing conditions, freezing rain and freezing drizzle were selected. A Liquid Water Content (LWC) calculation methodology for a rotor in hover was developed, and procedures to determine experimental LWC in the facility are presented in this paper. For the test setup, the cold chamber test section was adapted to fit the rotor and to control its ground clearance. Testing was aimed at studying the effect of rotor height h on aerodynamic performance, both with and without icing conditions. Results show no significant effect on the ground effect between h = 2 m and h = 4 m in dry runs, while the icing behavior can be largely influenced for certain conditions by the proximity of the precipitation source, which depend on the height of the rotor in these experiments.

Aeroacoustic modeling of helicopter transonic rotor noise

The noise propagation of an helicopter rotor with blade tip in transonic regime is numerically simulated. To calculate the acoustic characteristics of the helicopter rotor, a novel CFD/Penetrable Surface Ffowcs Williams-Hawkings (PS-FWH) hybrid approach is established. The approach is characterized by the solution of the Reynolds-Averaged Navier-Stokes (RANS) equations on a Chimera system of overset meshes, coupled with an in-house acoustic code (ROCAAP), which employs the PS-FWH method with a Marching-Cube Emission-Surface (MCES) integral algorithm, for the noise propagation. The hybrid approach is validated through two databases. The well-known Caradonna-Tung hovering rotor test is used to assess the pressure distribution at each cross-section along the blade. The transonic acoustic survey of the UH-1H rotor is adopted to evaluate the time domain noise signal with different blade tip Mach numbers during hover. The capabilities of the hybrid method to predict helicopter transonic rotor noise are demonstrated through the achieved good correlation with experimental data.

Numerical analysis of rotor aeroacoustic characteristics during collective pitch aperiodic variation in hover

Based on the FW-H equations and the CFD method, the rotor aeroacoustic characteristics during collective pitch aperiodic variation in hover are calculated and analyzed. First, a set of analysis method for aperiodic rotor aeroacoustic characteristics is developed. The aerodynamic and aeroacoustic characteristics of the BO-105 rotor in hover, the AH-1G rotor in forward flight and the NACA rotor in a ramp increase of collective pitch are calculated, and the employed numerical analysis method is validated through comparisons with experimental data. Then, the aeroacoustic characteristics of the NACA rotor during an oscillating collective pitch are analyzed, and the sound pressure peak at different collective pitch is discussed in detail. Finally, the BO-105 rotor in hover during collective pitch aperiodic variation is calculated. In addition, parameters, such as the collective pitch change rate, are quantified, and some conclusions are obtained. The configuration of collective pitch change rate will influence the directivity of rotor noise in the collective pitch increasing process, and the loading noise will be smaller if the change rate is decreasing.

A Modified Fifth-Order WENO-HLLC Riemann Solver for Transonic Viscous Flows around Helicopter Rotor in Hover

This paper presents a modified fifth-order WENO-HLLC Riemann solver for the transonic compressible viscous flows around the helicopter rotor in hover. The HLLC approximate Riemann solver is proposed to discrete the convection term involving grid velocity of the Navier-Stokes equations. In order to solve the interface flow accurately, a modified fifth-order WENO scheme is presented by designing the smoothness indicators based on L 1 norm measurement. The improved WENO scheme can provide the optimal approximation order even at critical points. Numerical accuracy and robustness are validated by several benchmark inviscid flow problems. Then the numerical properties of the WENO-HLLC solver in conjunction with the implicit LU-SGS time integration method with high efficiency are further validated by simulating transonic viscous flows over RAE2822 airfoil and ONERA-M6 wing. The results show that the accuracy of calculating shock, discontinuity, and the vortex is significantly improved. Finally, the method is developed to compute the transonic vortex flow around the helicopter rotor with a domain discretized by overset grids. The results indicate that the proposed method is very robust and effective in acquiring high resolution for vortex wake.

