Results for “computational fluid dynamics simulation” — Page 19

NASA NTRS2017

Transonic Semispan Aerodynamic Testing of the Hybrid Wing Body with Over Wing Nacelles in the National Transonic Facility

A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.

Chan, David T.↗

NASA NTRS2017

A Computational Fluid Dynamics Study of Swirling Flow Reduction by Using Anti-Vortex Baffle

An anti-vortex baffle is a liquid propellant management device placed adjacent to an outlet of the propellant tank. Its purpose is to substantially reduce or eliminate the formation of free surface dip and vortex, as well as prevent vapor ingestion into the outlet, as the liquid drains out through the flight. To design an effective anti-vortex baffle, Computational Fluid Dynamic (CFD) simulations were undertaken for the NASA Ares I vehicle LOX tank subjected to the simulated flight loads with and without the anti-vortex baffle. The Six Degree-Of-Freedom (6-DOF) dynamics experienced by the Crew Launch Vehicle (CLV) during ascent were modeled by modifying the momentum equations in a CFD code to accommodate the extra body forces from the maneuvering in a non-inertial frame. The present analysis found that due to large moments, the CLV maneuvering has a significant impact on the vortical flow generation inside the tank. Roll maneuvering and side loading due to pitch and yaw are shown to induce swirling flow. The vortical flow due to roll is symmetrical with respect to the tank centerline, while those induced by pitch and yaw maneuverings showed two vortices side by side. The study found that without the anti-vortex baffle, the swirling flow caused surface dip during the late stage of drainage and hence early vapor ingestion. The flow can also be non-uniform in the drainage pipe as the secondary swirling flow velocity component can be as high as 10% of the draining velocity. An analysis of the vortex dynamics shows that the swirling flow in the drainage pipe during the Upper Stage burn is mainly the result of residual vortices inside the tank due to the conservation of angular momentum. The study demonstrated that the swirling flow in the drainage pipe can be effectively suppressed by employing the anti-vortex baffle.

Yang, H. Q.↗

NASA NTRS2017

A Computational Fluid Dynamics Study of Swirling Flow Reduction by Using Anti-Vortex Baffle

An anti-vortex baffle is a liquid propellant management device placed adjacent to an outlet of the propellant tank. Its purpose is to substantially reduce or eliminate the formation of free surface dip and vortex, as well as prevent vapor ingestion into the outlet, as the liquid drains out through the flight. To design an effective anti-vortex baffle, Computational Fluid Dynamic (CFD) simulations were undertaken for the NASA Ares I vehicle LOX tank subjected to the simulated flight loads with and without the anti-vortex baffle. The Six Degree-Of-Freedom (6-DOF) dynamics experienced by the Crew Launch Vehicle (CLV) during ascent were modeled by modifying the momentum equations in a CFD code to accommodate the extra body forces from the maneuvering in a non-inertial frame. The present analysis found that due to large moments, the CLV maneuvering has a significant impact on the vortical flow generation inside the tank. Roll maneuvering and side loading due to pitch and yaw are shown to induce swirling flow. The vortical flow due to roll is symmetrical with respect to the tank centerline, while those induced by pitch and yaw maneuverings showed two vortices side by side. The study found that without the anti-vortex baffle, the swirling flow caused surface dip during the late stage of drainage and hence early vapor ingestion. The flow can also be non-uniform in the drainage pipe as the secondary swirling flow velocity component can be as high as 10% of the draining velocity. An analysis of the vortex dynamics shows that the swirling flow in the drainage pipe during the Upper Stage burn is mainly the result of residual vortices inside the tank due to the conservation of angular momentum. The study demonstrated that the swirling flow in the drainage pipe can be effectively suppressed by employing the anti-vortex baffle.

Yang, H. Q.↗

NASA NTRS2017

Numerical and Analytical Assessment of a Coupled Rotating Detonation Engine and Turbine Experiment

An analysis is presented of an experimental rig comprising a rotating detonation engine (RDE) with bypass ejector flow coupled to a downstream turbine. The analysis used a validated computational fluid dynamics RDE simulation combined with straightforward algebraic mixing equations for the bypass flow. The objectives of the analysis were to supplement and interpret the necessarily sparse measurements from the rig, and to assess the performance of the RDE itself, which was not instrumented in this installation. The analysis is seen to agree reasonably well with available data. It shows that the RDE is operating in an unusual fashion, with subsonic flow throughout the exhaust plane. The detonation event itself is producing a total pressure rise relative to the pre-detonative pressure; however, the length of the device and the substantial flow restriction at the inlet yield an overall pressure loss. This is not surprising since the objective of the rig test was primarily aimed at investigating RDE/turbine interactions, and not on performance optimization. Furthermore, the RDE was designed for fundamental detonation studies and not performance. Nevertheless, the analysis indicates that with some small alterations to the design, an RDE with an overall pressure rise is possible.

