Karen A. Deere

Efficient Unstructured Grid Adaptation Methods for Sonic Boom Prediction

This paper examines the use of two grid adaptation methods to improve the accuracy of the near-to-mid field pressure signature prediction of supersonic aircraft computed using the USM3D unstructured grid flow solver. The first method (ADV) is an interactive adaptation process that uses grid movement rather than enrichment to more accurately resolve the expansion and compression waves. The second method (SSGRID) uses an a priori adaptation approach to stretch and shear the original unstructured grid to align the grid with the pressure waves and reduce the cell count required to achieve an accurate signature prediction at a given distance from the vehicle. Both methods initially create negative volume cells that are repaired in a module in the ADV code. While both approaches provide significant improvements in the near field signature (< 3 body lengths) relative to a baseline grid without increasing the number of grid points, only the SSGRID approach allows the details of the signature to be accurately computed at mid-field distances (3-10 body lengths) for direct use with mid-field-to-ground boom propagation codes.

Computational investigation of fluidic counterflow thrust vectoring

A computational study of fluidic counterflow thrust vectoring has been conducted. Two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and linear Reynolds stress modeling. For validation, computational results were compared to experimental data obtained at the NASA Langley Jet Exit Test Facility. In general, computational results were in good agreement with experimental performance data, indicating that efficient thrust vectoring can be obtained with low secondary flow requirements (less than 1% of the primary flow). An examination of the computational flowfield has revealed new details about the generation of a countercurrent shear layer, its relation to secondary suction, and its role in thrust vectoring. In addition to providing new information about the physics of counterflow thrust vectoring, this work appears to be the first documented attempt to simulate the counterflow thrust vectoring problem using computational fluid dynamics.

A Grid Sourcing and Adaptation Study Using Unstructured Grids for Supersonic Boom Prediction

NASA created the Supersonics Project as part of the NASA Fundamental Aeronautics Program to advance technology that will make a supersonic flight over land viable. Computational flow solvers have lacked the ability to accurately predict sonic boom from the near to far field. The focus of this investigation was to establish gridding and adaptation techniques to predict near-to-mid-field (<10 body lengths below the aircraft) boom signatures at supersonic speeds using the USM3D unstructured grid flow solver. The study began by examining sources along the body the aircraft, far field sourcing and far field boundaries. The study then examined several techniques for grid adaptation. During the course of the study, volume sourcing was introduced as a new way to source grids using the grid generation code VGRID. Two different methods of using the volume sources were examined. The first method, based on manual insertion of the numerous volume sources, made great improvements in the prediction capability of USM3D for boom signatures. The second method (SSGRID), which uses an a priori adaptation approach to stretch and shear the original unstructured grid to align the grid and pressure waves, showed similar results with a more automated approach. Due to SSGRID s results and ease of use, the rest of the study focused on developing a best practice using SSGRID. The best practice created by this study for boom predictions using the CFD code USM3D involved: 1) creating a small cylindrical outer boundary either 1 or 2 body lengths in diameter (depending on how far below the aircraft the boom prediction is required), 2) using a single volume source under the aircraft, and 3) using SSGRID to stretch and shear the grid to the desired length.

PAB3D: Its History in the Use of Turbulence Models in the Simulation of Jet and Nozzle Flows

This is a review paper for PAB3D s history in the implementation of turbulence models for simulating jet and nozzle flows. We describe different turbulence models used in the simulation of subsonic and supersonic jet and nozzle flows. The time-averaged simulations use modified linear or nonlinear two-equation models to account for supersonic flow as well as high temperature mixing. Two multiscale-type turbulence models are used for unsteady flow simulations. These models require modifications to the Reynolds Averaged Navier-Stokes (RANS) equations. The first scheme is a hybrid RANS/LES model utilizing the two-equation (k-epsilon) model with a RANS/LES transition function, dependent on grid spacing and the computed turbulence length scale. The second scheme is a modified version of the partially averaged Navier-Stokes (PANS) formulation. All of these models are implemented in the three-dimensional Navier-Stokes code PAB3D. This paper discusses computational methods, code implementation, computed results for a wide range of nozzle configurations at various operating conditions, and comparisons with available experimental data. Very good agreement is shown between the numerical solutions and available experimental data over a wide range of operating conditions.