Effect of Turbulence Modeling Selection Within Helios for Small Quadrotor Aerodynamics

The high-fidelity rotorcraft simulation framework HPCMP CREATE-AV Helios was used to study the impact of turbulence modeling methods on the prediction of a small quadrotor unmanned aerial vehicle. First, the vehicle’s rotor was simulated as an isolated rotor in both hover and forward flight for turbulence model assessment. Predictions using a fully turbulent turbulence model were compared with predictions from several laminar–turbulent transition models. Both comparison with experiments and computational performance were considered. The laminar–turbulent transition models predicted flow separation near the tip for all flight conditions simulated, leading to lower torque and thrust than the fully turbulent model, which predicted attached turbulent flow. The fully turbulent model was shown to provide reasonable comparison with experiments with the lowest computational cost and was used for full vehicle simulations. The full quadrotor configuration calculations were performed in hover and forward flight. The forward flight simulations were performed with and without a detached eddy simulation method, and similar interactional aerodynamics were predicted. The Helios simulation suite is shown to be well-suited for prediction of aerodynamic performance of small-scale rotorcraft.

A novel high-order scheme for numerical simulation of wake flow over helicopter rotors in hover

Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor. A new high-order scheme (we call it WENO-K) proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors. WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction, and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities. Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid. To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid, a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid. Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation. Numerical results for Caradonna-Tung (C-T) rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%. Compared with WENO-JS scheme, WENO-K scheme achieves about 40% improvement on accuracy of predicting rotor thrust with only 4.1% extra computational cost. More importantly, WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270° compared to WENO-JS scheme.

Experimental and Computational Aeroacoustic Investigation of Small Rotor Interactions in Hover

This paper investigates the aeroacoustic interactions of small hovering rotors, using both experiments and computations. The experiments were conducted in an anechoic chamber with arrays of microphones setup to evaluate the azimuthal and polar directivity. The computational methodology consists of high fidelity detached eddy simulations coupled to the Ffowcs-Williams and Hawkings equation, supplemented by a trailing edge broadband noise code. The aerodynamics and aeroacoustics of a single rotor are investigated first. The simulations capture a Reynolds number effect seen in the performance parameters that results in the coefficient of thrust changing with the RPM. The acoustic analysis enables the identification of self-induced noise sources. Next, dual side-by-side rotors are studied in both counter-rotating and co-rotating configurations to quantify the impact of their interactions. Higher harmonics appear due to the interactions and it is verified that the counter-rotating case leads to more noise and a less uniform azimuthal directivity. Difficulties that arise when trying to validate small rotor calculations against experiments are discussed. Comparisons of computational and experimental results yield further insight into the noise mechanisms that are captured by each methodology.

Effect of Crossflow Transition on the Pressure-Sensitive-Paint Rotor in Hover

The amplification factor transport transition model has been extended to include crossflow transition and applied to a rotor in hover. Validation of the crossflow model implementation has been performed using standard test cases. The pressure-sensitive-paint rotor is simulated in hover for various collective pitch angles as a test case for the transition model. The OVERFLOW 2.2n solver is used with laminar–turbulent transition modeling and hybrid Reynolds–averaged Navier–Stokes/large-eddy simulation turbulence modeling enabled. Grid generation, numerical methods, and transition modeling are described. The effects of crossflow transition are qualitatively and quantitatively assessed through comparisons to experiments, to a transition model without crossflow, and to a fully turbulent model. Additionally, several techniques for evaluating unsteady boundary-layer transition are described and are applied to solutions for the pressure-sensitive-paint rotor in hover.

Aeroacoustic Characteristic Analyses of Coaxial Rotors in Hover and Forward Flight

Aeroacoustic Characteristic Analyses of Coaxial Rotors in Hover and Forward Flight

Using blade element momentum methods with gradient-based design optimization

Blade element momentum methods are widely used for initial aerodynamic analysis of propellers and wind turbines. A wide variety of correction methods exist, but common to all variations, a pair of residuals are converged to ensure compatibility between the two theories. This paper shows how to rearrange the sequence of calculations reducing to a single residual. This yields the significant advantage that convergence can be guaranteed and to machine precision. Both of these considerations are particularly important for gradient-based optimization where a wide variety of atypical inputs may be explored, and where tight convergence is necessary for accurate derivative computation. On a moderate-sized example optimization problem we show over an order of magnitude increase in optimization speed, with no changes to the physics. This is done by using the single residual form, providing numerically exact gradients using algorithmic differentiation with an adjoint, and by leveraging sparsity in the Jacobian using graph coloring techniques. Finally, we demonstrate a revised formulation for cases when no inflow exists in one of the directions (e.g., a hovering rotor or a parked rotor). These new residuals allow for robust convergence in optimization applications, avoiding the occasional numerical difficulties that exist with the standard formulation.

