| 1. |
- Shia-Hui, Peng, 1967, et al.
(författare)
-
CAA analysis of a wing section with flap side-edges based on hybrid RANS-LES computation
- 2015
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Ingår i: 21st AIAA/CEAS Aeroacoustics Conference, AIAA Aviation. - Reston, Virginia : American Institute of Aeronautics and Astronautics. - 9781624103674
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Konferensbidrag (refereegranskat)abstract
- Based on hybrid RANS-LES computation of the turbulent flow around a double-flapped wing section, CAA (Computational Aero-Acoustics) analysis was conducted using the Curle, the Kirchhoff and the FW-H acoustic analogy methods. The focus was placed on the flow-induced noise due to the flap side-edges (FSE). It was shown that the FSE has triggered extensive unsteady vortex motions and being the most potent noise-generating source with significant pressure fluctuations on the side-edge surface. In the CAA analysis, two integral surfaces, defined by the iso-surface of vorticity magnitude, were verified when using the Kirchhoff and the FW-H methods. The Kirchhoff method is more sensitive to the location of the integral surface. The analysis using the Curle method indicates that the pressure fluctuations on the surface of the main wing and the first flap have made similar contributions to the far-field noise level, while the second flap contributes slightly less. The Kirchhoff and FW-H methods have predicted overall higher noise levels comparing to the Curle method. In the comparison, the result obtained with a stochastic method based on a RANS solution was also involved. The result has clearly demonstrated that, to capture the most potential noise generation in the presence of flap side edges, turbulence-resolving simulations should be incorporated in hybrid CFD/CAA analysis.
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| 2. |
- Andersson, Niklas, 1976, et al.
(författare)
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A Study of Mach 0.75 Jets and Their Radiated Sound Using Large-Eddy Simulation
- 2004
-
Ingår i: AIAA 2004-3024, proc. of 10th AIAA/CEAS Aeroacoustics Conference, May 10-12, 2004.
-
Konferensbidrag (refereegranskat)abstract
- Large-Eddy Simulations (LES) of a compressible nozzle/jet configuration have been carried out. Two jets were simulated, an isothermal jet and a jet with a higher temperature than the quiescent surrounding air. The Mach number was in both cases 0.75 and the jet Reynolds number was 50,000. Sound pressure levels in far-field observer locations were evaluated using Kirchhoff surface integration. The Favre filtered Navier-Stokes equations were solved using a finite volume method solver with a low-dissipation third-order upwind scheme for the convective fluxes, a second-order centered difference approach for the viscous fluxes and a three-stage second-order Runge-Kutta technique in time. The computational domain was discretized using a block structured boundary fitted mesh with approximately 3,000,000 cells. The calculations were performed on a parallel computer, using message-passing interface (MPI). A compressible form of Smagorinsky's subgrid scale model was used for computation of the subgrid scale stresses. Absorbing boundary conditions based on characteristic variables were adopted for all free boundaries. Velocity components specified at the entrainment boundaries were estimated from corresponding Reynolds Averaged Navier-Stokes (RANS) calculations, which enable the use of a rather narrow domain. This, furthermore, ensures that the correct amount of fluid is entrained into the domain. Two-point space-time correlations were obtained for locations in the shear layer center, from which length and time scales of turbulence structures were evaluated. Predicted near-field flow statistics and far-field sound pressure levels (SPL) are both in good agreement with experiments. Predicted (SPL) are for all observers locations, where evaluated, within a 3.0 [dB] deviation from measured levels and for most locations within a 1.0 [dB] deviation. Experimental data used for validation were provided by Laboratoire dEtude Aeròdynamiques, Poitiers, France.
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| 3. |
- Andersson, Niklas, 1976, et al.
(författare)
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Effects of Inflow Conditions and Subgrid Model on LES for Turbulent Jets
- 2005
-
Ingår i: AIAA 2005-2925, proc. of 11th AIAA/CEAS Aeroacoustics Conference, May 23-25, 2005, Monterey, California.
-
Konferensbidrag (refereegranskat)abstract
- The turbulent mixing process prescribing the spreading rate of the jet and the length of the potential core region is influenced by a number of factors. Using large-eddy simulation (LES), the four factors that are believed to be the most important in this respect are: subgrid-scale properties, the accuracy of the numerical scheme, the entrainment boundary conditions, and the inflow conditions. In a previously performed study of a subsonic (Mach 0.75) jet, the turbulence mixing was found to be too efficient and hence the length of the potential core region was underpredicted. In that study indications were found of that the overpredicted mixing was due to the inflow conditions. For a model nozzle, capturing the initial turbulent shear flow might not be of that great importance for accurate prediction of radiated sound since most of these effects will appear in the high-frequency range. When dealing with real engine geometries, however, it becomes quite important. Moreover, methods for industrial use have to cope with complex geometries and high temperature and velocity ratios making the ability to capture the initial flow physics even more important. In the present work LES has been used for the same Mach 0.75 jet. The acoustic field is extracted to the far field using Kirchhoff surface integration. The effects of inflow conditions, Reynolds number, and subgrid-scale model on flowfield and acoustic signature are investigated.The Favre-filtered Navier-Stokes equations were solved using a finite-volume method solver with a low-dissipation third-order upwind scheme for the convective fluxes, a second-order centered difference approach for the viscous fluxes and a three-stage second-order Runge-Kutta technique in time. The computational domain was discretized using a block-structured boundary-fitted mesh with approximately 3,000,000 cells. The calculations were performed on a parallel computer, using message-passing interface (MPI). A compressible form of Smagorinsky's subgrid-scale model was used to compute the subgrid-scale stresses. Absorbing boundary conditions based on characteristic variables were adopted for all free boundaries.
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| 4. |
- Andersson, Niklas, 1976, et al.
