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NASA NTRS1983

Langley test highlights, 1982

A 20 ft vertical spin tunnel, a 30 by 60 ft tunnel, a 7 by 10 ft high speed tunnel, a 4 by 7 meter tunnel, an 8 ft transonic pressure tunnel, a transonic dynamics tunnel, a 16 ft transonic tunnel, a national transonic facility, a 0.3 meter transonic cryogenic tunnel, a unitary plan wind tunnel, a hypersonic facilities complex, an 8 ft high temperature tunnel, an aircraft noise reduction lab, an avionics integration research lab, a DC9 full workload simulator, a transport simulator, a general aviation simulator, an advanced concepts simulator, a mission oriented terminal area simulation (MOTAS), a differential maneuvering simulator, a visual/motion simulator, a vehicle antenna test facility, an impact dynamics research facility, and a flight research facility are all reviewed.

Source record↗

NASA NTRSDate not supplied

Validating uPSP with Transducers

Unsteady Pressure-Sensitive Paint (uPSP) technology provides simultaneously high-frequency and high-spatial resolution surface pressure fluctuation measurements. It has demonstrated value in wind tunnel test facilities to evaluate aerospace vehicle unsteady aerodynamics. The calibration of measurements from digital pixel counts into corresponding physical units relies on both lab-based and in-situ calibration steps, and its validation is both challenging and critical to adoption of the technology. We present a rigorous methodology for direct comparison of uPSP-derived surface pressure measurements to reference pressure transducer data based on a model of the camera-based measurement error, a model of the flow field pressure signal, and their propagation through the uPSP data reduction algorithm. The methodology is demonstrated on data collected in a small-scale, subsonic wind tunnel test as well as data collected from a Space Launch System unsteady aerodynamics characterization test at the NASA Ames Research Center Unitary Plan Wind Tunnel Complex.

Pressure-Sensitive Paint↗

NASA NTRSDate not supplied

Methodology for Validation of Unsteady Pressure-Sensitive Paint Measurements using Pressure Transducers

Pressure-Sensitive Paint↗

NASA NTRS1983

Aerodynamic characteristics of a series of bodies with variations in nose camber

The aerodynamic characteristics of a series of cambered forebody models having a systematic variation in nose droop angle were determined from tests in the Langley 8-Foot Transonic Pressure Tunnel at Mach numbers from 0.60 to 1.20 and in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.47, 1.80, and 2.16. The models were tested through an angle-of-attack range of about 0 deg to 12 deg in the 8-Foot Transonic Pressure Tunnel and -2 deg to 20 deg in the Unitary Plan Wind Tunnel. Static longitudinal aerodynamic characteristics of the models were determined for all Mach numbers, and lateral-directional characteristics were determined for Mach numbers of 1.47 to 2.16. The investigation indicated that the principal effect of varying nose droop was on pitching moment, with some secondary effects on lift and drag. The experimental data were also compared with theoretical estimates.

Shrout, B. L.↗

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 NTRS2011

Aerodynamic Analyses and Database Development for Lift-Off/Transition and First Stage Ascent of the Ares I A106 Vehicle

NASA Langley Research Center, in partnership with NASA Marshall Space Flight Center and NASA Ames Research Center, was involved in the aerodynamic analyses, testing, and database development for the Ares I A106 crew launch vehicle in support of the Ares Design and Analysis Cycle. This paper discusses the development of lift-off/transition and ascent databases. The lift-off/transition database was developed using data from tests on a 1.75% scale model of the A106 configuration in the NASA Langley 14x22 Subsonic Wind Tunnel. The power-off ascent database was developed using test data on a 1% A106 scale model from two different facilities, the Boeing Polysonic Wind Tunnel and the NASA Langley Unitary Plan Wind Tunnel. The ascent database was adjusted for differences in wind tunnel and flight Reynolds numbers using USM3D CFD code. The aerodynamic jet interaction effects due to first stage roll control system were modeled using USM3D and OVERFLOW CFD codes.

