Computational Fluid Dynamics (CFD) serves as an important tool for studying aerodynamics in turbomachinery and helps to achieve sustainable flying by improving engine designs. While experiments can assess engine's performance in real operation, simulation data complements the experimental tests by delivering a more comprehensive view of the flow. This paper studies an axial transonic compressor at different operating points through three-dimensional (3D) CFD simulations to understand flow physics in-depth. Different turbulence models, the hybrid LES/URANS and URANS, are carried out by PowerFLOW (R), a Lattice Boltzmann Method (LBM) based software. The flow field of the compressor, starting from 0D (compressor map), 1D (radial), 2D (cross-sectional), to 2D/3D (blade streamlines), is analyzed and validated with the experiment. The investigation shows the impact of turbulence modeling and how the simulations are sensitized to geometry modifications. Finally, stall analysis is performed by visualizing the propagation of the stall cell in 3D views and analyzing the pressure signal patterns. The validation of the high-fidelity CFD simulations provides insight into the compressor's behavior through different turbulence modeling approaches and represents a baseline for further stall investigation.

Assessing the ice accretion on aerodynamic surfaces due to aircraft operating at adverse weather conditions remains a challenging task. However, the impact on aerodynamic performance can be significant which makes ice accretion assessment an important aspect in the design and certification process. Simulation techniques can be employed to complement wind tunnel and flight tests, which are typically used to meet regulation requirements. This paper presents a novel, particle model based technique embedded in the Lattice Boltzmann Method (LBM) based flow solver PowerFLOW® to predict rime ice shapes. Validation results are presented for the NACA 0012 airfoil and for an inboard wing section of the NASA Common Research Model (CRM). Since realistic rime ice shapes can be obtained, scale-resolving simulations based on Very Large Eddy Simulations (VLES) can then be performed on the iced airfoil to assess the impact on aerodynamic performance and to analyze the turbulent flow field created by the ice topology.

Unsteady, scale-resolving aerodynamic simulations are presented that were conducted for the Fourth AIAA High-Lift Prediction Workshop on the high-lift version of the NASA Common Research Model using a Lattice Boltzmann Method. The method has been previously shown to capture the relevant flow physics of complex high-lift cases. Grid convergence studies for different angles of attack show the convergence behavior at distinct flow regimes and discuss challenges for time-accurate and scale-resolving simulations. Results of a flap deflection study show good qualitative agreement to experiments and demonstrate the capability to predict design changes, superior to standard RANS simulation tools as they are used in industry today. Dedicated maximum lift studies using free-air and wind tunnel boundary conditions, respectively, confirm the sensitivity of the aircraft model to wind tunnel installation effects. Free-air simulation results show good correlation to experiments within the linear range of the lift curve, but capture a critical break in the pitching moment curve later than seen in the experiments. The trend of an earlier pitch break, triggered by the wind tunnel floor boundary layer, is accurately captured in the simulations including the wind tunnel environment. This supports the notion that a precise representation of the experimental setup is important to accurately predict the high-lift performance of a test article.

This paper presents a sequential multi-physics/multi-scale simulation methodology for a landing gear that is retracting under aerodynamic loads. A main landing gear representative of a single-aisle aircraft has been designed and a mechanical model of the underlying kinematics has been established. This allows accessing the position of the landing gear at any intermediate position between fully deployed and fully retracted. Computational Fluid Dynamics (CFD) simulations are then executed for multiple landing gear positions at take-off flight conditions. The resulting aerodynamic loads are subsequently used inside a multi-body simulation of the landing gear. The engineering scenario considered here is to assess the actuator force that is required to retract the landing gear within a prescribed period of time. This force can then be used further in the design process, e.g. to dimension the actuator of the landing gear or the hydraulic system of the aircraft, which exemplarily shows the engineering value of the proposed workflow.