Best Practices for Aero-Database CFD Simulations of Ares V Ascent

In support of NASA’s next generation heavy lift launch vehicle (HLLV), a simulation protocol has been developed to generate databases of the aerodynamic force and moment coefficients for HLLV ascent. The simulation protocol has been established and validated with a series of computational analyses that ensure best practices are achieved. Results of the sensitivity analyses using a full-scale Ares V flight vehicle are next applied in a validation study with three scaled-down Ares V wind tunnel test articles. Three independent computational fluid dynamic (CFD) flow solvers were included in the study. These included OVERFLOW, a viscous Reynolds Averaged Navier-Stokes (RANS) solver for structured overset grids, USM3D, a viscous RANS solver for unstructured tetrahedral grids, and Cart3D, an inviscid Euler solver using unstructured Cartesian grids and adjoint-based adaptive mesh refinement. First, a series of tests was independently performed for each applicable CFD code, including a grid convergence study and sensitivity studies of turbulence models and convective flux discretization methods. Once the proper grid resolution, physical models, and numerical parameters were determined for each of the codes, the process was continued with a code-to-code comparison. Each CFD code was applied to the Ares V flight vehicle at several points in the ascent trajectory, with all three codes obtaining consistent force and moment predictions. Finally, an extensive validation of the CFD approach was performed, in which the three codes were used to generate aero-databases of force and moment coefficients for three distinct Ares V wind tunnel test articles. These computations were performed concurrent to the experimental databases generated in the 14-inch wind tunnel at Marshall Space Flight Center (MSFC). Comparisons of the CFD results with the experimental data are reported and the viscous flow results compare well.

Computational Analysis of a Wing Designed for the X-57 Distributed Electric Propulsion Aircraft

A computational study of the wing for the distributed electric propulsion X-57 Maxwell airplane configuration at cruise and takeoff/landing conditions was completed. Two unstructured-mesh, Navier-Stokes computational fluid dynamics methods, FUN3D and USM3D, were used to predict the wing performance. The goal of the X-57 wing and distributed electric propulsion system design was to meet or exceed the required lift coefficient 3.95 for a stall speed of 58 knots, with a cruise speed of 150 knots at an altitude of 8,000 ft. The X-57 Maxwell airplane was designed with a small, high aspect ratio cruise wing that was designed for a high cruise lift coefficient (0.75) at angle of attack of 0°. The cruise propulsors at the wingtip rotate counter to the wingtip vortex and reduce induced drag by 7.5 percent at an angle of attack of 0.6°. The unblown maximum lift coefficient of the high-lift wing (with the 30° flap setting) is 2.439. The stall speed goal performance metric was confirmed with a blown wing computed effective lift coefficient of 4.202. The lift augmentation from the high-lift, distributed electric propulsion system is 1.7. The predicted cruise wing drag coefficient of 0.02191 is 0.00076 above the drag allotted for the wing in the original estimate. However, the predicted drag overage for the wing would only use 10.1 percent of the original estimated drag margin, which is 0.00749. Nomenclature CD drag coefficient Vt,ratio ratio of tip speed to freestream velocity CD,HLN drag coefficient, high-lift nacelles contribution W aircraft weight, lb CD,pylons drag coefficient, pylons contribution y axis along the wing span, in. CD,TN drag coefficient, wingtip nacelles contribution y + nondimensional first node height in boundary layer CD,wing Cf drag coefficient, wing contribution skin friction coefficient yCC + nondimensional first cell centroid height in boundary layer CL lift coefficient Symbols cl sectional lift coefficient  angle of attack, degrees CL,eff effective lift coefficient: CL+ CL,prop Δ delta CL,max maximum lift coefficient ρ density CL,prop lift coefficient from the contribution of propeller thrust in lift direction Acronyms BSL Menter k-ω basic turbulence model Cm pitching moment coefficient CFL pseudo time advancement Courant-Friedrichs-Lewy Cp pressure coefficient DEP distributed electric propulsion Cref reference chord, in. HLN high-lift nacelles, including pylons CT thrust coefficient HP horse power CQ torque coefficient KCAS knots calibrated airspeed D drag force KEAS knots equivalent airspeed d propeller diameter, ft. KTAS knots true airspeed h altitude, ft. LM Langtry-Menter transition model KT normalized thrust coefficient mph miles per hour KQ normalized torque coefficient QCR quadratic constitutive relation M Mach number RPM revolutions per minute P pressure, lbf/in SA Spalart-Almaras one equation turbulence model q dynamic pressure SARC SA rotation and curvature correction Re S Reynolds number based on Cref wing reference area, ft SCEPTOR Scalable Convergent Electric Propulsion Technology and Operations Research T temperature, °F SST Menter’s Shear Stress Transport model V freestream velocity, ft/sec TN wingtip nacelles * Aerospace Engineer, Configuration Aerodynamics Branch, Mail Stop 499, AIAA Senior Member. † Aerospace Engineer, Aeronautics Systems Analysis Branch, Mail Stop 442, AIAA Senior Member. ‡ Aerospace Engineer, Configuration Aerodynamics Branch, Mail Stop 499, AIAA Associate Fellow. § Senior Researcher, GEOLAB, Mail Stop 128. ** Technical Group Lead, GEOLAB, Mail Stop 128. https://ntrs.nasa.gov/search.jsp?R=20170005883 2019-12-26T23:55:46+00:00Z

