The CFD solver OVERFLOW was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary PlanWind Tunnel as part of a CFD and wind-tunnel integration effort. Selection of numerical methods and resulting comparisons to selected tunnel conditions of interest are detailed. Comparisons of mean pressures with static taps and pressure-sensitive paint are provided for multiple SST-based turbulence models. URANS predictions of heatshield pressures at high thrust conditions are poor compared to TMRC and DES predictions. Comparisons to experimental measurements of pressure fluctuations indicate complex relationships between heatshield unsteadiness, flight Mach number, and engine thrust coefficient. Bow shock topology and shock cell structure between experimental and OVERFLOW data show good agreement in Schlieren imaging. Recommendations for future investigation into flow field unsteadiness and turbulence models are suggested.

Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flow-field predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, or when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at higher CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Simulations of unsteady supersonic retropropulsion (SRP) flow over a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) blunt-body vehicle were performed using the OVERFLOW Computational Fluid Dynamics (CFD) solver. High-fidelity flow solver techniques, including Detached Eddy Simulation (DES) turbulence modeling and Adaptive Mesh Refinement (AMR), were employed to obtain improved realism in CFD predictions. Simulation conditions and geometry configurations were designed to match specific runs in the Descent System Study (DSS) wind tunnel testing (WTT) campaign. The accuracy of each simulation is assessed by direct comparison to experimental data. Comparisons of computational predictions of the SRP flowfield and bow shock shape to experimental schlieren imaging show reasonable prediction of mean shock shape, with approximately 10% similarity in shock standoff distance for selected conditions, as well as similarity in local, time-varying fluctuations of the shock-plume interaction. Comparisons of discrete measurements of surface pressure coefficient (C-P) indicate CFD accuracy within approximately 10% of the experiment across the majority of the model heatshield, with larger variations at some of the heatshield edge locations with stronger flow unsteadiness. Simulated unsteadiness of these chaotic flows, which were highly dynamic and multi-modal, was shown to be within 20-40% of experimentally-measured pressure standard deviation (SD) for the majority of the sampled locations.

View Video Presentation: https://doi.org/10.2514/6.2022-0912.vid Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not appropriate for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. A test will be conducted in the NASA Langley Unitary Plan Wind Tunnel to begin addressing powered descent aerodynamics risks for large-scale human Mars entry concepts and to identify gaps in computational predictive capabilities. This paper covers pre-test computational flowfield predictions of two different models derived from full-scale reference vehicles: a blunt low lift-to-drag vehicle and a more slender geometry. Calculations of the blunt model include variations in nozzle configuration: nozzle location, size, area ratio, and pointing direction. There exist some significant differences between solvers, but some general trends are observed from simulations of the blunt model. First, aerodynamic axial force from the heatshield decreases with increasing thrust due to expanding plume blockage. Second, nozzles that point along the model axis result in lower aerodynamic axial force compared to nozzles that have a radial thrust component. Finally, placing the nozzles further away from the model nose preserves more heatshield axial force with increasing thrust compared to nozzles that are closer to the nose. For the slender model, the axial force from the heatshield is similar to the non-blowing axial force regardless of thrust magnitude, due to the nozzle arrangement on the heatshield. Once the test is completed, direct comparisons between the computations and test data will be made to determine computational uncertainties in a wind tunnel environment, to identify gaps in predictive capabilities, and to inform planning for future ground and flight test programs for Mars powered descent vehicles.

The CFD solver OVERFLOW was used to simulate the CobraMRV undergoing supersonic retropropulsion (SRP) in the Langley Unitary Plan Wind Tunnel as part of a pre-test study. Aerodynamics of the CobraMRV are summarized and the sensitivity of vehicle loads to CFD parameters at a subset of operating conditions are investigated. Specific numerical methods, boundary conditions, and turbulence modeling options have been selected after considering the complex phenomena in SRP flows. Proper shock-capturing methods and dynamic grid adaption are necessary to correctly capture vehicle loads and dynamics. Present data indicates that at particular conditions, upstream asymmetries originating from the tunnel inflow plane contribute to measurable asymmetries on the CobraMRV heatshield.