detonation↗

NASA NTRS2016

Passive Rocket Diffuser Testing: Reacting Flow Performance of Four Second-Throat Geometries

Second-throat diffusers serve to isolate rocket engines from the effects of ambient back pressure. As one of the nation's largest rocket testing facilities, the performance and design limitations of diffusers are of great interest to NASA's Stennis Space Center. This paper describes a series of tests conducted on four diffuser configurations to better understand the effects of inlet geometry and throat area on starting behavior and boundary layer separation. The diffusers were tested for a duration of five seconds with a 1455-pound thrust, LO2/GH2 thruster to ensure they each reached aerodynamic steady state. The effects of a water spray ring at the diffuser exits and a water-cooled deflector plate were also evaluated. Static pressure and temperature measurements were taken at multiple axial locations along the diffusers, and Computational Fluid Dynamics (CFD) simulations were used as a tool to aid in the interpretation of data. The hot combustion products were confirmed to enable the diffuser start condition with tighter second throats than predicted by historical cold-flow data or the theoretical normal shock method. Both aerodynamic performance and heat transfer were found to increase with smaller diffuser throats. Spray ring and deflector cooling water had negligible impacts on diffuser boundary layer separation. CFD was found to accurately capture diffuser shock structures and full-flowing diffuser wall pressures, and the qualitative behavior of heat transfer. However, the ability to predict boundary layer separated flows was not consistent.

Jones, Daniel R.↗

NASA NTRS2017

Numerical and Analytical Assessment of a Coupled Rotating Detonation Engine and Turbine Experiment

An analysis is presented of an experimental rig comprising a rotating detonation engine (RDE) with bypass ejector flow coupled to a downstream turbine. The analysis used a validated computational fluid dynamics RDE simulation combined with straightforward algebraic mixing equations for the bypass flow. The objectives of the analysis were to supplement and interpret the necessarily sparse measurements from the rig, and to assess the performance of the RDE itself, which was not instrumented in this installation. The analysis is seen to agree reasonably well with available data. It shows that the RDE is operating in an unusual fashion, with subsonic flow throughout the exhaust plane. The detonation event itself is producing a total pressure rise relative to the pre-detonative pressure; however, the length of the device and the substantial flow restriction at the inlet yield an overall pressure loss. This is not surprising since the objective of the rig test was primarily aimed at investigating RDE turbine interactions, and not on performance optimization. Furthermore, the RDE was designed for fundamental detonation studies and not performance. Nevertheless, the analysis indicates that with some small alterations to the design, an RDE with an overall pressure rise is possible.

propulsion↗

NASA NTRS2017

Two-Pendulum Model of Propellant Slosh in Europa Clipper PMD Tank

The objective of this fluids analysis is to model propellant slosh for the Europa Clipper mission using a two-pendulum model, such that controls engineers can predict slosh behavior during the mission. Propellant slosh causes shifts in center of mass and exerts forces and torques on the spacecraft which, if not adequately controlled, can lead to mission failure. The two-pendulum model provides a computationally simple model that can be used to predict slosh for the Europa Clipper tank geometry. The Europa Clipper tank is cylindrical with a domed top and bottom and includes a propellant management device (PMD). Due to the lack of experimental data in low gravity environments, computational fluid dynamics (CFD) simulation results were used as 'real' slosh behavior for two propellants at three fill fractions. Key pendulum parameters were derived that allow the pendulum model's center of mass, forces, and moments to closely match the CFD data. The parameter trends were examined as a function of tank fill fraction and compared with solutions to analytic equations that describe the frequency of slosh in tanks with simple geometries. The trends were monotonic as expected, and parameters resembled analytical predictions; any differences could be explained by the specific differences in the geometry of the tank. This paper summarizes the new method developed at Goddard Space Flight Center (GSFC) for deriving pendulum parameters for two-pendulum equivalent sloshing models. It presents the results of this method and discusses the validity of the results. This analysis is at a completed stage and will be applied in the immediate future to the evolving tank geometry as Europa Clipper moves past its preliminary design review (PDR) phase.

