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检索条件"主题词=Computational Fluid Dynamics Code"

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Powered Descent Aerodynamics for Low and Mid Lift-to-Drag Human Mars Entry, Descent, and Landing Vehicles

Powered Descent Aerodynamics for Low and Mid Lift-to-Drag Hu...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech)

作者： Korzun, Ashley M. Bibb, Karen L. Canabal, Francisco Childs, Robert E. Tang, Chun Y. Rizk, Yehia M. Van Norman, John W. Tynis, Jake A. NASA Langley Res Ctr Atmospher Flight & Entry Syst Branch Hampton VA 23681 USA NASA Langley Res Ctr Aerothermodynam Branch Hampton VA 23681 USA NASA Marshall Space Flight Ctr Aerosci Branch Huntsville AL 35812 USA Sci & Technol Corp Moffett Field CA 94035 USA NASA Ames Res Ctr Aerothermodynam Branch Moffett Field CA 94035 USA Analyt Mech Associates Atmospher Flight & Entry Syst Branch Hampton VA 23666 USA

ISBN: (数字)9781624105951

ISBN: (纸本)9781624105951

NASA human-scale Mars entry, descent, and landing vehicles currently under consideration have system requirements beyond the current state of the art. Specifically, heritage approaches for deceleration using parachutes and initiating engines at subsonic speeds are not adequate to land these large vehicles (50-60 metric ton entry mass) precisely (within 50 meters of a surface target). Therefore, the present approach for landing humans on Mars utilizes entry vehicles with large aerodynamic surface areas that transition to the descent and landing phase by initiating retropropulsion at supersonic speeds. Vehicle performance, including stability and control, depends on accurate characterization of the aerodynamic-propulsive interference effects, especially at engine initiation, when the vehicles are transitioning from an entry attitude to that suitable for descent and landing. This paper summarizes the human Mars entry vehicle configurations, retropropulsion integration, performance assumptions, and expected flight environment at engine initiation. Cases spanning supersonic, transonic, and subsonic conditions during powered descent are analyzed using four computational fluid dynamics codes, all of which have been previously applied to supersonic retropropulsion at varying scales and conditions. The results of the analysis are used to develop an aerodynamic-propulsive interference model that has since been applied in trajectory simulations to augment the fidelity of human-scale Mars entry vehicle performance.

关键词： Aerodynamic Interference Hypersonic Inflatable Aerodynamic Decelerator computational fluid dynamics code Supersonic Retropropulsion Supersonic Speed Reynolds Averaged Navier Stokes computational fluid dynamics Angle of Attack Turbulence Models Detached Shock

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Coupled Flight Mechanics based on Reduced Order Models for use in Tiltrotor Stability Analysis

Coupled Flight Mechanics based on Reduced Order Models for u...

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AIAA Aviation Forum and Exposition

作者： Wakefield, J. E. G. Cassillas, S. M. Gaitondet, A. L. Jones, D. P. Univ Bristol Dept Aerosp Engn Bristol BS8 1TR Avon England

ISBN: (数字)9781624105890

ISBN: (纸本)9781624105890

The development of two coupled stability models is presented. The models are based on the reduced order modelling of an Unsteady Vortex Lattice Method (UVLM) computational fluid dynamics code that provides an explicit wake model, coupled with classical rigid body equations of motion. These have been derived to include all 12 states to accommodate a full dynamic wake response. The model retains the entirety of the linear dynamics of the reduced order aerodynamics in the stability analysis, neglecting the need for aerodynamic derivatives. This allows for rapid and accurate force response and flow solution reconstruction. A new Linear Time Periodic balanced model reduction is developed, as well as an ensemble-averaging approach which may allow for a much lower computational cost than the full periodic stability. Further investigation is required to prove the equivalence of the two approaches. Validation studies have been performed to show the accuracy offered by the coupled model. An investigation into the stability of the XV-15 in comparison to an equivalent classical model was performed, showing similar eigenvalue distributions but with the ensemble-average coupled model producing larger magnitude instabilities. An investigation into the reconstruction of flow solutions during unstable motions was performed, allowing significant insight into the aerodynamics and showing potential for use as a design tool.

关键词： Aircraft Flight Mechanics Tiltrotor Aircraft Aerodynamic Force Vector computational fluid dynamics computational fluid dynamics code Eigensystem Realization Algorithm Reynolds Averaged Navier Stokes Controllability Gramian Singular Value Decomposition Aircraft Configurations

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Resolvent Analysis using Subspace Algorithms for Aerodynamic Applications

Resolvent Analysis using Subspace Algorithms for Aerodynamic...