Numerical investigation of effect of crossflow transition on rotor blade performance in hover

In the present study, the effect of crossflow transition on the rotor blade performance in hover was investigated numerically. The simulations were conducted for the Pressure Sensitive Paint (PSP) rotor in hover by using a Reynolds-averaged Navier-Stokes (RANS) solver based on unstructured meshes. The k−ωSST model was used for simulating the turbulent flow fields and for calculating the turbulent eddy viscosity. For predicting the laminar-turbulent onset phenomena involving crossflow-induced transition, the γ−Reθt−CF+ transition model was adopted. For the comparison of the transition locations without considering the crossflow transition effect, simulations were also carried out using the γ−Reθt transition model. The calculations were made for the PSP rotor at collective pitch angles from four to 12 degrees with an interval of two degrees. To investigate the effect of blade tip Mach number, two blade tip Mach numbers of 0.585 and 0.65 were tested. The predicted results such as the transition onset location and the rotor aerodynamic performance in terms of thrust coefficient, torque coefficient and figure of merit were compared with experimental data. It was found that proper consideration of the effect of crossflow transition is critical for the accurate prediction of the laminar-turbulent transition onset location on rotor blades. The figure of merit also compares better with the experimental data when the effect of laminar-turbulent transition is included. With the addition of the crossflow transition, the figure of merit is slightly decreased, particularly at low thrust levels.

Investigation on the aerodynamic interaction between a rotor and ship with opened hangar door

Unsteady aerodynamic interference between a rotorcraft and a ship occurs during shipboard launch and recovery operations and has a negative impact on the safety. An experiment of a reduced-scale model rotor and a CFD analysis in hover was carried out to investigate the performance and flow field of a rotor approaching a ship. In this paper, the thrust and pitching moment of the rotor hovering above the ground, deck were tested, and the influence of hangar door on the thrust, pitching moment, and flow field was also measured. A CFD method based on RANS and overset technology was used to investigate the flow field of the rotor operating on the model-scale ship. As the rotor approaches the deck, its thrust first decreases induced by a recirculation near the deck, and then increases induced by the effect of deck, and finally obviously decreases caused by the recirculation near hangar door. The deck and hangar door also affects the flow field to yield an intensive nose-down pitching moment. The status of the hangar door has a significant influence on the rotor thrust and pitching moment. The recirculation is weakened with an opened hangar door resulting in recovery of the rotor thrust and decrease of the nosed down pitching moment.

Experimental investigation of tip anhedral effects on the aerodynamics of a model helicopter rotor in hover

This study experimentally investigates the effects of tip anhedral on the rotor aerodynamic performance and the tip vortex characteristics in hovering flight. A five-bladed scaled helicopter rotor with blades that have either rectangular (baseline) or anhedral tip geometries was used as the experimental model. Thrust and torque measurements were performed at the tip Mach numbers (Mtip) of 0.3 and 0.4 at five different pitch angles. In addition, flow field measurements via phase-locked particle image velocimetry technique were conducted at Mtip = 0.4 for a selected pitch angle. Thrust and torque measurements reveal that the anhedral tip configuration reaches a maximum Figure of Merit (FOM) value of 0.67 yielding approximately 5% FOM improvement when compared to the rectangular blade at Mtip = 0.3. It is observed that the tip anhedral changes the tip vortex trajectory by increasing the radial contraction of the wake and decreasing the average axial convection speed in the range of the wake age considered in this study. The meandering amplitude of the tip vortex increases abruptly after the passage of the proceeding blade in both cases yet a greater increase is observed in the anhedral tip configuration. Moreover, the axial convection of the previously shed tip vortices in the downstream direction is significantly slowed down during the course of the blade passage such that they reside close to the rotor disc in the anhedral case. The resultant proximity of the tip vortices to the blade with anhedral tip configuration promotes the induced flow in the radially inwards direction, which probably attenuates the strength of the newly forming tip vortex. Accordingly, the maximum tangential velocity and circulation of the tip vortex are reduced by 20% and 13%, respectively, when compared to the rectangular blade case. The reduced tip-vortex circulation is consistent with the increased FOM value. A good agreement was found with the Vatistas vortex model using n=3 for the rectangular blade and n=2.2 for the blade with tip anhedral.