(författare)
-
Investigation of an Isothermal Mach 0.75 Jet and its Radiated sound Using Large-Eddy Simulation and Kirchhoff Surface Integration
- 2005
-
Ingår i: International Journal of Heat and Fluid Flow. ; 26, s. 393-410
-
Tidskriftsartikel (refereegranskat)abstract
- A large-eddy simulation (LES) of a compressible nozzle/jet configuration has been carried out. An isothermal Mach 0.75 jet was simulated. The Reynolds number based on the jet velocity at the nozzle exit plane and the nozzle diameter was 50,000. The Favre filtered Navier-Stokes equations were solved using a finite volume method solver with a low-dissipation third-order upwind scheme for the convective fluxes, a second-order centered difference approach for the viscous fluxes and a three-stage second-order Runge-Kutta time marching technique. A compressible form of Smagorinsky's subgrid scale model was used for computation of the subgrid scale stresses. The computational domain was discretized using a block structured boundary fitted mesh with approximately 3,000,000 cells. The calculations were performed on a parallel computer, using message-passing interface (MPI). Absorbing boundary conditions based on characteristic variables were adopted for all free boundaries. Velocity components specified at the entrainment boundaries were estimated from a corresponding Reynolds Averaged Navier-Stokes (RANS) calculation, which enabled the use of a rather narrow domain. In order to diminish disturbances caused by the outlet boundary, a buffer layer was added at the domain outlet. Kirchhoff surface integration using instantaneous pressure data from the LES was utilized to obtain far-field sound pressure levels in a number of observer locations. The predicted sound pressure levels were for all observer locations within a 3dB deviation from the measured levels and for most observer locations within a 1dB deviation. Aerodynamic results and predicted sound pressure levels are both in good agreement with experiments. Experimental data were provided by Laboratoire dEtude Aeròdynamiques, Poiters, France.
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| 5. |
- Andersson, Niklas, 1976, et al.
(författare)
-
Large-Eddy Simulation of a Mach 0.75 Jet
- 2003
-
Ingår i: AIAA 2003-3312, proc. of 9th AIAA/CEAS Aeroacoustic Conference, May 12-14, 2003.
-
Konferensbidrag (refereegranskat)abstract
- A large-eddy simulation (LES) of a compressible nozzle/jet configuration has been carried out. A cold Mach 0.75 jet was simulated. The Reynolds number based on the jet velocity at the nozzle exit plane and the nozzle diameter was 50,000. The Favre filtered Navier-Stokes equations were solved using a finite volume method with a low dissipative third order upwind scheme for the convective fluxes, a second order centered difference approach for the viscous fluxes and a three-stage second order Runge-Kutta time marching technique. A compressible form of Smagorinsky's sub-grid scale model was used for computation of the sub-grid scale stresses. The calculations were performed using a block structured boundary fitted mesh with approximately 3,000,000 cells. The calculations have been performed on a parallel computer, using message-passing interface (MPI). Absorbing boundary conditions based on characteristic variables were adopted for all free boundaries. Velocity components specified at the entrainment boundaries were estimated from a corresponding Reynolds Averaged Navier-Stokes (RANS) calculation. In order to diminish disturbances caused by the outlet boundary a buffer layer was added at the domain outlet. Kirchhoff surface integration has been utilized to obtain far-field sound pressure levels in a number of observer locations using instantaneous pressure from the LES. Aerodynamic results and predicted sound pressure levels are both in good agreement with experiments. Experimental data were provided by Laboratoire dEtude Aeròdynamiques, Poiters, France.
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| 6. |
- Andersson, Niklas, 1976, et al.
(författare)
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Large-Eddy Simulation of Subsonic Turbulent Jets and Their Radiated Sound
- 2005
-
Ingår i: AIAA Journal. ; 43:9, s. 1899-1912
-
Tidskriftsartikel (refereegranskat)abstract
- Large-Eddy Simulations (LES) of a compressible nozzle/jet configuration have been carried out. Two jets were simulated, an isothermal jet and a jet with a higher temperature than the quiescent surrounding air. The Mach number was in both cases 0.75 and the jet Reynolds number was 50,000. Sound pressure levels in far-field observer locations were evaluated using Kirchhoff surface integration. The Favre filtered Navier-Stokes equations were solved using a finite volume method solver with a low-dissipation third-order upwind scheme for the convective fluxes, a second-order centered difference approach for the viscous fluxes and a three-stage second-order Runge-Kutta technique in time. The computational domain was discretized using a block structured boundary fitted mesh with approximately 3,000,000 cells. The calculations were performed on a parallel computer, using message-passing interface (MPI). A compressible form of Smagorinsky's subgrid scale model was used for computation of the subgrid scale stresses. Absorbing boundary conditions based on characteristic variables were adopted for all free boundaries. Velocity components specified at the entrainment boundaries were estimated from corresponding Reynolds Averaged Navier-Stokes (RANS) calculations, which enable the use of a rather narrow domain. This, furthermore, ensures that the correct amount of fluid is entrained into the domain. Two-point space-time correlations were obtained for locations in the shear layer center, from which length and time scales of turbulence structures were evaluated. Predicted near-field flow statistics and far-field sound pressure levels (SPL) are both in good agreement with experiments. Predicted (SPL) are for all observers locations, where evaluated, within a 3.0 [dB] deviation from measured levels and for most locations within a 1.0 [dB] deviation. Experimental data used for validation were provided by Laboratoire dEtude Aeròdynamiques, Poiters, France.
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| 7. |
- Andersson, Niklas, 1976, et al.
(författare)
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LES Prediction of Flow and Acoustic Field of a Coaxial Jet
- 2005
-
Ingår i: AIAA 2005-2884, proc. of 11th AIAA/CEAS Aeroacoustics Conference, May 23-25, 2005, Monterey, California.
