| 1. |
- Shia-Hui, Peng, 1967, et al.
(author)
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CAA analysis of a wing section with flap side-edges based on hybrid RANS-LES computation
- 2015
-
In: 21st AIAA/CEAS Aeroacoustics Conference, AIAA Aviation. - Reston, Virginia : American Institute of Aeronautics and Astronautics. - 9781624103674
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Conference paper (peer-reviewed)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.
(author)
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A Study of Mach 0.75 Jets and Their Radiated Sound Using Large-Eddy Simulation
- 2004
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In: AIAA 2004-3024, proc. of 10th AIAA/CEAS Aeroacoustics Conference, May 10-12, 2004.
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Conference paper (peer-reviewed)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.
(author)
-
Effects of Inflow Conditions and Subgrid Model on LES for Turbulent Jets
- 2005
-
In: AIAA 2005-2925, proc. of 11th AIAA/CEAS Aeroacoustics Conference, May 23-25, 2005, Monterey, California.
-
Conference paper (peer-reviewed)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.
(author)
-
Investigation of an Isothermal Mach 0.75 Jet and its Radiated sound Using Large-Eddy Simulation and Kirchhoff Surface Integration
- 2005
-
In: International Journal of Heat and Fluid Flow. ; 26, s. 393-410
-
Journal article (peer-reviewed)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.
(author)
-
Large-Eddy Simulation of a Mach 0.75 Jet
- 2003
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In: AIAA 2003-3312, proc. of 9th AIAA/CEAS Aeroacoustic Conference, May 12-14, 2003.
-
Conference paper (peer-reviewed)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.
(author)
-
Large-Eddy Simulation of Subsonic Turbulent Jets and Their Radiated Sound
- 2005
-
In: AIAA Journal. ; 43:9, s. 1899-1912
-
Journal article (peer-reviewed)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.
(author)
-
LES Prediction of Flow and Acoustic Field of a Coaxial Jet
- 2005
-
In: AIAA 2005-2884, proc. of 11th AIAA/CEAS Aeroacoustics Conference, May 23-25, 2005, Monterey, California.
-
Conference paper (peer-reviewed)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.
(author)
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Acoustic Source Terms for the Linearized Euler Equations in Conservative Form
- 2005
-
In: AIAA journal. ; 43:4, s. 752-759
-
Journal article (peer-reviewed)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.
(author)
-
A Comprehensive Investigation of Pulsed Fluidic Injection for Active Control of Supersonic Jet Noise
- 2013
-
In: 51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition. - Reston, Virigina : American Institute of Aeronautics and Astronautics. ; 2013
-
Conference paper (peer-reviewed)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.
(author)
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Analysis of Supersonic Jet Thrust with Fluidic Injection
- 2014
-
In: 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
-
Conference paper (peer-reviewed)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.
(author)
-
Nozzle Throat Optimization on Acoustics and Performance of a Supersonic Jet
- 2012
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In: 18th AIAA/CEAS Aeroacoustics Conference (33rd AIAA Aeroacoustics Conference). - Reston, Virigina : American Institute of Aeronautics and Astronautics.
-
Conference paper (peer-reviewed)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.
(author)
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Assessment of high-lift concepts for a regional aircraft in the ALONOCO project
- 2012
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In: 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
-
Conference paper (peer-reviewed)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.
(author)
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Balancing efficiency and stability in the design of transonic compressor stages
- 2013
-
In: Proceedings of the ASME Turbo Expo 2013. - 9780791855232 ; 6 B, s. (Article number) GT2013-94838
-
Conference paper (peer-reviewed)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.
(author)
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CFD optimization of a transonic compressor using multiobjective GA and metamodels
- 2012
-
In: 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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Conference paper (peer-reviewed)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.
(author)
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Design of compressor blades considering efficiency and stability using CFD based optimization
- 2012
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In: ASME Turbo-expo 2012. - 9780791844748 ; 8:PARTS A, B, AND C, s. 371-382
-
Conference paper (peer-reviewed)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.
(author)
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Predictive Capability of CFD Models for Transonic Compressor Design
- 2014
-
In: ASME Turbo Expo: Turbine Technical Conference and Exposition, Dusseldorf, GERMANY. JUN 16-20, 2014. - 9780791845615 ; 2B, s. V02BT39A041-
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Conference paper (peer-reviewed)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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