The Design of Wind Tunnels and Wind Tunnel Propellers

NASA Technical Reports Server (NTRS)

Warner, Edward P; Norton, F H; Hebbert, C M

1919-01-01

Report discusses the theory of energy losses in wind tunnels, the application of the Drzewiecki theory of propeller design to wind tunnel propellers, and the efficiency and steadiness of flow in model tunnels of various types.

Experimental parametric studies of transonic T-tail flutter. [wind tunnel tests

NASA Technical Reports Server (NTRS)

Ruhlin, C. L.; Sandford, M. C.

1975-01-01

Wind-tunnel tests of the T-tail of a wide-body jet airplane were made at Mach numbers up to 1.02. The model consisted of a 1/13-size scaled version of the T-tail, fuselage, and inboard wing of the airplane. Two interchangeable T-tails were tested, one with design stiffness for flutter-clearance studies and one with reduced stiffness for flutter-trend studies. Transonic antisymmetric-flutter boundaries were determined for the models with variations in: (1) fin-spar stiffness, (2) stabilizer dihedral angle (-5 deg and 0 deg), (3) wing and forward-fuselage shape, and (4) nose shape of the fin-stabilizer juncture. A transonic symmetric-flutter boundary and flutter trends were established for variations in stabilizer pitch stiffness. Photographs of the test configurations are shown.

4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

4. VIEW NORTHWEST OF SUPERSONIC WIND TUNNEL BUILDING TO TRANSONIC WIND TUNNEL BUILDING - Naval Surface Warfare Center, Bounded by Clara Barton Parkway & McArthur Boulevard, Silver Spring, Montgomery County, MD

Persistence Characteristics of Wind-Tunnel Pressure Signatures From Two Similar Models

NASA Technical Reports Server (NTRS)

Mack, Robert J.

2004-01-01

Pressure signatures generated by two sonic-boom wind-tunnel models and measured at Mach 2 are presented, analyzed, and discussed. The two wind-tunnel models differed in length and span by a factor of fourteen, but were similar in wing-body planform shape. The geometry of the larger model had been low-boom tailored to generate a flat top ground pressure signature, and the nacelles-off pressure signatures from this model became more flattop in shape as the model-probe separation distances increased from 0.94 to 4.4 span lengths. The geometry of the smaller model had not been low-boom tailored, yet its measured pressure signatures had non-N-wave shapes that persisted as model-probe separation distances increased from 26.0 to 104.2 span lengths. Since the overall planforms of the two wind-tunnel models were so similar, it was concluded that the shape-persistence trends in the pressure signatures of the smaller, non-low-boom tailored model would also be present at very large distances in the pressure signatures of the larger, low-boom-tailored model.

Aeroelastic instability stoppers for wind tunnel models

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Ricketts, R. H. (Inventor)

1981-01-01

A mechanism for diverting the flow in a wind tunnel from the wing of a tested model is described. The wing is mounted on the wall of a tunnel. A diverter plate is pivotally mounted on the tunnel wall ahead of the model. An actuator fixed to the tunnel is pivotably connected to the diverter plate, by plunger. When the model is about to become unstable during the test the actuator moves the diverter plate from the tunnel wall to divert maintaining stable model conditions. The diverter plate is then retracted to enable normal flow.

Gottingen Wind Tunnel for Testing Aircraft Models

NASA Technical Reports Server (NTRS)

Prandtl, L

1920-01-01

Given here is a brief description of the Gottingen Wind Tunnel for the testing of aircraft models, preceded by a history of its development. Included are a number of diagrams illustrating, among other things, a sectional elevation of the wind tunnel, the pressure regulator, the entrance cone and method of supporting a model for simple drag tests, a three-component balance, and a propeller testing device, all of which are discussed in the text.

Propulsion simulator for magnetically-suspended wind tunnel models

NASA Technical Reports Server (NTRS)

Joshi, P. B.; Malonson, M. R.; Sacco, G. P.; Goldey, C. L.; Garbutt, Keith; Goodyer, M.

1992-01-01

In order to demonstrate the measurement of aerodynamic forces/moments, including the effects of exhaust jets in Magnetic Suspension and Balance System (MSBS) wind tunnels, two propulsion simulator models were developed at Physical Sciences Inc. (PSI). Both the small-scale model (1 in. diameter X 8 in. long) and the large-scale model (2.5 in. diameter X 15 in. long) employed compressed, liquefied carbon dioxide as a propellant. The small-scale simulator, made from a highly magnetizable iron alloy, was demonstrated in the 7 in. MSBS wind tunnel at the University of Southampton. It developed a maximum thrust of approximate 1.3 lbf with a 0.098 in. diameter nozzle and 0.7 lbf with a 0.295 in. diameter nozzle. The Southampton MSBS was able to control the simulator at angles-of attack up to 20 deg. The large-scale simulator was demonstrated to operate in both a steady-state and a pulse mode via a miniaturized solinoid valve. It developed a stable and repeatable thrust of 2.75 lbf over a period of 4s and a nozzle pressure ratio (NPR) of 5.

TWINTN4: A program for transonic four-wall interference assessment in two-dimensional wind tunnels

NASA Technical Reports Server (NTRS)

Kemp, W. B., Jr.

1984-01-01

A method for assessing the wall interference in transonic two-dimensional wind tunnel tests including the effects of the tunnel sidewall boundary layer was developed and implemented in a computer program named TWINTN4. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the equivalent free air flow around the model, and the perturbation attributable to the model. Required input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall-induced perturbation field is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.

Pressure distributions obtained on a 0.10-scale model of the space shuttle Orbiter's forebody in the AEDC 16T propulsion wind tunnel

NASA Technical Reports Server (NTRS)

Siemers, P. M., III; Henry, M. W.

1986-01-01

Pressure distribution test data obtained on a 0.10-scale model of the forward fuselage of the Space Shuttle Orbiter are presented without analysis. The tests were completed in the AEDC 16T Propulsion Wind Tunnel. The 0.10-scale model was tested at angles of attack from -2 deg to 18 deg and angles of side slip from -6 to 6 deg at Mach numbers from 0.25 to 1/5 deg. The tests were conducted in support of the development of the Shuttle Entry Air Data System (SEADS). In addition to modeling the 20 SEADS orifices, the wind-tunnel model was also instrumented with orifices to match Development Flight Instrumentation (DFI) port locations that existed on the Space Shuttle Orbiter Columbia (OV-102) during the Orbiter Flight Test program. This DFI simulation has provided a means of comparisons between reentry flight pressure data and wind-tunnel and computational data.

SOFIA 2 model telescope wind tunnel test report

NASA Technical Reports Server (NTRS)

Keas, Paul

1995-01-01

This document outlines the tests performed to make aerodynamic force and torque measurements on the SOFIA wind tunnel model telescope. These tests were performed during the SOFIA 2 wind tunnel test in the 14 ft wind tunnel during the months of June through August 1994. The test was designed to measure the dynamic cross elevation moment acting on the SOFIA model telescope due to aerodynamic loading. The measurements were taken with the telescope mounted in an open cavity in the tail section of the SOFIA model 747. The purpose of the test was to obtain an estimate of the full scale aerodynamic disturbance spectrum, by scaling up the wind tunnel results (taking into account differences in sail area, air density, cavity dimension, etc.). An estimate of the full scale cross elevation moment spectrum was needed to help determine the impact this disturbance would have on the telescope positioning system requirements. A model of the telescope structure, made of a light weight composite material, was mounted in the open cavity of the SOFIA wind tunnel model. This model was mounted via a force balance to the cavity bulkhead. Despite efforts to use a 'stiff' balance, and a lightweight model, the balance/telescope system had a very low resonant frequency (37 Hz) compared to the desired measurement bandwidth (1000 Hz). Due to this mechanical resonance of the balance/telescope system, the balance alone could not provide an accurate measure of applied aerodynamic force at the high frequencies desired. A method of measurement was developed that incorporated accelerometers in addition to the balance signal, to calculate the aerodynamic force.

Hot-bench simulation of the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Buttrill, Carey S.; Houck, Jacob A.

1990-01-01

Two simulations, one batch and one real-time, of an aeroelastically-scaled wind-tunnel model were developed. The wind-tunnel model was a full-span, free-to-roll model of an advanced fighter concept. The batch simulation was used to generate and verify the real-time simulation and to test candidate control laws prior to implementation. The real-time simulation supported hot-bench testing of a digital controller, which was developed to actively control the elastic deformation of the wind-tunnel model. Time scaling was required for hot-bench testing. The wind-tunnel model, the mathematical models for the simulations, the techniques employed to reduce the hot-bench time-scale factors, and the verification procedures are described.

Aeroservoelastic Wind-Tunnel Test of the SUGAR Truss Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Allen, Timothy J.; Funk, Christie J.; Castelluccio, Mark A.; Sexton, Bradley W.; Claggett, Scott; Dykman, John; Coulson, David A.; Bartels, Robert E.

2015-01-01

The Subsonic Ultra Green Aircraft Research (SUGAR) Truss-Braced Wing (TBW) aeroservoelastic (ASE) wind-tunnel test was conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) and was completed in April, 2014. The primary goals of the test were to identify the open-loop flutter boundary and then demonstrate flutter suppression. A secondary goal was to demonstrate gust load alleviation (GLA). Open-loop flutter and limit cycle oscillation onset boundaries were identified for a range of Mach numbers and various angles of attack. Two sets of control laws were designed for the model and both sets of control laws were successful in suppressing flutter. Control laws optimized for GLA were not designed; however, the flutter suppression control laws were assessed using the TDT Airstream Oscillation System. This paper describes the experimental apparatus, procedures, and results of the TBW wind-tunnel test. Acquired system ID data used to generate ASE models is also discussed.2 study.

The 12-foot pressure wind tunnel restoration project model support systems

NASA Technical Reports Server (NTRS)

Sasaki, Glen E.

1992-01-01

The 12 Foot Pressure Wind Tunnel is a variable density, low turbulence wind tunnel that operates at subsonic speeds, and up to six atmospheres total pressure. The restoration of this facility is of critical importance to the future of the U.S. aerospace industry. As part of this project, several state of the art model support systems are furnished to provide an optimal balance between aerodynamic and operational efficiency parameters. Two model support systems, the Rear Strut Model Support, and the High Angle of Attack Model Support are discussed. This paper covers design parameters, constraints, development, description, and component selection.

The self streamlining wind tunnel. [wind tunnel walls

NASA Technical Reports Server (NTRS)

Goodyer, M. J.

1975-01-01

A two dimensional test section in a low speed wind tunnel capable of producing flow conditions free from wall interference is presented. Flexible top and bottom walls, and rigid sidewalls from which models were mounted spanning the tunnel are shown. All walls were unperforated, and the flexible walls were positioned by screw jacks. To eliminate wall interference, the wind tunnel itself supplied the information required in the streamlining process, when run with the model present. Measurements taken at the flexible walls were used by the tunnels computer check wall contours. Suitable adjustments based on streamlining criteria were then suggested by the computer. The streamlining criterion adopted when generating infinite flowfield conditions was a matching of static pressures in the test section at a wall with pressures computed for an imaginary inviscid flowfield passing over the outside of the same wall. Aerodynamic data taken on a cylindrical model operating under high blockage conditions are presented to illustrate the operation of the tunnel in its various modes.

Portable Test And Monitoring System For Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Poupard, Charles A.