-
Konferensbidrag (refereegranskat)abstract
- A compressible high-subsonic coaxial jet has been simulated using large-eddy simulation (LES). The acoustic field was extended to the far field using Kirchhoff surface integration. The jet Mach number based on the local speed of sound is approximately 0.9 for both the primary and secondary stream. The static temperature in the primary stream is three times that of the secondary stream. In order to resolve the acoustic field, it is desirable to have a computational domain with a rather large radial extent and a mesh that is relatively fine even in the far-field regions. Furthermore, the mesh should be as equidistant as possible so as to minimize the introduction of numerical errors. In order to keep the number of cells down, the computational domain was divided into three regions: a well resolved near-wall LES region, a medium-resolution LES region optimized for propagation of acoustic waves, and a coarse LES region. Over the interfaces between these regions, the number of cells is increased by factor two in each direction. Special treatment of the interfaces between the regions is utilized in order to minimize undesirable numerical errors. The radial extent of the computational domain increases downstream such that the flow in the outer boundary region can be assumed to be irrotational and axisymmetric. Hence, the flow outside the three-dimensional computational domain can be represented by a less expensive two-dimensional axisymmetric calculation. The interface between the full 3D LES region and the 2D region is based on azimuthally averaged quantities and acts as an absorbing boundary condition. The Favre-filtered Navier-Stokes equations were solved using a finite-volume method solver with a low-dissipation third-order upwind scheme for the convective fluxes, a second-order centered difference approach for the viscous fluxes and a three-stage second-order Runge-Kutta technique in time. The computational domain was discretized using a block-structured boundary-fitted mesh with approximately 20,000,000 nodes. The calculations were performed on a parallel computer, using message-passing interface (MPI). A compressible form of Smagorinsky's subgrid-scale model was used to compute the subgrid-scale stresses.
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| 8. |
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| 9. |
- Billson, Mattias, 1973, et al.
(författare)
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Acoustic Source Terms for the Linearized Euler Equations in Conservative Form
- 2005
-
Ingår i: AIAA journal. ; 43:4, s. 752-759
-
Tidskriftsartikel (refereegranskat)abstract
- A rather novel approach to predict jet noise is the Stochastic Noise Generation and Radiation (SNGR) method. The SNGR method uses the linear Euler equations as an acoustic analogy together with source terms which are modeled. In other studies the Euler equations on primitive form are used. In the present work the linear Euler equations on conservative form are used. Due to this, new source terms have to be derived for the conservative set of equations. A formal derivation of the correct source terms for the linear Euler equations on conservative form is presented. Simplified versions of the derived source terms are also developed. To validate the derived source terms a direct simulation of a forced 2D mixing layer is carried out. The solutions to the linearized Euler equations with source terms are compared to the solution of the direct simulation and show a good agreement. All simulations are performed using Tam and Webb's fourth order DRP scheme and a four step fourth order Runge-Kutta time marching technique. Artificial selective damping introduced through the numerical scheme is used to avoid spurious waves. Absorbing boundary conditions based on characteristic variables, Engquist and Majda, are used at the free boundaries and a buffer layer is added at the outflow.
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| 10. |
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| 11. |
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| 12. |
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| 13. |
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| 14. |
- Cuppoletti, Dan, et al.
(författare)
-
A Comprehensive Investigation of Pulsed Fluidic Injection for Active Control of Supersonic Jet Noise
- 2013
-
Ingår i: 51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. - Reston, Virigina : American Institute of Aeronautics and Astronautics. ; 2013
-
Konferensbidrag (refereegranskat)abstract
- Fluidic injection for noise control of high Reynolds number jets has shown promise and recent tests have demonstrated improved noise reduction while decreasing the injection mass flow required. This investigation was an experimental and numerical study on the capability of pulsed fluidic injection to reduce noise on a Md = 1.56 supersonic jet. The effect of pulse frequency, duty cycle, injector phasing, and injection angle on the noise components were studied. The pulsed injectors were characterized with hot-wire measurements. Far-fleld acoustics was used to survey the noise reduction of pulsed injection (up to 400 Hz) in comparison to the baseline and steady injection cases. Injection angles θinj = 30° to 90° with respect to the primary jet axis were investigated. High-speed shadowgraph was used to quantify the time scales involved in response of the shock train and screech instabilities with pulsed fluidic injection. LES and CAA were compared with measurements to evaluate the capability of numerical simulation of the pulsed injection configurations. It was shown that reduction of turbulent mixing noise generally scales with the actual duty cycle of applied injection. For 30 Hz injection at 20% mass flow up to up to 80% of the steady flow {increment}OASPL is achieved, demonstrating that low frequency injection is capable of enhanced noise reduction at certain conditions. The shocks in the jet potential core respond in 1 ms when injection is removed, while the jet column instability requires up to 7 ms to redevelop after injection is removed. The results demonstrate the feasibility of using active control with pulsed fluidic actuators to provide at least steady flow noise reduction with significantly reduced injection mass flow.
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| 15. |
- Cuppoletti, Daniel, et al.
(författare)
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Analysis of Supersonic Jet Thrust with Fluidic Injection
- 2014
-
Ingår i: 52nd AIAA Aerospace Sciences Meeting - AIAA Science and Technology Forum and Exposition, SciTech 2014; National Harbor, MD; United States; 13 January 2014 through 17 January 2014. - Reston, Virginia : American Institute of Aeronautics and Astronautics. - 9781624102561
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Konferensbidrag (refereegranskat)abstract
- Considerable focus on noise abatement for aircraft has spawned various noise control devices, passive and active. Aircraft and propulsion system design now has the additional criteria of acoustic performance to consider among many other criteria in advanced flight vehicle design. It is essential to consider the effect that noise control methods have on the performance of the propulsion device and overall effect on system performance. Thrust calculated from measurements and LES are compared for a Md = 1.56 jet at various operating conditions for validation. Experimental measurements on the baseline supersonic jet are used to validate computational results for the pressure and momentum thrust components. Thrust for various fluidic injection configurations are evaluated using computational results from the highly three dimensional flowfield. Analysis and discussion of requirements for fluidic injection air are provided to develop a complete system approach to aid design of fluidic injection systems. Fluidic injection decreases momentum thrust by creating axial velocity deficits in the region of injection. Pressure thrust is increased from local pressure rise from the injectors and area control at the nozzle exit. Fluidic injection increases total thrust as the pressure thrust gains are greater than the momentum thrust deficits. Specific thrust is reduced slightly with 6 injectors being a more efficient use of the injection air with greater noise reduction.
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| 16. |
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| 17. |
- Daniel, Cuppoletti, et al.
(författare)
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Nozzle Throat Optimization on Acoustics and Performance of a Supersonic Jet
- 2012
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Ingår i: 18th AIAA/CEAS Aeroacoustics Conference (33rd AIAA Aeroacoustics Conference). - Reston, Virigina : American Institute of Aeronautics and Astronautics.
