National Transonic Facility Research Articles

Effect of Leading-Edge Cranks on Stability and Control of Active-Flow-Control-Enabled Tailless Aircraft

The Swept Wing Flow Test (SWIFT) is a tailless unmanned combat aerial vehicle (UCAV) model to be tested at high Reynolds numbers in NASA’s National Transonic Facility. The model is designed around a -shaped wing with a single, large crank at its leading edge (LE). It suffers from an unstable nose-up pitch departure resulting from flow separation augmented by the LE crank. A small-scale, modular wind tunnel model () was built that allowed for changes in the crank angle by increasing the outboard wing sweep. Eliminating the crank entirely increased the and changed the sign of pitch departure, thus exposing the significance of the LE crank. The model was equipped with sweeping jet actuators that could be individually enabled by valves located at the actuator inlets, allowing one to explore the role of active flow control (AFC) in expanding the model’s longitudinal stability margins and controlling its yaw while being cognizant of the coupling between changes in the model’s planform and their effect on AFC. Test results indicated that selective actuation depending on the model’s attitude modified the flow and dramatically increased the trimmed , while further suggesting that the actuation should dynamically change with incidence to improve AFC efficacy.

Numerical Simulation of the High-Speed Leg of the National Transonic Facility

Numerical simulations for the flow inside the high-speed leg of the National Transonic Facility were conducted. Simulations were conducted for three configurations: the empty tunnel, an installed body of revolution, and the NASA Common Research Model installed in the test section. Simulations were performed for a test-section Mach numbers of 0.7 and 0.85 and a corresponding Reynolds number of 8 million per ft. The numerical simulations were performed using the NASA USM3d-ME Navier-Stokes flow solver with the Spalart-Almaras one-equation turbulence model and were run in the steady-state mode. USM3D-ME supports solutions on mixed-element grids which provide improved numerical simulation using the flow-aligned anisotropic hexahedral or prismatic cells in the boundary layer and isotropic tetrahedral cells away from the boundary layer. A controller was developed that automated the outflow boundary and streamlined the process of running multiple simulations. The use of a dynamic outflow boundary was the key parameter to drive a simulation to the desired tunnel test section conditions. The numerical simulations captured the expected flow features in the NTF test section and in the tunnel plenum. The simulations revealed that the separation of the flow inside the diffuser is asymmetric and more extensive toward the bottom wall. Analysis of the flowfield was conducted, and the computed drag coefficient was compared to wind tunnel data.

Transport-Based Transition Prediction for the Common Research Model Natural Laminar Flow Configuration

This paper presents the application of a new -based one-equation transition transport model to the Common Research Model Natural Laminar Flow Configuration (CRM-NLF), which was the central test case of the 1st AIAA CFD Transition Modeling and Prediction Workshop. The new model uses a simplified formulation of the Arnal, Habiballah, and Delcourt transition criterion that is stability based. The model is formulated in a grid-point-local manner and is coupled to the Menter shear stress transport eddy-viscosity turbulence model. The flow conditions applied are exactly those that were reported from the National Transonic Facility wind tunnel test campaign. The comparisons of the results yielded by the new model in terms of the transition fronts along the wing span show a very good match with the transition fronts derived from the temperature-sensitive paint images taken during the wind tunnel test and with computational result from an -based Two--factor strategy using critical -factors for Tollmien–Schlichting and crossflow instabilities and applying a local linear stability code.

Application of femtosecond-laser tagging for unseeded velocimetry in a large-scale transonic cryogenic wind tunnel

Femtosecond laser electronic excitation tagging (FLEET) velocimetry was applied in the National Transonic Facility and assessed for its use in providing quantitative velocity measurements in a large-scale cryogenic wind tunnel. Comparisons of freestream results with theory and existing tunnel measurements indicate that FLEET velocity measurements agree within 1% of the tunnel reference, while the largest error among all conditions remained within 2.5%. For single-shot velocity measurements, binning of the FLEET intensity data improved the signal-to-noise ratio sufficiently to provide measurement precisions on the order of 1% of the freestream velocity. After confirmation of system performance, spatially resolved velocity profile measurements were obtained in the wake downstream of the Common Research Model wing. Effects of the model angle-of-attack on velocity deficit profiles were explored, and a two-dimensional, one-component velocity map resolving the wake region was constructed by scanning the laser’s position within the test section. The experimental campaign described herein represents the first non-intrusive, quantitative measurements of velocity made in this facility since its inception in 1984.

