Spalart-Allmaras Turbulence Research Articles

Rotorcraft thickness noise control by tip ventilation

The paper explains further developments of a new concept called TNC (Thickness Noise Control) of the application of surface ventilation to the reduction of helicopter rotor low-frequency in-plane harmonic (LF-IPH) noise. The TNC method is based on introduction of four cavities covered by perforated plates (connected to low and high pressure reservoirs) and positioned symmetrically at the front and rear extremities of the blade tip. Two operational modes are analyzed: constant (steady) and periodically varying (unsteady) transpiration mass-fluxes. For exemplary two-bladed model helicopter rotor of Boxwell et al. (Bell UH-1H Iroquois helicopter) in hover conditions, the results of numerical simulations, based on the CFD code SPARC (Spalart-Allmaras turbulence and Bohning-Doerffer transpiration models), suggest that the acoustic pressure fluctuations are significantly reduced in the near-field of the blade tip. Moreover, the unsteady approach (designed for forward located observers) is almost equally efficient to the basic (steady) operation mode, while substantially lowering penalties in terms of the aerodynamic performance and required transpiration flow intensity. Finally, it is proven that when TNC is activated, the twisted blade (operating in low-thrust conditions) exhibits lower deterioration of the aerodynamic performance compared to its straight (untwisted) counterpart.

Rotorcraft thickness noise control

The paper describes an innovative idea of Thickness Noise Control (TNC) based on adoption of a flow control strategy (i.e. surface ventilation) for acoustic attenuation of helicopter rotor periodic noise. The TNC method is relying on incorporation of multiple cavities (closed by perforated panels and linked to low- and high-pressure reservoirs) located in a symmetrical manner at front and rear portions of the blade tip. The efficiency of the new approach is verified using a two-bladed model rotor of Purcell (untwisted variant of the blade of Bell UH-1H Iroquois helicopter) in low-thrust hover conditions. The results of numerical simulations, obtained with CFD solver (Spalart–Allmaras turbulence and Bohning–Doerffer transpiration models), indicate that in the near-field of the blade tip, both the amplitude and spectral contents of pressure impulses of emitted thickness noise are significantly improved. The TNC method, in the proposed unsteady mode of operation, turns out to be a suitable means of thickness noise reduction in forward flight. Moreover, it is demonstrated that by proper azimuthal activation the efficiency is almost unaltered, while the rotor torque penalty and required transpiration mass-flux are decreased by a factor of 3–5 compared to a steady arrangement.

Hybrid Numerical Simulation of Jet Blast Distance of a Departing Aircraft

A hybrid numerical simulation method was established by combining the Spalart-Allmaras (SA) turbulence model and detached eddy simulation (DES). Numerical simulations were carried out to model cold and hot spray conditions of a nozzle without considering the internal flow of an engine to determine jet conditions. Analysis results show that the calculated hot spray results more in line with the reality. The jet effect of a typical aircraft engine was simulated numerically to determine the distance influenced by the jet blast from a departing aircraft engine.

Numerical modelling of shock wave-boundary layer interaction control by passive wall ventilation

The normal shock wave-boundary layer interaction (SBLI) phenomenon is known to constitute a main factor limiting the aerodynamic performance in many aeronautical applications (transonic wings, helicopter rotor blades, compressor and turbine cascades). The interaction process highly disturbs the boundary layer, often causing flow separation and onset of large scale unsteadiness (e.g. airfoil buffet or supersonic inlet buzz). In certain conditions it may also initiate a dramatic increase of acoustic emission levels (e.g. high-speed impulsive noise). To limit the negative impact of the phenomenon various flow control strategies are implemented, here in a form of a passive control system realised by placing a shallow cavity covered by a perforated plate just beneath the shock. Details of the flow structure obtained by this method are studied numerically. Three distinctive experimental set-ups are considered with the interaction taking place: on a flat wall (transonic nozzle, ONERA), on a convex wall (curved duct, University of Karlsruhe), and on an airfoil (NACA 0012, NASA Langley). Depending on the relative cavity length the ventilation process leads to a transformation of the normal shock topology into: a large λ-foot structure (classical, short cavity), a system of oblique waves (extended cavity), or a gradual compression (full-chord perforation). The reference and flow control cases are simulated with the SPARC code (RANS) with Spalart–Allmaras turbulence and Bohning–Doerffer transpiration models. The results are compared with the measurements, emphasizing the streamwise evolution of the boundary layer profiles and integral parameters during the interaction. The prediction capabilities of the solver in terms of the shock wave-boundary layer interaction control by wall ventilation are assessed and presented in details for the investigated range of flow configurations and conditions.