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Konferensbidrag (refereegranskat)abstract
- Nozzles used in supersonic flight applications have flow contours that cause the flow to differ from isentropic nozzle flow, resulting in less than ideal nozzle performance. The impact of nozzle contour on performance is well quantified, however it is less clear how the nozzle contour affects supersonic jet noise. This work investigates differences in noise characteristics of a sharp throat and contoured throat nozzle to identify the dependencies of supersonic noise components on the nozzle design. The nozzles are designed to be thrust matched at fully expanded conditions. The throat contour does not significantly affect the acoustics at fully expanded conditions, although the nozzle efficiency is increased for the contoured throat nozzle. Contouring the throat causes the nozzle to have screech instabilities over a broader range of operating conditions when imperfectly expanded. A detailed PIV and LES investigation was used to explain the acoustics behavior at all conditions. Reducing the throat shock strength increased the nozzle exit shock strength and periodicity, subsequently increasing the susceptibility to screech. Nozzle performance is increased at all operating conditions with the contoured throat nozzle.
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| 18. |
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| 19. |
- Eliasson, P., et al.
(författare)
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Assessment of high-lift concepts for a regional aircraft in the ALONOCO project
- 2012
-
Ingår i: 50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition, Nashville, 9-12 January 2012. - Reston, Virigina : American Institute of Aeronautics and Astronautics. - 9781600869365
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Konferensbidrag (refereegranskat)abstract
- This work introduces the work conducted in the EU JTI project ANOLOCO, which has aimed at an assessment of aerodynamic and aeroacoustic performance of several high-lift configurations of a regional aircraft. The high-lift designs are for a laminar and slat-less wing, including configurations with a double slotted flap, single slotted flap, drooped nose and a Krueger flap. The aerodynamic performance is assessed from steady state RANS calculations up to maximum lift. The aeroacoustic performance is based on hybrid RANS-LES calculations for flow-induced noise generation, and using acoustic analogy methods for far-field noise propagation. Three different analogy methods are evaluated and compared. The assessment shows that the configuration with a Krueger flap gives the best performance. The maximum lift is close to 20% higher than for any other configuration and the noise levels are also reduced, up to 10 dB lower than the configuration with a double slotted flap.
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| 20. |
- Ellbrant, Lars R, 1984, et al.
(författare)
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Balancing efficiency and stability in the design of transonic compressor stages
- 2013
-
Ingår i: Proceedings of the ASME Turbo Expo 2013. - 9780791855232 ; 6 B, s. (Article number) GT2013-94838
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Konferensbidrag (refereegranskat)abstract
- The paper describes an efficient design method for highly loaded transonic compressor stages which considers a balance between efficiency at the design point and stability at part-speed. Because of the high dimensionality of the problem, two levels of model complexity are included in the design method. The first level consists of optimizing the rotor and stator profiles positioned at three stream tubes along the span. The stream tube height and radius variations are included in the computational domain and it is analyzed using a 3D RANS solver incorporating a mixing plane between the components. Due to the relatively low complexity of this quasi-3D analysis, it is fast enough to explore a large design space. With the aid of the resulting pareto-fronts, the designer can select profiles with the appropriate trade between stability and efficiency. The initial 3D compressor stage is generated based on the selected 2D profiles and the method continues to the higher complexity mode where the 3D shapes of the rotor and stator are optimized to gain further performance improvements. To verify that the design method is feasible, it is used to re-design the first compressor stage of a three-stage highly loaded transonic compressor. The compressor stage designed with the presented design method has higher part-speed stability without a compromise in the efficiency compared to the original design. This is also verified when analyzing the new design in the full compressor module.
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| 21. |
- Ellbrant, Lars R, 1984, et al.
(författare)
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CFD optimization of a transonic compressor using multiobjective GA and metamodels
- 2012
-
Ingår i: 28th Congress of the International Council of the Aeronautical Sciences 2012, ICAS 2012; Brisbane; Australia; 23 September 2012 through 28 September 2012. - 9781622767540 ; 4:ICAS 2012 Paper no. 267, s. 2698-2712
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Konferensbidrag (refereegranskat)abstract
- This paper present an automatic design process where a non-deterministic, global search optimization is utilized to optimize a first stage rotor of a highly loaded transonic compressor. The first part is focused on finding the best suited metamodel that can be used to accelerate the design process. The second part presents the results of using the meta model within the design process for an industry relevant case. Using the radial basis functions as acceleration technique for the optimization was seen to be very successful. The meta model assisted optimization reduced the total design time from approximately 2 weeks to 3.5 days given that 8 designs could run in parallel on a cluster. The 3D optimization produced a pareto front where it was possible to select blades having either high efficiency or high stability.
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| 22. |
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| 23. |
- Ellbrant, Lars R, 1984, et al.
(författare)
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Design of compressor blades considering efficiency and stability using CFD based optimization
- 2012
-
Ingår i: ASME Turbo-expo 2012. - 9780791844748 ; 8:PARTS A, B, AND C, s. 371-382
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Konferensbidrag (refereegranskat)abstract
- To design a highly loaded axial transonic compressor several objectives need to be considered simultaneously. From an aerodynamic perspective, one of the major requirements is high efficiency at a specific operating condition where the fuel consumption is of main interest. Furthermore, the compressor needs to have a sufficient stall-margin along the entire flight envelope to ensure a stable operating range. This work is focused on creating an efficient design method which produces a trade-off between high stall margin and high efficiency. The design method is based on an automatic multiobjective optimization process divided into two steps. In the first step, 2D blade profiles are optimized where both efficiency and stall margin are considered. Once the optimization is finished the selected profiles are stacked together to be further optimized in 3D. When going to the second step, i.e. a 3D optimization, one can focus on a smaller set of design variables thereby reducing the time to get what is considered the optimal solution. The results show that it is possible to rate designs with potential of having high stall margin and high efficiency both in the 2D and 3D optimization. The main contribution in this work is the design method, which offers an efficient way of designing robust blades where the designer can decide the best trade off between stall margin at part speed and efficiency at the design point. Copyright © 2012 by ASME.
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| 24. |
- Ellbrant, Lars R, 1984, et al.
