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A brief history of the 8x6 Supersonic Wind Tunnel (SWT) and 9x15 Low Speed Wind Tunnel (LSWT) at NASA Glenn Research Center, Cleveland, Ohio is presented along with current capabilities and plans for future upgrades within the facility.
Time, money, and, personnel are becoming increasingly scarce resources within government agencies due to a reduction in funding and the desire to demonstrate responsible economic efficiency. The ability of an organization to plan and schedule resources effectively can provide the necessary leverage to improve productivity, provide continuous support to all projects, and insure flexibility in a rapidly changing environment. Without adequate internal controls the organization is forced to rely on external support, waste precious resources, and risk an inefficient response to change. Management systems must be developed and applied that strive to maximize the utility of existing resources in order to achieve the goal of "faster, cheaper, better". An area of concern within NASA Langley Research Center was the scheduling, planning, and resource management of the Wind Tunnel Enterprise operations. Nine wind tunnels make up the Enterprise. Prior to this research, these wind tunnel groups did not employ a rigorous or standardized management planning system. In addition, each wind tunnel unit operated from a position of autonomy, with little coordination of clients, resources, or project control. For operating and planning purposes, each wind tunnel operating unit must balance inputs from a variety of sources. Although each unit is managed by individual Facility Operations groups, other stakeholders influence wind tunnel operations. These groups include, for example, the various researchers and clients who use the facility, the Facility System Engineering Division (FSED) tasked with wind tunnel repair and upgrade, the Langley Research Center (LaRC) Fabrication (FAB) group which fabricates repair parts and provides test model upkeep, the NASA and LARC Strategic Plans, and unscheduled use of the facilities by important clients. Expanding these influences horizontally through nine wind tunnel operations and vertically along the NASA management structure greatly increases the complexity of developing a model that can be used for successfully implementing a standardized management planning tool. The objective of this study was to implement an Integrated Wind Tunnel Planning System to improve the operations within the aeronautics testing and research group, in particular Wind Tunnel Enterprise. The study included following steps: Conducted literature search and expert discussions (NASA and Old Dominion University faculty), Performed environmental scan of NASA Langley wind tunnel operations as foundation for problem definition. Established operation requirements and evaluation methodologies. Examined windtunnel operations to map out the common characteristics, critical components, and system structure. Reviewed and evaluated various project scheduling and management systems for implementation, Evaluated and implemented "Theory of Constraints (TOC)" project scheduling methodology at NASA Langley wind tunnel operations together with NASA staff.
At the 10- by 10-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center, a future full test section characterization generated an ideal opportunity to design and build new characterization hardware to improve the understanding of the flow field, including flow quality, uniformity, and uncertainty in primary variables of interest. An array of flow sensing probes, referred to as the Characterization Array, was designed and built to replace 1960’s-era test section characterization hardware. Many references exist to guide wind tunnel characterization practitioners in the design of new hardware to properly measure various aspects of the flow within their wind tunnel facilities. Although reliable sources of information, these references tend to be over 30 years old and are not exhaustive. In scenarios where design decisions needed to be validated, computational simulations of the flow field around the characterization hardware were used. Decisions regarding probe location, probe spacing, and performance of various probes were justified using computational fluid dynamic simulations and rules-of-thumb from the legacy resources available in literature. This paper is intended to serve as an example of the benefits from integrating CFD into the design of wind tunnel hardware, particularly hardware for wind tunnel characterization.
At the 10- by 10-Foot Supersonic Wind Tunnel at the NASA Glenn Research Center, a future full test section characterization generated an ideal opportunity to design and build new characterization hardware to improve the understanding of the flow field, including flow quality, uniformity, and uncertainty in primary variables of interest. An array of flow sensing probes, referred to as the Characterization Array, was designed and built to replace 1960’s-era test section characterization hardware. Many references exist to guide wind tunnel characterization practitioners in the design of new hardware to properly measure various aspects of the flow within their wind tunnel facilities. Although reliable sources of information, these references tend to be over 30 years old and are not exhaustive. In scenarios where design decisions needed to be validated, computational simulations of the flow field around the characterization hardware were used. Decisions regarding probe location, probe spacing, and performance of various probes were justified using computational fluid dynamic simulations and rules-of-thumb from the legacy resources available in literature. This paper is intended to serve as an example of the benefits from integrating CFD into the design of wind tunnel hardware, particularly hardware for wind tunnel characterization.
A remote control system for wind-tunnel model control surfaces was developed during the Space Shuttle program to make more efficient use of wind-tunnel occupancy time and to aid in gathering the large force test data base necessary for the definition of the Shuttle aerodynamic characteristics. This paper presents a history of the development of the remote system, details of the system and associated equipment, and results from wind-tunnel tests showing the effect of system improvements on experimental data. Wind-tunnel test rate and cost comparisons are made between conventional models with bracketed control surfaces and remote models.