1987-01-01

Portable system developed to test and monitor instrumentation used in wind-tunnel models. Self-contained and moves easily to model, either before or after model installed in wind tunnel. System is 44 1/2 in. high, 22 in. wide, and 17 in. deep and weighs 100 lb. Primary benefits realized with portable test and monitoring system associated with saving of time.

Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Ting, Eric; Lebofsky, Sonia; Nguyen, Nhan; Trinh, Khanh

2014-01-01

This paper presents an approach to the development of a scaled wind tunnel model for static aeroelastic similarity with a full-scale wing model. The full-scale aircraft model is based on the NASA Generic Transport Model (GTM) with flexible wing structures referred to as the Elastically Shaped Aircraft Concept (ESAC). The baseline stiffness of the ESAC wing represents a conventionally stiff wing model. Static aeroelastic scaling is conducted on the stiff wing configuration to develop the wind tunnel model, but additional tailoring is also conducted such that the wind tunnel model achieves a 10% wing tip deflection at the wind tunnel test condition. An aeroelastic scaling procedure and analysis is conducted, and a sub-scale flexible wind tunnel model based on the full-scale's undeformed jig-shape is developed. Optimization of the flexible wind tunnel model's undeflected twist along the span, or pre-twist or wash-out, is then conducted for the design test condition. The resulting wind tunnel model is an aeroelastic model designed for the wind tunnel test condition.

Modeling the Benchmark Active Control Technology Wind-Tunnel Model for Application to Flutter Suppression

NASA Technical Reports Server (NTRS)

Waszak, Martin R.

1996-01-01

This paper describes the formulation of a model of the dynamic behavior of the Benchmark Active Controls Technology (BACT) wind-tunnel model for application to design and analysis of flutter suppression controllers. The model is formed by combining the equations of motion for the BACT wind-tunnel model with actuator models and a model of wind-tunnel turbulence. The primary focus of this paper is the development of the equations of motion from first principles using Lagrange's equations and the principle of virtual work. A numerical form of the model is generated using values for parameters obtained from both experiment and analysis. A unique aspect of the BACT wind-tunnel model is that it has upper- and lower-surface spoilers for active control. Comparisons with experimental frequency responses and other data show excellent agreement and suggest that simple coefficient-based aerodynamics are sufficient to accurately characterize the aeroelastic response of the BACT wind-tunnel model. The equations of motion developed herein have been used to assist the design and analysis of a number of flutter suppression controllers that have been successfully implemented.

Mercury Capsule Model in the 1- by 1-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1959-10-21

National Aeronautics and Space Administration (NASA) researchers install a small-scale model of the capsule for Project Mercury in the 1- by 1-Foot Supersonic Wind Tunnel at the Lewis Research Center. NASA Lewis conducted a variety of tests for Project Mercury, including retrorocket calibration, escape tower engine performance, and separation of the capsule from simulated Atlas and Redstone boosters. The test of this capsule and escape tower model in the 1- by 1-foot tunnel were run in January and February 1960. The 1-by 1-Foot Supersonic Wind Tunnel had a 15-inch long test section, seen here, that was one foot wide and one foot high. The sides were made of glass to allow cameras to capture the supersonic air flow over the models. The tunnel could generate air flows from Mach 1.3 to 3.0. At the time, it was one of nine small supersonic wind tunnels at Lewis. These tunnels used the exhauster and compressor equipment of the larger facilities. The 1- by 1 tunnel, which began operating in the early 1950s, was built inside a test cell in the expansive Engine Research Building. During the 1950s the 1- by 1 was used to study a variety of inlets, nozzles, and cones for missiles and scramjets. The Mercury capsule tests were among the last at the facility for many years. The tunnel was mothballed in 1960. The 1- by 1 was briefly restored in 1972, then brought back online for good in 1979. The facility has maintained a brisk operating schedule ever since.

Propulsion simulation for magnetically suspended wind tunnel models

NASA Technical Reports Server (NTRS)

Joshi, Prakash B.; Beerman, Henry P.; Chen, James; Krech, Robert H.; Lintz, Andrew L.; Rosen, David I.

1990-01-01

The feasibility of simulating propulsion-induced aerodynamic effects on scaled aircraft models in wind tunnels employing Magnetic Suspension and Balance Systems. The investigation concerned itself with techniques of generating exhaust jets of appropriate characteristics. The objectives were to: (1) define thrust and mass flow requirements of jets; (2) evaluate techniques for generating propulsive gas within volume limitations imposed by magnetically-suspended models; (3) conduct simple diagnostic experiments for techniques involving new concepts; and (4) recommend experiments for demonstration of propulsion simulation techniques. Various techniques of generating exhaust jets of appropriate characteristics were evaluated on scaled aircraft models in wind tunnels with MSBS. Four concepts of remotely-operated propulsion simulators were examined. Three conceptual designs involving innovative adaptation of convenient technologies (compressed gas cylinders, liquid, and solid propellants) were developed. The fourth innovative concept, namely, the laser-assisted thruster, which can potentially simulate both inlet and exhaust flows, was found to require very high power levels for small thrust levels.

The aeolian wind tunnel

NASA Technical Reports Server (NTRS)

Iversen, J. D.

1991-01-01

The aeolian wind tunnel is a special case of a larger subset of the wind tunnel family which is designed to simulate the atmospheric surface layer winds to small scale (a member of this larger subset is usually called an atmospheric boundary layer wind tunnel or environmental wind tunnel). The atmospheric boundary layer wind tunnel is designed to simulate, as closely as possible, the mean velocity and turbulence that occur naturally in the atmospheric boundary layer (defined as the lowest portion of the atmosphere, of the order of 500 m, in which the winds are most greatly affected by surface roughness and topography). The aeolian wind tunnel is used for two purposes: to simulate the physics of the saltation process and to model at small scale the erosional and depositional processes associated with topographic surface features. For purposes of studying aeolian effects on the surface of Mars and Venus as well as on Earth, the aeolian wind tunnel continues to prove to be a useful tool for estimating wind speeds necessary to move small particles on the three planets as well as to determine the effects of topography on the evolution of aeolian features such as wind streaks and dune patterns.

V/STOL wind-tunnel testing

NASA Technical Reports Server (NTRS)

Koenig, D. G.

1984-01-01

Factors influencing effective program planning for V/STOL wind-tunnel testing are discussed. The planning sequence itself, which includes a short checklist of considerations that could enhance the value of the tests, is also described. Each of the considerations, choice of wind tunnel, type of model installation, model development and test operations, is discussed, and examples of appropriate past and current V/STOL test programs are provided. A short survey of the moderate to large subsonic wind tunnels is followed by a review of several model installations, from two-dimensional to large-scale models of complete aircraft configurations. Model sizing, power simulation, and planning are treated, including three areas is test operations: data-acquisition systems, acoustic measurements in wind tunnels, and flow surveying.

Visualizing Flutter Mechanism as Traveling Wave Through Animation of Simulation Results for the Semi-Span Super-Sonic Transport Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Christhilf, David M.

2014-01-01

It has long been recognized that frequency and phasing of structural modes in the presence of airflow play a fundamental role in the occurrence of flutter. Animation of simulation results for the long, slender Semi-Span Super-Sonic Transport (S4T) wind-tunnel model demonstrates that, for the case of mass-ballasted nacelles, the flutter mode can be described as a traveling wave propagating downstream. Such a characterization provides certain insights, such as (1) describing the means by which energy is transferred from the airflow to the structure, (2) identifying airspeed as an upper limit for speed of wave propagation, (3) providing an interpretation for a companion mode that coalesces in frequency with the flutter mode but becomes very well damped, (4) providing an explanation for bursts of response to uniform turbulence, and (5) providing an explanation for loss of low frequency (lead) phase margin with increases in dynamic pressure (at constant Mach number) for feedback systems that use sensors located upstream from active control surfaces. Results from simulation animation, simplified modeling, and wind-tunnel testing are presented for comparison. The simulation animation was generated using double time-integration in Simulink of vertical accelerometer signals distributed over wing and fuselage, along with time histories for actuated control surfaces. Crossing points for a zero-elevation reference plane were tracked along a network of lines connecting the accelerometer locations. Accelerometer signals were used in preference to modal displacement state variables in anticipation that the technique could be used to animate motion of the actual wind-tunnel model using data acquired during testing. Double integration of wind-tunnel accelerometer signals introduced severe drift even with removal of both position and rate biases such that the technique does not currently work. Using wind-tunnel data to drive a Kalman filter based upon fitting coefficients to

Turboprop Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1976-08-21

National Aeronautics and Space Administration (NASA) engineer Robert Jeracki prepares a Hamilton Standard SR-1 turboprop model in the test section of the 8- by 6-Foot Supersonic Wind Tunnel at the Lewis Research Center. Lewis researchers were analyzing a series of eight-bladed propellers in their wind tunnels to determine their operating characteristics at speeds up to Mach 0.8. The program, which became the Advanced Turboprop, was part of a NASA-wide Aircraft Energy Efficiency Program which was designed to reduce aircraft fuel costs by 50 percent. The ATP concept was different from the turboprops in use in the 1950s. The modern versions had at least eight blades and were swept back for better performance. After Lewis researchers developed the advanced turboprop theory and established its potential performance capabilities, they commenced an almost decade-long partnership with Hamilton Standard to develop, verify, and improve the concept. A series of 24-inch scale models of the SR-1 with different blade shapes and angles were tested in Lewis’ wind tunnels. A formal program was established in 1978 to examine associated noise levels, aerodynamics, and the drive system. The testing of the large-scale propfan was done on test rigs, in large wind tunnels, and, eventually, on aircraft.

Wind Tunnel Management and Resource Optimization: A Systems Modeling Approach

NASA Technical Reports Server (NTRS)

Jacobs, Derya, A.; Aasen, Curtis A.

2000-01-01

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

Aeroelastic instability stoppers for wind tunnel models

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Ricketts, R. H. (Inventor)

1981-01-01

A mechanism for constraining models or sections thereof, was wind tunnel tested, deployed at the onset of aeroelastic instability, to forestall destructive vibrations in the model is described. The mechanism includes a pair of arms pivoted to the tunnel wall and straddling the model. Rollers on the ends of the arms contact the model, and are pulled together against the model by a spring stretched between the arms. An actuator mechanism swings the arms into place and back as desired.

1/50 Scale Model Of The 80x120 Foot Wind Tunnel Model (NFAC) In The Test Section Of The 40x80 Wind Tunnel.

NASA Image and Video Library

1996-06-27

(03/12/1976) 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 wind tunnel. Model viewed from the west, mounted on a rotating ground board designed for this test. Ramp leading to ground board includes a generic building placed in front of the 80x120 inlet.

A76-0634. 1/50 Scale Model Of The 80X120 Foot Wind Tunnel Model (Nfac) In The Test Section Of The 40X80 Foot Wind Tunnel.

NASA Image and Video Library

1996-06-27

(03/12/1976) 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 foot wind tunnel. Model mounted on a rotating ground board designed for this test, viewed from the west, oriented for North wind.

Modeling the Benchmark Active Control Technology Wind-Tunnel Model for Active Control Design Applications

NASA Technical Reports Server (NTRS)

Waszak, Martin R.