(författare)
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Predictive Capability of CFD Models for Transonic Compressor Design
- 2014
-
Ingår i: ASME Turbo Expo: Turbine Technical Conference and Exposition, Dusseldorf, GERMANY. JUN 16-20, 2014. - 9780791845615 ; , s. V02BT39A041-
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Konferensbidrag (refereegranskat)abstract
- The primary focus of this work is to validate a CFD model intended to be used for transonic compressor design purposes. This design model includes a coarse grid using wall functions and mixing planes at interfaces connecting the compressor components. The computations are compared with experimental data from the transonic highly loaded 1.5 stage compressor test case Hulda. Additional comparisons are done with higher complexity CFD models accounting for the rotor-stator interaction. The performance of Hulda has been measured with both a small and a large tip clearance. These two configurations are used to investigate the necessity of resolving the tip clearance gap in the design model. The comparison is presented in terms of the overall performance at two rotational speeds as well as radial distribution of total pressure and total temperature at stations downstream of the rotor The predictive capability at these speeds is assessed in terms of mass flow, pressure ratio and efficiency. Furthermore, the response of the predicted radial flow distributions with respect to the throttle setting along the two rotational speeds is qualitatively compared with the measurements. The validation of the small tip clearance test shows that the design model, with or without tip gap modeling, is in good agreement with the measurements at both speeds. As for the large tip clearance test a design model resolving the tip clearance was able to predict trends but the penalty related to the increased tip gap was overestimated compared to the measured.
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| 25. |
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| 26. |
- Gustafsson, Bernhard, 1971, et al.
(författare)
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Nozzle throat optimization for supersonic jet noise reduction
- 2012
-
Ingår i: 50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition 2012. - Reston, Virigina : American Institute of Aeronautics and Astronautics.
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Konferensbidrag (refereegranskat)abstract
- Noise from engines that operate at supersonic conditions, especially high performance military aircraft, often utilize a converging-diverging nozzle with variable area control. This design usually includes a sharp nozzle throat which creates internal shock formation. Turbulent structure interaction with these shocks results in additional noise components other than turbulent mixing noise to be introduced to the jet noise spectrum. The present study investigates how weakening the internal shocks affects the flow and acoustics of a Mach 1.6 jet. RANS simulations were used to minimize internal shock formation and optimize the flow contours of the converging portion and throat of a C-D nozzle. A response surface methodology was used to evaluate 3000 possible designs using the RANS results as model inputs. An experimental investigation was conducted with a splined nozzle design that is virtually free of internal shocks. The flow field was measured using PIV for comparison with RANS and LES. Mean velocity and turbulence was captured well by the computations for the sharp throat and splined nozzles. Although the throat shocks were nearly eliminated, the overall shock strength was relatively unchanged. Far-field acoustic results showed little difference at thrust matched conditions since the overall shock strength was unchanged. The nozzle performance is greatly improved through throat optimization, providing equivalent thrust with 4% less pressure with no acoustic penalty.
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| 27. |
- Hafsteinsson, Haukur, 1984, et al.
(författare)
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Active Suppression of Supersonic Jet Noise Using Pulsating Micro-Jets
- 2012
-
Ingår i: 50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition, Nashville, 9-12 January 2012. - Reston, Virigina : American Institute of Aeronautics and Astronautics.
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Konferensbidrag (refereegranskat)abstract
- Noise suppression devices on military jet engines are motivated by the need to reduce community noise aswell as the acoustic load on airfield personnel during peacetime operation. They may also reduce problemswith sonic fatigue on the aircraft. Micro-jets have previously been shown as a promising tool for active noisesuppression. In the work presented here, compressible LES simulations have been done for slightly overexpandedconical C-D nozzle with a Mach number of 1.58 at NPR = 4.0 and a free stream flow Mach numberof 0.1. Two microjet configurations have been simulated. One with steady-state injection and an other withpulsating trailing-edge injection having a maximum mass flow-rate of mi/mj = 1.6%. The acoustic field isexpanded to the far field using the Kirchhoff integral method. The effect of injection frequency and pulsationcharacteristics on the flow-field and the radiated sound is investigated. Comparison is made between the LESand simulations and experiments for the steady-state and no injection cases and shows excellent agreement forthe screech tone frequency and the predictided OASPL is within 2 dB deviation from the measurements. Thepulsating injection cases investigated show that the frequency spectrum and the noise levels are sensitive to theinjection frequency as well as pulsation characteristics. It is shown that steady-state injection and pulsating injectionof equal max mass flow result in comparable reduction in terms of OASPL. The latter, however, comeswith the penalty of increased noise for the upstream observers.
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| 28. |
- Hafsteinsson, Haukur, 1984, et al.
(författare)
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Near-field and far-field spectral analyzis of supersonic jet with and without fluidic injection
- 2014
-
Ingår i: 52nd AIAA Aerospace Sciences Meeting - AIAA Science and Technology Forum and Exposition, SciTech 2014; National Harbor, MD; United States; 13 January 2014 through 17 January 2014. - Reston, Virginia : American Institute of Aeronautics and Astronautics. - 9781624102561
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Konferensbidrag (refereegranskat)abstract
- In the presented study the time-dependent flow features of a supersonic jet with and without steady microjet injection are investigated. The flow field is sampled at various axial and radial locations in the supersonic region and its near surroundings. The jet is emitted from a sharp-throat converging diverging nozzle operated at a nozzle pressure ratio (NPR) of 4.0, which gives a jet exitMach number of M = 1.56 and a Reynolds number of Re = 2.46×106 based on the jet exit diameter. Large Eddy Simulation (LES) is used to obtain the fully three dimensional instantenous turbulent flow field and the Kirchhoff surface integral method is applied to obtain the far-field radiated noise. Both the near-field flow dynamics and the far-field noise obtained from the LES are in good agreement with experimental data. The noise components in the far-field noise are identified and compared with the spectra obtained from the probe-locations within the jet. The effect of micro-jet injection on the spectral characteristics within the jet and the far-field noise is analyzed. The screech tone appearing in the far-field noise is clearly established also in the jet-plume. Two point cross-correlations within and outside the supersonic region of the jet-plume revealed two types of moving phenomenon. These where found to be turbulent structures and acoustic waves. The odd thing at first sight was that the acoustic waves appeared to be traveling upstream within the supersonic region, which sounds contradictory. However, it was showed that the acoustic wave was traveling in the form of a helical mode which allows the phase velocity of the pressure wave to be higher than the flow velocity, even at supersonic flow speeds. The fluidic injection was shown to disrupt and weaken the helical pattern which resulted in a lower far-field screech tone noise. Upon sufficient dissipation of the injection, a few nozzle diameters downstream of the nozzle exit, the helical pattern picks up strength again. However, the feed-back loop mechanism associated with the screech tone is still disabled.