Numerical study of potential application of active suction and blowing through blade tip perforation to reduction of helicopter rotor thickness noise

The article explains a new concept (called here ATNC – active thickness noise control) of the application of surface transpiration to the reduction of helicopter main rotor harmonic noise. The ATNC method is based on an introduction of four cavities covered by perforated plates (connected to pressure reservoirs) at the leading and trailing edges of the outer 20% of blade span. For an exemplary, two-bladed (NACA 0012 airfoil) model helicopter rotor of Purcell in low-lifting, hover conditions (tip Mach number MaT = 0.66 and collective θ = 1.5° ) the results of numerical simulations, based on the SPARC code (Spalart-Allmaras turbulence and Bohning-Doerffer transpiration models), prove that the acoustic pressure fluctuations are significantly reduced by ATNC in the near-field of blade tip. The peak amplitude is attenuated by more than 45%, with a reduction of the Overall Sound Pressure Level OASPL by 3.4 dB. It is shown also that the venting has not only a large impact on the acoustic radiation, but also on the aerodynamic performance of the rotor, expressed in terms of the torque penalty of 38%. The ATNC method proves to be a promising candidate as a mean of helicopter rotor thickness noise control, but not in the current arrangement.

Influence of wind speed, stay cable inclination angle and wind yaw angle on formation of rivulets

Combining lubrication theory and CFD technology, a finite element model is established to simulate the rain-wind-induced vibration (RWIV). Based on Spalart-Allmaras(S-A)turbulence type, COMSOL software is adopted to calculate the wind pressure coefficient and wind friction coefficient that vary with the location and time. To verify the veracity and rationality of this method, the formation and evolution of rivulets at different wind speeds are studied and compared with the existing experimental results. Furthermore, the time, location, height and width of the initial formation of rivulets are analyzed at different wind speeds, cable inclination angles and wind yaw angles. The results show that the three influencing factors mentioned above have great effect on the formation of rivulet, and the influencing tendency, range and degree are different from each other.

Analysis on aerodynamic characteristics of morphing wing with flexible trailing edge

The wing applying active variable camber trailing edge technology can improve aircraft's aerodynamic performance,and the study of aerodynamic characteristics is of great significance to the design of the morphing wing. Choosing the 2D variable camber wing with flexible trailing edge as the object of study,the compressible Navier-Stokes equation and Spalart-Allmaras( S-A) turbulence model were applied to investigate the aerodynamics of morphing wing numerically by using the software Fluent. Aerodynamic characteristics were studied from the perspective of pressure distribution,flow field structure and the deformation mode. The numerical results show that the morphing wing has almost the same slope of lift curve and maximum lift coefficient with the conventional wing with simple flap,but has lower stall angle of attack. It has higher lift coefficient,lift-to-drag ratio,and higher nose-down pitching moment coefficient than conventional wing before the occurrence of stall. When flexible trailing edge deflects downward to a special angle,trailing edge vortex is suppressed to expand forward. The lift coefficient slowly rises after the stall,which increases the range of effective angle of attack of morphing wing and leads to a good stall characteristic.

Effects of tip cavity geometries on the aerothermal performance of the transonic turbine blade with cavity tip

The inevitable transonic tip leakage flow in unshrouded high pressure gas turbines induces aerodynamic loss and gives rise to heavy thermal loads on the blade tip. In the pursuit of superior gas turbine engine, cavity tips are widely used in high-pressure turbines to reduce the over-tip leakage flow. The present research aims to numerically investigate the influence of the depth of cavity and thickness of squealer rim on the aerothemal performance of the transonic turbine blade. Three-dimensional (3D) Reynolds-Averaged Navier-Stokes (RANS) equations with Spalart-Allmaras turbulence were solved to perform the simulation. The obtained results indicate that the variations of the cavity depth and squealer rim thickness have significant effects on the aerothermal performance of the transonic turbine blade tip regions. The phenomena in this transonic setup are different from the results obtained in subsonic condition. This study quantified that the improvement of aerodynamic efficiency contributed to the enhancement of heat flux on the blade tip in transonic condition by changing the geometry of tip cavity.