1998-01-01

This report describes the formulation of a model of the dynamic behavior of the Benchmark Active Controls Technology (BACT) wind tunnel model for active control design and analysis applications. The model is formed by combining the equations of motion for the BACT wind tunnel model with actuator models and a model of wind tunnel turbulence. The primary focus of this report is the development of the equations of motion from first principles by using Lagrange's equations and the principle of virtual work. A numerical form of the model is generated by making use of parameters obtained from both experiment and analysis. Comparisons between experimental and analytical data obtained from the numerical model show excellent agreement and suggest that simple coefficient-based aerodynamics are sufficient to accurately characterize the aeroelastic response of the BACT wind tunnel model. The equations of motion developed herein have been used to aid in the design and analysis of a number of flutter suppression controllers that have been successfully implemented.

General Dynamics YF-16 Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1974-01-21

A model of the General Dynamics YF-16 Fighting Falcon in the test section of the 8- by 6-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. The YF-16 was General Dynamics response to the military’s 1972 request for proposals to design a new 20,000-pound fighter jet with exceptional acceleration, turn rate, and range. The aircraft included innovative design elements to help pilots survive turns up to 9Gs, a new frameless bubble canopy, and a Pratt and Whitney 24,000-pound thrust F-100 engine. The YF-16 made its initial flight in February 1974, just six weeks before this photograph, at Edwards Air Force Base. Less than a year later, the Air Force ordered 650 of the aircraft, designated as F-16 Fighting Falcons. The March and April 1974 tests in the 8- by 6-foot tunnel analyzed the aircraft’s fixed-shroud ejector nozzle. The fixed-nozzle area limited drag, but also limited the nozzle’s internal performance. NASA researchers identified and assessed aerodynamic and aerodynamic-propulsion interaction uncertainties associated the prototype concept. YF-16 models were also tested extensively in the 11- by 11-Foot Transonic Wind Tunnel and 9- by 7-Foot Supersonic Wind Tunnel at Ames Research Center and the 12-Foot Pressure Wind Tunnel at Langley Research Center.

Wind-Tunnel Modeling of Flow Diffusion over an Urban Complex.

DTIC Science & Technology

URBAN AREAS, *ATMOSPHERIC MOTION, *AIR POLLUTION, ATMOSPHERIC MOTION, WIND TUNNEL MODELS, HEAT, DIFFUSION , TURBULENT BOUNDARY LAYER, WIND, SKIN FRICTION, MATHEMATICAL MODELS, URBAN PLANNING, INDIANA.

An experimental study of several wind tunnel wall configurations using two V/STOL model configurations. [low speed wind tunnels

NASA Technical Reports Server (NTRS)

Binion, T. W., Jr.

1975-01-01

Experiments were conducted in the low speed wind tunnel using two V/STOL models, a jet-flap and a jet-in-fuselage configuration, to search for a wind tunnel wall configuration to minimize wall interference on V/STOL models. Data were also obtained on the jet-flap model with a uniform slotted wall configuration to provide comparisons between theoretical and experimental wall interference. A test section configuration was found which provided some data in reasonable agreement with interference-free results over a wide range of momentum coefficients.

Effects of vibration on inertial wind-tunnel model attitude measurement devices

NASA Technical Reports Server (NTRS)

Young, Clarence P., Jr.; Buehrle, Ralph D.; Balakrishna, S.; Kilgore, W. Allen

1994-01-01

Results of an experimental study of a wind tunnel model inertial angle-of-attack sensor response to a simulated dynamic environment are presented. The inertial device cannot distinguish between the gravity vector and the centrifugal accelerations associated with wind tunnel model vibration, this situation results in a model attitude measurement bias error. Significant bias error in model attitude measurement was found for the model system tested. The model attitude bias error was found to be vibration mode and amplitude dependent. A first order correction model was developed and used for estimating attitude measurement bias error due to dynamic motion. A method for correcting the output of the model attitude inertial sensor in the presence of model dynamics during on-line wind tunnel operation is proposed.

A directional microphone array for acoustic studies of wind tunnel models

NASA Technical Reports Server (NTRS)

Soderman, P. T.; Noble, S. C.

1974-01-01

An end-fire microphone array that utilizes a digital time delay system has been designed and evaluated for measuring noise in wind tunnels. The directional response of both a four- and eight-element linear array of microphones has enabled substantial rejection of background noise and reverberations in the NASA Ames 40- by 80-foot wind tunnel. In addition, it is estimated that four- and eight-element arrays reject 6 and 9 dB, respectively, of microphone wind noise, as compared with a conventional omnidirectional microphone with nose cone. Array response to two types of jet engine models in the wind tunnel is presented. Comparisons of array response to loudspeakers in the wind tunnel and in free field are made.

1/50 Scale Model Of The 80X120 Foot Wind Tunnel Model (NFAC) In The Test Section Of The 40X80 Wind Tunnel At Nasa Ames.

NASA Image and Video Library

1976-03-12

(03/12/1976) Overhead view of 1/50 scale model of the 80x120 foot wind tunnel model (NFAC) in the test section of the 40x80 wind tunnel at NASA Ames. Model mounted on a rotating ground board designed for this test.

Space Shuttle Pressure Data Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1978-04-21

Technicians examine a scale model of the space shuttle used to obtain pressure data during tests in the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers used the 10- by 10 tunnel extensively in the 1970s to study shuttle configurations in order to forecast conditions during an actual flight. These tests included analysis of the solid rocket boosters’ aerodynamics, orbiter forebody angle -of -attack and air speed, base heating for entire shuttle, and engine-out loads. The test seen in this photograph used a 3.5- percent scale aluminum alloy model of the entire launch configuration. The program was designed to obtain aerodynamic pressure data. The tests were part of a larger program to study possible trouble areas for the shuttle’s new Advanced Flexible Reusable Surface Insulation. The researchers obtained aeroacoustic data and pressure distributions from five locations on the model. Over 100 high-temperature pressure transducers were attached to the model. Other portions of the test program were conducted at Lewis’ 8- by 6-Foot Supersonic Wind Tunnel and the 11- by 11-Foot Transonic Wind Tunnel at Ames Research Center.

A lumped parameter mathematical model for simulation of subsonic wind tunnels

NASA Technical Reports Server (NTRS)

Krosel, S. M.; Cole, G. L.; Bruton, W. M.; Szuch, J. R.

1986-01-01

Equations for a lumped parameter mathematical model of a subsonic wind tunnel circuit are presented. The equation state variables are internal energy, density, and mass flow rate. The circuit model is structured to allow for integration and analysis of tunnel subsystem models which provide functions such as control of altitude pressure and temperature. Thus the model provides a useful tool for investigating the transient behavior of the tunnel and control requirements. The model was applied to the proposed NASA Lewis Altitude Wind Tunnel (AWT) circuit and included transfer function representations of the tunnel supply/exhaust air and refrigeration subsystems. Both steady state and frequency response data are presented for the circuit model indicating the type of results and accuracy that can be expected from the model. Transient data for closed loop control of the tunnel and its subsystems are also presented, demonstrating the model's use as a control analysis tool.

Cryogenic Wind Tunnel Models. Design and Fabrication

NASA Technical Reports Server (NTRS)

Young, C. P., Jr. (Compiler); Gloss, B. B. (Compiler)

1983-01-01

The principal motivating factor was the National Transonic Facility (NTF). Since the NTF can achieve significantly higher Reynolds numbers at transonic speeds than other wind tunnels in the world, and will therefore occupy a unique position among ground test facilities, every effort is being made to ensure that model design and fabrication technology exists to allow researchers to take advantage of this high Reynolds number capability. Since a great deal of experience in designing and fabricating cryogenic wind tunnel models does not exist, and since the experience that does exist is scattered over a number of organizations, there is a need to bring existing experience in these areas together and share it among all interested parties. Representatives from government, the airframe industry, and universities are included.

Effect of thermal stability/complex terrain on wind turbine model(s): a wind tunnel study to address complex atmospheric conditions

NASA Astrophysics Data System (ADS)

Guala, M.; Hu, S. J.; Chamorro, L. P.

2011-12-01

Turbulent boundary layer measurements in both wind tunnel and in the near-neutral atmospheric surface layer revealed in the last decade the significant contribution of the large scales of motions to both turbulent kinetic energy and Reynolds stresses, for a wide range of Reynolds number. These scales are known to grow throughout the logarithmic layer and to extend several boundary layer heights in the streamwise direction. Potentially, they are a source of strong unsteadiness in the power output of wind turbines and in the aerodynamic loads of wind turbine blades. However, the large scales in realistic atmospheric conditions deserves further study, with well controlled boundary conditions. In the atmospheric wind tunnel of the St. Anthony Falls Laboratory, with a 16 m long test section and independently controlled incoming flow and floor temperatures, turbulent boundary layers in a range of stability conditions, from the stratified to the convective case, can be reproduced and monitored. Measurements of fluctuating temperature, streamwise and wall normal velocity components are simultaneously obtained by an ad hoc calibrated and customized triple-wire sensor. A wind turbine model with constant loading DC motor, constant tip speed ratio, and a rotor diameter of 0.128m is used to mimic a large full scale turbine in the atmospheric boundary layer. Measurements of the fluctuating voltage generated by the DC motor are compared with measurements of the blade's angular velocity by laser scanning, and eventually related to velocity measurements from the triple-wire sensor. This study preliminary explores the effect of weak stability and complex terrain (through a set of spanwise aligned topographic perturbations) on the large scales of the flow and on the fluctuations in the wind turbine(s) power output.

Dry wind tunnel system

NASA Technical Reports Server (NTRS)

Chen, Ping-Chih (Inventor)

2013-01-01

This invention is a ground flutter testing system without a wind tunnel, called Dry Wind Tunnel (DWT) System. The DWT system consists of a Ground Vibration Test (GVT) hardware system, a multiple input multiple output (MIMO) force controller software, and a real-time unsteady aerodynamic force generation software, that is developed from an aerodynamic reduced order model (ROM). The ground flutter test using the DWT System operates on a real structural model, therefore no scaled-down structural model, which is required by the conventional wind tunnel flutter test, is involved. Furthermore, the impact of the structural nonlinearities on the aeroelastic stability can be included automatically. Moreover, the aeroservoelastic characteristics of the aircraft can be easily measured by simply including the flight control system in-the-loop. In addition, the unsteady aerodynamics generated computationally is interference-free from the wind tunnel walls. Finally, the DWT System can be conveniently and inexpensively carried out as a post GVT test with the same hardware, only with some possible rearrangement of the shakers and the inclusion of additional sensors.

Measurement of deformations of models in a wind tunnel

NASA Astrophysics Data System (ADS)

Charpin, F.; Armand, C.; Selvaggini, R.

Techniques used at the ONERA Modane Center to monitor geometric variations in scale-models in wind tunnel trials are described. The methods include: photography of reflections from mirrors embedded in the model surface; laser-based torsiometry with polarized mirrors embedded in the model surface; predictions of the deformations using numerical codes for the model surface mechanical characteristics and the measured surface stresses; and, use of an optical detector to monitor the position of luminous fiber optic sources emitting from the model surfaces. The data enhance the confidence that the wind tunnel aerodynamic data will correspond with the in-flight performance of full scale flight surfaces.