Propellant Slosh↗

NASA NTRS2017

Numerical and Analytical Assessment of a Coupled Rotating Detonation Engine and Turbine Experiment

An analysis is presented of an experimental rig comprising a rotating detonation engine (RDE) with bypass flow coupled to a downstream turbine. The analysis used a validated computational fluid dynamics RDE simulation combined with straightforward algebraic mixing equations for the bypass flow. The objectives of the analysis were to supplement and interpret the necessarily sparse measurements from the rig, and to assess the performance of the RDE itself (which was not instrumented in this installation). The analysis is seen to agree reasonably well with available data. It shows that the RDE is operating in an unusual fashion, with subsonic flow throughout the exhaust plane. The detonation event itself is producing a total pressure rise relative to the pre-detonative pressure; however, the length of the device and the substantial flow restriction at the inlet yield an overall pressure loss. This is not surprising since the objective of the rig test was primarily aimed at investigating RDEturbine interactions, and not on performance optimization. Furthermore, the RDE was designed for fundamental detonation studies and not performance. Nevertheless, the analysis indicates that with some small alterations to the design, an RDE with an overall pressure rise is possible.

detonation↗

NASA NTRS2018

High-Fidelity Computational Aerodynamics of Multi-Rotor Unmanned Aerial Vehicles

High-fidelity Computational Fluid Dynamics (CFD) simulations have been carried out for several multi-rotor Unmanned Aerial Vehicles (UAVs). Three vehicles have been studied: the classic quadcopter DJI Phantom 3, an unconventional quadcopter specialized for forward flight, the SUI Endurance, and an innovative concept for Urban Air Mobility (UAM), the Elytron 4S UAV. The three-dimensional unsteady Navier-Stokes equations are solved on overset grids using high-order accurate schemes, dual-time stepping, and a hybrid turbulence model. The DJI Phantom 3 is simulated with different rotors and with both a simplified airframe and the real airframe including landing gear and a camera. The effects of weather are studied for the DJI Phantom 3 quadcopter in hover. The SUI En- durance original design is compared in forward flight to a new configuration conceived by the authors, the hybrid configuration, which gives a large improvement in forward thrust. The Elytron 4S UAV is simulated in helicopter mode and in airplane mode. Understanding the complex flows in multi-rotor vehicles will help design quieter, safer, and more efficient future drones and UAM vehicles.

High-Fidelity↗

NASA NTRS2017

Building Aerodynamic Databases for the SLS Design Process

NASA's new Space Launch System (SLS) will be the first rocket since the Saturn V (1967-1973) to carry astronauts beyond low earth orbit-and will carry 10% more payload than Saturn V and three times the payload of the space shuttle. The SLS configuration consists of a center core and two solid rocket boosters that separate from the core as their fuel is exhausted two minutes after lift-off. During these first two minutes of flight, the vehicle powers its way through strong shock waves as it accelerates past the speed of sound, then pushes beyond strong aerodynamic loads at the maximum dynamic pressure, and is ultimately enveloped by gaseous plumes from the booster-separation motors. The SLS program relies on computational fluid dynamic (CFD) simulations to provide much of the data needed to build aerodynamic databases describing the structural load distribution, surface pressures, and aerodynamic forces on the vehicle.

Aerodynamic↗

NASA NTRS2018

Examination of Wave Speed in Rotating Detonation Engines Using Simplified Computational Fluid Dynamics

A simplified, two-dimensional, computational fluid dynamic (CFD) simulation, with a reactive Euler solver is used to examine possible causes for the low detonation wave propagation speeds that are consistently observed in air breathing rotating detonation engine (RDE) experiments. Intense, small-scale turbulence is proposed as the primary mechanism. While the solver cannot model this turbulence, it can be used to examine the most likely, and profound effect of turbulence. That is a substantial enlargement of the reaction zone, or equivalently, an effective reduction in the chemical reaction rate. It is demonstrated that in the unique flowfield of the RDE, a reduction in reaction rate leads to a reduction in the detonation speed. A subsequent test of reduced reaction rate in a purely one-dimensional pulsed detonation engine (PDE) flowfield yields no reduction in wave speed. The reasons for this are explained. The impact of reduced wave speed on RDE performance is then examined, and found to be minimal. Two other potential mechanisms are briefly examined. These are heat transfer, and reactive mixture non-uniformity. In the context of the simulation used for this study, both mechanisms are shown to have negligible effect on either wave speed or performance.