Pamadi, Bandu N.↗

NASA NTRS2001

Isolated Performance at Mach Numbers From 0.60 to 2.86 of Several Expendable Nozzle Concepts for Supersonic Applications

Investigations have been conducted in the Langley 16-Foot Transonic Tunnel (at Mach numbers from 0.60 to 1.25) and in the Langley Unitary Plan Wind Tunnel (at Mach numbers from 2.16 to 2.86) at an angle of attack of 0 deg to determine the isolated performance of several expendable nozzle concepts for supersonic nonaugmented turbojet applications. The effects of centerbody base shape, shroud length, shroud ventilation, cruciform shroud expansion ratio, and cruciform shroud flap vectoring were investigated. The nozzle pressure ratio range, which was a function of Mach number, was between 1.9 and 11.8 in the 16-Foot Transonic Tunnel and between 7.9 and 54.9 in the Unitary Plan Wind Tunnel. Discharge coefficient, thrust-minus-drag, and the forces and moments generated by vectoring the divergent shroud flaps (for Mach numbers of 0.60 to 1.25 only) of a cruciform nozzle configuration were measured. The shortest nozzle had the best thrust-minus-drag performance at Mach numbers up to 0.95 but was approached in performance by other configurations at Mach numbers of 1.15 and 1.25. At Mach numbers above 1.25, the cruciform nozzle configuration having the same expansion ratio (2.64) as the fixed geometry nozzles had the best thrust-minus-drag performance. Ventilation of the fixed geometry divergent shrouds to the nozzle external boattail flow generally improved thrust-minus-drag performance at Mach numbers from 0.60 to 1.25, but decreased performance above a Mach number of 1.25.

Re, Richard J.↗

NASA NTRS1998

Model Deformation Measurements at NASA Langley Research Center

Only recently have large amounts of model deformation data been acquired in NASA wind tunnels. This acquisition of model deformation data was made possible by the development of an automated video photogrammetric system to measure the changes in wing twist and bending under aerodynamic load. The measurement technique is based upon a single view photogrammetric determination of two dimensional coordinates of wing targets with a fixed third dimensional coordinate, namely the spanwise location. A major consideration in the development of the measurement system was that use of the technique must not appreciably reduce wind tunnel productivity. The measurement technique has been used successfully for a number of tests at four large production wind tunnels at NASA and a dedicated system is nearing completion for a fifth facility. These facilities are the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel at NASA Langley, and the 12-FT Pressure Tunnel at NASA Ames. A dedicated system for the Langley 16-Foot Transonic Tunnel is scheduled to be used for the first time for a test in September. The advantages, limitations, and strategy of the technique as currently used in NASA wind tunnels are presented. Model deformation data are presented which illustrate the value of these measurements. Plans for further enhancements to the technique are presented.

Burner, A. W.↗

NASA NTRS1992

Supersonic dynamic stability characteristics of the test technique demonstrator NASP configuration

Wind tunnel tests of a National Aero-Space Plane (NASP) configuration were conducted in both test sections of the Langley Unitary Plan Wind Tunnel. The model used is a Langley designed blended body NASP configuration. Dynamic stability characteristics were measured on this configuration at Mach numbers of 2.0, 2.5, 3.5, and 4.5. In addition to tests of the baseline configuration, component buildup tests were conducted. The test results show that the baseline configuration generally has positive damping about all three axes with only isolated exceptions. In addition, there was generally good agreement between the in-pulse dynamic parameters and the corresponding static data which were measured during another series of tests in the Unitary Plan Wind Tunnel. Also included are comparisons of the experimental damping parameters with results from the engineering predictive code APAS (Aerodynamic Preliminary Analysis System). These comparisons show good agreement at low angles of attack; however, the comparisons are generally not as good at the higher angles of attack.

Dress, David A.↗

NASA NTRS1996

Automated Wing Twist And Bending Measurements Under Aerodynamic Load

An automated system to measure the change in wing twist and bending under aerodynamic load in a wind tunnel is described. The basic instrumentation consists of a single CCD video camera and a frame grabber interfaced to a computer. The technique is based upon a single view photogrammetric determination of two dimensional coordinates of wing targets with a fixed (and known) third dimensional coordinate, namely the spanwise location. The measurement technique has been used successfully at the National Transonic Facility, the Transonic Dynamics Tunnel, and the Unitary Plan Wind Tunnel at NASA Langley Research Center. The advantages and limitations (including targeting) of the technique are discussed. A major consideration in the development was that use of the technique must not appreciably reduce wind tunnel productivity.

