Jeremy Bain

Computational Modeling of Variable-Droop Leading Edge in Forward Flight

In recent years, there has been significant interest in on-blade concepts that expand the operating envelope of helicopters without compromising the performance characteristics of the baseline rotor. The variable-droop leading-edge concept is explored in a modified version of the Navier-Stokes solver OVERFLOW. Modifications were made to the solver to allow deforming-grid capability. This concept was explored in dynamic stall tests of the VR-12 and SC1095 helicopter airfoils. The variable-droop leading-edge airfoils have dramatically reduced drag and moment rises associated with dynamic stall. Using the results of these tests, a modified UH-60A rotor incorporating variable-droop leading-edge airfoils was analyzed using loosely coupled computational fluid dynamics and comprehensive structural dynamics with OVERFLOW and DYMORE and compared with the standard UH-60A rotor for high-thrust-case flight 9017. Results show a reduction in the peak negative pitching moment. The rotor efficiency was shown to improve by 2.9% and the 4/rev component of vertical force reduced by 8%. These performance improvements can be improved with an improved droop schedule and by incorporating improved transonic airfoils.

A Surrogate Based Approach to Reduced-Order Dynamic Stall Modeling

The surrogate-based recurrence framework (SBRF) approach to reduced-order nonlinear unsteady aerodynamic modeling associated with pitching/plunging airfoils subject to xed or time-varying freestream Mach numbers is described. Using full-order solutions generated by the OVERFLOW CFD code, the SBRF reduced-order modeling approach is shown to eectively mimic full-order solutions of unsteady lift, moment, and drag under dynamic stall conditions, but at a fraction of the computational cost. In addition to accounting for realistic helicopter rotor blade dynamics, it is shown that the SBRF can model advancing rotor blade stall due to shock induced separation, as well as retreating blade stall associated with excessive angles of attack. Therefore, the SBRF reduced-order modeling approach is ideally suited for a variety of aeroelasticity and active/passive design optimization studies that require high delity aerodynamic response solutions with minimal computational expense.

Prediction of Rotor Blade Ice Shedding using Empirical Methods

A methodology that couples computational fluid dynamics, computational structural dynamics, ice accretion models and ice shedding models is developed and applied to both isolated airfoils and rotors in forward flight, the latter with and without shedding. The individual modules are coupled to each other through industry-standard open file I/O methods, allowing the replacement of the individual modules with more advanced modules as technology matures. Ice shape results are presented and correlated with test data for a range of icing conditions. The torque rise associated with ice build-up on a UH-60A rotor in forward flight is modeled. Finally, results are presented for ice shedding phenomena for a small-scale model rotor. Reasonable correlation with test data is observed in the cases studied. BACKGROUND Despite decades of research on the phenomenon, rotor icing remains a major in-flight concern for many civilian and military helicopter operators. One particular facet of rotor blade icing receiving additional attention is shedding, more specifically self shedding. Self shedding occurs on a rotor blade when the aerodynamic and centrifugal forces on a section of ice exceed the structural adhesion forces holding the ice onto the blade. At the point when the adhesion force is exceeded, the ice is said to “shed” and separates from the rotor blade. After separation, the shed ice acts as a projectile and has the potential to strike components of the helicopter such as the tail rotor. In order to better evaluate the risks associated with self shedding, an accurate model must be developed for determining the conditions under which ice shedding will occur. This paper focuses on self-shedding mechanics and the development of a model, based on a combination of computational and empirical methods, for predicting shedding phenomena. The model presented is being developed as part of a larger initiative to investigate the trajectories of shed ice pieces and their potential for striking vehicle components. In order to meet this larger objective, the present model will first be used to predict shedding characteristics at particular operating conditions. At each operating condition analyzed, CFD and six degree-of-freedom modeling would then be used to determine the trajectories of shed ice pieces, as well as possible impact forces.

Reduced-Order Dynamic Stall Modeling with Swept Flow Effects Using a Surrogate-Based Recurrence Framework

locally optimizing the high-dimensional likelihood function in the vicinity of the stationary solution. The resulting nonstationary covariance structures are shown to significantly improve the accuracy of the surrogate-based recurrence framework predicted moment stall characteristics compared to a stationary model. It is shown that the nonstationary surrogate-based recurrence framework approach is better able to adapt to abrupt changes in airload behavior caused by the underlying dynamic stall vortex dynamics. The results indicate that the surrogate-based recurrence framework approach based on nonstationary Gaussian process models is a promising alternative to widelyusedsemiempiricalrotorcraftdynamicstallmodelsthatcannotaccountfortheeffectsoftime-varyingvelocity components associated with forward flight.

Rotorcraft Fuselage Drag Reduction using Combustion Powered Actuators

Separation control of the 3-D flow over the aft body of a scale model of the NASA ROBIN mod7 rotorcraft fuselage is investigated in wind tunnel experiments. Pulsed actuation is effected by arrays of momentary, combustion-based actuator jets having a characteristic time scale O[1 ms] that is an order of magnitude shorter than the convective time scale of the flow. The actuators are placed upstream of the transition region between the fuselage and the tail boom and their interactions with the massively separated cross flow in this domain and effects on the global aerodynamic forces and moments are measured using an onboard six-axis load cell and high resolution PIV that is acquired phase-locked to the actuation waveform. The present investigations have demonstrated that the actuation can significantly mitigate separation, and lead to a reduction in drag (although flow attachment is accompanied by some lift penalty). It also is shown that transitory aft flow attachment using burst-modulated actuation can be exploited for effecting significant steering aerodynamic side forces for improved flight maneuverability.