detonation↗

NASA NTRS2018

Examination of Wave Speed in Rotating Detonation Engines Using Simplified Computational Fluid Dynamics

A simplified, two-dimensional, computational fluid dynamic (CFD) simulation, with a reactive Euler solver is used to examine possible causes for the low detonation wave propagation speeds that are consistently observed in air breathing rotating detonation engine (RDE) experiments. Intense, small-scale turbulence is proposed as the primary mechanism. While the solver cannot model this turbulence, it can be used to examine the most likely, and profound effect of turbulence. That is a substantial enlargement of the reaction zone, or equivalently, an effective reduction in the chemical reaction rate. It is demonstrated that in the unique flowfield of the RDE, a reduction in reaction rate leads to a reduction in the detonation speed. A subsequent test of reduced reaction rate in a purely one-dimensional pulsed detonation engine (PDE) flowfield yields no reduction in wave speed. The reasons for this are explained. The impact of reduced wave speed on RDE performance is then examined, and found to be minimal. Two other potential mechanisms are briefly examined. These are heat transfer, and reactive mixture non-uniformity. In the context of the simulation used for this study, both mechanisms are shown to have negligible effect on either wave speed or performance.

combustion↗

NASA NTRS2018

Europa Clipper Preliminary Design Review Propellant Slosh Analysis

This presentation outlines the preliminary design review (PDR) propellant slosh analysis done for the Europa Clipper Mission. It provides sample results for both high acceleration pendulum-damper models and low acceleration pendulum-spring-damper models. The high acceleration pendulum-damper models were derived from STAR-CCM+ computational fluid dynamic (CFD) simulations and the low acceleration pendulum-spring-damper models were derived from Surface Evolver models.

slosh↗

NASA NTRS2018

A Physics-Based Approach to Urban Air Mobility

High-fidelity Computational Fluid Dynamics (CFD) simulations for multi-rotor vehicles have been carried out. The three-dimensional unsteady Navier-Stokes equations are solved on overset grids employing high order accurate schemes, dual-time stepping, and a hybrid turbulence model using NASA's CFD code Over- flow. The vehicles studied consist of small to medium sized drones, and bigger vehicles for future Urban Air Mobility (UAM) applications. The performances for different configurations and rotor mounting are calculated in hover and in forward flight. Understanding the complex flows and the interactions between rotors and with other elements will help design the future multi-rotor vehicles to be quieter, safer, and more efficient.

Urban↗

NASA NTRS2018

Generation of Performance Model for the Aeolian Wind Tunnel (AWT) Rotor at Reduced Pressure

The NASA Jet Propulsion Laboratory (JPL) designed the Mars Helicopter (MH) in collaboration with AeroVironment Inc., NASA Ames Research Center, and NASA Langley Research Center to explore the possibility of a vertical takeoff and landing (VTOL) Unmanned Aerial Vehicle (UAV) for flight on Mars. A 40-inch-diameter Aeolian Wind Tunnel (AWT) rotor, roughly approximating the proposed MH design by JPL, was tested in forward flight at Mars atmospheric pressure at the NASA Ames Planetary Aeolian Laboratory (PAL) in support of MH research efforts. This report describes the generation of the rotor model used to correlate with that experimental effort as reported by Ament and Koning. The 40-inch-diameter rotor was 3D-scanned and transformed into an airfoil deck. The scanned rotor airfoil sections are analyzed using C81 Generator (C81Gen) to generate the sectional aerodynamic coefficients for comprehensive analyses. A mid-fidelity computational fluid dynamics (CFD) simulation using Rotorcraft CFD (RotCFD) is pursued to efficiently estimate rotor hover and forward flight performance. Simulations at two pressures, 7 mbar (approximate Martian atmospheric pressure) and 1018 mbar (1 atmosphere), are performed to gain an understanding of the performance differences and Reynolds number effects observed. Experimental 1-atmosphere thrust for single- and dual-rotor isolated hover cases correlate well with the modeled rotor. Performance results at reduced pressure (7 mbar) show a drastic decrease in lift for equivalent RPMs tested at 1 atmosphere. Although this is primarily due to pressure reduction, Reynolds number effects also contribute to this decrease, as airfoil lift and drag coefficients are affected when compared with 1-atmosphere results. Further, simulated rotor power coefficient shows drastic increases at reduced pressures, attributed to laminar boundary layer separation, as described in Koning et al. for the MH rotor analysis. PAL experimental Martian Surface Wind Tunnel (MARSWIT) results are presented in the paper by Ament and Koning. The very low Reynolds number range is currently not well understood and presents various challenges for both experimentation and simulation.