Burner, A. W.↗

NASA NTRS1986

Current wind tunnel capability and planned improvements at the NASA Langley Research Center

NASA Langley's major wing wind tunnels and the projected facilities planned to be completed by 1988 are presented. Special capabilities, uses, and the improvements done during the Langley's tunnel revitalization program are described for the following facilities: (1) the 30 x 60-ft subsonic tunnel, (2) the 4 x 7-m low-speed testing tunnel, (3) the Low-Turbulence Tunnel, (4) the Spin Tunnel, (5) the National Transonic Pressure Tunnel, (6) Transonic Cryogenic Tunnel, (7) the National Transonic Facility, (8) the 16-ft Transonic Tunnel, (9) the Transonic Dynamic Tunnel, (10) the Unitary Plan Wind Tunnel, and (11) a new 20-inch supersonic wind tunnel which is currently undergoing final checkout. The design concept of an extremely-low disturbance level supersonic tunnel, the upgrading plans for the hypersonic aerothermal complex, and the present and planned capabilities for testing the hydrogen-fueled Scramjet engines are also presented. In addition, uses of inexpensive simple-to-operate research wind tunnels are discussed. Tunnel diagrams and graphs of upgrade results are included.

Bower, R. E.↗

NASA NTRS1992

Study of optical techniques for the Ames unitary wind tunnels. Part 3: Angle of attack

A review of optical sensors that are capable of accurate angle of attack measurements in wind tunnels was conducted. These include sensors being used or being developed at NASA Ames and Langley Research Centers, Boeing Airplane Company, McDonald Aircraft Company, Arnold Engineering Development Center, National Aerospace Laboratory of the Netherlands, National Research Council of Canada, and the Royal Aircraft Establishment of England. Some commercial sensors that may be applicable to accurate angle measurements were also reviewed. It was found that the optical sensor systems were based on interferometers, polarized light detector, linear or area photodiode cameras, position sensing photodetectors, and laser scanners. Several of the optical sensors can meet the requirements of the Ames Unitary Plan Wind Tunnel. Two of these, the Boeing interferometer and the Complere lateral effect photodiode sensors are being developed for the Ames Unitary Plan Wind Tunnel.

Lee, George↗

NASA NTRS2020

Wind Tunnel Investigation of the Supersonic Stage Separation Aerodynamics of a Generic 0.0175-Scale Bimese Two-Stage-to-Orbit Reusable Launch Vehicle Configuration

A wind tunnel investigation was conducted of the supersonic stage separation aerodynamics of a generic two-stage-to-orbit bimese wingbody configuration in the NASA Langley Research Center Unitary Plan Wind Tunnel. Proximity and isolated model testing was conducted at Mach numbers of 2.3, 3.0, and 4.5 and a unit Reynolds number of 2.0 million per foot using 0.0175-scale models of the Langley Glide-Back Booster concept designated as the orbiter and booster in belly-to-belly and back-to-belly configurations. Longitudinal forces and moments were obtained on both models and surface static pressure measurements were obtained on the orbiter model at 328 relative proximity locations and at relative angles of attack of 0 degrees and 5 degrees. The test results supported a larger effort to develop and validate experimental and computational tools applicable to the design and simulation of stage separation and abort procedures for reusable launch vehicles composed of multiple bodies, including winged bodies. An initial proof-of-concept experiment featuring low-cost uninstrumented models was conducted to verify an emerging automated model control system and new support system hardware, and to identify potential model and support system blockage and unsteady aerodynamics/model dynamics prior to committing to higher-fidelity instrumented models. This investigation led to upgrades in the facility stage separation hardware, calibration and testing techniques and capabilities, and data analysis and documentation methodologies that have been extended to the more recent NASA Constellation and Space Launch System crew and cargo launch vehicle programs. A virtual diagnostics interface methodology was used to facilitate the design of the stage separation support hardware, to position the models in the test section, and to define the experimental test space. Advances in the facility automated model positioning system established a foundation for the development of a continuous-sweep data acquisition technique that is responsible for significant productivity improvements to the current NASA Space Launch System test program. The automated model positioning capability was leveraged to conduct a companion statistically-designed stage separation experiment requiring randomization of the relative proximity positions of the orbiter and booster models. The respective zones of influence and interference effects of the orbiter and booster were identified from three-dimensional scatter plots, contour and influence maps, and two-dimensional plotting methods. The highly-nonlinear, shock-dominated aerodynamic characteristics of the orbiter and booster in the Unitary Plan Wind Tunnel exhibited good agreement with independent test data obtained in a NASA Marshall Space Flight Center wind tunnel and with computational fluid dynamics predictions using a compressible, three-dimensional flow solver and an inviscid, unstructured Cartesian method.