Construction of the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1948-06-21

The 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory was the nation’s largest supersonic facility when it began operation in April 1949. The emergence of new propulsion technologies such as turbojets, ramjets, and rockets during World War II forced the NACA and the aircraft industry to develop new research tools. In late 1945 the NACA began design work for new large supersonic wind tunnels at its three laboratories. The result was the 4- by 4-Foot Supersonic Wind Tunnel at Langley Memorial Aeronautical Laboratory, 6- by 6-foot supersonic wind tunnel at Ames Aeronautical Laboratory, and the largest facility, the 8- by 6-Foot Supersonic Wind Tunnel in Cleveland. The two former tunnels were to study aerodynamics, while the 8- by 6 facility was designed for supersonic propulsion. The 8- by 6-Foot Supersonic Wind Tunnel was used to study propulsion systems, including inlets and exit nozzles, combustion fuel injectors, flame holders, exit nozzles, and controls on ramjet and turbojet engines. Flexible sidewalls alter the tunnel’s nozzle shape to vary the Mach number during operation. A seven-stage axial compressor, driven by three electric motors that yield a total of 87,000 horsepower, generates air speeds from Mach 0.36 to 2.0. A section of the tunnel is seen being erected in this photograph.

Evaluation of a micro-scale wind model's performance over realistic building clusters using wind tunnel experiments

NASA Astrophysics Data System (ADS)

Zhang, Ning; Du, Yunsong; Miao, Shiguang; Fang, Xiaoyi

2016-08-01

The simulation performance over complex building clusters of a wind simulation model (Wind Information Field Fast Analysis model, WIFFA) in a micro-scale air pollutant dispersion model system (Urban Microscale Air Pollution dispersion Simulation model, UMAPS) is evaluated using various wind tunnel experimental data including the CEDVAL (Compilation of Experimental Data for Validation of Micro-Scale Dispersion Models) wind tunnel experiment data and the NJU-FZ experiment data (Nanjing University-Fang Zhuang neighborhood wind tunnel experiment data). The results show that the wind model can reproduce the vortexes triggered by urban buildings well, and the flow patterns in urban street canyons and building clusters can also be represented. Due to the complex shapes of buildings and their distributions, the simulation deviations/discrepancies from the measurements are usually caused by the simplification of the building shapes and the determination of the key zone sizes. The computational efficiencies of different cases are also discussed in this paper. The model has a high computational efficiency compared to traditional numerical models that solve the Navier-Stokes equations, and can produce very high-resolution (1-5 m) wind fields of a complex neighborhood scale urban building canopy (~ 1 km ×1 km) in less than 3 min when run on a personal computer.

Wind Tunnel Test of NASA’s Most Powerful Rocket (360° Animation)

NASA Image and Video Library

2018-01-08

What are wind tunnels? And how do they help researchers design and test next-generation aircraft and spacecraft? This interactive 360° animation takes you inside the Unitary Plan Wind Tunnel at NASA’s Ames Research Center in Silicon Valley. The facility is one of seven wind tunnels located at Ames for exploring the complex physics of flight. The video features a four percent scale model of NASA’s most powerful rocket, the Space Launch System, or SLS. Two SLS models--one silver and one pink--appear in the video. The latter is coated with a special paint to track surface pressure readings during testing. Once built, the SLS rocket will be capable of sending astronauts on bold new missions into deep space. How to watch 360 content in VR? YouTube and Google Cardboard 1. Open YouTube on your mobile device and select the video. 2. Click the Google Cardboard icon on the bottom right. 3. Insert the mobile device into the Google Cardboard device. 4. Watch through the headset. Samsung Gear VR 1. Download the 360 mp4 video file. 2. Create a folder in the root directory of your device or SD Card called “MilkVR” 3. Put the video file in that folder. 4. Open the Samsung VR application from the Oculus App 5. Insert the phone into the Gear VR 6. Put on the VR headset. 7. Navigate to the section called “Sideloaded” 8. Select the video from “Storage 1”. 9. The optimal viewing format is 360 x 360. Change the format by selecting thing format icon on the bottom right. PlayStation VR 1. Download the 360 mp4 video file from NASA.gov. 2. Create a folder on a USB drive, formatted in FAT32 or exFat. 3. Copy the video file into that folder. 4. Insert the USB drive in the PlayStation 4 5. Connect the PlayStation VR headset to the PlayStation 4 and turn on the power. 6. Put on the VR headset. 7. Open the PlayStation Media Player (updated to v2.50 or higher). 8. Be sure the Media Player is set to “VR Mode” by holding the “Option” button to enable it. 9. Open the video

Wind tunnel measurements of the power output variability and unsteady loading in a micro wind farm model

NASA Astrophysics Data System (ADS)

Bossuyt, Juliaan; Howland, Michael; Meneveau, Charles; Meyers, Johan

2015-11-01

To optimize wind farm layouts for a maximum power output and wind turbine lifetime, mean power output measurements in wind tunnel studies are not sufficient. Instead, detailed temporal information about the power output and unsteady loading from every single wind turbine in the wind farm is needed. A very small porous disc model with a realistic thrust coefficient of 0.75 - 0.85, was designed. The model is instrumented with a strain gage, allowing measurements of the thrust force, incoming velocity and power output with a frequency response up to the natural frequency of the model. This is shown by reproducing the -5/3 spectrum from the incoming flow. Thanks to its small size and compact instrumentation, the model allows wind tunnel studies of large wind turbine arrays with detailed temporal information from every wind turbine. Translating to field conditions with a length-scale ratio of 1:3,000 the frequencies studied from the data reach from 10-4 Hz up to about 6 .10-2 Hz. The model's capabilities are demonstrated with a large wind farm measurement consisting of close to 100 instrumented models. A high correlation is found between the power outputs of stream wise aligned wind turbines, which is in good agreement with results from prior LES simulations. Work supported by ERC (ActiveWindFarms, grant no. 306471) and by NSF (grants CBET-113380 and IIA-1243482, the WINDINSPIRE project).

On the aero-elastic design of the DTU 10MW wind turbine blade for the LIFES50+ wind tunnel scale model

NASA Astrophysics Data System (ADS)

Bayati, I.; Belloli, M.; Bernini, L.; Mikkelsen, R.; Zasso, A.

2016-09-01

This paper illustrates the aero-elastic optimal design, the realization and the verification of the wind tunnel scale model blades for the DTU 10 MW wind turbine model, within LIFES50+ project. The aerodynamic design was focused on the minimization of the difference, in terms of thrust coefficient, with respect to the full scale reference. From the Selig low Reynolds database airfoils, the SD7032 was chosen for this purpose and a proper constant section wing was tested at DTU red wind tunnel, providing force and distributed pressure coefficients for the design, in the Reynolds range 30-250 E3 and for different angles of attack. The aero-elastic design algorithm was set to define the optimal spanwise thickness over chord ratio (t/c), the chord length and the twist to match the first flapwise scaled natural frequency. An aluminium mould for the carbon fibre was CNC manufactured based on B-Splines CAD definition of the external geometry. Then the wind tunnel tests at Politecnico di Milano confirmed successful design and manufacturing approaches.

A numerical study of the effects of wind tunnel wall proximity on an airfoil model

NASA Technical Reports Server (NTRS)

Potsdam, Mark; Roberts, Leonard

1990-01-01

A procedure was developed for modeling wind tunnel flows using computational fluid dynamics. Using this method, a numerical study was undertaken to explore the effects of solid wind tunnel wall proximity and Reynolds number on a two-dimensional airfoil model at low speed. Wind tunnel walls are located at varying wind tunnel height to airfoil chord ratios and the results are compared with freestream flow in the absence of wind tunnel walls. Discrepancies between the constrained and unconstrained flows can be attributed to the presence of the walls. Results are for a Mach Number of 0.25 at angles of attack through stall. A typical wind tunnel Reynolds number of 1,200,000 and full-scale flight Reynolds number of 6,000,000 were investigated. At this low Mach number, wind tunnel wall corrections to Mach number and angle of attack are supported. Reynolds number effects are seen to be a consideration in wind tunnel testing and wall interference correction methods. An unstructured grid Navier-Stokes code is used with a Baldwin-Lomax turbulence model. The numerical method is described since unstructured flow solvers present several difficulties and fundamental differences from structured grid codes, especially in the area of turbulence modeling and grid generation.

Static and wind tunnel model tests for the development of externally blown flap noise reduction techniques

NASA Technical Reports Server (NTRS)

Pennock, A. P.; Swift, G.; Marbert, J. A.

1975-01-01

Externally blown flap models were tested for noise and performance at one-fifth scale in a static facility and at one-tenth scale in a large acoustically-treated wind tunnel. The static tests covered two flap designs, conical and ejector nozzles, third-flap noise-reduction treatments, internal blowing, and flap/nozzle geometry variations. The wind tunnel variables were triple-slotted or single-slotted flaps, sweep angle, and solid or perforated third flap. The static test program showed the following noise reductions at takeoff: 1.5 PNdB due to treating the third flap; 0.5 PNdB due to blowing from the third flap; 6 PNdB at flyover and 4.5 PNdB in the critical sideline plane (30 deg elevation) due to installation of the ejector nozzle. The wind tunnel program showed a reduction of 2 PNdB in the sideline plane due to a forward speed of 43.8 m/s (85 kn). The best combination of noise reduction concepts reduced the sideline noise of the reference aircraft at constant field length by 4 PNdB.

Evaluation of an I-box wind tunnel model for assessment of behavioral responses of blow flies.

PubMed

Moophayak, Kittikhun; Sukontason, Kabkaew L; Kurahashi, Hiromu; Vogtsberger, Roy C; Sukontason, Kom

2013-11-01

The behavioral response of flies to olfactory cues remains the focus of many investigations, and wind tunnels have sometimes been employed for assessment of this variable in the laboratory. In this study, our aim was to design, construct, and operate a new model of I-box wind tunnel with improved efficacy, highlighting the use of a new wind tunnel model to investigate the behavioral response of the medically important blow fly, Chrysomya megacephala (Fabricius). The I-box dual-choice wind tunnel designed for this study consists of seven conjoined compartments that resulted in a linear apparatus with clear glass tunnel of 30 × 30 × 190 cm ended both sides with wooden "fan compartments" which are equipped with adjustable fans as wind source. The clear glass tunnel consisted of two "stimulus compartments" with either presence or absence (control) of bait; two "trap compartments" where flies were attracted and allowed to reside; and one central "release compartment" where flies were introduced. Wind tunnel experiments were carried out in a temperature-controlled room, with a room light as a light source and a room-ventilated fan as odor-remover from tunnel out. Evaluation of testing parameters revealed that the highest attractive index was achieved with the use of 300 g of 1-day tainted pork scrap (pork meat mixed with offal) as bait in wind tunnel settings wind speed of 0.58 m/s, during 1.00-5.00 PM with light intensity of 341.33 lux from vertical light and 135.93 lux from horizontal light for testing a group of 60 flies. In addition, no significant response of well-fed and 24 h staved flies to this bait under these conditions was found. Results of this study supported this new wind tunnel model as a suitable apparatus for investigation of behavioral response of blow flies to bait chemical cues in the laboratory.

Water tunnel flow visualization and wind tunnel data analysis of the F/A-18. [leading edge extension vortex effects

NASA Technical Reports Server (NTRS)

Erickson, G. E.

1982-01-01

Six degree of freedom studies were utilized to extract a band of yawing and rolling moment coefficients from the F/A-18 aircraft flight records. These were compared with 0.06 scale model data obtained in a 16T wind tunnel facility. The results, indicate the flight test yawing moment data exhibit an improvement over the wind tunnel data to near neutral stability and a significant reduction in lateral stability (again to anear neutral level). These data are consistent with the flight test results since the motion was characterized by a relatively slo departure. Flight tests repeated the slow yaw departure at M 0.3. Only 0.16 scale model wind tunnel data showed levels of lateral stability similar to the flight test results. Accordingly, geometric modifications were investigated on the 0.16 scale model in the 30x60 foot wind tunnel to improve high angle of attack lateral stability.