Experimental Investigation of Convoluted Contouring for Aircraft Afterbody Drag Reduction

An experimental investigation was performed in the NASA Langley 16-Foot Transonic Tunnel to determine the aerodynamic effects of external convolutions, placed on the boattail of a nonaxisymmetric nozzle for drag reduction. Boattail angles of 15 and 22 were tested with convolutions placed at a forward location upstream of the boattail curvature, at a mid location along the curvature and at a full location that spanned the entire boattail flap. Each of the baseline nozzle afterbodies (no convolutions) had a parabolic, converging contour with a parabolically decreasing corner radius. Data were obtained at several Mach numbers from static conditions to 1.2 for a range of nozzle pressure ratios and angles of attack. An oil paint flow visualization technique was used to qualitatively assess the effect of the convolutions. Results indicate that afterbody drag reduction by convoluted contouring is convolution location, Mach number, boattail angle, and NPR dependent. The forward convolution location was the most effective contouring geometry for drag reduction on the 22 afterbody, but was only effective for M < 0.95. At M = 0.8, drag was reduced 20 and 36 percent at NPRs of 5.4 and 7, respectively, but drag was increased 10 percent for M = 0.95 at NPR = 7. Convoluted contouring along the 15 boattail angle afterbody was not effective at reducing drag because the flow was minimally separated from the baseline afterbody, unlike the massive separation along the 22 boattail angle baseline afterbody.

NASA ERA Integrated CFD for Wind Tunnel Testing of Hybrid Wing-Body Configuration

The NASA Environmentally Responsible Aviation (ERA) Project explored enabling technologies to reduce impact of aviation on the environment. One project research challenge area was the study of advanced airframe and engine integration concepts to reduce community noise and fuel burn. To address this challenge, complex wind tunnel experiments at both the NASA Langley Research Center’s (LaRC) 14’x22’ and the Ames Research Center’s 40’x80’ 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 vehicle aerodynamics. This paper presents a high-level summary of the CFD simulations that NASA performed in support of the model integration hardware design as well as the development of some CFD simulation guidelines 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.

Computational Analysis of Ares I Roll Control System Jet Interaction Effects on Rolling Moment

*† ‡ The computational flow solver USM3D was used to investigate the jet interaction effects from the roll control system on the rolling moment of the Ares I full protuberance configuration at wind tunnel Reynolds numbers. Solutions were computed at freestream Mach numbers from M = 0.5 to M = 5 at the angle of attack 0°, at the angle of attack 3.5° for a roll angle of 120°, and at the angle of attack 7° for roll angles of 120° and 210°. Results indicate that the RoCS housing provided a beneficial jet interaction effect on vehicle rolling moment for M ≥ 0.9. Most of the components downstream of the roll control system housing contributed to jet interaction penalties on vehicle rolling moment.

Transonic Investigation of Two-Dimensional Nozzles Designed for Supersonic Cruise

An experimental and computational investigation has been conducted to determine the off-design uninstalled drag characteristics of a two-dimensional convergent-divergent nozzle designed for a supersonic cruise civil transport. The overall objectives were to: (1) determine the effects of nozzle external flap curvature and sidewall boattail variations on boattail drag; (2) develop an experimental data base for 2D nozzles with long divergent flaps and small boattail angles and (3) provide data for correlating computational fluid dynamic predictions of nozzle boattail drag. The experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.80 to 1.20 at nozzle pressure ratios up to 9. Three-dimensional simulations of nozzle performance were obtained with the computational fluid dynamics code PAB3D using turbulence closure and nonlinear Reynolds stress modeling. The results of this investigation indicate that excellent correlation between experimental and predicted results was obtained for the nozzle with a moderate amount of boattail curvature. The nozzle with an external flap having a sharp shoulder (no curvature) had the lowest nozzle pressure drag. At a Mach number of 1.2, sidewall pressure drag doubled as sidewall boattail angle was increased from 4deg to 8deg. Reducing the height of the sidewall caused large decreases in both the sidewall and flap pressure drags. Summary