Numerical Investigation for Aerodynamic Derivatives of Bridge Deck Using DES

Detached Eddy Simulation（DES）is quite a new approach for the treatment of turbulence, which unites the efficiency of Reynolds Averaged Navier-Stokes Simulation (RANS) and the accuracy of Large Eddy Simulation (LES) into one framework. In this paper, DES method based on Spalart-Allmaras (S-A) turbulence model is employed to simulate the incompressible viscous flow around bridge decks. In order to obtain the aerodynamic forces, the forced motion simulations of the bridge decks are implemented by self-developed codes combined with FLUENT software. After obtaining the aerodynamic forces, aerodynamic derivatives are determined based on the least square algorithm. As the examples, the thin flat plate and the Great Belt East Bridge suspended spans cross-section are investigated to calculate their aerodynamic derivatives. Finally, the simulation results are compared to the data reported in other studies. The comparisons show that the present method gives much better prediction of the aerodynamic derivatives than RANS method and discrete vortex method (DVM).

Numerical study on the standing morphology of an oblique detonation wave under the influence of an incoming boundary layer

Abstract The influence of an incoming boundary layer to the standing morphology of an oblique detonation wave (ODW) induced by a compression ramp is numerically studied in this paper. The Spalart-Allmaras (SA) turbulence model is used to perform simulation of detonationboundary- layer interactions. Three different wall conditions are applied to realize control on the boundary-layer separation scales. Accordingly, different standing morphologies of the ODWs are obtained, including smooth ODW (without transverse wave) under no-slip, adiabatic wall condition with large-scale separation, abrupt ODW (with transverse wave) under no-slip, cold wall condition with moderate-scale separation, and bow-shaped detached ODW under slipwall condition without a boundary layer.

Reynolds Number Effects Investigation of Supercritical Airfoil Based on EFD and CFD

This study aimed to investigate the transonic Reynolds number effects of supercritical airfoil by EFD and CFD method. An experiment was conducted in NF-6 wind tunnel, to obtain the pressure distribution and aerodynamic coefficients of a typical supercritical airfoil through pressure measuring, with Reynolds numbers varied from 3.5×10 6 to 1.0×10 7 per airfoil chord, Mach numbers from 0.6 to 0.8, angles of attack from 0° to 8°. Also, flows over the supercritical airfoil were numerically studied; the two-dimensional Navier-Stokes equations were solved with structure grids by utilizing the Spalart-Allmaras (S-A) turbulence model, with Reynolds numbers varied from 2.0×10 6 to 50×10 6 per airfoil chord and Mach numbers from 0.6 to 0.8. Computational results compared well with experimental results. It is shown that the upper surface pressure distribution of supercritical airfoil including the location and intensity of shock wave and trailing-edge pressure coefficient, changed apparently with variable Reynolds numbers, when shock-induced trailing-edge separation existed. It is also noticed that the lift coefficient increased, drag and pitching moment coefficient decreased as Reynolds number increasing. Results implied that Reynolds number effects should be considered during the early designing stage and optimization of large aircrafts applied supercritical airfoil.

A New Wall Function Model for RANS Equations Based on Overlapping Grids

This paper presents a new numerical method for the modeling of turbulent flows based on a new wall model for computing Reynolds-Averaged-Navier-Stokes (RANS) equations with the Spalart-Allmaras (SA) turbulence model. The basic objective is the reduction of the total central processing unit (CPU) cost of the numerical simulation without harming the accuracy of the results. The main idea of this study is based on the use of two overlapping computational grids covering the two distinct regions of the flow (i.e., the boundary layer and the outer region), and the implementation of appropriate (different) numerical schemes in each case. The seamless cooperation of the grids in the iterative algorithm is achieved by defining an alternative wall function concept. The unstructured grid (UG) covering the outer region consists of mixed type elements (i.e., quadrilaterals and triangles), with relatively small degrees of anisotropy, on which the full set of Navier-Stokes (NS) along with the turbulent model (TM) equations are relaxed. The inner structured grid (SG), which aims at resolving the boundary layer, is a body-fitted mesh with high element density in the normal to the wall direction. The slow relaxation of the governing equations on anisotropic SGs is alleviated by using the Tridiagonal Matrix Algorithm (TDMA) and a block Lower Upper Method (LU). These prove to be quite suitable for the relaxation of the discretized equations on SGs, which consist of banded arrays in tensor form. The application of the proposed algorithm in a couple of benchmark cases proves its superiority over the High-Reynolds SA model with standard wall functions when both methods are compared with the (more costly) Low-Reynolds SA turbulence model and experimental results.