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AIAA Aviation Forum

作者： Vevek, U. S. Timme, Sebastian Univ Liverpool Sch Engn Liverpool L69 3GH Merseyside England

ISBN: (数字)9781624107047

ISBN: (纸本)9781624107047

Resolvent analysis has become ubiquitous in aerodynamic studies to identify strong modal behaviour through optimal forcing/response pairs and their associated gains. Such dominant modes can be used, for instance, for the understanding of flow physics and the construction of reduced order models. When interest is in non-canonical three-dimensional test cases of practical interest, earlier computational limitations, when solving the underlying truncated singular value decomposition for a small number of the most dominant modes using matrix forming methods with direct matrix factorisation, have been overcome with either efficient time-stepping methods or sparse iterative solver technology within linear harmonic incarnations of production computational fluid dynamics codes. For the latter iterative methods, besides the so-called randomised singular value decomposition, either single-vector iterations with deflation to identify next-in-line optimal modes or subspace methods with random starting vectors have shown great potential. When using subspace methods, the optimal modes are often computed by adapting algorithms designed for solving eigenvalue problems, but these algorithms yield either the forcing or response modes leaving the other to be determined separately incurring additional computational cost. Herein we will investigate two specialised algorithms for singular value decomposition that allows computation of both sets of modes simultaneously. One is a Krylov subspace method referred to as the Krylov-Golub-Kahan algorithm that benefits from a simpler and more robust deflation procedure compared to the widely used implicitly restarted Arnoldi method. The other is an inverse subspace iteration method referred to as the inexact subspace iteration algorithm that uses cheap, low-accuracy linear solutions. The subspace methods were applied for a two-dimensional circular cylinder and the NASA Common Research Model. Both subspace methods were able to compute the three m

关键词： Common Research Model Reduced Order Modelling computational fluid dynamics code Reynolds Averaged Navier Stokes Lift Coefficient Aerodynamic Loads Flow Conditions Vortex Shedding Eddy Viscosity European Transonic Windtunnel

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Visualizations of Hypersonic Boundary-Layer Transition on a Variable Bluntness Cone

Visualizations of Hypersonic Boundary-Layer Transition on a ...

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AIAA Aviation Forum and Exposition

作者： Kennedy, Richard E. Jagde, Eric K. Laurence, Stuart J. Jewell, Joseph S. Kimmel, Roger L. Univ Maryland Aerosp Engn College Pk MD 20742 USA US Air Force Res Lab AFRL RQHF Wright Patterson AFB OH 45433 USA

ISBN: (数字)9781624105890

ISBN: (纸本)9781624105890

The effect of nosetip bluntness on instability development within the hypersonic boundary layer on a 7 degrees half-angle cone is experimentally studied using high-speed calibrated schlieren visualizations and PCB pressure measurements. Experiments are performed in the Air Force Research Lab (AFRL) Mach-6 Ludwieg Tube at unit Reynolds numbers ranging from 13.72 x 10(6) to 22.71 x 10(6) per meter. The visualizations capture second-mode instability waves within the boundary layer for nosetip radii up to 2.54 mm. Quantitative data on second-mode instability waves including disturbance frequencies and integrated amplification rates (N factors) are calculated from the visualizations and PCB pressure measurements, and compared to parabolized stability equation (PSE) computations. The visualized instability features change dramatically for nosetip radii larger than 2.54mm where nonmodal growth is expected. Unsteady features that extend above the edge of the boundary layer are observed and the visualizations are used to track individual instances throughout their development process. PCB measurements reveal high-frequency (150 - 250 kHz) pressure disturbances associated with the features.

关键词： Boundary Layer Transition Parabolized Stability Equations Freestream Mach Number Air Force Research Laboratory Stagnation Temperature Air Forces Short Time Fourier Transform Hypersonic Vehicles Steady Flow computational fluid dynamics code

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Numerical Simulation of An Aspect Ratio 13.5 Common Research Model with Trailing Edge Mini-Flaps

Numerical Simulation of An Aspect Ratio 13.5 Common Research...

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AIAA Aviation Forum

作者： Xiong, Juntao Nguyen, Nhan Bartels, Robert E. KBR Wyle Inc Intelligent Syst Div Moffett Field CA 94035 USA NASA Ames Res Ctr Intelligent Syst Div Moffett Field CA 94035 USA NASA Langley Res Ctr MS 340 Hampton VA 23681 USA

ISBN: (数字)9781624107047

ISBN: (纸本)9781624107047

This paper presents an aeroelastic modeling study of an aspect ratio 13.5 wind tunnel scale Common Research Model (CRM) with trailing edge mini-flaps. FUN3D is used for the aeroelastic steady and unsteady simulations at Mach 0.85 and dynamic pressure 230 psf. The aeroelastic simulation results are compared with the in-house developed VSPAERO aeroelastic model results. The incremental aerodynamic coefficients predicted by FUN3D and VSPAERO are close except when the mini-flap deflection angle is 10 degrees. The control reversal behavior is observed for the wing tip mini-flap. A series of dynamic pressure sweep simulations are performed to determine the control reversal onset point. Unsteady simulation results show that there is no buffet phenomenon observed at angle of attacks 0 degrees and 2 degrees when the mini-flap is deflected 10 degrees. The mini-flap oscillation simulations show the nonlinear behavior appears when the deflection angle is about 4 degrees.