Evaluation of Cryogenic Force Balance Calibration Methodologies Relative to Wind Tunnel Results

Independent tests of the NASA Common Research Model at NASA’s National Transonic Facility and the European Transonic Windtunnel reveal differences at low operating temperatures and high Reynolds numbers that warrant further investigation. Since each facility used their own wind tunnel balance for their tunnel entry, one suggestion for the differences was the temperature compensation methodology developed and applied for each balance. This hypothesis is explored through simulation and experiment. Independent calibrations of NASA’s NTF-118A balance at NASA Langley Research Center and European Transonic Windtunnel reveal differences in the thermal compensation of the normal force and pitching moment primary sensitivities with temperature, while the axial force primary sensitivities are in good agreement. The application of the balance calibrations performed at NASA and European Transonic Windtunnel to the prior wind tunnel data suggests that the thermal compensation differences are an order of magnitude lower than the differences observed between the wind tunnel aerodynamic coefficients. Thus, the balance temperature compensation methodologies used by NASA and European Transonic Windtunnel are not a major contributor to the wind tunnel differences.

A parametric sensitivity study by numerical simulations on plume dispersion of the exhaust from a cryogenic wind tunnel

The low temperature plume exhausted from a cryogenic wind tunnel may sink down, posing a severe threat to public health and safety. Quantitative risk assessment of cryogenic plume flow behavior therefore plays an important role in the design and optimization of a cryogenic wind tunnel. A numerical model with a modified Hertz-Knudsen relation considering the phase change physics of the small quantity of water involved is applied to analyze the dispersion of the low temperature nitrogen plume exhausted from a 0.3 m cryogenic wind tunnel. The homogeneous multiphase flow is modeled using the single-fluid mixture model. A model validation is presented for the exhaust plume from the US National Transonic Facility (NTF). The predicted results are found to be better than those predicted by National Aeronautics and Space Administration (NASA)’s two-stage analytical model. The influences of the environmental wind speed, the environmental wind temperature, the relative humidity, and the exhaust flow rate, on low temperature nitrogen plume dispersion are obtained. In particular, the parametric sensitivities of different influence factors are analyzed. The environmental wind temperature and the exhaust flow rate of the nitrogen gas have greater impact on the temperature of the plume near the ground than do the environmental wind speed and the relative humidity. The exhaust flow rate of the nitrogen gas has greater impact on the oxygen concentration near the ground than does the environmental wind speed, while the environmental wind temperature and the relative humidity have negligible impacts. The results provide guidance on the operation and design improvement of a cryogenic gaseous nitrogen discharge system to avoid its potential hazards.

Retrospective on the Common Research Model for Computational Fluid Dynamics Validation Studies

The development of a wing–body–nacelle–pylon–horizontal-tail configuration for a common research model is presented, with the focus on the aerodynamic design of the wing. A contemporary transonic supercritical wing design is developed with aerodynamic characteristics that are well behaved and of high performance for configurations with and without the nacelle–pylon group. The horizontal tail is robustly designed for dive Mach number conditions and is suitably sized for typical stability and control requirements. The fuselage is representative of a wide-body commercial transport aircraft; it includes a wing–body fairing, as well as a scrubbing seal for the horizontal tail. The nacelle is a single-cowl high-bypass-ratio flowthrough design with an exit area sized to achieve a natural unforced mass flow ratio typical of commercial aircraft engines at cruise. The simplicity of this un-bifurcated unpowered nacelle geometry facilitates experimental data collection and grid generation efforts for computational fluid dynamics (CFD) validations. Detailed aerodynamic performance data are provided; however, this information was originally presented in such a manner as to not bias CFD predictions planned for the fourth AIAA CFD Drag Prediction Workshop in 2009 (DPW-IV); the DPW-IV incorporated this common research model into its blind test cases. The fifth AIAA CFD Drag Prediction Workshop in 2012 and the sixth AIAA CFD Drag Prediction Workshop in 2016 also used this configuration as the basis of their test cases. The CFD results presented include wing pressure distributions with and without the nacelle–pylon Mach lift-to-drag ratio (ML/D) trend lines, as well as the drag-divergence curves. The design point for the wing–body configuration is within 1% of its maximum ML/D. The original plans to test the common research model in the National Transonic Facility at NASA Langley Research Center and the 11-by-11 ft wind tunnel at NASA Ames Research Center are discussed. Furthermore, wind-tunnel models have been replicated by the Japan Aerospace Exploration Agency, ONERA–The French Aerospace Lab, the University of Washington, as well as by numerous other laboratories. A Web site has been established to collect, archive, and freely provide wind-tunnel data from these various experimental campaigns. The geometry definition has been extended to include a high-lift system. It has been used to form the basis for experimental swept-wing icing studies. Finite element models have been developed for the wind-tunnel environment, as well as to be more representative of a full-scale aircraft in flight. A simple search for technical papers related to this configuration yields thousands of results. Needless to say, with a decade in retrospect, the NASA Common Research Model has significantly exceeded the initial most-optimistic goals, and it has become the de facto standard test case for present-day transonic CFD validations.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel test data to NASA National Transonic Facility test data

