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NASA NTRS2001

Revalidation of the NASA Ames 11-by 11-Foot Transonic Wind Tunnel with a Commercial Airplane Model

The 11-By 11-Foot Transonic leg of the Unitary Plan Wind Tunnel (UPWT) was modernized to improve tunnel performance, capability, productivity, and reliability. Wind tunnel tests to demonstrate the readiness of the tunnel for a return to production operations included an Integrated Systems Test (IST), calibration tests, and airplane validation tests. One of the two validation tests was a 0.037-scale Boeing 777 model that was previously tested in the 11-By 11-Foot tunnel in 1991. The objective of the validation tests was to compare pre-modernization and post-modernization results from the same airplane model in order to substantiate the operational readiness of the facility. Evaluation of within-test, test-to-test, and tunnel-to-tunnel data repeatability were made to study the effects of the tunnel modifications. Tunnel productivity was also evaluated to determine the readiness of the facility for production operations. The operation of the facility, including model installation, tunnel operations, and the performance of tunnel systems, was observed and facility deficiency findings generated. The data repeatability studies and tunnel-to-tunnel comparisons demonstrated outstanding data repeatability and a high overall level of data quality. Despite some operational and facility problems, the validation test was successful in demonstrating the readiness of the facility to perform production airplane wind tunnel%, tests.

Kmak, Frank J.↗

NASA NTRS2006

Advanced Visualization of Experimental Data in Real Time Using LiveView3D

LiveView3D is a software application that imports and displays a variety of wind tunnel derived data in an interactive virtual environment in real time. LiveView3D combines the use of streaming video fed into a three-dimensional virtual representation of the test configuration with networked communications to the test facility Data Acquisition System (DAS). This unified approach to real time data visualization provides a unique opportunity to comprehend very large sets of diverse forms of data in a real time situation, as well as in post-test analysis. This paper describes how LiveView3D has been implemented to visualize diverse forms of aerodynamic data gathered during wind tunnel experiments, most notably at the NASA Langley Research Center Unitary Plan Wind Tunnel (UPWT). Planned future developments of the LiveView3D system are also addressed.

Schwartz, Richard J.↗

NASA NTRS2007

Virtual Diagnostics Interface: Real Time Comparison of Experimental Data and CFD Predictions for a NASA Ares I-Like Vehicle

Virtual Diagnostics Interface technology, or ViDI, is a suite of techniques utilizing image processing, data handling and three-dimensional computer graphics. These techniques aid in the design, implementation, and analysis of complex aerospace experiments. LiveView3D is a software application component of ViDI used to display experimental wind tunnel data in real-time within an interactive, three-dimensional virtual environment. The LiveView3D software application was under development at NASA Langley Research Center (LaRC) for nearly three years. LiveView3D recently was upgraded to perform real-time (as well as post-test) comparisons of experimental data with pre-computed Computational Fluid Dynamics (CFD) predictions. This capability was utilized to compare experimental measurements with CFD predictions of the surface pressure distribution of the NASA Ares I Crew Launch Vehicle (CLV) - like vehicle when tested in the NASA LaRC Unitary Plan Wind Tunnel (UPWT) in December 2006 - January 2007 timeframe. The wind tunnel tests were conducted to develop a database of experimentally-measured aerodynamic performance of the CLV-like configuration for validation of CFD predictive codes.

Schwartz, Richard J.↗

NASA NTRS2007

Virtual Diagnostics Interface: Real Time Comparison of Experimental Data and CFD Predictions for a NASA Ares I-Like Vehicle

Schwartz, Richard J.↗

NASA NTRS2007

Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

Erickson, Gary E.↗

NASA NTRS2008

Unified Application Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack (alpha). The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Erickson, Gary E.↗