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| 29. |
- Hafsteinsson, Haukur, 1984, et al.
(författare)
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Supersonic Jet Excitation using Flapping Injection
- 2013
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Ingår i: American Physical Society, Division of Fluid Dynamics 66th Annual Meeting, November 24–26, 2013 Pittsburgh, Pennsylvania.
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Konferensbidrag (övrigt vetenskapligt/konstnärligt)abstract
- Supersonic jet noise reduction is important for high speed military aircraft. Lower acoustic levels would reduce structural fatigue leading to longer lifetime of the jet aircraft. It is not solely structural aspects which are of importance, health issues of the pilot and the airfield personnel are also very important, as high acoustic levels may result in severe hearing damage. It remains a major challenge to reduce the overall noise levels of the aircraft, where the supersonic exhaust is the main noise source for near ground operation. Fluidic injection into the supersonic jet at the nozzle exhaust has been shown as a promising method for noise reduction. It has been shown to speed up the mixing process of the main jet, hence reducing the kinetic energy level of the jet and the power of the total acoustic radiation. Furthermore, the interaction mechanism between the fluidic injection and the shock structure in the jet exhaust plays a crucial role in the total noise radia- tion. In this study, LES is used to investigate the change in flow structures of a supersonic (M=1.56) jet from a converging-diverging nozzle. Six fluidic actuators, evenly distributed around the nozzle exit, inject air in a radial direction towards the main flow axis with a total mass flow ratio of 3%. Steady injection is compared with flapping injection. With flapping injection turned on, the injection angle of each injector is varied sinusoidally in the nozzle exit plane and the variation is the same for all injectors. This fluid dynamics video is submitted to the APS DFD Gallery of Fluid Motion 2013 at the 66 the Annual Meeting of the American Physical Society, Division of Fluid Dynamics (24-26 November, Pittsburgh, PA, USA).
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| 30. |
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| 31. |
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| 32. |
- Lejon, Marcus, 1986, et al.
(författare)
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Simulation of Tip-Clearance Effects in a Transonic Compressor
- 2015
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Ingår i: ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. - 9780791856635 ; 2A
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Konferensbidrag (refereegranskat)abstract
- Tip clearance flow can have a large impact on compressor performance and is therefore important to consider in compressor design. A wide range of computational approaches can be used for modeling tip clearance flow. The different modeling strategies reflect compromises between accuracy and computational time. The present study investigates a large tip gap configuration (1.65% of the rotor tip chord) of the highly loaded transonic 1.5 stage compressor Hulda. The aim of the study is to evaluate different modeling approaches and find a method that can predict the effects of a large tip clearance on compressor performance. The modeling approaches investigated are the SAS-SST model in CFX and the k-ε turbulence model using a wall function and Chien’s low-Reynolds model in the in-house CFD solver VolSol++. Results obtained using the chosen solvers and turbulence models are compared with experimental data, and all approaches are shown to predict the overall performance trends. However, the turbulence kinetic energy in the tip clearance flow and the trajectory of the vortex are shown to vary depending on the method. The SAS-SST model predicts the performance well in terms of total pressure ratio, polytropic efficiency as well as radial distribution of total pressure downstream the rotor. Based on the results from the study, SAS-SST is concluded to be a good candidate for detailed studies of transonic compressors with large tip gaps.
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| 33. |
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| 34. |
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| 35. |
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| 36. |
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| 37. |
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| 38. |
- Olander Burak, Markus, 1979, et al.
(författare)
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Experimental and Numerical Investigation of a Supersonic Convergent-Divergent Nozzle
- 2012
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Ingår i: AIAA Journal. - 1533-385X .- 0001-1452. ; 50:7, s. 1462-1475
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Tidskriftsartikel (refereegranskat)abstract
- Experiments and large eddy simulations of a conical convergent divergent nozzle, similar to the nozzles employed on high-performance tactical jets, have been carried out. The nozzle has been studied at a slightly overexpanded condition with a nozzle pressure ratio of 4.0 (a design nozzle pressure ratio 4.1). The primary nozzle is surrounded by a secondary jet operated at three different conditions: Mach 0.1, 0.3, and 0.8. This secondary jet provides a rough forward-flight simulation over a limited axial range. It has been found that the shock pattern in the jet plume is directly influenced by the secondary jet Mach number due to the outer geometric shape of the nozzle. As the flow follows the boat-tail, it locally increases the pressure, making the nozzle operate in a more overexpanded condition. Large eddy simulations of a primary jet with forward flight fully simulated are in very good agreement with the experiments regarding both flow and acoustics. A numerical investigation of screech mode character for the case with the lowest secondary flow Mach number has been made. This shows that the screech mode is mainly built up by two counter-rotating modes with a tangential mode number of I. In the numerical simulations, additional flow-adaptive dissipation has been added in order to stabilize the solution around the shocks. The effect on the flow and acoustics is found to be very small if a shock sensor combined with a damping threshold is used.
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| 39. |
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| 40. |
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| 41. |
- Strawbridge, Rona J., et al.