关键词： Lift to Drag Ratio Aspect Ratio Common Research Model Numerical Simulation Lift Coefficient Aerodynamic Coefficients computational fluid dynamics code Pressure Coefficient Trailing Edge Flap Control Reversal

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Application of DeepOnet to model inelastic scattering probabilities in air mixtures

Application of DeepOnet to model inelastic scattering probab...

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AIAA Aviation Forum and Exposition

作者： Sharma, Maitreyee P. Venturi, Simone Panesi, Marco Univ Illinois Ctr Hyperson & Entry Syst Studies CHESS Urbana IL 61801 USA Univ Illinois Dept Aerosp Engn Talbot Lab 302 E104 S Wright St Urbana IL 61801 USA Univ Illinois Dept Aerosp Engn 306 Talbot Lab104 S Wright St Urbana IL 61801 USA

ISBN: (数字)9781624106101

ISBN: (纸本)9781624106101

Hypersonic conditions during (re-)entry flight demand accurate representation of the nonequilibrium flow physics to enable predictive models. However, the large number of possible kinetic mechanisms for even small molecular systems makes enumerating all the possible reaction coefficients and quantifying their uncertainties a computationally challenging task. Addressing this concern, in this work we develop surrogate models for molecular scattering or quasi-classical trajectory calculations. Deep Operator network (DeepOnet), recently developed my Lu et al. [1], is employed to model the inelastic reaction rates. The state-to-state (StS) rates are obtained as a function of temperature and properties of the molecular diatomic potential. The training states used by the NN are randomly sampled from groups constructed using a novel diatomic potential-based grouping strategy. As an example, the surrogate QCT model is tested for the O-2+O system due to its interest in hypersonic earth re-entry flight. We have demonstrated a reduction in the computational costs of QCT calculations by 95% while maintaining a good accuracy on the predicted StS rates. To further validate the rates, we carried out isothermal heat bath simulations at translational temperatures between 1,500 K and 20,000 K. The observables from the master equation analysis were reproduced within an accuracy of 5%.

关键词： Scattering Surrogate Model Earth Neural Networks Hypersonic Flows Vibrational Energy Direct Simulation Monte Carlo Computing FORTRAN computational fluid dynamics code

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Aeroelastic Simulation of Transonic Truss-Braced Wing Aircraft for Flight Dynamic Stability Analysis

Aeroelastic Simulation of Transonic Truss-Braced Wing Aircra...

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AIAA SciTech Forum

作者： Xiong, Juntao Nguyen, Nhan KBR Wyle Inc Intelligent Syst Div Moffett Field CA 94035 USA NASA Ames Res Ctr Intelligent Syst Div Moffett Field CA 94035 USA

ISBN: (数字)9781624107115

ISBN: (纸本)9781624107115

This paper presents aeroelastic simulation of the Mach 0.8 Transonic Truss-BracedWing (TTBW) aeroelastic model using CFD solver FUN3D for the flight dynamic and control stability analysis. A jig twist optimization is performed to obtain a flight optimized jig twist for the flexible Mach 0.8 TTBW model. The developed aeroelastic model steady-state stability derivatives with respect to the angle of attack, angle of sideslip, and airspeed are calculated with perturbations in the angle of attack, angle of sideslip, and Mach number, respectively. A series of unsteady simulations is conducted for the developed aeroelastic model in pitch oscillation at various reduced frequencies. The dynamic stability derivatives are estimated using a frequency domain estimation method. The control derivatives of the elevator, rudder, and ailerons of the developed aeroelastic model are estimated for control stability analysis.

关键词： Truss Braced Wing Aircraft Aerodynamic Coefficients Aerodynamic Performance Flight Control Surfaces Angle of Sideslip Aeroelastic Analysis computational fluid dynamics code Frequency Domain Trusses Control Stability

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Multi-point Jig Twist Optimization of Mach 0.745 Transonic Truss-Braced Wing Aircraft and High-Fidelity CFD Validation

Multi-point Jig Twist Optimization of Mach 0.745 Transonic T...