NASA NTRS2009

Investigation of Seal-to-Floor Effects on Semi-Span Transonic Models

In an effort to achieve the maximum possible Reynolds number (Re) when conducting production testing for flight loads aerodynamic databases, it has been the preferred practice of The Boeing Company / Commercial Airplanes (BCA) -- Loads and Dynamics Group since the early 1990's to test large scale semi-span models in the 11- By 11-Foot Transonic Wind Tunnel (TWT) leg of the Unitary Plan Wind Tunnel (UPWT) at the NASA Ames Research Center (ARC). There are many problems related to testing large scale semi-span models of high aspect ratio flexible transport wings, such as; floor boundary layer effects, wing spanwise wall effects, solid blockage buoyancy effects, floor mechanical interference effects, airflow under the model effects, or tunnel flow gradient effects. For most of these issues, BCA has developed and implemented either standard testing methods or numerical correction schemes and these will not be discussed in this document. Other researchers have reported on semi-span transonic testing correction issues, however most of the reported research has been for low Mach testing. Some of the reports for low Mach testing address the difficult problem of preventing undesirable airflow under a semi-span model while ensuring unrestricted main balance functionality, however, for transonic models this issue has gone unresolved. BCA has been cognizant for sometime that there are marked differences in wing pressure distributions from semi-span transonic model testing than from full model or flight testing. It has been suspected that these differences are at least in part due to airflow under the model. Previous efforts by BCA to address this issue have proven to be ineffective or inconclusive and in one situation resulted in broken hardware. This paper reports on a Boeing-NASA collaborative investigation based on a series of small tests conducted between June 2006 and November 2007 in the 11 by 11 foot Transonic Wind Tunnel at NASA Ames on three large commercial jet transport configurations to assess the effects of sealing a semi-span model to the floor and to investigate efficient sealing and testing techniques. This document will show how sealing the model to the floor has a small but remarkably far reaching spanwise effect on wing pressures, wing local section forces and wing force summations.

Sleppy, Mark A.↗

NASA NTRS2010

Unified Application of Vapor Screen Flow Visualization and Pressure Sensitive Paint Measurement Techniques to Vortex- and Shock Wave-Dominated Flow Fields

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at high angles of attack. The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel (8-Foot TPT).

Erickson, Gary E.↗

NASA NTRS2011

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

Throughout three full design analysis cycles, the Ares I project within the Constellation program has consistently relied on the Boeing Polysonic Wind Tunnel (PSWT) for aerodynamic testing of the subsonic, transonic and supersonic portions of the atmospheric flight envelope (Mach=0.5 to 4.5). Each design cycle required the development of aerodynamic databases for the 6 degree-of-freedom (DOF) forces and moments, as well as distributed line-loads databases covering the full range of Mach number, total angle-of-attack, and aerodynamic roll angle. The high fidelity data collected in this facility has been consistent with the data collected in NASA Langley s Unitary Plan Wind Tunnel (UPWT) at the overlapping condition ofMach=1.6. Much insight into the aerodynamic behavior of the launch vehicle during all phases of flight was gained through wind tunnel testing. Important knowledge pertaining to slender launch vehicle aerodynamics in particular was accumulated. In conducting these wind tunnel tests and developing experimental aerodynamic databases, some challenges were encountered and are reported as lessons learned in this paper for the benefit of future crew launch vehicle aerodynamic developments.

Pinier, Jeremy T.↗

NASA NTRS2011

Wedge and Conical Probes for the Instantaneous Measurement of Free-Stream Flow Quantities at Supersonic Speeds

Wedge and conical shaped probes for the measurement of free-stream flow quantities at supersonic speeds have been tested in both wind tunnel and flight. These probes have improved capabilities over similar ones used in the past. Through the use of miniature pressure sensors, that are located inside the probes, they are able to provide instantaneous measurements of a time-varying environment. Detailed herein are the results of the tests in NASA Langley Researcher Center s Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8 and 2.0, as well as flight tests carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15 aircraft up to Mach numbers of 1.9. In the flight tests the probes were attached to a fixture on the underside of the F-15 fuselage. Problems controlling the velocity of the flow through the conical probe, required for accurate temperature measurements, are noted, as well as some calibration problems of the miniature pressure sensors that impact the accuracy of the measurements.