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the European Transonic Wind Tunnel. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the wing/body/tail = 0° incidence configuration in the National Transonic Facility and in the European Transonic Wind Tunnel. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in both facilities but only the force and moment, and surface pressure data are presented herein.

Experimental Investigations on Common Research Model at ONERA-S1MA–Drag Prediction Workshop Numerical Results

This paper aims to present some of the experimental and numerical results obtained with the NASA–Boeing Common Research Model at ONERA. The wind tunnel model used in the present study is the ONERA Large Reference Model (1/16.835), which has the same geometry as the Common Resarch Model considered in the latest AIAA Drag Prediction Workshops. Experimental data have been collected from the ONERA-S1MA wind tunnel at Mach numbers between 0.30 and 0.95 and a mean aerodynamic chord Reynolds number of 5 million for four different configurations: wing/body only and wing/body with horizontal or vertical tails or both. Force and moment and pressure and surface flow measurements have been performed. Concerning the numerical study, all the Reynolds-averaged Navier–Stokes computations have been completed with the structured solver elsA, and the Spalart–Allmaras and with shear stress transport turbulence models have been used in addition to the quadratic constitutive relation. In this paper, configuration effects (increments due to horizontal and/or vertical tails) are assessed both numerically and experimentally for several Mach numbers and angles of attack. The delicate issue of flow separation at the wing/body junction is also addressed with the support of oil flow visualizations. elsA and S1MA skin pressure distributions are presented; the agreement is satisfactory except for some outboard wing sections at high lift levels. Finally, comparisons of drag and moment values including computational fluid dynamics and test data from different wind tunnels are proposed (S1MA, NASA Ames Research Center and National Transonic Facility, and European Transonic Wind Tunnel).

Model Deformation Corrections of NASA Common Research Model Using Computational Fluid Dynamics

Computational fluid dynamics-aided wind-tunnel data corrections for the static model deformation effect of the NASA-Common Research Model are performed for wind-tunnel data obtained at the Transonic Wind Tunnel of Japan Aerospace Exploration Agency, the National Transonic Facility, and the European Transonic Wind Tunnel. Although it is important to use the same configuration when comparing wind-tunnel data obtained at various facilities, this has not yet been performed with an appropriate validation. In this paper, the model deformation effects are estimated from the difference between the computational fluid dynamics results for a designed configuration and that for a deformed configuration considering support systems. Initially, the techniques for the model deformation correction are validated with two sets of National Transonic Facility wind-tunnel data at a Reynolds number of with different dynamic pressures. Then, the validated model deformation corrections are performed on the wind-tunnel data obtained at the Transonic Wind Tunnel, National Transonic Facility, and European Transonic Wind Tunnel. Differences in the drag coefficients among the three sets of wind-tunnel data fall within about 10 drag counts for the model deformation corrections and the Reynolds number corrections applied to the Transonic Wind Tunnel data to align the Reynolds number to .