(författare)
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Identification of a novel proinsulin-associated SNP and demonstration that proinsulin is unlikely to be a causal factor in subclinical vascular remodelling using Mendelian randomisation
- 2017
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Ingår i: Atherosclerosis. - : Elsevier BV. - 0021-9150 .- 1879-1484. ; 266, s. 196-204
-
Tidskriftsartikel (refereegranskat)abstract
- Background and aims: Increased proinsulin relative to insulin levels have been associated with subclinical atherosclerosis (measured by carotid intima-media thickness (cIMT)) and are predictive of future cardiovascular disease (CVD), independently of established risk factors. The mechanisms linking proinsulin to atherosclerosis and CVD are unclear. A genome-wide meta-analysis has identified nine loci associated with circulating proinsulin levels. Using proinsulin-associated SNPs, we set out to use a Mendelian randomisation approach to test the hypothesis that proinsulin plays a causal role in subclinical vascular remodelling.Methods: We studied the high CVD-risk IMPROVE cohort (n = 3345), which has detailed biochemical phenotyping and repeated, state-of-the-art, high-resolution carotid ultrasound examinations. Genotyping was performed using Illumina Cardio-Metabo and Immuno arrays, which include reported proinsulin-associated loci. Participants with type 2 diabetes (n = 904) were omitted from the analysis. Linear regression was used to identify proinsulin-associated genetic variants.Results: We identified a proinsulin locus on chromosome 15 (rs8029765) and replicated it in data from 20,003 additional individuals. An 11-SNP score, including the previously identified and the chromosome 15 proinsulin-associated loci, was significantly and negatively associated with baseline IMTmean and IMTmax (the primary cIMT phenotypes) but not with progression measures. However, MR-Eggers refuted any significant effect of the proinsulin-associated 11-SNP score, and a non-pleiotropic SNP score of three variants (including rs8029765) demonstrated no effect on baseline or progression cIMT measures.Conclusions: We identified a novel proinsulin-associated locus and demonstrated that whilst proinsulin levels are associated with cIMT measures, proinsulin per se is unlikely to have a causative effect on cIMT.
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| 42. |
- Wollblad, Christian, 1978, et al.
(författare)
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Investigation of large scale shock movement in transonic flow
- 2010
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Ingår i: International Journal of Heat and Fluid Flow. - : Elsevier BV. - 0142-727X. ; 31:4, s. 528-535
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Tidskriftsartikel (refereegranskat)abstract
- Large eddy simulations were made of transonic flow over a two-dimensional bump where shock wave/turbulent boundary layer interaction takes place. Different flow conditions were investigated to find conditions for large scale shock movement. The innermost part of the shock was found to be moving for sufficiently strong shocks. None of the cases display large scale movement of the whole shock. (c) 2010 Elsevier Inc. All rights reserved.
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| 43. |
- Wollblad, Christian, 1978, et al.
(författare)
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Large Eddy Simulation of Transonic Flow with Shock Wave/Turbulent Boundary Layer Interaction
- 2006
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Ingår i: AIAA Journal. - : American Institute of Aeronautics and Astronautics (AIAA). - 1533-385X .- 0001-1452. ; 44:10, s. 2340-2353
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Tidskriftsartikel (refereegranskat)abstract
- Large eddy simulation was made of transonic flow over a two-dimensional bump where shock wave turbulent boundary layer interaction takes place. Grid refinement and the effect of the domain width were investigated. Special care was taken to ensure physically correct inlet boundary conditions. The shock wave turbulent boundary layer interaction induces strong separation of the boundary layer and events such as bursting events in the incoming boundary layer, and creation of large flow scale structures behind the shock are detected. However, the shock features no large scale movement even though, according to several sources, the prerequisites for such movement are fulfilled.
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| 44. |
- Wollblad, Christian, 1978, et al.
(författare)
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Semi-implicit Preconditioning for Wall-bounded Flow
- 2004
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Ingår i: 34th AIAA Fluid Dynamic Conference and Exhibit. - Reston, Virigina : American Institute of Aeronautics and Astronautics. - 9781624100314
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Konferensbidrag (övrigt vetenskapligt/konstnärligt)abstract
- Large eddy simulations (LES) of high speed compressible viscous flow require high space resolution as well as short time steps if explicit time stepping is used. A semi-implicit preconditioning scheme specially suitable for wall-bounded flow is developed which, when combined with an explicit time stepping algorithm, makes much larger time steps possible than with the explicit scheme alone. The method is validated for developing turbulent channel flow and applied to a transonic flow with shock wave turbulent boundary layer interaction. In both cases, the computational time saved is measured. Finally possible developments of the method are discussed.
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| 45. |
- Yao, Huadong, 1982, et al.
(författare)
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Aero-acoustic analysis and assessment of conceptual low-noise airframe configurations
- 2015
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Ingår i: 2nd GRAIN2 Open Workshop and Mid-term Review 'Greening Aviation – A Global Challenge' Xi'an, China, 5th - 8th May 2015.
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Konferensbidrag (övrigt vetenskapligt/konstnärligt)abstract
- High-lift (HL) configuration represents a significant part of airframe noise. Two different low-noise HL designs have been assessed with respect to a baseline HL configuration, which is a full-scale regional aircraft wing with a double-slotted part-span flap. The first design uses a single slotted flap. The second design, with a single slotted flap, has invoked a Krueger flap. Hybrid CFD/CAA analysis has been performed using acoustic analogies (the Kirchhoff, FW-H and Curle methods) based on hybrid RANS/-LES computations. It is shown that the Krueger flap has the best aerodynamic and aero-acoustic performances, which haves led to a noise reduction by 6-10dB. One of energetic airframe noise sources is associated to turbulent flow motion over the flap side-edge (FSE). To reduce FSE noise, a fence is attached onto the side edge. In the analysis of this configuration, an effective stochastic noise generation and radiation (SNGR) method was used, which formulates the noise sources based on the RANS solutions computations. The Lighthill acoustic analogy and boundary element method (BEM) in frequency domain have been employed for noise prediction. The analysis shows that the fence is able to effectively reduce noise by about 7dB. Another dominant airframe noise source is the turbulent flow induced by a landing gear (LG). Four low-noise conceptual designs are investigated for a full-scale main landing gear (MLG). The LG-bay door is opened to explore its noise scattering effect. Several different concepts measures for noise reduction have been investigated, including fairing, shallow LG-bay and acoustic liners. The aero-acoustic analysis has been demonstrated that, with a lined bay rear wall, the noise level has been was reduced by 1.817dBA. In the presentation, the aero-acoustic analysis and assessment on the aforementioned airframe configurations will be reporteds and discussed.
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| 46. |
- Yao, Huadong, 1982, et al.
(författare)
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Aeroacoustic Assessment of Conceptual Low-Noise High-Lift Wing Configurations
- 2012
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Ingår i: 50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. - Reston, Virigina : American Institute of Aeronautics and Astronautics.