The combined effects of variable density, curvature, and convective Mach number on turbu- lence in time-developing free shear layers are studied using turbulence-resolving simulations. The range of parameters considered include the convective Mach number at 0.2 and 0.8, the density ratio at 1/7, 1 and 7, and the level of streamwise curvature which is measured by the shear layer thickness ratio to the radius of curvature at 0.001 and 0.01. Simulations at the limit of zero curvature exhibit mean profiles and turbulent stresses comparable with previously simulated planar (non-curved) shear layers. Growth rate changes depending on compressibility, curvature and density ratio are compared with expected behavior and unexpected trends due to combined effects are investigated. Standard turbulent shear layer statistics are presented, including thickness growth rates, turbulent stresses, and stress budgets to characterize individual and combined effects of the selected physical parameters.

The ability of CFD simulations to serve as a surrogate for wind tunnel testing at high supersonic speeds has been evaluated for a sub-scale model of the Co-Optimization Blunt-body Re-entry Analysis-Mid-lift-to-drag Rigid Vehicle (CobraMRV) human Mars entry vehicle concept. The vehicle was tested at the Unitary Plan Wind Tunnel (UPWT) at the NASA Langley Research Center under flow conditions and surface control configurations relevant to the entry stage of a flight mission. The CFD simulations were performed prior to gaining access to test results in order to assess how blind predictions obtained using best practices compare to experiments. Solutions of empty tunnel simulations were used as inflow boundary condition for the CobraMRV simulations in a truncated portion of the test section. Different solvers and turbulence models were used by separate teams to assess sensitivity to numerical methods, physics, and users. After release of the test results, the pre-test computations were compared to the experimental results, and additional analyses have been conducted to explain observed discrepancies. The amount of time and resources dedicated to each phase of the computational work was logged for comparison to that required for wind tunnel tests, and to inform planning of future CFD data base development projects.

As part of studying the ability of Computational Fluid Dynamics (CFD) to accurately model important flow physics in the high-supersonic Mach number range, control-surface effectiveness on an entry vehicle for Mars exploration was examined. As with several other important flow regimes studied under the CFD as Surrogate for Wind Tunnel Testing at High Supersonic Speeds Project, a combined CFD and wind-tunnel study of a proposed Mars landing configuration was undertaken. The testing was performed in high-speed test section of the NASA Langley Research Center’s Unitary Plan Wind Tunnel. The CFD team was an integral part of the overall evaluation team throughout the model development and test planning process and performed pre-test computations predicting the results of the testing. For the CFD predictions of the model in the wind tunnel, the flow into the test section was imposed as a boundary condition. The imposed inflow was based on a previous flow characterization study and companion CFD simulating the flow from the settling chamber through the test section. This paper presents a description of the control-surface effectiveness testing and the resulting aerodynamic data. A summary of comparisons of the experimental and CFD results will also be presented as well as a cost comparison of the two approaches to generating data on this previously untested vehicle shape.

Flow in the empty Unitary Planform Wind Tunnel at the NASA Langley Research Center is simulated with computational fluid dynamics methods. The objectives are to assess CFD’s true predictive capability and to generate flow-field maps upstream of the tunnel’s test section for use in associated vehicle tests. Multiple CFD solvers, grid adaption methods, and turbulence models are used by five teams doing the simulations. The simulation domain is as large as practical: from the start of the settling chamber, just downstream of the last set of turning vanes, to well downstream of the test section. Simulations are done at three Mach numbers and several Reynolds numbers that cover most of the tunnel’s operating range. The sensitivity of the CFD solutions to simulation process details, including turbulence model, spatial resolution and settling-chamber inflow velocity profile, is characterized. Significant sensitivities in some details of the predicted test section flow are described. CFD predicts that details of flow in the settling chamber can affect the flow in the test section. Turning of the flow through the tunnel contraction and past nozzle-block hardware generates streamwise vortices, some of which survive into the core of the test section, most significantly at low Mach number. Preliminary comparisons to experimental measurements are given. The major flow characteristics and the dynamical vortices predicted by CFD are observed in the experiment. Several details in CFD and experiment differ, including the effects of the vortices and features that appear to be Mach waves in the test section. Sensitivity to grid resolution and turbulence modeling is noted in the vortices. No probably cause of the discrepancies in Mach waves has yet been identified in CFD sensitivity studies.

