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NASA NTRS2001

Recent National Transonic Facility Test Process Improvements (Invited)

This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feedforward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

Kilgore, W. A.↗

NASA NTRS2001

Recent National Transonic Facility Test Process Improvements (Invited)

This paper describes the results of two recent process improvements; drag feed-forward Mach number control and simultaneous force/moment and pressure testing, at the National Transonic Facility. These improvements have reduced the duration and cost of testing. The drag feed-forward Mach number control reduces the Mach number settling time by using measured model drag in the Mach number control algorithm. Simultaneous force/moment and pressure testing allows simultaneous collection of force/moment and pressure data without sacrificing data quality thereby reducing the overall testing time. Both improvements can be implemented at any wind tunnel. Additionally the NTF is working to develop and implement continuous pitch as a testing option as an additional method to reduce costs and maintain data quality.

Kilgore, W. A.↗

NASA NTRS2017

Transonic Semispan Aerodynamic Testing of the Hybrid Wing Body with Over Wing Nacelles in the National Transonic Facility

A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.

Chan, David T.↗

NASA NTRSDate not supplied

Single-Component Average Velocity Profiles in the Wake of the Orion Crew Capsule at the National Transonic Facility

A minimally intrusive molecular tagging instrument measured single-component average velocity profiles in two planes in the wake of the Orion capsule with different heat shield configurations in the National Transonic Facility at the NASA Langley Research Center. Reynolds number effects at subsonic conditions have proven difficult to predict due to the largely separated wake flow. Therefore two measurement planes in the wake of the model were probed to measure the wake profile for different heat shield configurations and act as a validation reference for computational tools. Air testing included Mach numbers of 0.3, 0.5, and 0.7 at Reynolds numbers of 5.3 and 7.5 million. In cryogenic nitrogen, the instrument was employed under transient conditions during the facility warm up at M = 0.3 with decreasing Reynolds numbers from 16 million. An exhaustive list of the results is shown and discussed here. For free transition heat shield configurations, the size of the wake was found to increase with Mach number, yet remain constant with Reynolds number for low (M = 0.3) and high (M = 0.7) subsonic Mach numbers. However, intermediate Mach numbers (M = 0.5) showed that the wake was smaller at higher Reynolds numbers for the IDAT heat shield. The addition of surface roughness in the form of grit, known as the fixed transition cases, negated any Mach number dependence to the wake profile and increased the size of the wake for all cases.

Velocimetry↗

NASA NTRSDate not supplied

Single-Component Average Velocity Profiles in the Wake of the Orion Crew Capsule at the National Transonic Facility

Velocimetry↗

NASA NTRS1991

Mechanical properties of the fiberglass prepreg system used for the National Transonic Facility replacement blade set

The results are presented of mechanical and physical properties characterization testing for the fiber glass prepreg system used to fabricate 15 of the replacement set of 25 fan blades for the National Transonic Facility. The fan blades were fabricated to be identical to the original blade set with the exception that the 7576 style E glass cloth used for the replacement set has a different surface finish than the original 7576 cloth. The 7781 E glass cloth and resin system were unchanged. The data are presented for elevated, room, and cryogenic temperatures. The results are compared with data from the original blade set and evaluated against selected structural design criteria. Test experience is described along with recommendations for future testing of these materials if required.

Young, Clarence P., Jr.↗

NASA NTRS2009

The Development and Implementation of a Cryogenic Pressure Sensitive Paint System in the National Transonic Facility

The Pressure Sensitive Paint (PSP) method was used to measure global surface pressures on a model at full-scale flight Reynolds numbers. In order to achieve these conditions, the test was carried out at the National Transonic Facility (NTF) operating under cryogenic conditions in a nitrogen environment. The upper surface of a wing on a full-span 0.027 scale commercial transport was painted with a porous PSP formulation and tested at 120K. Data was acquired at Mach 0.8 with a total pressure of 200 kPa, resulting in a Reynolds number of 65 x 106/m. Oxygen, which is required for PSP operation, was injected using dry air so that the oxygen concentration in the flow was approximately 1535 ppm. Results show qualitative agreement with expected results. This preliminary test is the first time that PSP has been successfully deployed to measure global surface pressures at cryogenic condition in the NTF. This paper will describe the system as installed, the results obtained from the test, as well as proposed upgrades and future tests.

