CFD Analysis of a Slatted UH-60 Rotor in Hover

The effect of leading-edge slats (LE) on the performance of a UH-60A rotor in hover was studied using the OverTURNS CFD solver. The objective of the study was to quantify the effect of LE slats on the hover stall boundary and analyze the reasons for any potential improvement/penalty. CFD predictions of 2-D slatted airfoil aerodynamics were validated against available wind tunnel measurements for steady angle of attack variations. The 3-D CFD framework was validated by comparing predictions for the baseline UH-60A rotor against available experimental values. Subsequent computations were performed on a slatted UH-60A rotor blade with a 40\%-span slatted airfoil section and two different slat configurations. The effect of the slat root and tip vortices as they convect over the main blade element was captured using appropriately refined main element meshes and their impact on the slatted rotor performance was analyzed. It was found that the accurate capture of the slat root and tip vortices using the refined meshes made a significant difference to the performance predictions for the slatted rotors. The calculations were performed over a range of thrust values and it was observed that the slatted rotor incurred a slight performance penalty at lower thrust and was comparable to the baseline rotor at higher thrust conditions. It was also found that shock induced separation near the blade tip was the limiting factor for the baseline UH-60 rotor in hover, causing an increase in rotor power and resulting in a reduction of figure of merit for the baseline rotor at higher thrust values. The shock induced separation occured outboard of the slat tip and therefore limited the performance of the slatted rotors as well. Overall, the study provides useful insights into effects of leading edge slats on rotor hover performance, aerodynamics and wake behavior

Some recent applications of Navier-Stokes codes to rotorcraft

Many operational limitations of helicopters and other rotary-wing aircraft are due to nonlinear aerodynamic phenomena incuding unsteady, three-dimensional transonic and separated flow near the surfaces and highly vortical flow in the wakes of rotating blades. Modern computational fluid dynamics (CFD) technology offers new tools to study and simulate these complex flows. However, existing Euler and Navier-Stokes codes have to be modified significantly for rotorcraft applications, and the enormous computational requirements presently limit their use in routine design applications. Nevertheless, the Euler/Navier-Stokes technology is progressing in anticipation of future supercomputers that will enable meaningful calculations to be made for complete rotorcraft configurations

Empirical Evaluation of Ground, Ceiling, and Wall Effect for Small-Scale Rotorcraft

Ground effect refers to the apparent increase in lift that an aircraft experiences when it flies close to the ground. For helicopters, this effect has been modeled since the 1950\u27s based on the work of Cheeseman and Bennett, perhaps the most common method for predicting hover performance due to ground effect. This model, however, is based on assumptions that are often not realistic for small-scale rotorcraft because it was developed specifically for conventional helicopters. It is clear that the Cheeseman-Bennett model cannot be applied to today\u27s multirotor UAVs. Experimental findings suggest that some of the conventional thinking surrounding helicopters cannot be applied directly to rotorcraft using fixed propellers at variable speeds (e.g. multirotors). A parametric multirotor-specific ground effect model is developed and presented to overcome some of the limitations in classical helicopter theory. Likewise, ceiling effect refers to the apparent increase in lift that a rotorcraft experiences when flying close to a ceiling or any similar surface that is present above the rotor(s). Ceiling effect is similar in principle to ground effect, and can be explained using a similar equation. Ceiling effect, however, was never explored in detail for conventional helicopters because large manned aircraft do not operate in enclosed spaces. For multirotors, the work presented in this dissertation suggests that the classical helicopter theory adequately describes ceiling effect performance. Wall effect is the phenomena that occurs when a rotorcraft flies near a vertical wall, and has the tendency to pitch towards the wall and be drawn into it. Wall effect is the least-understood of these three areas of interest. Wall effect has not been explored in great detail for any aircraft, and is addressed in detail in this dissertation. The recent widespread use of small-scale UAVs and the demand for increased autonomy when flying in enclosed environments has created a need for detailed studies of ground effect, ceiling effect and wall effect. Ultimately, this work provides foundations for the development of an improved UAV flight controller that can accurately account for various aerodynamic disturbances that occur near surfaces and structures to improve flight stability

Empirical Evaluation of Ground, Ceiling, and Wall Effect for Small-Scale Rotorcraft