A remote control system for wind-tunnel model control surfaces was developed during the Space Shuttle program to make more efficient use of wind-tunnel occupancy time and to aid in gathering the large force test data base necessary for the definition of the Shuttle aerodynamic characteristics. A history of the development of the remote system, details of the system and associated equipment, and results from wind-tunnel tests showing the effect of system improvements on experimental data are given. Wind-tunnel test rate and cost comparisons are made between conventional models with bracketed control surfaces and remote models.
Porous wall wind tunnels have been used for several decades and have proven effective in reducing wall interference effects in both low speed and transonic testing. They allow for testing through Mach 1, reduce blockage effects and reduce shock wave reflections in the test section. Their usefulness in developing computational fluid dynamics (CFD) codes has been limited, however, by the difficulties associated with modelling the effect of a porous wall in CFD codes. Previous approaches to modelling porous wall effects have depended either upon a simplified linear boundary condition, which has proven inadequate, or upon detailed measurements of the normal velocity near the wall, which require extensive wind tunnel time. The current work was initiated in an effort to find a simple, accurate method of modelling a porous wall boundary condition in CFD codes. The development of such a method would allow data from porous wall wind tunnels to be used more readily in validating CFD codes. This would be beneficial when transonic validations are desired, or when large models are used to achieve high Reynolds numbers in testing. A computational and experimental study was undertaken to investigate a new method of modelling solid and porous wall boundary conditions in CFD codes. The method utilized experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution was then imposed as a pressure boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the effect of porosity changes on the wall was investigated. Also, the effectiveness of this method when only sparse experimental measurements were available has been investigated. The current work demonstrated this approach for low speed flows and compared the results with experimental data obtained from a heavily instrumented variable porosity test section. The approach developed was simple, computationally inexpensive, and did not require extensive or intrusive measurements of the boundary conditions during the wind tunnel test. It may be applied to both solid and porous wall wind tunnel tests.
Optical measurement techniques have become a standard option for wind tunnel tests. Pressure-sensitive paint (PSP) is a mature test technique and a common experimental technique in many wind tunnels to measure the global mean static pressure on a model. PSP is a valuable tool when a more detailed distribution of the pressure is needed rather than the conventional pressure taps alone. Planning for a test with optical-based techniques can present new challenges even for experienced customer. The purpose of this paper is to provide a resource to the wind tunnel testing community and customers interested in obtaining PSP measurements on a wind tunnel model at the NASA Ames Research Center’s Unitary Plan Wind Tunnel. An overview of PSP mechanics, a list of requirements for ones considering PSP measurements, and PSP deliverable details are specified.
The cryogenic wind tunnel and its potential for advancing maneuvering aircraft technology is discussed. A brief overview of the cryogenic wind tunnel concept and the capabilities and status of the Langley cryogenic facilities is given, as is a review of the considerations leading to the selection of the cryogenic concept such as capital and operating costs of the tunnel, model and balance construction implications, and test condition. Typical viscous, compressibility and aeroelastic effects encountered by maneuvering aircraft are illustrated and the unique ability of the cryogenic wind tunnels to isolate and investigate these parameters while simulating full scale conditions is discussed. The status of the Langley cryogenic wind tunnel facilities is reviewed and their operating envelopes described in relation to maneuvering aircraft research and development requirements. The status of cryogenic testing technology specifically related to aircraft maneuverability studies including force balances and buffet measurement techniques is discussed. Included are examples of research carried out in the Langley 0.3 meter transonic cryogenic wind tunnel to verify the various techniques.
A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.
A unique, low-disturbance (quiet) supersonic wind tunnel has been commissioned at the NASA-Ames Fluid Mechanics Laboratory (FML) to support Supersonic Laminar Flow Control (SLFC) research. Known as the Laminar Flow Supersonic Wind Tunnel (LFSWT), this tunnel is designed to operate at potential cruise Mach numbers and unit Reynolds numbers (Re) of the High Speed Civil Transport (HSCT). The need to better understand the receptivity of the transition phenomena on swept (HSCT) wings to attachment-line contamination and cross-flows has provided the impetus for building the LFSWT. Low-disturbance or "quiet" wind tunnels are known to be an essential part of any meaningful boundary layer transition research. In particular, the receptivity of supersonic boundary layers to wind tunnel disturbances can significantly alter the transition phenomena under investigation on a test model. Consequently, considerable effort has gone into the design of the LFSWT to provide quiet flow. The paper describes efforts to quantify the low-disturbance flows in the LFSWT operating at Mach 1.6, as a precursor to transition research on wing models. The research includes: (1) Flow measurements in both the test section and settling chamber of the LFSWT, using a full range of measurement techniques; (2) Study of the state of the test section boundary layer so far by using a single hot-wire mounted above the floor centerline, with and without boundary layer trips fitted at the test section entrance; (3) The effect of flow quality of unsteady supersonic diffuser flow, joint steps and gaps, and wall vibration.