Watkins, A. Neal↗

NASA NTRS2011

Results From a Pressure Sensitive Paint Test Conducted at the National Transonic Facility on Test 197: The Common Research Model

This report will serve to present results of a test of the pressure sensitive paint (PSP) technique on the Common Research Model (CRM). This test was conducted at the National Transonic Facility (NTF) at NASA Langley Research Center. PSP data was collected on several surfaces with the tunnel operating in both cryogenic mode and standard air mode. This report will also outline lessons learned from the test as well as possible approaches to challenges faced in the test that can be applied to later entries.

Watkins, A. Neal↗

NASA NTRS2001

Development of a Semi-Span Test Capability at the National Transonic Facility

A need for low-speed, high Reynolds number test capabilities has been identified for the design and development of advanced subsonic transport high-lift systems. In support of this need, multiple investigations have been conducted in the National Transonic Facility (NTF) at the NASA Langley Research Center to develop a semi-span testing capability that will provide the low-speed, flight Reynolds number data currently unattainable using conventional sting-mounted, full-span models. Although a semi-span testing capability will effectively double the Reynolds number capability over full-span models, it does come at the expense of contending with the issue of the interaction of the flow over the model with the windtunnel wall boundary layer. To address this issue the size and shape of the semi-span model mounting geometry have been investigated, and the results are presented herein. The cryogenic operating environment of the NTF produced another semi-span test technique issue in that varying thermal gradients have developed on the large semi-span balance. The suspected cause of these thermal gradients and methods to eliminate them are presented. Data are also presented that demonstrate the successful elimination of these varying thermal gradients during cryogenic operations.

Gatlin, G. M.↗

NASA NTRS1976

LaRC design analysis report for National Transonic Facility for 9% nickel tunnel shell. Volume 1: Finite difference analysis of cone/cylinder junction

The results of extensive computer (finite element, finite difference and numerical integration), thermal, fatigue, and special analyses of critical portions of a large pressurized, cryogenic wind tunnel (National Transonic Facility) are presented. The computer models, loading and boundary conditions are described. Graphic capability was used to display model geometry, section properties, and stress results. A stress criteria is presented for evaluation of the results of the analyses. Thermal analyses were performed for major critical and typical areas. Fatigue analyses of the entire tunnel circuit are presented.

Ramsey, J. W., Jr.↗

NASA NTRS1985

Design and construction of models for the National Transonic Facility, part 2

This lecture presents the results of fastener load and retention systems tests which were carried out as a part of the cryogenic models technology development program a the NASA Langley Research Center (LaRC). Various design concepts for the National Transonic Facility (NTF) developmental and production models are discussed. A number of NTF models are described with emphasis on materials used, uniqueness of design and design drivers. Design and fabrication experience is presented in terms of the primary thermal and mechanical considerations required for design as well as fabrication. Cost considerations are addressed in terms of factors influencing costs for NTF models and cost data comparisons which are taken from both NASA Langley and industry experience.

Young, C. P., Jr.↗

NASA NTRS1987

Design study of advanced model support systems for the National Transonic Facility (NTF)

It has long been recognized that the sting (or support system) is a very critical part of the model system. The designer is frequently faced with the tradeoff of minimizing sting size, thereby compromising facility and model safety, against a larger sting and the subsequent problems of sting interference effects. In the NASA Langley Research Center National Transonic Facility (NTF), this problem is accentuated by the severe environment of high pressure/low temperature, designed into the facility to provide the desired high Reynolds number. Compromises in the configuration geometry and/or limiting the test envelope are therefore contrary to the purposes and goals of the NTF and are unacceptable. The results of an investigation aimed at improvements of 25% in both strength and Young's modulus of elasticity as compared to high strength cryogenically acceptable steels currently being used are presented. Various materials or combinations of materials were studied along with different design approaches. Design concepts were developed which included conventional material stings, advanced composites, and hybrid configurations. Candidate configurations are recommended.

Source record↗

NASA NTRS1976

LaRC design analysis report for National Transonic Facility for 304 stainless steel tunnel shell. Volume 1S: Finite difference analysis of cone/cylinder junction

Ramsey, J. W., Jr.↗

NASA NTRS2011

Experimental Investigations of the NASA Common Research Model in the NASA Langley National Transonic Facility and NASA Ames 11-Ft Transonic Wind Tunnel (Invited)

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 gave an increase in drag, 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, 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, while the drag shows a difference of less than ten 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.