Accessing Wind Tunnels From NASA's Information Power Grid

NASA Technical Reports Server (NTRS)

Becker, Jeff; Biegel, Bryan (Technical Monitor)

2002-01-01

The NASA Ames wind tunnel customers are one of the first users of the Information Power Grid (IPG) storage system at the NASA Advanced Supercomputing Division. We wanted to be able to store their data on the IPG so that it could be accessed remotely in a secure but timely fashion. In addition, incorporation into the IPG allows future use of grid computational resources, e.g., for post-processing of data, or to do side-by-side CFD validation. In this paper, we describe the integration of grid data access mechanisms with the existing DARWIN web-based system that is used to access wind tunnel test data. We also show that the combined system has reasonable performance: wind tunnel data may be retrieved at 50Mbits/s over a 100 base T network connected to the IPG storage server.

Introduction to cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Goodyer, M. J.

1985-01-01

The background to the evolution of the cryogenic wind tunnel is outlined, with particular reference to the late 60's/early 70's when efforts were begun to re-equip with larger wind tunnels. The problems of providing full scale Reynolds numbers in transonic testing were proving particularly intractible, when the notion of satisfying the needs with the cryogenic tunnel was proposed, and then adopted. The principles and advantages of the cryogenic tunnel are outlined, along with guidance on the coolant needs when this is liquid nitrogen, and with a note on energy recovery. Operational features of the tunnels are introduced with reference to a small low speed tunnel. Finally the outstanding contributions are highlighted of the 0.3-Meter Transonic Cryogenic Tunnel (TCT) at NASA Langley Research Center, and its personnel, to the furtherance of knowledge and confidence in the concept.

The Langley Wind Tunnel Enterprise

NASA Technical Reports Server (NTRS)

Paulson, John W., Jr.; Kumar, Ajay; Kegelman, Jerome T.

1998-01-01

After 4 years of existence, the Langley WTE is alive and growing. Significant improvements in the operation of wind tunnels have been demonstrated and substantial further improvements are expected when we are able to truly address and integrate all the processes affecting the wind tunnel testing cycle.

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

DOE Office of Scientific and Technical Information (OSTI.GOV)

Parker, M.J.; Addis, R.P.

1991-04-04

The Department of Energy (DOE) Environment, Safety and Health Compliance Assessment (Tiger Team) of the Savannah River Site (SRS) questioned the method by which wind speed sensors (cup anemometers) are calibrated by the Environmental Technology Section (ETS). The Tiger Team member was concerned that calibration data was generated by running the wind tunnel to only 26 miles per hour (mph) when speeds exceeding 50 mph are readily obtainable. A wind tunnel experiment was conducted and confirmed the validity of the practice. Wind speeds common to SRS (6 mph) were predicted more accurately by 0--25 mph regression equations than 0--50 mphmore » regression equations. Higher wind speeds were slightly overpredicted by the 0--25 mph regression equations when compared to 0--50 mph regression equations. However, the greater benefit of more accurate lower wind speed predictions accuracy outweight the benefit of slightly better high (extreme) wind speed predictions. Therefore, it is concluded that 0--25 mph regression equations should continue to be utilized by ETS at SRS. During the Department of Energy Tiger Team audit, concerns were raised about the calibration of SRS cup anemometers. Wind speed is measured by ETS with Teledyne Geotech model 1564B cup anemometers, which are calibrated in the ETS wind tunnel. Linear regression lines are fitted to data points of tunnel speed versus anemometer output voltages up to 25 mph. The regression coefficients are then implemented into the data acquisition computer software when an instrument is installed in the field. The concern raised was that since the wind tunnel at SRS is able to generate a maximum wind speed higher than 25 mph, errors may be introduced in not using the full range of the wind tunnel.« less

Wind tunnel test of Teledyne Geotech model 1564B cup anemometer

NASA Astrophysics Data System (ADS)

Parker, M. J.; Addis, R. P.

1991-04-01

The Department of Energy (DOE) Environment, Safety, and Health Compliance Assessment (Tiger Team) of the Savannah River Site (SRS) questioned the method by which wind speed sensors (cup anemometers) are calibrated by the Environmental Technology Section (ETS). The Tiger Team member was concerned that calibration data was generated by running the wind tunnel to only 26 miles per hour (mph) when speeds exceeding 50 mph are readily obtainable. A wind tunnel experiment was conducted and confirmed the validity of the practice. Wind speeds common to SRS (6 mph) were predicted more accurately by 0-25 mph regression equations than 0-50 mph regression equations. Higher wind speeds were slightly overpredicted by the 0-25 mph regression equations when compared to 0-50 mph regression equations. However, the greater benefit of more accurate lower wind speed predictions accuracy outweigh the benefit of slightly better high (extreme) wind speed predictions. Therefore, it is concluded that 0-25 mph regression equations should continue to be utilized by ETS at SRS. During the Department of Energy Tiger Team audit, concerns were raised about the calibration of SRS cup anemometers. Wind speed is measured by ETS with Teledyne Geotech model 1564B cup anemometers, which are calibrated in the ETS wind tunnel. Linear regression lines are fitted to data points of tunnel speed versus anemometer output voltages up to 25 mph. The regression coefficients are then implemented into the data acquisition computer software when an instrument is installed in the field. The concern raised was that since the wind tunnel at SRS is able to generate a maximum wind speed higher than 25 mph, errors may be introduced in not using the full range of the wind tunnel.

Wind tunnel tests on a one-foot diameter SR-7L propfan model

NASA Technical Reports Server (NTRS)

Aljabri, Abdullah S.

1987-01-01

Wind tunnel tests have been conducted on a one-foot diameter model of the SR-7L propfan in the Langley 16-Foot and 4 x 7 Meter Wind Tunnels as part of the Propfan Test Assessment (PTA) Program. The model propfan was sized to be used on a 1/9-scale model of the PTA testbed aircraft. The model propeller was tested in isolation and wing-mounted on the aircraft configuration at various Mach numbers and blade pitch angles. Agreement between data obtained from these tests and data from Hamilton Standard validate that the 1/9-scale propeller accurately simulates the aerodynamics of the SR-7L propfan. Predictions from an analytical computer program are presented and show good agreement with the experimental data.

Videometric Applications in Wind Tunnels

NASA Technical Reports Server (NTRS)

Burner, A. W.; Radeztsky, R. H.; Liu, Tian-Shu

1997-01-01

Videometric measurements in wind tunnels can be very challenging due to the limited optical access, model dynamics, optical path variability during testing, large range of temperature and pressure, hostile environment, and the requirements for high productivity and large amounts of data on a daily basis. Other complications for wind tunnel testing include the model support mechanism and stringent surface finish requirements for the models in order to maintain aerodynamic fidelity. For these reasons nontraditional photogrammetric techniques and procedures sometimes must be employed. In this paper several such applications are discussed for wind tunnels which include test conditions with Mach number from low speed to hypersonic, pressures from less than an atmosphere to nearly seven atmospheres, and temperatures from cryogenic to above room temperature. Several of the wind tunnel facilities are continuous flow while one is a short duration blowdown facility. Videometric techniques and calibration procedures developed to measure angle of attack, the change in wing twist and bending induced by aerodynamic load, and the effects of varying model injection rates are described. Some advantages and disadvantages of these techniques are given and comparisons are made with non-optical and more traditional video photogrammetric techniques.

Near real time wind energy forecasting incorporating wind tunnel modeling

NASA Astrophysics Data System (ADS)

Lubitz, William David

A series of experiments and investigations were carried out to inform the development of a day-ahead wind power forecasting system. An experimental near-real time wind power forecasting system was designed and constructed that operates on a desktop PC and forecasts 12--48 hours in advance. The system uses model output of the Eta regional scale forecast (RSF) to forecast the power production of a wind farm in the Altamont Pass, California, USA from 12 to 48 hours in advance. It is of modular construction and designed to also allow diagnostic forecasting using archived RSF data, thereby allowing different methods of completing each forecasting step to be tested and compared using the same input data. Wind-tunnel investigations of the effect of wind direction and hill geometry on wind speed-up above a hill were conducted. Field data from an Altamont Pass, California site was used to evaluate several speed-up prediction algorithms, both with and without wind direction adjustment. These algorithms were found to be of limited usefulness for the complex terrain case evaluated. Wind-tunnel and numerical simulation-based methods were developed for determining a wind farm power curve (the relation between meteorological conditions at a point in the wind farm and the power production of the wind farm). Both methods, as well as two methods based on fits to historical data, ultimately showed similar levels of accuracy: mean absolute errors predicting power production of 5 to 7 percent of the wind farm power capacity. The downscaling of RSF forecast data to the wind farm was found to be complicated by the presence of complex terrain. Poor results using the geostrophic drag law and regression methods motivated the development of a database search method that is capable of forecasting not only wind speeds but also power production with accuracy better than persistence.

An evaluation of proposed acoustic treatments for the NASA LaRC 4 x 7 meter wind tunnel

NASA Technical Reports Server (NTRS)

Abrahamson, A. L.

1985-01-01

The NASA LaRC 4 x 7 Meter Wind Tunnel is an existing facility specially designed for powered low speed (V/STOL) testing of large scale fixed wing and rotorcraft models. The enhancement of the facility for scale model acoustic testing is examined. The results are critically reviewed and comparisons are drawn with a similar wind tunnel (the DNW Facility in the Netherlands). Discrepancies observed in the comparison stimulated a theoretical investigation using the acoustic finite element ADAM System, of the ways in which noise propagating around the tunnel circuit radiates into the open test section. The reasons for the discrepancies noted above are clarified and assists in the selection of acoustic treatment options for the facility.

The 6-foot-4-inch Wind Tunnel at the Washington Navy Yard

NASA Technical Reports Server (NTRS)

Desmond, G L; Mccrary, J A

1935-01-01

The 6-foot-4-inch wind tunnel and its auxiliary equipment has proven itself capable of continuous and reliable output of data. The real value of the tunnel will increase as experience is gained in checking the observed tunnel performance against full-scale performance. Such has been the case of the 8- by 8-foot tunnel, and for that reason the comparison in the calibration tests have been presented.

Dynamic response tests of inertial and optical wind-tunnel model attitude measurement devices

NASA Technical Reports Server (NTRS)

Buehrle, R. D.; Young, C. P., Jr.; Burner, A. W.; Tripp, J. S.; Tcheng, P.; Finley, T. D.; Popernack, T. G., Jr.

1995-01-01

Results are presented for an experimental study of the response of inertial and optical wind-tunnel model attitude measurement systems in a wind-off simulated dynamic environment. This study is part of an ongoing activity at the NASA Langley Research Center to develop high accuracy, advanced model attitude measurement systems that can be used in a dynamic wind-tunnel environment. This activity was prompted by the inertial model attitude sensor response observed during high levels of model vibration which results in a model attitude measurement bias error. Significant bias errors in model attitude measurement were found for the measurement using the inertial device during wind-off dynamic testing of a model system. The amount of bias present during wind-tunnel tests will depend on the amplitudes of the model dynamic response and the modal characteristics of the model system. Correction models are presented that predict the vibration-induced bias errors to a high degree of accuracy for the vibration modes characterized in the simulated dynamic environment. The optical system results were uncorrupted by model vibration in the laboratory setup.

