Melissa B. Carter

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.

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.

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

USM3D Predictions of Supersonic Nozzle Flow

This study focused on the NASA Tetrahedral Unstructured Software System CFD code (USM3D) capability to predict supersonic plume flow. Previous studies, published in 2004 and 2009, investigated USM3Ds results versus historical experimental data. This current study continued that comparison however focusing on the use of the volume souring to capture the shear layers and internal shock structure of the plume. This study was conducted using two benchmark axisymmetric supersonic jet experimental data sets. The study showed that with the use of volume sourcing, USM3D was able to capture and model a jet plumes shear layer and internal shock structure.

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.

USM3D Analysis of Low Boom Configuration (Invited)

In the past few years considerable improvement was made in NASA’s in house boom prediction capability. As part of this improved capability, the USM3D Navier-Stokes flow solver, when combined with a suitable unstructured grid, went from accurately predicting boom signatures at 1 body length to 10 body lengths. Since that time, the research emphasis has shifted from analysis to the design of supersonic configurations with boom signature mitigation In order to design an aircraft, the techniques for accurately predicting boom and drag need to be determined. This paper compares CFD results with the wind tunnel experimental results conducted on a Gulfstream reduced boom and drag configuration. Two different wind-tunnel models were designed and tested for drag and boom data. The goal of this study was to assess USM3D capability for predicting both boom and drag characteristics. Overall, USM3D coupled with a grid that was sheared and stretched was able to reasonably predict boom signature. The computational drag polar matched the experimental results for a lift coefficient above 0.1 despite some mismatch in the predicted lift-curve slope.

Computational Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Takeoff and Landing Conditions

As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Design of the Cruise and Flap Airfoil for the X-57 Maxwell Distributed Electric Propulsion Aircraft

A computational and design study on an airfoil and high-lift flap for the X-57 Maxwell Distributed Electric Propulsion (DEP) testbed aircraft was conducted. The aircraft wing sizing study resulted in a wing area of 66.67 ft2 and aspect ratio of 15 with a design requirement of Vstall = 58 KEAS, at a gross weight of 3,000 lb. To meet this goal an aircraft CL,max of 4.0 was required. The design cruise condition is 150 KTAS at 8,000 ft. This resulted in airfoil requirements of cl ~ 0.90 for the cruise condition at Re = 2.35 x 106. A flapped airfoil with a cl,max of approximately 2.5 or greater, at Re = 1.0 x 106, was needed to have enough lift to meet the stall requirement with the DEP system. MSES computational analyses were conducted on the GAW-1, GAW-2, and the NACA 5415 airfoil sections, however they had limitations in either high drag or low cl,max on the cruise airfoil, which was the impetus for a new design. A design was conducted to develop a low drag airfoil for the X-57 cruise conditions with high cl,max. The final design was the GNEW5BP93B airfoil with a minimum drag coefficient of cd = 0.0053 at cl = 0.90 and achieved laminar flow back to 69% chord on the upper surface and 62% chord on the lower surface. With fully turbulent flow, the drag increases to cd = 0.0120. The predicted maximum lift with turbulent flow is a cl,max of 1.95 at  = 19°. The airfoil is characterized by relatively flat pressure gradient regions on both surfaces at  = 0°, and aft camber to get extra lift out of the lower surface concave region. A 25% chord slotted flap was designed and analyzed with MSES for a 30° flap deflection. Additional 30° and 40° flap deflection analyses for two flap positions were conducted with USM3D using several turbulence models, for two angles of attack, to assess near cl,max with varied flap position. The maximum cl varied between 2.41 and 3.35. An infinite-span powered high-lift study was conducted on a GAW-1 constant chord 40° flapped airfoil section with FUN3D to quantify the airfoil lift increment that can be expected from a DEP system. The 16.7 hp/propeller blown wing increases the maximum CL from 3.45 to CL = 6.43, which is an effective q ratio of 1.86. This indicates that if the unblown high-lift flapped airfoil of the X-57 airplane achieves a cl,max of 2.78, then the high-lift augmentation blowing could yield a sectional lift coefficient of approximately 4.95 at cl,max. Finally, a computational study was conducted with FUN3D on an infinite-span constant chord GAW-1 cruise airfoil to determine the impact of high-lift propeller diameter to wing chord ratio on the lift increment of the DEP system. A constant diameter propeller and nacelle size were used in the study. Three computational grids were made with airfoil chords of 0.5*chord, 1.0*chord, and 2.0*chord. Results of the propeller diameter to wing chord ratio study indicated that the blown to unblown CL ratio increased as the chord was decreased. However, because of the increase in relative size of the high-lift nacelle to the wing, which impacted wing lift performance, the study indicated that a propeller diameter to wing chord ratio of 1.0 gives the overall best maximum lift on the wing with the DEP system.

Numerical Examination of Shock Generator Geometries and Nozzle Plume Effects on Pressure Signature

The NASA Advanced Air Vehicles Program, Commercial Supersonics Technology Project seeks to advance tools and techniques to make over-land supersonic flight feasible. In this study, preliminary computational results are presented for future tests in the NASA Ames 9’ x 7’ supersonic wind tunnel to be conducted in early 2016. Shock-plume interactions and their effect on pressure signature are examined for six model geometries. Near-field pressure signatures are assessed using the CFD code USM3D to model the proposed test geometries in free-air. Additionally, results obtained using the commercial grid generation software Pointwise R

Computational and Experimental Study of Supersonic Nozzle Flow and Shock Interactions

This study focused on the capability of NASA Tetrahedral Unstructured Software System’s CFD code USM3D capability to predict the interaction between a shock and supersonic plume flow. Previous studies, published in 2004, 2009 and 2013, investigated USM3D’s supersonic plume flow results versus historical experimental data. This current study builds on that research by utilizing the best practices from the early papers for properly capturing the plume flow and then adding a wedge acting as a shock generator. This computational study is in conjunction with experimental tests conducted at the Glenn Research Center 1’x1’ Supersonic Wind Tunnel. The comparison of the computational and experimental data shows good agreement for location and strength of the shocks although there are vertical shifts between the data sets that may be do to the measurement technique.