Bobbitt, Percy J.↗

NASA NTRS2012

Numerical Predictions of Sonic Boom Signatures for a Straight Line Segmented Leading Edge Model

A sonic boom wind tunnel test was conducted on a straight-line segmented leading edge (SLSLE) model in the NASA Langley 4- by 4- Foot Unitary Plan Wind Tunnel (UPWT). The purpose of the test was to determine whether accurate sonic boom measurements could be obtained while continuously moving the SLSLE model past a conical pressure probe. Sonic boom signatures were also obtained using the conventional move-pause data acquisition method for comparison. The continuous data acquisition approach allows for accurate signatures approximately 15 times faster than a move-pause technique. These successful results provide an incentive for future testing with greatly increased efficiency using the continuous model translation technique with the single probe to measure sonic boom signatures. Two widely used NASA codes, USM3D (Navier-Stokes) and CART3D-AERO (Euler, adjoint-based adaptive mesh), were used to compute off-body sonic boom pressure signatures of the SLSLE model at several different altitudes below the model at Mach 2.0. The computed pressure signatures compared well with wind tunnel data. The effect of the different altitude for signature extraction was evaluated by extrapolating the near field signatures to the ground and comparing pressure signatures and sonic boom loudness levels.

Elmiligui, Alaa A.↗

NASA NTRS2012

Wind-Tunnel and Flight Test Results for the Measurements of Flow Variables at Supersonic Speeds Using Improved Wedge and Conical Probes

The results of supersonic wind-tunnel tests on three probes at nominal Mach numbers of 1.6, 1.8 and 2.0 and flight tests on two of these probes up to a Mach number of 1.9 are described. One probe is an 8 deg. half-angle wedge with two total-pressure measurements and one static. The second, a conical probe, is a cylinder that has a 15 deg., semi-angle cone tip with one total-pressure orifice at the apex and four static-pressure orifices on the surface of the cone, 90 deg. apart, and about two-thirds of the distance from the cone apex to the base of the cone. The third is a 2 deg. semi-angle cone that has two static ports located 180 deg. apart about 1.5 inches behind the apex of the cone. The latter probe was included since it has been the "probe of choice" for wind-tunnel flow-field pressure measurements (or one similar to it) for the past half-century. The wedge and 15 deg. conical probes used in these tests were designed for flight diagnostic measurements for flight Mach numbers down to 1.35 and 1.15 respectively, and have improved capabilities over earlier probes of similar shape. The 15. conical probe also has a temperature sensor that is located inside the cylindrical part of the probe that is exposed to free-stream flow through an annulus at the apex of the cone. It enables the determination of free-stream temperature, density, speed of sound, and velocity, in addition to free-stream pressure, Mach number, angle of attack and angle of sideslip. With the time-varying velocity, acceleration can be calculated. Wind-tunnel tests of the two probes were made in NASA Langley Research Center fs Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8, and 2.0. Flight tests were carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15B aircraft up to Mach numbers of 1.9. The probes were attached to a fixture, referred to as the Centerline Instrumented Pylon (CLIP), under the fuselage of the aircraft. Problems controlling the velocity of the flow through the conical probe required for accurate temperature measurements are noted, as well as some calibration problems of the miniature pressure sensors that required a re-calculation of the flow variables. Data are presented for angle of attack, pressure and Mach number obtained in the wind tunnel and in flight. In the wind tunnel some transient data were obtained by translating the probes through the shock flow field created by a bump on the wind-tunnel wall.