Measurement of Model Noise in a Hard-Wall Wind Tunnel

NASA Technical Reports Server (NTRS)

Soderman, Paul T.

2006-01-01

Identification, analysis, and control of fluid-mechanically-generated sound from models of aircraft and automobiles in special low-noise, semi-anechoic wind tunnels are an important research endeavor. Such studies can also be done in aerodynamic wind tunnels that have hard walls if phased microphone arrays are used to focus on the noise-source regions and reject unwanted reflections or background noise. Although it may be difficult to simulate the total flyover or drive-by noise in a closed wind tunnel, individual noise sources can be isolated and analyzed. An acoustic and aerodynamic study was made of a 7-percent-scale aircraft model in a NASA Ames 7-by-10-ft (about 2-by-3-m) wind tunnel for the purpose of identifying and attenuating airframe noise sources. Simulated landing, takeoff, and approach configurations were evaluated at Mach 0.26. Using a phased microphone array mounted in the ceiling over the inverted model, various noise sources in the high-lift system, landing gear, fins, and miscellaneous other components were located and compared for sound level and frequency at one flyover location. Numerous noise-alleviation devices and modifications of the model were evaluated. Simultaneously with acoustic measurements, aerodynamic forces were recorded to document aircraft conditions and any performance changes caused by geometric modifications. Most modern microphone-array systems function in the frequency domain in the sense that spectra of the microphone outputs are computed, then operations are performed on the matrices of microphone-signal cross-spectra. The entire acoustic field at one station in such a system is acquired quickly and interrogated during postprocessing. Beam-forming algorithms are employed to scan a plane near the model surface and locate noise sources while rejecting most background noise and spurious reflections. In the case of the system used in this study, previous studies in the wind tunnel have identified noise sources up to 19 d

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions

NASA Technical Reports Server (NTRS)

Cliff, Susan E.; Denison, Marie; Sozer, Emre; Moini-Yekta, Shayan

2016-01-01

NASA and Industry are performing vehicle studies of configurations with low sonic boom pressure signatures. The computational analyses of modern configuration designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty in the aft signatures with often greater boundary layer effects and nozzle jet pressures. Wind tunnel testing at significantly lower Reynolds numbers than in flight and without inlet and nozzle jet pressures make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel from Mach 1.6 to 2.0 will be used to assess the effects of shocks from components passing through nozzle jet plumes on the sonic boom pressure signature and provide datasets for comparison with CFD codes. A large number of high-fidelity numerical simulations of wind tunnel test models with a variety of shock generators that simulate horizontal tails and aft decks have been studied to provide suitable models for sonic boom pressure measurements using a minimally intrusive pressure rail in the wind tunnel. The computational results are presented and the evolution of candidate wind tunnel models is summarized and discussed in this paper.

Thermal and Pressure Characterization of a Wind Tunnel Force Balance Using the Single Vector System. Experimental Design and Analysis Approach to Model Pressure and Temperature Effects in Hypersonic Wind Tunnel Research

NASA Technical Reports Server (NTRS)

Lynn, Keith C.; Commo, Sean A.; Johnson, Thomas H.; Parker, Peter A,

2011-01-01

Wind tunnel research at NASA Langley Research Center s 31-inch Mach 10 hypersonic facility utilized a 5-component force balance, which provided a pressurized flow-thru capability to the test article. The goal of the research was to determine the interaction effects between the free-stream flow and the exit flow from the reaction control system on the Mars Science Laboratory aeroshell during planetary entry. In the wind tunnel, the balance was exposed to aerodynamic forces and moments, steady-state and transient thermal gradients, and various internal balance cavity pressures. Historically, these effects on force measurement accuracy have not been fully characterized due to limitations in the calibration apparatus. A statistically designed experiment was developed to adequately characterize the behavior of the balance over the expected wind tunnel operating ranges (forces/moments, temperatures, and pressures). The experimental design was based on a Taylor-series expansion in the seven factors for the mathematical models. Model inversion was required to calculate the aerodynamic forces and moments as a function of the strain-gage readings. Details regarding transducer on-board compensation techniques, experimental design development, mathematical modeling, and wind tunnel data reduction are included in this paper.

The Effect of Wall Interference Upon the Aerodynamic Characteristics of an Airfoil Spanning a Closed-Throat Circular Wind Tunnel

DTIC Science & Technology

1946-01-01

tunnel-wall correctliois is divided conveniently into two gen - wind tunnels of various types has been the subject of numer- era] sections. First, the...1 1-2 7. I n77~ 8 - ----------------- -/00 -80 -60 -40 -20 0 20 40 60 80 /00 0/s fance from t0he tunnel center/,ne In percent radius 1 4

Sting Dynamics of Wind Tunnel Models

DTIC Science & Technology

1976-05-01

Patterson AFB, AFFDL, Ohio, October 1964. 17. Brunk, James E. "Users Manual: Extended Capability Magnus Rotor and Ballistic Body 6-DOF Trajectory...measure "second-order" aerodynamic effects resulting, for example, from Reynolds number in- fluence. Consequently, all wind tunnel data systems are...sting-model interference effects , sting configurations normally consist of one or more linearly tapered sections combined with one or more untapered

Optimization of transonic wind tunnel data acquisition and control systems for providing continuous mode tests

NASA Astrophysics Data System (ADS)

Petronevich, V. V.

2016-10-01

The paper observes the issues related to the increase of efficiency and information content of experimental research in transonic wind tunnels (WT). In particular, questions of optimizing the WT Data Acquisition and Control Systems (DACS) to provide the continuous mode test method are discussed. The problem of Mach number (M number) stabilization in the test section of the large transonic compressor-type wind tunnels at subsonic flow conditions with continuous change of the aircraft model angle of attack is observed on the example of T-128 wind tunnel. To minimize the signals distortion in T-128 DACS measurement channels the optimal MGCplus filter settings of the data acquisition system used in T-128 wind tunnel to measure loads were experimentally determined. As a result of the tests performed a good agreement of the results of balance measurements for pitch/pause and continuous test modes was obtained. Carrying out balance tests for pitch/pause and continuous test methods was provided by the regular data acquisition and control system of T-128 wind tunnel with unified software package POTOK. The architecture and functional abilities of POTOK software package are observed.

Detailed Uncertainty Analysis for Ares I Ascent Aerodynamics Wind Tunnel Database

NASA Technical Reports Server (NTRS)

Hemsch, Michael J.; Hanke, Jeremy L.; Walker, Eric L.; Houlden, Heather P.

2008-01-01

A detailed uncertainty analysis for the Ares I ascent aero 6-DOF wind tunnel database is described. While the database itself is determined using only the test results for the latest configuration, the data used for the uncertainty analysis comes from four tests on two different configurations at the Boeing Polysonic Wind Tunnel in St. Louis and the Unitary Plan Wind Tunnel at NASA Langley Research Center. Four major error sources are considered: (1) systematic errors from the balance calibration curve fits and model + balance installation, (2) run-to-run repeatability, (3) boundary-layer transition fixing, and (4) tunnel-to-tunnel reproducibility.

Models of Lift and Drag Coefficients of Stalled and Unstalled Airfoils in Wind Turbines and Wind Tunnels

NASA Technical Reports Server (NTRS)

Spera, David A.

2008-01-01

Equations are developed with which to calculate lift and drag coefficients along the spans of torsionally-stiff rotating airfoils of the type used in wind turbine rotors and wind tunnel fans, at angles of attack in both the unstalled and stalled aerodynamic regimes. Explicit adjustments are made for the effects of aspect ratio (length to chord width) and airfoil thickness ratio. Calculated lift and drag parameters are compared to measured parameters for 55 airfoil data sets including 585 test points. Mean deviation was found to be -0.4 percent and standard deviation was 4.8 percent. When the proposed equations were applied to the calculation of power from a stall-controlled wind turbine tested in a NASA wind tunnel, mean deviation from 54 data points was -1.3 percent and standard deviation was 4.0 percent. Pressure-rise calculations for a large wind tunnel fan deviated by 2.7 percent (mean) and 4.4 percent (standard). The assumption that a single set of lift and drag coefficient equations can represent the stalled aerodynamic behavior of a wide variety of airfoils was found to be satisfactory.

Air Flow Modeling in the Wind Tunnel of the FHWA Aerodynamics Laboratory at Turner-Fairbank Highway Research Center

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sitek, M. A.; Lottes, S. A.; Bojanowski, C.

Computational fluid dynamics (CFD) modeling is widely used in industry for design and in the research community to support, compliment, and extend the scope of experimental studies. Analysis of transportation infrastructure using high performance cluster computing with CFD and structural mechanics software is done at the Transportation Research and Analysis Computing Center (TRACC) at Argonne National Laboratory. These resources, available at TRACC, were used to perform advanced three-dimensional computational simulations of the wind tunnel laboratory at the Turner-Fairbank Highway Research Center (TFHRC). The goals were to verify the CFD model of the laboratory wind tunnel and then to use versionsmore » of the model to provide the capability to (1) perform larger parametric series of tests than can be easily done in the laboratory with available budget and time, (2) to extend testing to wind speeds that cannot be achieved in the laboratory, and (3) to run types of tests that are very difficult or impossible to run in the laboratory. Modern CFD software has many physics models and domain meshing options. Models, including the choice of turbulence and other physics models and settings, the computational mesh, and the solver settings, need to be validated against measurements to verify that the results are sufficiently accurate for use in engineering applications. The wind tunnel model was built and tested, by comparing to experimental measurements, to provide a valuable tool to perform these types of studies in the future as a complement and extension to TFHRC’s experimental capabilities. Wind tunnel testing at TFHRC is conducted in a subsonic open-jet wind tunnel with a 1.83 m (6 foot) by 1.83 m (6 foot) cross section. A three component dual force-balance system is used to measure forces acting on tested models, and a three degree of freedom suspension system is used for dynamic response tests. Pictures of the room are shown in Figure 1-1 to Figure 1-4. A

Wind Tunnel Complex at the Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-09-21

This aerial photograph shows the entire original wind tunnel complex at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The large Altitude Wind Tunnel (AWT) at the center of the photograph dominates the area. The Icing Research Tunnel to the right was incorporated into the lab’s design to take advantage of the AWT’s powerful infrastructure. The laboratory’s first supersonic wind tunnel was added to this complex just prior to this September 1945 photograph. The AWT was the nation’s only wind tunnel capable of studying full-scale engines in simulated flight conditions. The AWT’s test section and control room were within the two-story building near the top of the photograph. The exhauster equipment used to thin the airflow and the drive motor for the fan were in the building to the right of the tunnel. The unique refrigeration equipment was housed in the structure to the left of the tunnel. The Icing Research Tunnel was an atmospheric tunnel that used the AWT’s refrigeration equipment to simulate freezing rain inside its test section. A spray bar system inside the tunnel was originally used to create the droplets. The 18- by 18-inch supersonic wind tunnel was built in the summer of 1945 to take advantage of the AWT’s powerful exhaust system. It was the lab’s first supersonic tunnel and could reach Mach 1.91. Eventually the building would house three small supersonic tunnels, referred to as the “stack tunnels” because of the vertical alignment. The two other tunnels were added to this structure in 1949 and 1951.

The Beginner's Guide to Wind Tunnels with TunnelSim and TunnelSys

NASA Technical Reports Server (NTRS)

Benson, Thomas J.; Galica, Carol A.; Vila, Anthony J.