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AIAA Science and Technology Forum and Exposition (AIAA SciTech Forum)

作者： Nguyen, Nhan Xiong, Juntao Fugate, Jason NASA Ames Res Ctr Intelligent Syst Div Moffett Field CA 94035 USA KBR Wyle Inc Intelligent Syst Div Moffett Field CA 94035 USA

ISBN: (数字)9781624106095

ISBN: (纸本)9781624106095

This paper presents a multi-point jig twist optimization study of the Mach 0.745 Transonic Truss-Braced Wing (TTBW) aircraft using an in-house developed aero-structural analysis solver VSPAERO coupled to BEAM3D. A vortex-lattice model of the TTBW model is developed, and a transonic small disturbance /integral boundary layer correction method is implemented in the VSPAERO model to account for transonic and viscous flow effects. A correction method for the wing-strut interference aerodynamics is developed and applied to the VSPAERO solver. Also, a structural dynamic finite-element model of the TTBW aircraft is developed. This finite-element model includes the geometric nonlinear effect due to the tension in the struts which causes a deflection-dependent nonlinear stiffness. The VSPAERO model is coupled to the corresponding finite-element model to provide a rapid aero-structural analysis. The VSPAERO model has been validated against wind tunnel test data and FUN3D CFD simulation data. Three different flight conditions corresponding to Mach 0.745 at 40,000 ft are selected for the TTBW aircraft jig twist optimization to reduce the drag coefficients at the design and off-design cruise lift coefficients. The preliminary optimization results show that the TTBW aircraft with the optimized jig twist achieves a drag reduction of about 9 counts, 7 counts, and 8 counts corresponding to the lift coefficients 0.68, 0.73, and 0.78, respectively. A high-fidelity CFD solver FUN3D is used to validate the jig twist optimization.

关键词： Transport Aircraft Trusses computational fluid dynamics Aerodynamic Design Structural Finite Elements Aerodynamic Performance Transonic Wind Tunnel computational fluid dynamics code Flow Conditions Structural Analysis

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Simulations of Ducted and Coaxial Rotors for Air Taxi Operations

Simulations of Ducted and Coaxial Rotors for Air Taxi Operat...

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AIAA Aviation Forum and Exposition

作者： Diaz, Patricia Ventura Rubio, Rafael Caracuel Yoonz, Seokkwan NASA Ames Res Ctr Moffett Field CA 94035 USA NASA Ames Res Ctr Adv Supercomputing Div Moffett Field CA USA

ISBN: (数字)9781624105890

ISBN: (纸本)9781624105890

Effects of ducts on coaxial rotors in hover have been studied using a high-fidelity computational fluid dynamics code that solves the three-dimensional unsteady Navier-Stokes equations on overset structured grids. High-order accurate schemes, dual-time stepping, and the delayed detached-eddy simulation model have been employed. Performance of coaxial rotors with various duct placement has been assessed. A qualitative analysis of aeroacoustics has been conducted. Results for both counter-rotating and co-rotating rotors with or without ducts are presented.

关键词： Counter Rotating Rotor computational fluid dynamics Navier Stokes Equations Aeroacoustics computational fluid dynamics code Reynolds Averaged Navier Stokes Blade Vortex Interaction Urban Air Mobility Power Coefficient Rotor Systems

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Flow and Noise Predictions Around Tandem Cylinders using DDES approach with SU2

Flow and Noise Predictions Around Tandem Cylinders using DDE...

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AIAA Science and Technology (SciTech) Forum and Exposition

作者： Molina, Eduardo S. Alonso, Juan J. Zhou, Beckett Y. Righi, Marcello da Silva, Roberto Gil A. Stanford Univ Dept Aeronaut & Astronaut Stanford CA 94305 USA Tech Univ Kaiserslautern Chair Sci Comp Kaiserslautern Germany Zurich Univ Appl Sci Sch Engn Winterthur Switzerland Aeronaut Inst Technol Dept Aeronaut & Mech Sao Jose Dos Campos Brazil

ISBN: (数字)9781624105784

ISBN: (纸本)9781624105784

Flowand noise predictions of the tandem cylinder benchmark are performed using the open-source computational fluid dynamics code (CFD) Stanford University Unstructured (SU2). The numerical results were obtained using the Delayed Detached Eddy Simulation (DDES) approach with the new shear-layer adapted sub-grid length scale (SGS) for faster transition between RANS and LES. The FfowcsWilliams-Hawkings (FWH) analogy is used to propagate the pressure fluctuations to the farfield. Both flow and noise results are compared with experimental measurements from the Basic Aerodynamic Research Tunnel (BART) and Quiet Flow Facility (QFF) at NASA Langley Research Center. The compressible simulations are carried out on the mandatory 2-D grid from the ATAAC project with a spanwise extension of 3.0D, two different spanwise discretization were used to analyze the discretization effect. The comparison shows that the shear-layer adapted (Delta(SLA)) is less grid dependent and more accurate than the standard (Delta(max)) SGS.

关键词： Tandem Cylinders Laminar to Turbulent Transition Detached Eddy Simulation NASA Langley Research Center computational fluid dynamics code Kelvin Helmholtz Instability Vortex Structure Power Spectral Density Overall Sound Pressure Level Eddy Viscosity

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