A surrogate based approach to reduced-order dynamic stall modeling

The surrogate-based recurrence framework (SBRF) approach to reduced-order nonlinear unsteady aerodynamic modeling associated with pitching/plunging airfoils subject to xed or time-varying freestream Mach numbers is described. Using full-order solutions generated by the OVERFLOW CFD code, the SBRF reduced-order modeling approach is shown to e ectively mimic full-order solutions of unsteady lift, moment, and drag under dynamic stall conditions, but at a fraction of the computational cost. In addition to accounting for realistic helicopter rotor blade dynamics, it is shown that the SBRF can model advancing rotor blade stall due to shock induced separation, as well as retreating blade stall associated with excessive angles of attack. Therefore, the SBRF reduced-order modeling approach is ideally suited for a variety of aeroelasticity and active/passive design optimization studies that require high delity aerodynamic response solutions with minimal computational expense.

Assessment of the Effects of Computational Parameters on Physics-based Models of Ice Accretion

A state of the art CFD analysis (OVERFLOW 2.1y) has been coupled to an industry standard ice accretion analysis (LEWICE3D) and grid generator (Chimera Grid Tools) so that temporal growth of ice over 2D airfoils may be computed in a robust automated fashion. Results are presented for a SC1095 airfoil for two conditions. The coupled methodology was used to systematically study the effects of grid density and establish grid convergence. The effects on the icing phenomena attributable to turbulence modeling and the frequency of update of the flow field were investigated. Grid convergence was established at low Mach numbers, and it was concluded that approximately 300 points in the wrap around direction, and a normal y

Assessment of a Kinetic-Eddy Simulation Turbulence Model for 3D Unsteady Transonic Flows

School of Aerospace Engineering, Georgia Institute of Technology The Kinetic-Eddy Simulation (KES) turbulence model has been evaluated for a variety of unsteady and oscillating wing transonic flows. KES is a two equation VLES/LES model that solves for the unresolved kinetic energy and the local turbulent length scales without any dependency on grid spacing. KES was used in OVERFLOW and compared to standard models. KES showed improved ability to capture the negative lift curve slope and separated flow of the NACA0036. In dynamic stall, KES improved prediction of peak lift, drag, and pitching moment. In transonic flows, KES was able to improve modeling of separated flow at the wing-body connection of the DLR-F6 wing-body configuration but over predicted drag. Unsteady oscillating transonic flow over a scaled F-5 wing also showed improved prediction with KES.

Modeling dynamic stall of SC-1095 airfoil at high mach number

The Leishman-Beddoes method of determining airloads for an airfoil undergoing dynamic stall is studied over a wide range of Mach numbers. To validate the method for higher Mach numbers where there is less available experimental data, a Computational Fluid Dynamics solver is utilized to provide airload predictions for comparison to the Leishman-Beddoes results. It is found that even for high Mach numbers the Leishman-Beddoes method provides reliable predictions for lift coefficient. However, at the higher Mach numbers pitching moment is sometimes overpredicted at high angle of attack. This is seemingly due to an inability to accurately determine the center of pressure in the high speed unsteady flow environment.

Neural Network Models for the OnBoard and Individu al Blade Control of Helicopter Rotors

Neural Networks are evaluated to model OnBoard Bla de Control concepts for helicopter rotors. Computational Fluid Dynamic simulations of active flap and active twist concepts have been run using the compressible Navier Stokes solver OVERFLOW. Neural Network models were then made as function of Mach number and mean angle of attack for the change in lift, drag, and pitching moment and their associated time constants. These models show the ability to capture the nonlinear effects of stall and shocks. These models are now suitable for incorporation into rotorcraft flight simulation software. I. Introduction elicopters are versatile vehicles that can vertically take off and land, hover, and perform maneuvers at very low forward speeds. These characteristics make them unique for a number of civilian and military applications. However, the radial and azimuthal variation of dynamic pressure causes rotors to experience adverse phenomena such as transonic shocks and 3-D dynamic stall. Adverse interactions such as blade vortex interaction and rotor- airframe interaction may also occur. These phenomena contribute to noise and vibrations. A variety of techniques have been proposed for reducing the noise and vibrations and for improving handling qualities. These include on-board control (OBC) devices, individual blade control (IBC), and higher harmonic control (HHC). Addition of these devices adds to the weight, cost, and complexity of the rotor system and reduces reliability of operations. Simpler OBC concepts will greatly alleviate these drawbacks and enhance the operating envelope of vehicles. IBC and OBC concepts may be modeled using physics-based computational fluid dynamics (CFD) and computational structural dynamics (CSD) tools that are coupled to each other. However, these approaches are expensive and not suitable for the design of controllers. Reduced order models are highly desirable for designing these devices and for assessing the handling qualities of helicopters that employ these devices. Development of reduced order models is hampered by the fact that the flow field is highly non-linear and spans a wide range of Mach numbers and mean angles of attack. II. Approach A three-step approach is under development for modeling IBC and OBC devices, in particular active twist and trailing edge flap devices mounted on rotor blades. The first step involves conducting 2-D and 3-D non-linear simulations to generate a large database of steady state lift, drag, and pitching moments as a function of angle of attack, Reynolds number, and Mach number. The second step involves modeling the evolution of lift, drag, and pitching moment as a function of time for the active twist and trailing edge control concepts. The third step involves the use of neural network curve fits of the computed static and unsteady flow data. The fourth and final step is the incorporation of these neural network based models in rotorcraft flight simulation software and comparisons with high fidelity simulations. In this work, results from the first three steps are discussed. Incorporation of the neural network models in flight simulation software are also being done. For details, the reader is referred to Ref. 1. The computational fluid dynamics analyses used in this study were all done using the compressible Navier-Stokes computational fluid dynamics code OVERFLOW 2 Versions 2.0y and 2.1y. OVERFLOW, developed by NASA,