Bobbitt, Percy J.↗

NASA NTRS2013

Effectiveness of a Wedge Probe to Measure Sonic Boom Signatures in a Supersonic Wind Tunnel

A wind tunnel investigation was conducted in the Langley Unitary Plan Wind Tunnel (UPWT) to determine the effectiveness of a wedge probe to measure sonic boom pressure signatures compared to a slender conical probe. A generic business jet model at a constant angle of attack and at a single model to probe separation distance was used to generate a sonic boom signature. Pressure signature data were acquired with both the wedge probe and a slender conical probe for comparison. The test was conducted at a Mach number of 2.0 and a free-stream unit Reynolds number of 2 million per foot. The results showed that the wedge probe was not effective in measuring the sonic boom pressure signature of the aircraft model in the supersonic wind tunnel. Data plots and a discussion of the results are presented. No tabulated data or flow visualization photographs are included.

Wilcox, Floyd J., Jr.↗

NASA NTRS2015

11 Foot Unitary Plan Tunnel Facility Optical Improvement Large Window Analysis

The test section of the 11 by 11-foot Unitary Plan Transonic Wind Tunnel (11-foot UPWT) may receive an upgrade of larger optical windows on both the North and South sides. These new larger windows will provide better access for optical imaging of test article flow phenomena including surface and off body flow characteristics. The installation of these new larger windows will likely produce a change to the aerodynamic characteristics of the flow in the Test Section. In an effort understand the effect of this change, a computational model was employed to predict the flows through the slotted walls, in the test section and around the model before and after the tunnel modification. This report documents the solid CAD model that was created and the inviscid computational analysis that was completed as a preliminary estimate of the effect of the changes.

computer aided design↗

NASA NTRS2017

Recent Advancements in the Infrared Flow Visualization System for the NASA Ames Unitary Plan Wind Tunnels

The following details recent efforts undertaken at the NASA Ames Unitary Plan wind tunnels to design and deploy an advanced, production-level infrared (IR) flow visualization data system. Highly sensitive IR cameras, coupled with in-line image processing, have enabled the visualization of wind tunnel model surface flow features as they develop in real-time. Boundary layer transition, shock impingement, junction flow, vortex dynamics, and buffet are routinely observed in both transonic and supersonic flow regimes all without the need of dedicated ramps in test section total temperature. Successful measurements have been performed on wing-body sting mounted test articles, semi-span floor mounted aircraft models, and sting mounted launch vehicle configurations. The unique requirements of imaging in production wind tunnel testing has led to advancements in the deployment of advanced IR cameras in a harsh test environment, robust data acquisition storage and workflow, real-time image processing algorithms, and evaluation of optimal surface treatments. The addition of a multi-camera IR flow visualization data system to the Ames UPWT has demonstrated itself to be a valuable analyses tool in the study of new and old aircraft/launch vehicle aerodynamics and has provided new insight for the evaluation of computational techniques.

Infrared Thermography↗

NASA NTRS2017

Retroreflective Background Oriented Schlieren Imaging Results from the NASA Plume/Shock Interaction Test

In support of NASA's Commercial Supersonics Technology (CST) project, a test was conducted in the 9-by-7 ft. supersonic section of the NASA Ames Unitary Plan Wind Tunnel (UPWT). The tests were designed to study the interaction of shocks with a supersonic jet characteristic of those that may occur on a commercial supersonic aircraft. Multiple shock generating geometries were tested to examine the interaction dynamics as they pertain to sonic boom mitigation. An integral part of the analyses of these interactions are the interpretation of the data generated from the retroreflective Background Oriented Schlieren (RBOS) imaging technique employed for this test. The regularization- based optical flow methodology used to generate these data is described. Sample results are compared to those using normalized cross-correlation. The reduced noise, additional feature detail, and fewer false artifacts provided by the optical flow technique produced clearer time-averaged images, allowing for better interpretation of the underlying flow phenomena. These images, coupled with pressure signatures in the near field, are used to provide an overview of the detailed interaction flowfields.

optical flow↗

NASA NTRSDate not supplied

Evaluation of CFD as a Surrogate for Wind-Tunnel Testing for Mach 2.4 to 4.6 - Project Overview