-
Konferensbidrag (refereegranskat)abstract
- Aeroacoustics performance is assessed for three conceptual low-noise and high-lift wings using the surface integral methods of acoustic analogy, including the Kirchhoff method, the Ffowcs Williams and Hawkings method and the Curle method. A new way is proposed to define the integral surface that encloses the core flow region for the first two methods on the basis of the vorticity magnitudes. Both of the first two methods are used to compute the noise generated by the core region of the flow, which is responsible for most of the noise generation. The results obtained by the FWH method support those given by the Kirchhoff method. The Curle analogy approach is adopted to calculate the noise by the boundary layer flow attached on the walls inside the core flow region. Through comparison of the results of the three approaches, the levels of the contribution to the noise by the volume flow and the boundary layers are studied for the configurations. Finally, the mechanism of efficiency on the noise reduction is presented for the three configurations.
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| 47. |
- Yao, Huadong, 1982, et al.
(författare)
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Assessment of Flap Side-Edge Fence Noise using SNGR Method
- 2016
-
Ingår i: 21st AIAA/CEAS Aeroacoustics Conference, 2015; Dallas; United States; 22 June 2015 through 26 June 2015. - Reston, Virginia : American Institute of Aeronautics and Astronautics.
-
Konferensbidrag (refereegranskat)abstract
- In this paper, the noise reduction efficiency of a flap side-edge fence is revisited and investigated using the stochastic noise generation and radiation (SNGR) method coupled with the Reynolds averaged Navier-Stokes equations (RANS). The configurations are of full scale. The baseline configuration is slotted with double flaps. The fence is attached onto both side edges of the flaps for the purpose of suppressing turbulent vortices induced by the side edge. Furthermore, the efficiency of the SNGR method is assessed as a fast prediction approach for the flap side-edge noise. The noise generation in the SNGR method employs the stochastic model to construct a synthetic turbulent field and takes advantage of the acoustic analogy to formulate the noise sources. The wave equation for the noise radiation is computed by means of the boundary element method (BEM) in frequency space. The BEM takes into account the noise scattered by the surfaces. The fences are found to be efficient for depressing the wake produced by the side edges of the double slotted flaps. Since the energetic vortices contained in the wake are the major contributor of the noise generation, the fence enable an effective noise reduction, particularly in the low frequencies below 200 Hz.
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| 48. |
- Yao, Huadong, 1982, et al.
(författare)
-
Noise radiated by low-Reynolds number flows past a hemisphere at Ma=0.3
- 2017
-
Ingår i: Physics of Fluids. - : AIP Publishing. - 1070-6631 .- 1089-7666. ; 29:7, s. Article no 076102 -
-
Tidskriftsartikel (refereegranskat)abstract
- Flows past a hemisphere and their noise generation are investigated at the Reynolds numbers (Re) of 1000 and 5000. The Mach number is 0.3. The computational method of the flows is large eddy simulation. The noise is computed using the Ffowcs Williams and Hawkings Formulation 1C (F1C). An integral surface with an open end is defined for the F1C. The end surface is removed to reduce the numerical contamination that is introduced by vortices passing this surface. However, the contamination cannot be completely reduced since a discontinuity of the flow quantities still exists at the open surface boundary. This problem is solved using a surface correction method, in which a buffer zone is set up at the end of the integral surface. The transformation of flow structures due to Re is explored. Large coherent structures are observable at low Re, whereas they diminish at high Re. A large amount of small-scale turbulent vortices occur in the latter case. It is found that these characteristics of the flows have an important influence on the noise generation in regard to the noise spectra. In the flows studied in this work, the fluctuating pressure on the walls is a negligible noise contributor as compared with the wake.
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| 49. |
- Yao, Huadong, 1982, et al.
(författare)
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SNGR Method for Assessing Noise Induced by Flap Side-Edges
- 2013
-
Ingår i: NUMECA International 2013 Users Meeting.
-
Konferensbidrag (övrigt vetenskapligt/konstnärligt)abstract
- We investigated the stochastic noise generation and radiation (SNGR) method [1] for assessing acoustic performances of flap side-edges, with the assistance of FINE/Acoustics produced by NUMECA. The flap side-edges include the baseline configuration, and a configuration that is formed by attaching a noise-reduction facility onto the baseline configuration.In the SNGR method, the time-averaged field required by this method comes from a Reynolds-averaged Navier-Stokes (RANS) computation. The kinetic energy and dissipation rate are the basic variables for constructing a synthetic turbulent field on the basis of the stochastic method. This turbulent field transforms to the noise sources using the Lighthill acoustic analogy. After that, the noise radiation is computed in the frequency space using the boundary element method (BEM). Since FINE/Acoustics include packages for the SNGR method and BEM, it is an ideal choice for realizing the computational procedure.FINE/Acoustics is proven for its sufficient capability on predicting the flap side-edge noise.
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| 50. |
- Yao, Huadong, 1982, et al.
(författare)
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Surface integral analogy approaches for predicting noise from 3D high-lift low-noise wings
- 2014
-
Ingår i: Acta Mechanica Sinica/Lixue Xuebao. - : Springer Science and Business Media LLC. - 0567-7718. ; 30:3, s. 326-338
-
Tidskriftsartikel (refereegranskat)abstract
- Three surface integral approaches of the acoustic analogies are studied to predict the noise from three conceptual configurations of three-dimensional high-lift low-noise wings. The approaches refer to the Kirchhoff method, the Ffowcs Williams and Hawkings (FW-H) method of the permeable integral surface and the Curle method that is known as a special case of the FW-H method. The first two approaches are used to compute the noise generated by the core flow region where the energetic structures exist. The last approach is adopted to predict the noise specially from the pressure perturbation on the wall. A new way to construct the integral surface that encloses the core region is proposed for the first two methods. Considering the local properties of the flow around the complex object-the actual wing with high-lift devices-the integral surface based on the vorticity is constructed to follow the flow structures. The surface location is discussed for the Kirchhoff method and the FW-H method because a common surface is used for them. The noise from the core flow region is studied on the basis of the dependent integral quantities, which are indicated by the Kirchhoff formulation and by the FW-H formulation. The role of each wall component on noise contribution is analyzed using the Curle formulation. Effects of the volume integral terms of Lighthill's stress tensors on the noise prediction are then evaluated by comparing the results of the Curle method with the other two methods.
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