Ice Accretion Prediction for Single Airfoil and Multi-Element Airfoil

Ice accretion on aircraft components is an enormous threat to flight safety. In this paper, ice accretions on the leading edge of the NACA 0012 airfoil and the NLR 7301 multi-element airfoil with flap are predicted using the icing code developed by us. This code mainly contains five modules which are grid module, airflow module, droplet module, heat module, and boundary reconstruction module. The effectiveness and robustness of this code are tested by executing the five modules orderly and repeatedly. The Spalart-Allmaras one-equation turbulence model is adopt to calculate the viscous airflow field and the four-order Runge-Kutta method is used to solve the droplet trajectory equations. In order to enhance the efficiency of the icing calculations, the multi-block grid technique is integrated into the grid module. Based on the above methods, numerical results in both two cases are presented and the necessary comparisons with the experimental data are given in corresponding chapters. The computational results show that performance of the icing code is very good for the wide range of icing conditions.

NUMERICAL SIMULATION OF PARACHUTE INFLATION PROCESS USING AN OVERSET DEFORMING GRIDS METHOD

A numerical method for the simulation of parachute inflation process is presented in this paper. The unsteady compressible N-S equations are fully coupled with MSD (Mass Spring Damper) structure model and integrated forward in time. The CFD solver is based on an unstructured finite volume algorithm and the preconditioning technique is applied to alleviate the stiffness caused by low Mach number. The Spalart-Allmaras one-equation turbulence model is implemented to evaluate the turbulent viscosity. The whole system (fluid equations and structural model equations) is marched implicitly in time using a dual time stepping method. An overset deforming grids method is adopted in this paper to deal with the very large domain deformation during the parachute inflation process. Finally numerical test is performed to validate the robustness of this method.

NUMERICAL SIMULATION OF UNSTEADY FLOW FIELD AROUND HELICOPTER IN FORWARD FLIGHT USING A PARALLEL DYNAMIC OVERSET UNSTRUCTURED GRIDS METHOD

A parallel Navier-Stokes solver based on dynamic overset unstructured grids method is presented to simulate the unsteady turbulent flow field around helicopter in forward flight. The grid method has the advantages of unstructured grid and Chimera grid and is suitable to deal with multiple bodies in relatively moving. Unsteady Navier-Stokes equations are solved on overset unstructured grids by an explicit dual time-stepping, finite volume method. Preconditioning method applied to inner iteration of the dual-time stepping is used to speed up the convergence of numerical simulation. The Spalart-Allmaras one-equation turbulence model is used to evaluate the turbulent viscosity. Parallel computation is based on the dynamic domain decomposition method in overset unstructured grids system at each physical time step. A generic helicopter Robin with a four-blade rotor in forward flight is considered to validate the method presented in this paper. Numerical simulation results show that the parallel dynamic overset unstructured grids method is very efficient for the simulation of helicopter flow field and the results are reliable.

A Comparison of Approximate Versus Exact Geometrical Representations of Roughness for CFD Calculations of cf and St

Skin friction (cf) and heat transfer (St) predictions were made for a turbulent boundary layer over randomly rough surfaces at Reynolds number of 1×106. The rough surfaces are scaled models of actual gas turbine blade surfaces that have experienced degradation after service. Two different approximations are used to characterize the roughness in the computational model: the discrete element model and full 3D discretization of the surface. The discrete element method considers the total aerodynamic drag on a rough surface to be the sum of shear drag on the flat part of the surface and the form drag on the individual roughness elements. The total heat transfer from a rough surface is the sum of convection on the flat part of the surface and the convection from each of the roughness elements. Correlations are used to model the roughness element drag and heat transfer, thus avoiding the complexity of gridding the irregular rough surface. The discrete element roughness representation was incorporated into a two-dimensional, finite difference boundary layer code with a mixing length turbulence model. The second prediction method employs a viscous adaptive Cartesian grid approach to fully resolve the three-dimensional roughness geometry. This significantly reduces the grid requirement compared to a structured grid. The flow prediction is made using a finite-volume Navier-Stokes solver capable of handling arbitrary grids with the Spalart-Allmaras (S‐A) turbulence model. Comparisons are made to experimentally measured values of cf and St for two unique roughness characterizations. The two methods predict cf to within ±8% and St within ±17%, the RANS code yielding slightly better agreement. In both cases, agreement with the experimental data is less favorable for the surface with larger roughness features. The RANS simulation requires a two to three order of magnitude increase in computational time compared to the DEM method and is not as readily adapted to a wide variety of roughness characterizations. The RANS simulation is capable of analyzing surfaces composed primarily of roughness valleys (rather than peaks), a feature that DEM does not have in its present formulation. Several basic assumptions employed by the discrete element model are evaluated using the 3D RANS flow predictions, namely: establishment of the midheight for application of the smooth wall boundary condition; cD and Nu relations employed for roughness elements; and flow three dimensionality over and around roughness elements.