Erickson, Gary E.↗

NASA NTRS2016

Overview of ERA Integrated Technology Demonstration (ITD) 51A Ultra-High Bypass (UHB) Integration for Hybrid Wing Body (HWB)

The NASA Environmentally Responsible Aircraft Project (ERA) was a ve year project broken into two phases. In phase II, high N+2 Technical Readiness Level demonstrations were grouped into Integrated Technology Demonstrations (ITD). This paper describes the work done on ITD-51A: the Vehicle Systems Integration, Engine Airframe Integration Demonstration. Refinement of a Hybrid Wing Body (HWB) aircraft from the possible candidates developed in ERA Phase I was continued. Scaled powered, and unpowered wind- tunnel testing, with and without acoustics, in the NASA LARC 14- by 22-foot Subsonic Tunnel, the NASA ARC Unitary Plan Wind Tunnel, and the 40- by 80-foot test section of the National Full-Scale Aerodynamics Complex (NFAC) in conjunction with very closely coupled Computational Fluid Dynamics was used to demonstrate the fuel burn and acoustic milestone targets of the ERA Project.

Flamm, Jeffrey D.↗

NASA NTRS2010

Aerothermal Protuberance Heating Design and Test Configurations for Ascent Vehicle Design

A series of tests were conducted to evaluate protuberance heating for the purposes of vehicle design and modification. These tests represent a state of the art approach to both testing and instrumentation for defining aerothermal protuberance effects on the protuberance and surrounding areas. The testing was performed with a number of wind tunnel entries beginning with the proof of concept "pathfinder" test in the Test Section 1 (TS1) tunnel in the Langley Unitary Plan Wind Tunnel (UPWT). The TS1 section (see Figures 1a and 1b) is a lower Mach number tunnel and the Test Section 2 (TS2) has overlapping and higher Mach number capability as showin in Figure 1c. The pathfinder concept was proven and testing proceeded for a series of protuberance tests using an existing splitter aluminum protuberance mounting plate, Macor protuberances, thin film gages, total temperature and pressure gages, Kulite pressure transducers, Infra-Red camera imaging, LASER velocimetry evaluations and the UPWT data collection system. A boundary layer rake was used to identify the boundary layer profile at the protuberance locations for testing and helped protuberance design. This paper discusses the techniques and instrumentation used during the protuberance heating tests performed in the UPWT in TS1 and TS2. Runs of the protuberances were made Mach numbers of 1.5, 2.16, 2.65, and 3.51. The data set generated from this testing is for ascent protuberance effects and is termed Protuberance Heating Ascent Data (PHAD) and this testing may be termed PHAD-1 to distinguish it from future testing of this type.

Martin, Charles E.↗

NASA NTRS1973

A Wind-Tunnel Investigation of the Application of the NASA Supercritical Airfoil to a Variable-Wing-Sweep Fighter Airplane

An investigation was conducted in the Langley 8 foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel to evaluate the effectiveness of three variations of the NASA supercritical airfoil as applied to a model of a variable wing sweep fighter airplane. Wing panels incorporating conventional NACA 64A series airfoil with 0.20 and 0.40 camber were used as bases of reference for this evaluation. Static force and moment measurements were obtained for wing leading edge sweep angles of 26, 33, 39, and 72.5 degrees. Fluctuating wing root bending moment data were obtained at subsonic speeds to determine buffet characteristics. Subsonic data were also obtained for determining the effects of wing transition location and spoiler deflection. Limited lateral directional data are included for the conventional 0.20 cambered wing and the supercritical wing.

Ayers, T. G.↗

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↗

NASA NTRS2001

Characteristics of Pressure Sensitive Paint Intrusiveness Effects on Aerodynamic Data

One effect of using pressure sensitive paint (PSP) is the potential intrusiveness to the aerodynamic characteristics of the model. The paint thickness and roughness may affect the pressure distribution, and therefore, the forces and moments on the wind tunnel model. A study of these potential intrusive effects was carried out at NASA Langley Research Center where a series of wind tunnel tests were conducted using the Modem Design of Experiments (MDOE) test approach. The PSP effects on the integrated forces were measured on two different models at different test conditions in both the Low Turbulence Pressure Tunnel (LTPT) and the Unitary Plan Wind Tunnel (UPWT) at Langley. The paint effect was found to be very small over a range of Reynolds numbers, Mach numbers and angles of attack. This is due to the very low surface roughness of the painted surface. The surface roughness, after applying the NASA Langley developed PSP, was lower than that of the clean wing. However, the PSP coating had a localized effects on the pressure taps, which leads to an appreciable decrease in the pressure tap reading.

Amer, Tahani R.↗