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.

This work is a simulation technology demonstrator, of sweep jet flow control used to suppress boundary layer separation and increase the maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate streamwise direction. It also generates turbulent eddies at the oscillation frequency, which are typically large relative to the scales of boundary layer turbulence, and which augment mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from a nozzle downstream of the oscillator, and an array of sweep jets which suppresses boundary layer separation are performed. Simulation results are compared to data from a dedicated validation experiment of a single oscillator and its sweep jet, and from a wind tunnel test of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets. A critical step in the work is the development of realistic time-dependent sweep jet inflow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the computational fluid dynamics (CFD) solver Overow, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

A hybrid transition trip-dot sizing and placement test technique was developed in support of recent experimental research on a hybrid wing-body configuration under study for the NASA Environmentally Responsible Aviation project. The approach combines traditional methods with Computational Fluid Dynamics. The application had three-dimensional boundary layers that were simulated with either fully turbulent or transitional flow models using established Reynolds-Averaged Navier-Stokes methods. Trip strip effectiveness was verified experimentally using infrared thermography during a low-speed wind tunnel test. Although the work was performed on one specific configuration, the process was based on fundamental flow physics and could be applicable to other configurations.

NASAs Environmentally Responsible Aviation (ERA) Project explores enabling technologies to reduce aviations impact on the environment. One research challenge area for the project has been to study advanced airframe and engine integration concepts to reduce community noise and fuel burn. In order to achieve this, complex wind tunnel experiments at both the NASA Langley Research Centers (LaRC) 14x22 and the Ames Research Centers 40x80 low-speed wind tunnel facilities were conducted on a Boeing Hybrid Wing Body (HWB) configuration. These wind tunnel tests entailed various entries to evaluate the propulsion airframe interference effects including aerodynamic performance and aeroacoustics. In order to assist these tests in producing high quality data with minimal hardware interference, extensive Computational Fluid Dynamic (CFD) simulations were performed for everything from sting design and placement for both the wing body and powered ejector nacelle systems to the placement of aeroacoustic arrays to minimize its impact on the vehicles aerodynamics. This paper will provide a high level summary of the CFD simulations that NASA performed in support of the model integration hardware design as well as some simulation guideline development based on post-test aerodynamic data. In addition, the paper includes details on how multiple CFD codes (OVERFLOW, STAR-CCM+, USM3D, and FUN3D) were efficiently used to provide timely insight into the wind tunnel experimental setup and execution.

The accuracy of Computational Fluid Dynamics predictions of subsonic capsule aerodynamics is examined by comparison against recent NASA wind-tunnel data at high-Reynolds-number flight conditions. Several aspects of numerical and physical modeling are considered, including inviscid numerical scheme, mesh adaptation, rough-wall modeling, rotation and curvature corrections for eddy-viscosity models, and Detached-Eddy Simulations of the unsteady wake. All of these are considered in isolation against relevant data where possible. The results indicate that an improved predictive capability is developed by considering physics-based approaches and validating the results against flight-relevant experimental data.

The National Aeronautics and Space Administrations Environmentally Responsible Aviation (ERA) Program is currently investigating the use of sweeping jet actuators as active flow control devices to improve the aerodynamic performances of vertical tails on commercial transporters. Computational Fluid Dynamics (CFD) simulations have shown that the motion of the jet is not a simple sinusoid, but lingers at the extremes of jet deflection. As part of an effort to better understand this non-sinusoidal behavior and validate the CFD, a sweeping jet actuator was tested in the 48-by-32-inch wind tunnel in the Fluid Mechanics Laboratory (FML) at NASA Ames Research Center. The jet was visualized at very high frequencies using a new technique: laser speckle retroreflective background oriented schlieren (RBOS). These measurements confirmed the non-sinusoidal nature of the jet motion. Although measurements were also made by Particle Image Velocimetry (PIV) that resolved the flow velocities in the jet, only the new RBOS technique could provide high enough frequencies to both spatially and temporally resolve the non-sinusoidal motion. This paper presents the laser speckle RBOS method and visualization, as well as a brief comparison to CFD simulations.