The flow through a 2.5D axial compressor cascade blade-row is simulated using a compressible Lattice-Boltzmann Method. The simulations are carried out for a nominal flow condition that yields a Reynolds number of 300,000 based on the axial chord and an inlet Mach number of 0.66. The simulations incorporate incoming wakes generated by moving bars with an equivalent reduced frequency of 1.9. A URANS simulation is carried out first, which reveals deficiencies in predicting time-averaged isentropic surface Mach numbers and wall shear stresses when compared against published DNS results. In order to increase the accuracy of the simulation, an implicit LES is run on a wall-resolved mesh. These results match the reference data because transition effects are correctly accounted for. A new, transition enabled hybrid RANS-LES approach is then carried out which shows that similar accuracy as the wall resolved LES can be obtained.

The NASA 2030 vision study illustrates the need for multi-physics simulations during the design process. Fluid-Structure Interaction (FSI) for airframes is an important example of multi-physics, multi-disciplinary simulation. The NASA Common Research Model (CRM) mock-up demonstrates the importance of FSI with displacements of up to 15.5mm, which is equivalent to 8.25% of the mean aerodynamic chord. The change in wing twist at cruise conditions yields values of up to -1.1° at the wingtip, which influences the local loading of the wing and the overall aerodynamic performance. The CRM mock-up serves as a validation case for the coupling of a Lattice-Boltzmann Method (LBM) based fluid solver with an implicit, structural mechanics solver. Since the coupling requires an iterative process to achieve the final wing deformation, an automated workflow was created that takes care of the run management, surface data mapping and geometry deformation. The predicted wing deformation is compared to the wing deformation measured in the wind tunnel for multiple angles of attack. The change in behavior of wing bending and twisting is captured correctly when transitioning from the linear to the non-linear aerodynamic regime. The accuracy lies within 0.25° for twisting and 5mm for bending with the largest differences occurring at low angles of attack.

This paper investigates the capability of the Lattice-Boltzmann Method (LBM) based solver SIMULIA PowerFLOW® to predict the flow separation at a wing-fuselage junction. Turbulence is resolved using a Very Large Eddy Simulation (VLES) turbulence modeling approach in the juncture region. Geometrical trips are applied to the wing and the fuselage to create dynamic boundary layers. The juncture flow separation is investigated at a negative angle of attack of -2.5° and results are compared to high fidelity wind tunnel measurements. A comparison of velocity profiles in the separation region with the previously published results at an angle of 5.0° are included as well.

This is an invited paper for the SciTech 2020 special session: NASA Juncture Flow. It summarizes Lattice Boltzmann Very Large Eddy Simulations of the NASA Juncture Flow Model using SIMULIA PowerFLOW. The approach followed here is to trigger dynamically resolved boundary layers in the region of interest by adding small geometrical tripping dots to the geometry. Two setup strategies were created where the first one follows a pragmatic and economical approach while the rationale of the second one is to replicate the experimental setup as accurately as possible.

Full-scale simulations of a Gulfstream G-III aircraft, performed in support of the NASA Acoustic Research Measurements flights, are presented to complement results discussed in earlier studies. The flow solver employed in those studies, Dassault Systemes’ lattice Boltzmann PowerFLOW®, was also used during this investigation to conduct time-dependent simulations of the entire aircraft in landing configuration with a fully dressed landing gear. The high-fidelity simulations, performed at a Mach number of 0.23 and a Reynolds number of 10.5 × 106 based on mean aerodynamic chord, captured all relevant airframe noise sources. The computations were used to assess the aeroacoustic performance of the main landing gear, with and without noise reduction fairings installed, of a G-III aircraft equipped with Adaptive Compliant Trailing Edge technology and conventional Fowler flaps. To facilitate comparison of predicted noise signatures with effective perceived noise levels obtained from flight test measurements, the “as-flown” nose landing gear geometry, missing in our earlier studies, was added to the simulated G-III aircraft configurations. The high fidelity, synthetic data were post-processed using a Ffowcs-Williams and Hawkings integral approach to estimate farfield acoustic behavior, with pressures on the model solid surface or pressure and velocity components on a permeable surface enveloping the acoustic near field used as input.