Koning, Witold J. F.↗

NASA NTRS2018

Rotor CFD Analysis at Terrestrial and Martian Atmospheric Densities

Much effort has been made to enhance exploration on Mars. In addition to a rover and Mars-orbiting satellites, a Mars helicopter (MH) was proposed in order to augment planetary research. Computational Fluid Dynamics (CFD) simulations have been performed to have a better understanding of the behavior and performance of vertical lift Planetary Aerial Vehicles (PAV). Due to the large differences in atmospheric conditions between Mars and Earth, predicting and testing rotorcraft performance is a complex task. The goal of this project is to understand the capability of the mid-fidelity CFD software RotCFD to predict rotor performance in terms of thrust at 1013.25 milibar and 14 milibar corresponding to Terrestrial and Martian conditions, respectively. Also, in order to characterize the wind tunnel wall effects free field and wind tunnel simulations were performed, analyzed and compared. Different analytical tools have been used in order to aid with the design process for the future vertical lift planetary aerial vehicles. One of them includes experimental tests performed on a rotor in the Aeolian Wind Tunnel (AWT) facility at NASA Ames Research Center under different pressure conditions ranging from Terrestrial to Martian atmospheric conditions. Other software was used as well in order to capture the aerodynamic coefficients of the corresponding rotor sections based on the Mach and Reynolds numbers used for the experimental tests. The aerodynamic coefficients were input into RotCFD, and various simulations were performed under Terrestrial and Martian conditions in order to mimic the experimental test. Then, the obtained results from RotCFD were compared with the AWT collected data.

Atmospheric Pressures↗

NASA NTRS2016

Effect of Aft Rotor on the Inter-Rotor Flow of an Open Rotor Propulsion System

The effects of the aft rotor on the inter-rotor flow field of an open rotor propulsion rig were examined. A Particle Image Velocimetry (PIV) dataset that was acquired phase locked to the front rotor position has been phase averaged based on the relative phase angle between the forward and aft rotors. The aft rotor phase was determined by feature tracking in raw PIV images through an image processing algorithm. The effects of the aft rotor potential field on the inter-rotor flow were analyzed and shown to be in reasonably good agreement with Computational Fluid Dynamics (CFD) simulations. The aft rotor position was shown to have a significant upstream effect, with implications for front rotor interaction noise. It was found that the aft rotor had no substantial effect on the position of the forward rotor tip vortex but did have a small effect on the circulation strength of the vortex when the rotors were highly loaded.

Counter Rotating↗

NASA NTRS2019

Surface Flow Visualization of the High-Lift Common Research Model

A 10% scale version of the High-Lift Common Research Model (CRM-HL) was tested in the NASA Langley 14- by 22-Foot Subsonic Tunnel (14x22) in support of the NASA Advanced Air Transport Technology (AATT) Project. The CRM-HL experiment included various configurations such as conventional and simple-hinged flaps, with and without engine nacelle/pylon, with and without nacelle chine, different Active Flow Control (AFC) methods (sweeping jets, alternating pulsed jets, and preconditioned boundary layer blowing), and their various parameters. This particular study is focused on the surface flow visualization of the conventional CRM-HL model at landing configuration. The conventional CRM-HL model with the single-slotted Fowler flap system serves as a baseline for the AFC-enabled simplified high-lift configuration as well as a high-lift technology development platform due to its publicly open geometry. Surface flow visualizations were performed using fluorescent minitufts, which were found to be nonintrusive to the aerodynamic performance. Tuft flow visualizations are supplemented with the relevant pressure and force measurements in order to understand the flow characteristics developed on the conventional CRM- HL model. In addition, three dimensional, unsteady, compressible Computational Fluid Dynamic (CFD) simulations were performed for selective cases. The surface streamlines and transverse velocity fluctuations obtained by the CFD simulations are qualitatively compared to the tuft direction and tuft unsteadiness, respectively. Force measurements of the CRM-HL model show performance degradation at higher angles of attack. Surface flow visualizations revealed the performance loss due to the nacelle/pylon wake that grows with angle of attack and eventually promotes flow separation over the inboard wing. This performance loss was successfully recovered by placing a chine on the engine nacelle.

Koklu, Mehti↗