This work is a simulation technology demonstrator, of sweep jet flow control used to suppress boundary layer separation and increase the maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate streamwise direction. It also generates turbulent eddies at the oscillation frequency, which are typically large relative to the scales of boundary layer turbulence, and which augment mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from a nozzle downstream of the oscillator, and an array of sweep jets which suppresses boundary layer separation are performed. Simulation results are compared to data from a dedicated validation experiment of a single oscillator and its sweep jet, and from a wind tunnel test of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets. A critical step in the work is the development of realistic time-dependent sweepjet inflow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the computational fluid dynamics (CFD) solver Overflow, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

Characterization of the launch abort system of the Multi-purpose Crew Vehicle (MPCV) for control design and accurate simulation has provided a significant challenge to aerodynamicists and design engineers. The design space of the launch abort vehicle (LAV) includes operational altitudes from ground level to approximately 300,000 feet, Mach numbers from 0-9, and peak dynamic pressure near 1300psf during transonic flight. Further complicating the characterization of the aerodynamics and the resultant vehicle controllability is the interaction of the vehicle flowfield with the plumes of the two solid propellant motors that provide attitude control and the main propulsive impulse for the LAV. These interactions are a function of flight parameters such as Mach number, altitude, dynamic pressure, vehicle attitude, as well as parameters relating to the operation of the motors themselves - either as a function of time for the AM, or as a result of the flight control system requests for control torque from the ACM. This paper discusses the computational aerodynamic modeling of the aerodynamic interaction caused by main abort motor and the attitude control motor of the MPCV LAV, showing the effects of these interactions on vehicle controllability.

This work was motivated by the need to accurately characterize the aerodynamic behavior of the Orion Launch Abort Vehicle (LAV). The abort system is powered by solid rocket motors, and the vehicle’s aerodynamics involves complex interactions between multiple rocket plumes and the freestream. The LAV’s aerodynamic behavior is predicted using a combination of wind tunnel testing and CFD analyses. The goal is to rely on wind tunnel data to the extent possible. However, numerous issues, like the di culties of using solid rocket motors in a wind tunnel, make realistic and accurate experiments di cult to perform. CFD is used to reach the conditions which are di cult or impossible to achieve experimentally. CFD also has limitations, and one of the most signi cant is turbulence modeling. This work addresses turbulence modeling for rocket motor plumes, and particularly, the e ects of high-temperature jets. This is a signi cant concern because the majority of wind tunnel tests use ambient temperature compressed air for plumes, while the stagnation temperature of ight plumes is an order of magnitude higher. Issues of rocket plume chemistry are beyond the scope of this study. The scope of this work is both to validate that an available CFD model can correctly simulate the experimental results, and to select an appropriate CFD model for future hot plume predictions. The CFD model selection and validation involves comparisons of several turbulence model results to data from a single underexpanded supersonic jet inclined to create a 25 or 40 angle between the jet and a M = 0:3 freestream. Jets with stagnation temperatures of 530 R and 1350 R were studied. The work focuses on the SST turbulence model, but results from other 1and 2-equation models are also given. A compressibility correction which reduces the spreading rate of high-speed mixing layers, and a temperature correction (AbolHamid) which increases the spreading rate of hot jets were studied. Comparisons of CFD and experiment are given for PIV data of the hot-jet velocity eld, and for surface pressures on an LAV-like capsule for hot and cold jets. Water-vapor condensation precluded PIV measurements in the cold-jet cases. The experimental work has been given the designation \85-AA within the Orion Crew Exploration Vehicle (CEV) Aerosciences Project and hereafter will be referred to as 85-AA.