Enhancements to the National Transonic Facility Semispan Force Measurement System

Recent wind-tunnel tests at the NASA Langley Research Center National Transonic Facility used high-pressure bellows to route air to the semispan model for evaluating aircraft circulation control testing techniques. The introduction of these bellows within the sidewall model support system impacted the performance of the sidewall mounted force measurement system. A capability has been developed to facilitate system-level calibrations that characterize the performance of the force measurement system under influence of static pressure tare and thermal effects. From the highest system-level perspective, the aerodynamic research being conducted using this system would benefit from a system-level calibration at conditions that most accurately simulate testlike operating conditions. Detail is given on the recent improvements and design modifications to improve performance and calibration of the system as well as recommendations for future improvements. Experimental data from recent testing using the force measurement system are presented, with the results supporting the necessity for future system enhancements to improve system performance.

Normalization of Wind-Tunnel Data for NASA Common Research Model

A wind-tunnel test of an 80%-scale copy of the NASA Common Research Model was performed by the Japan Aerospace Exploration Agency using its transonic wind tunnel. The wind-tunnel model was fabricated by the Japan Aerospace Exploration Agency in consultation with NASA Langley Research Center and AIAA Drag Prediction Workshop committee members. The static aerodynamic forces and surface pressure distributions on the wing of the Japan Aerospace Exploration Agency model were measured at a relatively low Reynolds number of due to tunnel capability limitations, where boundary-layer transition was simulated using optimized roughness. Measured data were compared with those of wind-tunnel tests of the Common Research Model obtained from NASA Langley Research Center’s National Transonic Facility as well as computational fluid dynamics predictions, both at a Reynolds number of . The comparison among these datasets required data normalization to the designed shape aligned at a reference condition, because the wing deformation of the wind-tunnel models deformed due to the dynamic pressure and the Reynolds number was different between the Japan Aerospace Exploration Agency and the National Transonic Facility tests. In this paper, the normalization procedures, in addition to basic wind-tunnel corrections, are proposed and the corrected results are examined.

Experimental Investigations of the NASA Common Research Model

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11 ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure, and surface flow visualization data were obtained in both facilities, but only the force and moment data are presented herein. Nacelle/pylon, tail effects, and tunnel-to-tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon produced an increase in drag, a decrease in lift, and a less nose-down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, nearly all of the data shown fall within the 2-sigma limits for repeatability. The tunnel-to-tunnel differences are negligible for lift and pitching moment, whereas the drag shows a difference of less than 10 counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

Turbulence Model Study for the Flow Around the NASA Common Research Model

A turbulence model study was performed for the wing–body configuration of NASA’s common research model on common hybrid grids provided by the Fifth Drag Prediction Workshop Committee and on an in-house-generated hybrid grid. Here, eddy-viscosity models with and without quadratic constitution relation extension, which accounts for the anisotropy of normal turbulent stresses, and a differential Reynolds stress model were applied. In this paper, first, the influence of turbulence models on the prediction of side-of-body separation in the wing–body region and on the overall aerodynamic force and moment coefficients is discussed, and later, the influence of grid refinement on the predictions is presented. Finally, the effect of the quadratic constitution relation extension on the predictions of eddy-viscosity models is demonstrated and performance of different turbulence models is discussed by comparing numerical predictions with the experimental data from NASA’s National Transonic Facility.

Investigation of Aeroelastic Effects on the NASA Common Research Model

Static fluid-structure coupled simulations were performed on NASA’s Common Research Model to assess the influence of aeroelastic effects on the numerical prediction of the overall aerodynamic coefficients and wing static pressure distributions. The numerical results of both rigid steady-state computational fluid dynamics and static aeroelastic coupled simulations were compared to the experimental data from wind tunnel test campaigns at NASA’s National Transonic Facility and the NASA Ames Research Center’s 11-Foot Transonic Wind Tunnel Facility. Coupled analyses were performed using an in-house simulation procedure built around the German Aerospace Research Center’s flow solver TAU and the commercial finite element analysis code NASTRAN®. The results show a considerable reduction of deviations between the computational results obtained during the fourth and fifth AIAA Computational Fluid Dynamics Drag Prediction Workshops and the measured data when aeroelastic wing deformations are taken into account.