2010-01-01

The Beginner's Guide to Wind Tunnels is a Web-based, on-line textbook that explains and demonstrates the history, physics, and mathematics involved with wind tunnels and wind tunnel testing. The Web site contains several interactive computer programs to demonstrate scientific principles. TunnelSim is an interactive, educational computer program that demonstrates basic wind tunnel design and operation. TunnelSim is a Java (Sun Microsystems Inc.) applet that solves the continuity and Bernoulli equations to determine the velocity and pressure throughout a tunnel design. TunnelSys is a group of Java applications that mimic wind tunnel testing techniques. Using TunnelSys, a team of students designs, tests, and post-processes the data for a virtual, low speed, and aircraft wing.

WIND TUNNEL INVESTIGATION OF THE RESPONSE OF A SONIC ANEMOMETER

EPA Science Inventory

An Applied Technology Inc. (ATI) sonic of the type used by J. C. Kaimal at the Boulder Tower was tested in the large wind tunnel at the U.S. EPA Fluid Modeling Facility. The wind tunnel is approximately 6 ft high, 10 ft wide with a test section bed 60 ft long. The air speed in th...

Automated Boundary Conditions for Wind Tunnel Simulations

NASA Technical Reports Server (NTRS)

Carlson, Jan-Renee

2018-01-01

Computational fluid dynamic (CFD) simulations of models tested in wind tunnels require a high level of fidelity and accuracy particularly for the purposes of CFD validation efforts. Considerable effort is required to ensure the proper characterization of both the physical geometry of the wind tunnel and recreating the correct flow conditions inside the wind tunnel. The typical trial-and-error effort used for determining the boundary condition values for a particular tunnel configuration are time and computer resource intensive. This paper describes a method for calculating and updating the back pressure boundary condition in wind tunnel simulations by using a proportional-integral-derivative controller. The controller methodology and equations are discussed, and simulations using the controller to set a tunnel Mach number in the NASA Langley 14- by 22-Foot Subsonic Tunnel are demonstrated.

Computation of wind tunnel wall effects for complex models using a low-order panel method

NASA Technical Reports Server (NTRS)

Ashby, Dale L.; Harris, Scott H.

1994-01-01

A technique for determining wind tunnel wall effects for complex models using the low-order, three dimensional panel method PMARC (Panel Method Ames Research Center) has been developed. Initial validation of the technique was performed using lift-coefficient data in the linear lift range from tests of a large-scale STOVL fighter model in the National Full-Scale Aerodynamics Complex (NFAC) facility. The data from these tests served as an ideal database for validating the technique because the same model was tested in two wind tunnel test sections with widely different dimensions. The lift-coefficient data obtained for the same model configuration in the two test sections were different, indicating a significant influence of the presence of the tunnel walls and mounting hardware on the lift coefficient in at least one of the two test sections. The wind tunnel wall effects were computed using PMARC and then subtracted from the measured data to yield corrected lift-coefficient versus angle-of-attack curves. The corrected lift-coefficient curves from the two wind tunnel test sections matched very well. Detailed pressure distributions computed by PMARC on the wing lower surface helped identify the source of large strut interference effects in one of the wind tunnel test sections. Extension of the technique to analysis of wind tunnel wall effects on the lift coefficient in the nonlinear lift range and on drag coefficient will require the addition of boundary-layer and separated-flow models to PMARC.

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raynold S.

2010-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1 1 SWT for Schlieren photography and comparison to CFD analysis.

Verification and Calibration of a Reduced Order Wind Farm Model by Wind Tunnel Experiments

NASA Astrophysics Data System (ADS)

Schreiber, J.; Nanos, E. M.; Campagnolo, F.; Bottasso, C. L.

2017-05-01

In this paper an adaptation of the FLORIS approach is considered that models the wind flow and power production within a wind farm. In preparation to the use of this model for wind farm control, this paper considers the problem of its calibration and validation with the use of experimental observations. The model parameters are first identified based on measurements performed on an isolated scaled wind turbine operated in a boundary layer wind tunnel in various wind-misalignment conditions. Next, the wind farm model is verified with results of experimental tests conducted on three interacting scaled wind turbines. Although some differences in the estimated absolute power are observed, the model appears to be capable of identifying with good accuracy the wind turbine misalignment angles that, by deflecting the wake, lead to maximum power for the investigated layouts.

Aeroservoelastic Testing of a Sidewall Mounted Free Flying Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer

2008-01-01

A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three j wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the rst of these three tests, a semispan, aeroelastically scaled, wind-tunnel model of a ying wing SensorCraft vehi- cle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree-of-freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid-body modes. Gust Load Alleviation (GLA) and Body Freedom Flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Wind tunnel

NASA Technical Reports Server (NTRS)

Wilson, E. M. (Inventor)

1969-01-01

A supersonic wind wind tunnel is described for testing several air foils mounted in a row. A test section of a wind tunnel contains means for mounting air foil sections in a row, means for rotating each section about an axis so that the angle of attack of each section changes with the other sections, and means for rotating the row with respect to the air stream so that the row forms an oblique angle with the air stream.

Probabilistic Design of a Wind Tunnel Model to Match the Response of a Full-Scale Aircraft

NASA Technical Reports Server (NTRS)

Mason, Brian H.; Stroud, W. Jefferson; Krishnamurthy, T.; Spain, Charles V.; Naser, Ahmad S.

2005-01-01

approach is presented for carrying out the reliability-based design of a plate-like wing that is part of a wind tunnel model. The goal is to design the wind tunnel model to match the stiffness characteristics of the wing box of a flight vehicle while satisfying strength-based risk/reliability requirements that prevents damage to the wind tunnel model and fixtures. The flight vehicle is a modified F/A-18 aircraft. The design problem is solved using reliability-based optimization techniques. The objective function to be minimized is the difference between the displacements of the wind tunnel model and the corresponding displacements of the flight vehicle. The design variables control the thickness distribution of the wind tunnel model. Displacements of the wind tunnel model change with the thickness distribution, while displacements of the flight vehicle are a set of fixed data. The only constraint imposed is that the probability of failure is less than a specified value. Failure is assumed to occur if the stress caused by aerodynamic pressure loading is greater than the specified strength allowable. Two uncertain quantities are considered: the allowable stress and the thickness distribution of the wind tunnel model. Reliability is calculated using Monte Carlo simulation with response surfaces that provide approximate values of stresses. The response surface equations are, in turn, computed from finite element analyses of the wind tunnel model at specified design points. Because the response surface approximations were fit over a small region centered about the current design, the response surfaces were refit periodically as the design variables changed. Coarse-grained parallelism was used to simultaneously perform multiple finite element analyses. Studies carried out in this paper demonstrate that this scheme of using moving response surfaces and coarse-grained computational parallelism reduce the execution time of the Monte Carlo simulation enough to make the

Aeroacoustic research in wind tunnels: A status report

NASA Technical Reports Server (NTRS)

Bender, J.; Arndt, R. E. A.

1973-01-01

The increasing attention given to aerodynamically generated noise brings into focus the need for quality experimental research in this area. To meet this need several specialized anechoic wind tunnels have been constructed. In many cases, however, budgetary constraints and the like make it desirable to use conventional wind tunnels for this work. Three basic problems are inherent in conventional facilities: (1) high background noise, (2) strong frequency dependent reverberation effects, and (3) unique instrumentation problems. The known acoustic characteristics of several conventional wind tunnels are evaluated and data obtained in a smaller 4- x 5-foot wind tunnel which is convertible from a closed jet to an open jet mode are presented. The data from these tunnels serve as a guideline for proposed modifications to a 7- x 10-foot wind tunnel. Consideration is given to acoustic treatment in several different portions of the wind tunnel.

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated Exhaust Nozzles

NASA Technical Reports Server (NTRS)

Castner, Raymond S.

2009-01-01

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nose cone and an upper plenum. Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls in the 1x1 SWT for Schlieren photography and comparison to CFD analysis.

Rudolf Hermann, wind tunnels and aerodynamics

NASA Astrophysics Data System (ADS)

Lundquist, Charles A.; Coleman, Anne M.

2008-04-01

Rudolf Hermann was born on December 15, 1904 in Leipzig, Germany. He studied at the University of Leipzig and at the Aachen Institute of Technology. His involvement with wind tunnels began in 1934 when Professor Carl Wieselsberger engaged him to work at Aachen on the development of a supersonic wind tunnel. On January 6, 1936, Dr. Wernher von Braun visited Dr. Hermann to arrange for use of the Aachen supersonic wind tunnel for Army problems. On April 1, 1937, Dr. Hermann became Director of the Supersonic Wind Tunnel at the Army installation at Peenemunde. Results from the Aachen and Peenemunde wind tunnels were crucial in achieving aerodynamic stability for the A-4 rocket, later designated as the V-2. Plans to build a Mach 10 'hypersonic' wind tunnel facility at Kochel were accelerated after the Allied air raid on Peenemunde on August 17, 1943. Dr. Hermann was director of the new facility. Ignoring destruction orders from Hitler as WWII approached an end in Europe, Dr. Hermann and his associates hid documents and preserved wind tunnel components that were acquired by the advancing American forces. Dr. Hermann became a consultant to the Air Force at its Wright Field in November 1945. In 1951, he was named professor of Aeronautical Engineering at the University of Minnesota. In 1962, Dr. Hermann became the first Director of the Research Institute at the University of Alabama in Huntsville (UAH), a position he held until he retired in 1970.

Free-to-Roll Testing of Airplane Models in Wind Tunnels

NASA Technical Reports Server (NTRS)

Capone, Francis J.; Owens, D. Bruce; Hall, Robert M.

2007-01-01

A free-to-roll (FTR) test technique and test rig make it possible to evaluate both the transonic performance and the wingdrop/ rock behavior of a high-strength airplane model in a single wind-tunnel entry. The free-to-roll test technique is a single degree-of-motion method in which the model is free to roll about the longitudinal axis. The rolling motion is observed, recorded, and analyzed to gain insight into wing-drop/rock behavior. Wing-drop/rock is one of several phenomena symptomatic of abrupt wing stall. FTR testing was developed as part of the NASA/Navy Abrupt Wing Stall Program, which was established for the purposes of understanding and preventing significant unexpected and uncommanded (thus, highly undesirable) lateral-directional motions associated with wing-drop/rock, which have been observed mostly in fighter airplanes under high-subsonic and transonic maneuvering conditions. Before FTR testing became available, wingrock/ drop behavior of high-performance airplanes undergoing development was not recognized until flight testing. FTR testing is a reliable means of detecting, and evaluating design modifications for reducing or preventing, very complex abrupt wing stall phenomena in a ground facility prior to flight testing. The FTR test rig was designed to replace an older sting attachment butt, such that a model with its force balance and support sting could freely rotate about the longitudinal axis. The rig (see figure) includes a rotary head supported in a stationary head with a forward spherical roller bearing and an aft needle bearing. Rotation is amplified by a set of gears and measured by a shaft-angle resolver; the roll angle can be resolved to within 0.067 degrees at a rotational speed up to 1,000 degrees/s. An assembly of electrically actuated brakes between the rotary and stationary heads can be used to hold the model against a rolling torque at a commanded roll angle. When static testing is required, a locking bar is used to fix the rotating

Avrocar Test in Ames 40x80 Foot Wind Tunnel.