Simulation of Supersonic Jet Noise with Nonlinear CFD and Kirchhoff Surface Integral

An acoustic prediction capability for supersonic axisymmetric jets was developed on the basis of the OVERFLOW Navier-Stokes CFD (Computational Fluid Dynamics) code. Reynolds-averaged turbulent stresses in the flow field are modeled with the aid of Spalart-Allmaras one-equation turbulence model. Appropriate acoustic and outflow boundary conditions were implemented to compute time-dependent acoustic pressure in the nonlinear source-field. Based on the specification of acoustic pressure, its temporal and normal derivatives on the Kirchhoff surface, the near-field and the far-field sound pressure levels are computed via Kirchhoff surface integral, with the Kirchhoff surface chosen to enclose the nonlinear sound source region described by the CFD code. The method is validated by a comparison of the predictions of sound pressure levels with the available data for axisymmetric turbulent supersonic perfectly expanded jets.

Assessment of Rotor Hover Performance Using a Node-based Flow Solver

A three-dimensional viscous flow solver has been developed for the prediction of the aerodynamic performance of hovering helicopter rotor blades using unstructured hybrid meshes. The flow solver utilized a vertex-centered finite-volume scheme that is based on the Roe's flux-difference splitting with an implicit Jacobi/Gauss-Seidel time integration. The eddy viscosity are estimated by the Spalart-Allmaras one-equation turbulence model. Calculations were performed at three operating conditions with varying tip Mach number and collective pitch setting for the Caradonna-Tung rotor in hover. Additional computations are made for the UH-60A rotor in hover. Reasonable agreements were obtained between the present results and the experiment in both blade loading and overall rotor performance. It was demonstrated that the present vertex-centered flow solver is an efficient and accurate tool for the assessment of rotor performance in hover.

A numerical analysis of turbulent compressible radial channel flow with particular reference to pneumatic controllers

A numerical analysis of turbulent axisymmetric radial compressible channel flow between a nozzle and a flat plate is presented in this paper. In particular, the application of this type of flow situation in pneumatic dimensional control systems is considered. The Spalart-Allmaras one-equation turbulence model is used. The resulting highly coupled PDE system has been solved using the control volume based numerical approach where the power-law scheme was used extensively to compute the diffusive and convective fluxes of momentum. Results show that local Mach numbers can easily achieve and surpass unity for typical industrial configurations. Also, in the case of a standard industrial nozzle geometry, the presence of a toroidal recirculation zone that moves radially outward is clearly identified in most cases. Separated flow areas are of particular concern as it has been shown previously that they can cause nozzle fouling in industrial applications. It has been shown that the size of this region is dependant on feed pressure. Considerable differences between results obtained using the Spalart-Allmaras and standard k- e turbulence models have also been noticed.

Numerical and experimental investigation of unsteady flow interaction in a low-pressure multistage turbine

The paper presents results of unsteady viscous flow calculations and corresponding cold flow experiments on a three-stage low-pressure turbine. The investigation emphasizes the study of unsteady flow interaction. A time-accurate, Reynolds-averaged Navier-Stokes solver is applied for the computations. Turbulence is modelled using the Spalart-Allmaras one-equation turbulence model. The influence of modern transition models on the unsteady flow predictions is investigated. Integration of the governing equations in time is performed with a four-stage Runge-Kutta scheme, which is accelerated by a two-grid method in the viscous boundary layer around the blades. At the inlet and outlet, non-reflecting boundary conditions are used. The quasi-three-dimensional calculations are conducted on a stream surface around mid-span, allowing a varying stream tube thickness. A three-stage, low-pressure turbine rig of a modern commercial jet engine is used for a study of the unsteady flow interaction. The numerical method is able to capture important unsteady effects found in the experiments, i.e. unsteady transition as well as the bladerow interaction. In particular, the flowfield with respect to time-averaged and unsteady quantities such as surface pressure, vorticity and turbulence intensity is compared with the experiments conducted in the cold airflow test rig.