Microphone phased-array and pole-mounted microphone data gathered during the NASA Acoustics Research Measurements flight tests were used to benchmark results from companion full-scale aeroacoustics simulations. Conducted with the lattice Boltzmann solver PowerFLOW®, the simulations predicted the acoustic behavior of various tested aircraft configurations. Emphasis was placed on those flown during the third flight test - a Fowler flap-equipped Gulfstream G-III with and without noise abatement technology on the main landing gear. Direct comparisons between experimental and synthetic microphone phasedarray data were achieved by applying the same processing and deconvolution technique to both sets of data. To extend the validation of the computations to the metric used for noise certification, the Effective Perceived Noise Level, a high-fidelity digital model of the nose landing gear, which was excluded from earlier computations, was developed and integrated into the G-III aircraft geometry. The acoustic study presented here demonstrates that the simulated beamform maps and corresponding integrated farfield spectra accurately predict the locations and strengths of the prominent airframe noise sources present on the G-III aircraft.

Numerical simulations have been performed for a simplified high-lift (SHL) version of the Common Research Model (CRM) configuration, where the Fowler flaps of the conventional high-lift (CRM-HL) configuration are replaced by a set of simple hinged flaps. These hinged flaps are equipped with integrated modular active flow control (AFC) cartridges on the suction surface, and the resulting geometry is known as the CRM-SHL-AFC configuration. The main objective is to make use of AFC devices on the CRM-SHL-AFC configuration to recover the aerodynamic performance (lift) of the CRM-HL configuration. In the current paper, a Lattice Boltzmann method-based computational fluid dynamics (CFD) code, known as PowerFLOW R © is used to simulate the entire flow field associated with the CRM-SHL-AFC configuration equipped with several different types of AFC devices. The transonic version of the PowerFLOW R © code that has been validated for high speed flows is used to accurately simulate the flow field generated by the high-momentum actuators required to mitigate reversed flow regions on the suction surfaces of the main wing and the flap. The numerical solutions predict the expected trends in aerodynamic forces as the actuation levels are increased. More efficient AFC systems and actuator arrangements emerged based on the parametric studies performed prior to a Fall 2018 wind tunnel test. Preliminary comparisons of the numerical solutions for lift and surface pressures are presented here with the experimental data, demonstrating the usefulness of CFD for predicting the flow field and lift characteristics of AFC-enabled high-lift configurations.

Computational results for a full-scale simulation of a Gulfstream G-III aircraft are presented. In support of a NASA airframe noise flight test campaign, Exa Corporation’s lattice Boltzmann PowerFLOW® solver was used to perform time-accurate simulations of the flow around a highly detailed, full-scale aircraft model. Free-air boundary conditions were used at a Mach number of 0.23 and a Reynolds number of 10.5 × 10(exp 6) based on mean aerodynamic chord. This paper documents the simulation campaign for the baseline aircraft configuration at several flight conditions, including multiple flap deflections and main landing gear deployed or retracted. The high-fidelity, synthetic data were post-processed using a Ffowcs-Williams and Hawkings integral approach to estimate farfield acoustic behavior, with pressures on the model solid surface or a permeable surface enveloping the acoustic near field used as input. The numerical approach, simulation attributes, and the effects of grid resolution, gear deployment, and multiple flap deflections, are discussed as well.

Numerical simulations have been performed for a simplified high-lift configuration that is representative of a modern transport airplane. This configuration includes a leading-edge slat, fuselage, wing, nacelle-pylon and a simple hinged flap. The suction surface of the flap is embedded with multiple rows of fluidic actuators to reduce the extent of reversed flow regions and improve the aerodynamic performance of the configuration with flap in a deployed state. In the current paper, a Lattice Boltzmann Method based high-fidelity computational fluid dynamics (CFD) code, known as PowerFLOW® is used to simulate the entire flow field associated with this configuration, including the flow inside the actuators. A fully compressible version of the PowerFLOW® code that has been validated for high speed flows is used for the present simulations to accurately represent the transonic flow regimes that are encountered in the flow field generated by the actuators operating at higher mass flow (momentum) rates required to mitigate reverse flow regions on the suction surfaces of the main wing and the flap. The numerical solutions predict the expected trends in aerodynamic forces as the actuation levels are increased. More efficient active flow control (AFC) systems and actuator arrangement for lift augmentation are emerging based on the parametric studies conducted here prior to wind tunnel tests. These numerical solutions will be compared with experimental data, once such data becomes available.