Ground effect refers to the apparent increase in lift that an aircraft experiences when it flies close to the ground. For helicopters, this effect has been modeled since the 1950\u27s based on the work of Cheeseman and Bennett, perhaps the most common method for predicting hover performance due to ground effect. This model, however, is based on assumptions that are often not realistic for small-scale rotorcraft because it was developed specifically for conventional helicopters. It is clear that the Cheeseman-Bennett model cannot be applied to today\u27s multirotor UAVs. Experimental findings suggest that some of the conventional thinking surrounding helicopters cannot be applied directly to rotorcraft using fixed propellers at variable speeds (e.g. multirotors). A parametric multirotor-specific ground effect model is developed and presented to overcome some of the limitations in classical helicopter theory. Likewise, ceiling effect refers to the apparent increase in lift that a rotorcraft experiences when flying close to a ceiling or any similar surface that is present above the rotor(s). Ceiling effect is similar in principle to ground effect, and can be explained using a similar equation. Ceiling effect, however, was never explored in detail for conventional helicopters because large manned aircraft do not operate in enclosed spaces. For multirotors, the work presented in this dissertation suggests that the classical helicopter theory adequately describes ceiling effect performance. Wall effect is the phenomena that occurs when a rotorcraft flies near a vertical wall, and has the tendency to pitch towards the wall and be drawn into it. Wall effect is the least-understood of these three areas of interest. Wall effect has not been explored in great detail for any aircraft, and is addressed in detail in this dissertation. The recent widespread use of small-scale UAVs and the demand for increased autonomy when flying in enclosed environments has created a need for detailed studies of ground effect, ceiling effect and wall effect. Ultimately, this work provides foundations for the development of an improved UAV flight controller that can accurately account for various aerodynamic disturbances that occur near surfaces and structures to improve flight stability

The flow physics of helicopter brownout

The formation of the dust cloud that is associated with low-level helicopter operations in desert environments has been simulated using the Vorticity Transport Model together with a coupled model to represent the entrainment and subsequent transport of particulate matter through the flow. A simple thin-layer theory, supported by simulations performed using the more physically-representative numerical model, is used to explain the formation of characteristic sheet- and filament-like structures in the dust cloud in terms of the interactions between individual vortical filaments and the ground. In parts of the flow, for instance near the ground vortex that is formed under the leading edge of the rotor when in forward flight, the dust cloud becomes more space-filling than sheet-like in character, and the theory suggests that this is a result of the dust distribution having been processed by multiple vortices over a significant period of time. The distribution of the regions on the ground plane from which significant entrainment of dust into the flow takes place is shown to be influenced strongly by the unstable nature of the vortical structures within the flow. It is suggested that the effect of this vortical instability, when integrated over the timescales that are characteristic of the formation of the dust cloud, is to de-sensitize the gross characteristics of the dustcloud to the details of the wake structure at its inception on the rotor blades. This suggests that the formation of the brownout cloud may be relatively insensitive to the detailed design of the blades of the rotors and may thus be influenced only by less subtle characteristics of the helicopter system

Prediction & Active Control of Multi-Rotor Noise

Significant developments have been made in designing and implementation of Advanced Air Mobility Vehicles (AAMV). However, wider applications in urban areas require addressing several challenges, such as safety and quietness. These vehicles differ from conventional helicopter in that they operate at a relatively lower Reynolds number. More chiefly, they operate with multiples of rotors, which may pose some issues aerodynamically, as well as acoustically. The aim of this research is to first investigate the various noise sources in multi-rotor systems. High-fidelity simulations of two in-line counter-rotating propellers in hover, and in forward flight conditions are performed. Near field flow and acoustic properties were resolved using Hybrid LES-Unsteady RANS approach. Far-field sound predictions were performed using Ffowcs-Williams-Hawkings formulation. The two-propeller results in hovering are compared with that of the single propeller. This enabled us to identify the aerodynamic changes resulting from the proximity of the two propellers to each other and to understand the mechanisms causing the changes in the radiated sound. It was discovered that there is a dip in the thrust due to the relative proximity of the rotors. Owing to this, there is also some acoustic banding above the rotors mainly because they operate at the same rotational rate. We then considered the forward flight case and compared it with the corresponding hovering case. This enabled us to identify the aerodynamic changes resulting from the incoming stream. By examining the near acoustic field, the far-field spectra, the Spectral Proper Orthogonal Decomposition, and by conducting periodic averaging, we were able to identify the sources of the changes to the observed tonal and broadband noise. The convection of the oncoming flow was seen to partially explain the observed enhancement in the tonal and BBN, compared to the hovering case. It is well known that High fidelity methods are critical in predicting the full spectrum of rotor acoustics. However, these methods can be prohibitively expensive. We present here an investigation of the feasibility of reduction methods such as Proper Orthogonal Decomposition as well as Dynamic Mode decomposition for reduction of data obtained via Hybrid Large-Eddy – Unsteady Reynolds Averaged Navier Stokes approach (HLES) to be used further to obtain additional parameters. Specifically, we investigate how accurate reduced models of the high-fidelity computations can be used to predict the far-field noise. It was found that POD was capable of reconstructing accurately the parameters of interest with 15-40% of the total mode energies, whereas the DMD could only reconstruct primitive parameters such as velocity and pressure loosely. A rank truncation convergence criterion \u3e 99.8% was needed for better performance of the DMD algorithm. In the far-field spectra, DMD could only predict the tonal contents in the lower- mid frequencies whiles the POD could reproduce all frequencies of interest. Lastly, we develop an active rotor noise control technology to reduce the in-plane thickness noise associated with multi-rotor Advanced Air Mobility Vehicles (AAMV). An actuation signal is determined via the Ffowcs-Williams-Hawking (FWH) formula. Two in-line rotors are considered and we showed that the FWH-determined actuation signal can produce perfect cancellation at a point target. However, the practical need is to achieve noise reduction over an azimuthal zone, not just a single point. To achieve this zonal noise reduction, an optimization technique is developed to determine the required actuation signal produced by the on-blade distribution of embedded actuators on the two rotors. For the specific geometry considered here, this produced about 9 dB reduction in the in-plane thickness noise during forward flight of the two rotors. We further developed a technology that replaces using a point actuator on each bladed by distributed micro actuators system to achieve the same noise reduction goal with significantly reduced loading amplitudes per actuator. Overall, this research deepens the knowledge base of multi-rotor interaction. We utilize several techniques for extracting various flow and acoustic features that help understand the dynamics of such systems. Additionally, we provide a more practical approach to active rotor noise control without a performance penalty to the rotor system