Rivers, S. M.↗

NASA NTRS2010

Assessment of the National Transonic Facility for Laminar Flow Testing

A transonic wing, designed to accentuate key transition physics, is tested at cryogenic conditions at the National Transonic Facility at NASA Langley. The collaborative test between Boeing and NASA is aimed at assessing the facility for high-Reynolds number testing of configurations with significant regions of laminar flow. The test shows a unit Reynolds number upper limit of 26 M/ft for achieving natural transition. At higher Reynolds numbers turbulent wedges emanating from the leading edge bypass the natural transition process and destroy the laminar flow. At lower Reynolds numbers, the transition location is well correlated with the Tollmien-Schlichting-wave N-factor. The low-Reynolds number results suggest that the flow quality is acceptable for laminar flow testing if the loss of laminar flow due to bypass transition can be avoided.

Crouch, Jeffrey D.↗

NASA NTRSDate not supplied

Computational Analysis of a Boundary-Layer Ingesting Tailcone Thruster Configuration Within the National Transonic Facility Using the LAVA Solver

Boundary-layer ingesting (BLI) propulsion systems are one of the many technologies currently under investigation within the aerospace community to meet NASA’s Advanced Air Transport Technology project goal of a sustainable future in aviation. To that end, an experimental campaign was conducted in the National Transonic Facility (NTF) to study the propulsion-airframe integration effects of the Boundary-Layer Ingesting Tailcone System installed in the aft portion of a 2.7% scale design of the NASA Common Research Model (CRM). This test was accompanied by a numerical simulation effort using the Launch, Ascent, and Vehicle Aerodynamics (LAVA) solver, to validate the applicability of current best-practice Reynolds-averaged Navier Stokes (RANS) models in accurately predicting the relevant flow-physics in free-air. A total of 205 steady RANS simulations were performed using LAVA, covering the range of conditions present in the NTF test matrix. Flow conditions ranged from 5- to 15-million Reynolds number, Mach numbers of 0.75, 0.80 and 0.85, and angles-of-attack between –3° and 4°. Four different mass-flow plugs were also tested to assess propulsor operating condition effects. Sensitivity to these flow conditions are analyzed in detail, especially in terms of the nacelle flow distortion which was the main subject of the study. Aerodynamic load coefficients, aftbody boundary-layer measurements and fuselage pressure tap results are also discussed, providing a wide range of validation results from this experimental campaign. This validation effort shows that current best-practices using RANS models within the LAVA solver flow solver are well suited for predicting the inlet flow distortion characteristics of novel aircraft configurations employing BLI at the aft end of the fuselage.

AATT↗

NASA NTRS2000

A Factorial Data Rate and Dwell Time Experiment in the National Transonic Facility

This report is an introductory tutorial on the application of formal experiment design methods to wind tunnel testing, for the benefit of aeronautical engineers with little formal experiment design training. It also describes the results of a Study to determine whether increases in the sample rate and dwell time of the National Transonic Facility data system Would result in significant changes in force and moment data. Increases in sample rate from 10 samples per second to 50 samples per second were examined, as were changes in dwell time from one second per data point to two seconds. These changes were examined for a representative aircraft model in a range of tunnel operating conditions defined by angles of attack from 0 deg to 3.8 degrees, total pressure from 15.0 psi to 24.1 psi, and Mach numbers from 0.52 to 0.82. No statistically significant effect was associated with the change in sample rate. The change in dwell time from one second to two seconds affected axial force measurements, and to a lesser degree normal force measurements. This dwell effect comprises a "rectification error" caused by incomplete cancellation of the positive and negative elements of certain low frequency dynamic components that are not rejected by the one-Hz low-pass filters of the data system. These low frequency effects may be due to tunnel circuit phenomena and other sources. The magnitude of the dwell effect depends on dynamic pressure, with angle of attack and Mach number influencing the strength of this dependence. An analysis is presented which suggests that the magnitude of the rectification error depends on the ratio of measurement dwell time to the period of the low-frequency dynamics, as well as the amplitude of the dynamics The essential conclusion of this analysis is that extending the dwell time (or, equivalently, replicating short-dwell data points) reduces the rectification error.

DeLoach, R.↗

NASA NTRS1984

Design of advanced technology maneuvering aircraft models for the National Transonic Facility

The need for a large High-Reynolds-Number Transonic Wind Tunnel which will provide a tool to study phenomena sensitive to Reynolds number is discussed. The National Transonic Facility (NTF), is in the calibration phase and the desired capability. Its usefulness, however, will be influenced by the ability of industry to develop model systems capable of withstanding the severe operating environment of the facility so necessary to achieve full scale Reynolds number, without degradation of accuracy, and at reasonable cost. The feasibility of designing models of advanced aerodynamic technology maneuvering aircraft and to achieve full scale Reynolds number for each configuration in the NTF are determined. It is concluded that the facility does not offer the potential for making tunnel to full scale data correlations for this type of aircraft configuration.

Griffin, S. A.↗