The debate over when wind-tunnel testing (WTT) will be replaced by Computational Fluid Dynamics (CFD) comes and goes. More recently the debate has subsided with a more collaborative spirit between practitioners of these two disciplines resulting in significant improvements in the outcomes of both. There may come a time, however, when CFD has sufficient accuracy to supplant WTT as the dominant or perhaps only tool for aerodynamic simulation. If and/or when that happens, financial pressure result in efforts to close or severely limit the operations of wind tunnels. Presumably additional resources will go toward CFD in order to generate aerodynamic databases, load environments, and new aero/fluid-dynamic knowledge. It is therefore important to develop appropriate processes by which wind-tunnel closure decisions are made to ensure that facilities critical to industry and government research and development aren’t closed prematurely without ensuring that the existing CFD tools have sufficient accuracy and low-enough cost (and enough experts and computational facilities) to take on the traditional role of wind tunnels. This paper will describe a project intended to answer the specific question of whether CFD can replace WTT for the limited Mach-number range 2.4 to 4.6. The wind tunnel being examined in this context is the high-speed leg of the Unitary Plan Wind Tunnel at NASA’s Langley Research Center (LaRC UPWT).

James C. Ross↗

NASA NTRSDate not supplied

Analysis and Comparison of Surface Roughness Effects on Pressure Data from SLS Wind Tunnel Test

Unsteady flows are becoming a more characterized field of study as advancements in technology is allowing for more high-speed, time-resolved data acquisition. It is important to understand these flows as the shock-wave boundary-layer interactions and separation they are involved with can generate extreme loads caused by high pressure and temperature on the body of a flight vehicle. It is vital to characterize these loads, as they can impede upon the structural integrity of a high-speed system. A common instrument utilized today to characterize pressure distributions caused by unsteady flows are unsteady pressure transducers. Set flush with the surface of a flight vehicle or wind tunnel model, these transducers can provide a pressure distribution, capturing the frequency of unsteady loads that could impinge upon the test article. A couple negative aspects of this instrumentation is that they are expensive, have long lead times, and a pressure distribution can only be determined in one discrete location where the transducer is installed. An optical diagnostic known as unsteady pressure-sensitive paint (uPSP) can help to satisfy these deficiencies. This is a spray paint that can be applied to a surface of a model to provide a global pressure distribution across the whole painted region.2 A complete picture of the fluid dynamic pressures occurring on the surface of a test article can be analyzed, especially in locations where it may have been impossible to install a transducer. Pressure-sensitive paint is also less expensive than traditional pressure transducers, which are, on average, about $2,000 each including the cost of installation.3 Unsteady PSP will not replace unsteady pressure transducers, but the transducers can be used to validate the PSP response to the flow and measure frequencies higher than the frequency response of the paint. One issue that can arise using these tools simultaneously, is that the surface roughness of the uPSP can affect the pressure readings of the transducers. Applying the uPSP to the surface of the model is going to affect the flow in some way as the topography of the model is now different than if it was not painted. One study with steady PSP conducted by Amer, Obara, and Liu at NASA Langley concluded that the PSP surface roughness was minimally intrusive at different test conditions with various lift, drag, etc., however, it was noted that data from some pressure transducers had decreased, likely due to the uneven application of the PSP in the localized region of the transducers.9 Sugioka et. al. tested various formulations of uPSP to study surface roughness effects on a Common Research Model (CRM) in transonic flow. The influence of uPSP on the surface of the model, depending on how it was formulated, was noted to have the possibility of relocating the shock wave present in the flow. This study was conducted to examine and compare transducer data acquired in December 2017 of a 4% scale model of the Space Launch System (SLS) performed at NASA Ames Research Center (ARC) Unitary Plan Wind Tunnel (UPWT). One goal of the test was to demonstrate to the customer the potential of the uPSP system developed at NASA ARC. The effect of uPSP on transducer signals also became an interest in this experiment and is the reason for this data analysis. The results of this study were discussed with the SLS team to help them make decisions regarding the next test on an SLS model with the PSP and Unitary Plan Wind Tunnel teams.

pressure-sensitive paint↗