Discussed here is a new type of wind tunnel, its advantages, the difficulties attendant upon its use, and the special methods required for its operation. The main difference between the new type of wind tunnel and the ones now in operation is the use of a different fluid. The idea is to diminish the effect of viscosity If air is compressed, it becomes a fluid with new properties - a fluid that is best suited for reliable and exact tests on models. When air is compressed, its density increases, but its viscosity does not. It is argued that the increase of pressure greatly increases the range and value of wind tunnel tests. Reynolds number, deductions from the Reynolds law, the causes of errors that result in differences between tests on models and actual flights, and the dimensions of a compressed air wind tunnel are covered.
A wind tunnel test of an 0.030 scale model of the Rockwell International Configuration 140A/B Space Shuttle Vehicle Orbiter were conducted in the Ames Research Center 11- by 11-Foot Transonic Wind Tunnel. This part (part A) of test series 0A53 was conducted at Mach numbers of 0.6, 0.8, 0.9, 1.05, and 1.20, and at Reynolds numbers from 1.8 x to 6.5 million per foot. The objective of this test was to establish and verify longitudinal and lateral-directional aerodynamic performance, stability, and control characteristics for the Configuration 140A/B SSV Orbiter. Reynolds number studies were performed for certain nominal control-settings. An alternate leading-edge wing configuration and sealed elevon-split arrangement were tested. Bodyflap, elevon, speedbrake, and rudder hinge moments were measured in addition to standard six-component forces and moments and base pressure data. Furthermore, six-component force and moment data were measured for the vertical tail assembly.
Wind-tunnel tests were conducted to determine the low speed two dimensional aerodynamic characteristics of a 17-percent-thick supercritical airfoil. The results were compared with three dimensional wind-tunnel and flight data. The tests were conducted over a Mach number range from 0.15 to 0.30. Reynolds numbers based on the airfoil chord varied from 2.0x10 to the 6th power to 15.0x10 to the 6th power.
Development of the space shuttle necessitated an extensive wind tunnel test program, with the cooperation of all the major wind tunnels in the United States. The result was approximately 100,000 hours of space shuttle wind tunnel testing conducted for aerodynamics, heat transfer, and structural dynamics. The test results were converted into Chrysler DATAMAN computer program format to facilitate use by analysts, a very cost effective method of collecting the wind tunnel test results from many test facilities into one centralized location. This report provides final documentation of the space shuttle wind tunnel program. The two-volume set covers evolution of space shuttle aerodynamic configurations and gives wind tunnel test data, titles of wind tunnel data reports, sample data sets, and instructions for accessing the digital data base.
A test program was conducted in a 40 by 80 foot wind tunnel to evaluate the effect of relative velocity on the jet noise signature of a conical ejector, auxiliary inlet ejector, 32 spokes and 104 tube nozzle with and without an acoustically treated shroud. The freestream velocities in the wind tunnel were varied from 0 to 103.6 m/sec (300 ft/sec) for exhaust jet velocities of 259.1 m/sec (850 ft/sec) to 609.6 m/sec (2000 ft/sec). Reverberation corrections for the wind tunnel were developed and the procedure is explained. In conjunction with wind tunnel testing the nozzles were also evaluated on an outdoor test stand. The wind tunnel microphone arrays were duplicated during the outdoor testing. The data were then extrapolated for comparisons with data measured using a microphone array placed on a 30.5 meter (100 ft) arc. Using these data as a basis, farfield to nearfield arguments are presented with regards to the data measured in the wind tunnel. Finally, comparisons are presented between predictions made using existing methods and the measured data.
An obvious extension of wind tunnel laser Doppler velocimetry (LDV), which is currently mainly limited to two-component measurements, would involve the measurement of three simultaneous velocity components. The present paper is concerned with an approach to reduce the degree of optical complexity involved in the design of a three-dimensional (3D) LDV system, taking into account the use of polarization separation. Such a system, utilizing polarization as well as color separation, has been designed and tested in a 6 x 6 foot supersonic wind tunnel. The considered instrument was designed for the on line measurement of three mean velocity components, turbulence levels, and shear stresses on a number of models under a wide variety of test conditions. Attention is given to optical details, data reduction, and sample application.
High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is reported. The LAP is a 2.74 meter (9.0 ft) diameter, 8-bladed tractor type rated for 4475 KW (6000 SHP) at 1698 rpm. It was designated and built by Hamilton Standard under contract to the NASA Lewis Research Center. The LAP employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Testing was conducted in the ONERA S1-MA Atmospheric Wind Tunnel in Modane, France. The test objectives were to confirm that the LAP is free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers. The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer prediction codes used for the design. There were no signs of classical flutter throughout all phases of the test up to and including the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were successfully measured for a wide range of Mach numbers, inflow angles, rotational speeds and blade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.