Flow Disturbance Measurements in the National Transonic Facility

Recent flow measurements have been acquired in the National Transonic Facility to assess the test-section unsteady flow environment. The primary purpose of the test is to determine the feasibility of the facility to conduct laminar-flow-control testing and boundary-layer transition-sensitive testing at flight-relevant operating conditions throughout the transonic Mach number range. The facility can operate in two modes, warm and cryogenic test conditions for testing full and semispan-scaled models. Data were acquired for Mach and unit Reynolds numbers ranging from and collectively at air and cryogenic conditions. Measurements were made in the test section using a survey rake that was populated with 19 probes. Roll polar data at selected conditions were obtained to look at the uniformity of the flow disturbance field in the test section. Data acquired included mean total temperatures, mean and fluctuating static/total pressures, and mean and fluctuating hot-wire measurements. This paper focuses primarily on the unsteady pressure and hot-wire results. Based on the current measurements and previous data, an assessment was made that the facility may be a suitable facility for ground-based demonstrations of laminar-flow technologies at flight-relevant conditions in the cryogenic mode.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106 to 30×106 per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106 was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

Engineering Extrapolation to Flight Reynolds Number for Supercritical Airfoil Pressure Distribution Based on CFD Results

Pressure distribution of supercritical airfoil at flight Reynolds number could not be fully simulated except in cryogenic wind tunnel such as NTF (National Transonic Facility) and ETW (European Transonic Wind tunnel), which is costly and time resuming. This paper aimed to explore an engineering extrapolation to flight Reynolds number from low Reynolds number wind tunnel data for supercritical airfoil pressure distribution. However, the extrapolation method requiring plenty of data was investigated based on the CFD results for the reason of low cost and short period. Flows over a typical supercritical airfoil were numerically simulated by solving the two dimensional Navier-Stokes equations, with applications of ROE scheme spatial discretization and LU-SGS time march. Influence of computational grids convergence and turbulent models were investigated during the process of simulation. The supercritical airfoil pressure distribution were obtained with Reynolds numbers varied from 3.0×106to 30×106per airfoil chord, angles of attack from 0 degree to 6 degree and Mach numbers from 0.74 to 0.8. Simulated results indicated that weak shock existed on the upper surface of supercritical airfoil at cruise condition, that the shock location, shock strength and trailing edge pressure were dependent of Reynolds number, attack angles and Mach numbers. A similar parameter describing the Reynolds number effects factors was obtained by analyzing the relationship of shock wave location, shock front pressure and trailing edge pressure. Based on the similar parameter, airfoil pressure distribution at Reynolds number 30×106was obtained by extrapolation. It was shown that extrapolated result compared well with simulated result at Reynolds number 30×106, implying that the engineering method was at least promising applying to the extrapolation of low Reynolds number wind tunnel data.

National Transonic Facility Model and Tunnel Vibrations

Since coming online in 1984, the National Transonic Facility (NTF) cryogenic wind tunnel at the NASA Langley Research Center has provided unique high Reynolds number testing capability. While turbulence levels in the tunnel, expressed in terms of percent dynamic pressure, are typical of other transonic wind tunnels, the significantly increased load levels utilized to achieve flight Reynolds numbers, in conjunction with the unique structural design requirements for cryogenic operation, have brought forward the issue of model and model support structure vibrations. This paper reports new experimental measurements documenting aerodynamic and structural dynamics processes involved in such vibrations experienced in the NTF. In particular, evidence of local un-steady airloads developed about the model support strut is shown and related to well documented acoustic features known as Parker modes. Two-dimensional unsteady viscous computations illustrate this model support structure loading mechanism.

Investigations for Supersonic Transports at Transonic and Supersonic Conditions

Several computational studies were conducted as part of NASA s High Speed Research Program. Results of turbulence model comparisons from two studies on supersonic transport configurations performed during the NASA High-Speed Research program are given. The effects of grid topology and the representation of the actual wind tunnel model geometry are also investigated. Results are presented for both transonic conditions at Mach 0.90 and supersonic conditions at Mach 2.48. A feature of these two studies was the availability of higher Reynolds number wind tunnel data with which to compare the computational results. The transonic wind tunnel data was obtained in the National Transonic Facility at NASA Langley, and the supersonic data was obtained in the Boeing Polysonic Wind Tunnel. The computational data was acquired using a state of the art Navier-Stokes flow solver with a wide range of turbulence models implemented. The results show that the computed forces compare reasonably well with the experimental data, with the Baldwin-Lomax with Degani-Schiff modifications and the Baldwin-Barth models showing the best agreement for the transonic conditions and the Spalart-Allmaras model showing the best agreement for the supersonic conditions. The transonic results were more sensitive to the choice of turbulence model than were the supersonic results.