Modeling and Simulation of Coaxial Helicopter Rotor Aerodynamics

A framework is developed for the computational fluid dynamics (CFD) analyses of a series of helicopter rotor flowfields in hover and in forward flight. The methodology is based on the unsteady solutions of the three-dimensional, compressible Navier-Stokes equations recast in a rotating frame of reference. The simulations are carried out by solving the developed mathematical model on hybrid meshes that aim to optimally exploit the benefits of both the structured and the unstructured grids around complex configurations. The computer code is prepared for parallel processing with distributed memory utilization in order to significantly reduce the computational time and the memory requirements. The developed model and the simulation methodology are validated for single-rotor-in-hover flowfields by comparing the present results with the published experimental data. The predictive merit of different turbulence models for complex helicopter aerodynamics are tested extensively. All but the κ-ω and LES results demonstrate acceptable agreement with the experimental data. It was deemed best to use the one-equation Spalart-Allmaras turbulence model for the subsequent rotor flowfield computations. First, the flowfield around a single rotor in forward flight is simulated. These time—accurate computations help to analyze an adverse effect of increasing the forward flight speed. A dissymmetry of the lift on the advancing and the retreating blades is observed for six different advance ratios. Since the coaxial rotor is proposed to mitigate the dissymmetry, it is selected as the next logical step of the present investigation. The time—accurate simulations are successfully obtained for the flowfields generated by first a hovering then a forward-flying coaxial rotor. The results for the coaxial rotor in forward flight verify the aerodynamic balance proposed by the previously published advancing blade concept. The final set of analyses aims to investigate if the gap between the two rotors of the coaxial configuration has any significant effect on the generated forces. The present results indicate either little or no such effect on the lift

Method for calculating rotors with active gurney flaps

This paper builds on the Helicopter Multi-Block version 2 computational-fluid-dynamics solver of the University of Liverpool and demonstrates the implementation and use of Gurney flaps on wings and rotors. The idea is to flag any cell face within the computational mesh with a solid, no-slip boundary condition. Hence, the infinitely thin Gurney can be approximated by “blocking cells” in the mesh. Comparison between thick Gurney flaps and infinitely thin Gurneys showed no difference on the integrated loads; the same flow structure was captured and the same vortices were identified ahead and behind the Gurney. The results presented for various test cases suggest that the method is simple and efficient, and it can therefore be used for routine analysis of rotors with Gurney flaps. Moreover, the current method adds to the flexibility of the solver because no special grids are required, and Gurney flaps can be easily implemented. Simple two-dimensional aerofoil, three-dimensional wing, and rotors in hover and forward flight were tested with fixed, linearly actuated, and swinging Gurneys, and the ability of the code to deploy a Gurney flap within the multiblock mesh is highlighted. The need for experimental data suitable for validation of computational-fluid-dynamics methods for cases of rotors with Gurney flaps is also highlighted

A CFD study of tilt rotor flowfields

The download on the wing produced by the rotor wake of a tilt rotor vehicle in hover is of major concern because of its severe impact on payload-carrying capability. In a concerted effort to understand the fundamental fluid dynamics that cause this download, and to help find ways to reduce it, computational fluid dynamics (CFD) is employed to study this problem. The thin-layer Navier-Stokes equations are used to describe the flow, and an implicit, finite difference numerical algorithm is the method of solution. The methodology is developed to analyze the tilt rotor flowfield. Included are discussions of computations of an airfoil and wing in freestream flows at -90 degrees, a rotor alone, and wing/rotor interaction in two and three dimensions. Preliminary results demonstrate the feasibility and great potential of the present approach. Recommendations are made for both near-term and far-term improvements to the method