The PowerFLOW R (cid:13) code has been used to perform numerical simulations of the high-lift version of the Common Research Model (HL-CRM) that will be used for experimental testing of airframe noise. Time-averaged surface pressure results from PowerFLOW R (cid:13) are found to be in reasonable agreement with those from steady-state computations using FUN3D. Surface pressure fluctuations are highest around the slat break and nacelle/pylon region, and synthetic array beamforming results also indicate that this region is the dominant noise source on the model. The gap between the slat and pylon on the HL-CRM is not realistic for modern aircraft, and most nacelles include a chine that is absent in the baseline model. To account for those effects, additional simulations were completed with a chine and with the slat extended into the pylon. The case with the chine was nearly identical to the baseline, and the slat extension resulted in higher surface pressure fluctuations but slightly reduced radiated noise. The full-span slat geometry without the nacelle/pylon was also simulated and found to be around 10 dB quieter than the baseline over almost the entire frequency range. The current simulations are still considered preliminary as changes in the radiated acoustics are still being observed with grid refinement, and additional simulations with finer grids are planned.

Active flow control technology is increasingly used in aerospace applications to control flow separation and to improve aerodynamic performance. In this paper, PowerFLOW is used to simulate the flow through a sweeping jet actuator at two different pressure ratios. The lower pressure ratio leads to a high subsonic flow, whereas the high pressure ratio produces a choked flow condition. Comparison of numerical results with experimental data is shown, which includes qualitatively good agreement of pressure histories and spectra. PIV measurements are also available but the simulation overestimates mean and fluctuation quantities outside the actuator. If supply pressure is matched at one point inside the mixing chamber a good qualitative agreement is achieved at all other monitor points.

High-fidelity simulations focused on full-scale evaluation of new technologies for mitigating flap and landing gear noise are presented. These noise reduction concepts were selected because of their superior acoustic performance, as demonstrated during NASA wind tunnel tests of an 18%-scale, semi-span model of a Gulfstream aircraft. The full-scale, full-aircraft, time-accurate simulations were performed with the lattice Boltzmann PowerFLOW solver for free air at a Mach number of 0.2. Three aircraft configurations (flaps deflected at 39 without and with main gear deployed, and 0 flaps with main gear extended) were used to determine the aeroacoustic performance of the concepts on component-level (individually) and system-level (concurrent application) bases. Farfield noise spectra were obtained using a Ffowcs-Williams and Hawkings acoustic analogy approach. Comparison of the predicted spectra without (baseline) and with the noise treatments applied showed that noise reduction benefits between 2-3 dB for the flap and 1.3-1.7 dB for the main landing gear are obtained. It was also found that the full extent of the benefits is being masked by the noise generated from the flap brackets and main gear cavities, which act as prominent secondary sources.

Unsteady flow computations are presented for a Gulfstream aircraft model in landing configuration, i.e., flap deflected 39deg and main landing gear deployed. The simulations employ the lattice Boltzmann solver PowerFLOW(Trademark) to simultaneously capture the flow physics and acoustics in the near field. Sound propagation to the far field is obtained using a Ffowcs Williams and Hawkings acoustic analogy approach. Two geometry representations of the same aircraft are analyzed: an 18% scale, high-fidelity, semi-span model at wind tunnel Reynolds number and a full-scale, full-span model at half-flight Reynolds number. Previously published and newly generated model-scale results are presented; all full-scale data are disclosed here for the first time. Reynolds number and geometrical fidelity effects are carefully examined to discern aerodynamic and aeroacoustic trends with a special focus on the scaling of surface pressure fluctuations and farfield noise. An additional study of the effects of geometrical detail on farfield noise is also documented. The present investigation reveals that, overall, the model-scale and full-scale aeroacoustic results compare rather well. Nevertheless, the study also highlights that finer geometrical details that are typically not captured at model scales can have a non-negligible contribution to the farfield noise signature.