High-Fidelity Aeromechanical Analysis of Coaxial Mars Helicopter

A high-fidelity coupled computational fluid dynamics and comprehensive analysis solver is developed for helicopter flight on Mars. The objectives are accurate prediction of flight loads and fundamental understanding of Martian aeromechanics. Performance, structural loads, pitch link loads, rotor wake, and interrotor blade separation are studied for hingeless and flap articulated coaxial rotors. In the absence of any validation data, pressure chamber tests are carried out at Mars-like conditions on a single rotor in hover. The data provided an elementary assessment of rotor behavior and a baseline for model validation. Predictions of the high-fidelity solver are also compared with a lower-fidelity lifting-line comprehensive analysis. It is generally concluded that lower-fidelity tools are not adequate for capturing the complex flow phenomena encountered on Mars. A motion and loads comparison of two fundamental hub types revealed that 1) an articulated hub has similar interrotor blade separation as hingeless on Mars, 2) a hingeless hub experiences only marginally higher (6–7%) flap bending moments, and 3) an articulated hub can lower steady pitch link loads (by an order of magnitude) at a marginal expense of oscillatory loads, 15% higher.

Actuator Surface Model with Computational-Fluid-Dynamics-Convected Wake Model for Rotorcraft Applications

An actuator surface model has been developed that couples an aerodynamics model to a computational fluid dynamics solver. The model is designed for simulation of rotors in complex operating environments, and it advances a recently developed actuator surface model through the addition of two new elements: a computational fluid dynamics (CFD)-convected wake model, and an improved coupling algorithm. The wake model uses Lagrangian particle tracking to determine the geometry of the rotor wake from the CFD solution. Prior study of the influence of the maximum wake age in the model led to the conclusion that only 45 deg of wake is required for accurate resolution of air loads. The S-76 rotor in hover and the UH-60A rotor in forward flight are used for validation. Thrust coefficients, torque coefficients, and figures of merit over a range of collective settings agree well with experimental data for the S-76 rotor. Sectional normal force coefficients in forward flight are compared with experimental data, results from resolved-blade simulations, and other rotor models: again showing good agreement. These results suggest the new actuator surface model can predict time-accurate rotor loads where accurate resolution of the near-blade flowfield and high-frequency loads, such as blade–vortex interactions, are not critical.

Investigation of Bioinspired Flow for a Nano Coaxial Rotor in Hover

AbstractA computational study of a nano coaxial rotor with bioinspired blade motion was conducted at an ultralow Reynolds number to determine the principle of improving rotor performance via an ind...

Comparative Study of Coaxial Main Rotor Aerodynamics in the Hover with the Usage of Two Methods of Computational Fluid Dynamics

The paper is focused on numerical modeling of the aerodynamic characteristics of a full-scale coaxial main rotor in hover. The simulation was performed using two approaches of computational fluid dynamics (CFD): the original free wake model (FWM) developed by the authors and the unsteady Reynolds-averaged Navier–Stokes (URANS) equations method based on the Ansys Fluent software. The structure of the rotor vortex wake, flow images, vorticity and induced velocity fields, total and distributed aerodynamic characteristics of the rotor, including the rotor performance and figure of merit diagrams, have been analyzed. A comparison between FWM/URANS based calculations and calculated/experimental data by other authors has been performed. A satisfactory match of these data confirms reliability of used methods when modeling the aerodynamic characteristics of coaxial rotors. The FMW demonstrates a significant advantage in speed and resources intensity when calculating the total aerodynamic characteristics of a coaxial rotor. The URANS method makes it possible to model with significant accuracy the effects associated with the blade vortex interactions and pressure distribution over the blade surface. Finally, conclusions about most effectiveness of joint application of considered methods in solving complex problems of coaxial rotor aerodynamics has been made.