NASA Image and Video Library

1961-04-03

Rear view of the Avrocar with tail, mounted on variable height struts. Overhead doors of the wind tunnel test section open. The first Avrocar, S/N 58-7055 (marked AV-7055), after tethered testing, became the "wind tunnel" test model at NASA Ames, where it remained in storage from 1961 until 1966, when it was donated to the National Air and Space Museum, in Suitland, Maryland.

Pressure distribution on the roof of a model low-rise building tested in a boundary layer wind tunnel

NASA Astrophysics Data System (ADS)

Goliber, Matthew Robert

validate the tube length, tube diameter, port diameter, and pressure transducer used in the field. Also, Model 2 was used to study the effects of surface roughness on pressure readings. A partial roof and wall of the low-rise building in the field was used as the third model. Similar to the second model, Model 3 was tested in the AABL Wind Tunnel. Initially, the objectives of the third model were to validate the pressure port protection device (PPPD) being used in the field and test the possibility of interpolating between pressure ports. But in the end, Model 3 was best used to validate the inconsistencies of the full-scale PPPD, validate the transducers used in the field, and prove the importance of scaling either all or none of the model. Fourthly, Model 4 was a 1:16 model of the low-rise building itself. Based on the three previous model tests, Model 4 was instrumented with 202 pressure transducers to better understand: (1) the pressure distribution on the roof of the structure, (2) the affects of the fundamental test variables such as tube length, tube diameter, port diameter, transducer type, and surface roughness, (3) the affects of a scaled PPPD, (4) the importance of wind angle of attack, and (5) the possibility of measuring pressure data and load data simultaneously. In the end, the combination of all four model tests proved to be helpful in understanding the pressure data gathered on the roof of the low-rise building in the field. The two main recommendations for the field structure are for reevaluation of the PPPD design and slight redistribution of the pressure ports. The wind tunnel model tests show a need for these two modifications in order to gather more accurate field pressure data. Other than these two adjustments, the model tests show that the remaining data gathering system is currently accurate.

Review of Aeronautical Wind Tunnel Facilities

NASA Technical Reports Server (NTRS)

1988-01-01

The nation's aeronautical wind tunnel facilities constitute a valuable technological resource and make a significant contribution to the global supremacy of U.S. aircraft, both civil and military. At the request of NASA, the National Research Council's Aeronautics and Space Engineering Board organized a commitee to review the state of repair, adequacy, and future needs of major aeronautical wind tunnel facilities in meeting national goals. The comittee identified three main areas where actions are needed to sustain the capability of NASA's aeronautical wind tunnel facilities to support the national aeronautical research and development activities: tunnel maintenance and upgrading, productivity enhancement, and accommodation of new requirements (particularly in hypersonics). Each of these areas are addressed and the committee recommendations for appropriate actions presented.

Altitude Wind Tunnel Control Room

NASA Image and Video Library

1945-05-21

Researchers at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory monitor a ramjet's performance in the Altitude Wind Tunnel from the control room. The soundproof control room was just a few feet from the tunnel’s 20-foot-diameter test section. In the control room, the operators could control all aspects of the tunnel’s operation, including the air density, temperature, and speed. They also operated the engine or test article in the test section by controlling the angle-of-attack, speed, power, and other parameters. The men in this photograph are monitoring the engine’s thrust and lift. A NACA-designed 20-inch-diameter ramjet was installed in the tunnel in May 1945. Thrust figures from these runs were compared with drag data from tests of scale models in small supersonic tunnels to verify the ramjet’s feasibility. The tunnel was used to analyze the ramjet’s overall performance up to altitudes of 47,000 feet and speeds to Mach 1.84. The researchers found that an increase in altitude caused a reduction in the engine’s horsepower and identified optimal flameholder configurations.

Static and wind tunnel near-field/far field jet noise measurements from model scale single-flow baseline and suppressor nozzles. Volume 2: Forward speed effects

NASA Technical Reports Server (NTRS)

Jaeck, C. L.

1976-01-01

A model scale flight effects test was conducted in the 40 by 80 foot wind tunnel to investigate the effect of aircraft forward speed on single flow jet noise characteristics. The models tested included a 15.24 cm baseline round convergent nozzle, a 20-lobe and annular nozzle with and without lined ejector shroud, and a 57-tube nozzle with a lined ejector shroud. Nozzle operating conditions covered jet velocities from 412 to 640 m/s at a total temperature of 844 K. Wind tunnel speeds were varied from near zero to 91.5 m/s. Measurements were analyzed to (1) determine apparent jet noise source location including effects of ambient velocity; (2) verify a technique for extrapolating near field jet noise measurements into the far field; (3) determine flight effects in the near and far field for baseline and suppressor nozzles; and (4) establish the wind tunnel as a means of accurately defining flight effects for model nozzles and full scale engines.

Assessing Videogrammetry for Static Aeroelastic Testing of a Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Spain, Charles V.; Heeg, Jennifer; Ivanco, Thomas G.; Barrows, Danny A.; Florance, James R.; Burner, Alpheus W.; DeMoss, Joshua; Lively, Peter S.

2004-01-01

The Videogrammetric Model Deformation (VMD) technique, developed at NASA Langley Research Center, was recently used to measure displacements and local surface angle changes on a static aeroelastic wind-tunnel model. The results were assessed for consistency, accuracy and usefulness. Vertical displacement measurements and surface angular deflections (derived from vertical displacements) taken at no-wind/no-load conditions were analyzed. For accuracy assessment, angular measurements were compared to those from a highly accurate accelerometer. Shewhart's Variables Control Charts were used in the assessment of consistency and uncertainty. Some bad data points were discovered, and it is shown that the measurement results at certain targets were more consistent than at other targets. Physical explanations for this lack of consistency have not been determined. However, overall the measurements were sufficiently accurate to be very useful in monitoring wind-tunnel model aeroelastic deformation and determining flexible stability and control derivatives. After a structural model component failed during a highly loaded condition, analysis of VMD data clearly indicated progressive structural deterioration as the wind-tunnel condition where failure occurred was approached. As a result, subsequent testing successfully incorporated near- real-time monitoring of VMD data in order to ensure structural integrity. The potential for higher levels of consistency and accuracy through the use of statistical quality control practices are discussed and recommended for future applications.

Comparison of Store Trajectory and Aerodynamic Loads, and Model Flow- Field Characteristics Obtained in the AEDC PWT/4T and VFK/A Wind Tunnels at Mach Number 1.63

DTIC Science & Technology

1979-02-01

8217 -13.2 i 2 Lt 5 8 10 D Zp . I i 2 u, 6 8 Zp b. Xp = O, Yp = -1 in., 0 = O, ~ = 0 Figure 17. Continued. 10 44 5YM TUNNEL REXI ~ e 0 4" 5.0 0 A...TUNNEL REXI ~ s 0 ~T 3.~ 0 R 5.0 CN Cm 6 ,u, 2 0 -2 -u, 6 I i𔃻 i I I I, t~ l I P -2 i i - 6 i Cy 2 i I ~ , r - - - I -2

Wind Tunnel to Atmospheric Mapping for Static Aeroelastic Scaling

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Spain, Charles V.; Rivera, J. A.

2004-01-01

Wind tunnel to Atmospheric Mapping (WAM) is a methodology for scaling and testing a static aeroelastic wind tunnel model. The WAM procedure employs scaling laws to define a wind tunnel model and wind tunnel test points such that the static aeroelastic flight test data and wind tunnel data will be correlated throughout the test envelopes. This methodology extends the notion that a single test condition - combination of Mach number and dynamic pressure - can be matched by wind tunnel data. The primary requirements for affecting this extension are matching flight Mach numbers, maintaining a constant dynamic pressure scale factor and setting the dynamic pressure scale factor in accordance with the stiffness scale factor. The scaling is enabled by capabilities of the NASA Langley Transonic Dynamics Tunnel (TDT) and by relaxation of scaling requirements present in the dynamic problem that are not critical to the static aeroelastic problem. The methodology is exercised in two example scaling problems: an arbitrarily scaled wing and a practical application to the scaling of the Active Aeroelastic Wing flight vehicle for testing in the TDT.

A New Global Regression Analysis Method for the Prediction of Wind Tunnel Model Weight Corrections

NASA Technical Reports Server (NTRS)

Ulbrich, Norbert Manfred; Bridge, Thomas M.; Amaya, Max A.

2014-01-01

A new global regression analysis method is discussed that predicts wind tunnel model weight corrections for strain-gage balance loads during a wind tunnel test. The method determines corrections by combining "wind-on" model attitude measurements with least squares estimates of the model weight and center of gravity coordinates that are obtained from "wind-off" data points. The method treats the least squares fit of the model weight separate from the fit of the center of gravity coordinates. Therefore, it performs two fits of "wind- off" data points and uses the least squares estimator of the model weight as an input for the fit of the center of gravity coordinates. Explicit equations for the least squares estimators of the weight and center of gravity coordinates are derived that simplify the implementation of the method in the data system software of a wind tunnel. In addition, recommendations for sets of "wind-off" data points are made that take typical model support system constraints into account. Explicit equations of the confidence intervals on the model weight and center of gravity coordinates and two different error analyses of the model weight prediction are also discussed in the appendices of the paper.

Advanced optical position sensors for magnetically suspended wind tunnel models

NASA Technical Reports Server (NTRS)

Lafleur, S.

1985-01-01

A major concern to aerodynamicists has been the corruption of wind tunnel test data by model support structures, such as stings or struts. A technique for magnetically suspending wind tunnel models was considered by Tournier and Laurenceau (1957) in order to overcome this problem. This technique is now implemented with the aid of a Large Magnetic Suspension and Balance System (LMSBS) and advanced position sensors for measuring model attitude and position within the test section. Two different optical position sensors are discussed, taking into account a device based on the use of linear CCD arrays, and a device utilizing area CID cameras. Current techniques in image processing have been employed to develop target tracking algorithms capable of subpixel resolution for the sensors. The algorithms are discussed in detail, and some preliminary test results are reported.

Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet Flows with Shock Interactions

NASA Technical Reports Server (NTRS)

Cliff, Susan E.; Denison, Marie; Moini-Yekta, Shayan; Morr, Donald E.; Durston, Donald A.

2016-01-01

NASA and the U.S. aerospace industry are performing studies of supersonic aircraft concepts with low sonic boom pressure signatures. The computational analyses of modern aircraft designs have matured to the point where there is confidence in the prediction of the pressure signature from the front of the vehicle, but uncertainty remains in the aft signatures due to boundary layer and nozzle exhaust jet effects. Wind tunnel testing without inlet and nozzle exhaust jet effects at lower Reynolds numbers than in-flight make it difficult to accurately assess the computational solutions of flight vehicles. A wind tunnel test in the NASA Ames 9- by 7-Foot Supersonic Wind Tunnel is planned for February 2016 to address the nozzle jet effects on sonic boom. The experiment will provide pressure signatures of test articles that replicate waveforms from aircraft wings, tails, and aft fuselage (deck) components after passing through cold nozzle jet plumes. The data will provide a variety of nozzle plume and shock interactions for comparison with computational results. A large number of high-fidelity numerical simulations of a variety of shock generators were evaluated to define a reduced collection of suitable test models. The computational results of the candidate wind tunnel test models as they evolved are summarized, and pre-test computations of the final designs are provided.

Ramjet Model and Technicians in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1952-02-21

A researcher at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory checks the setup of a RJM-2 ramjet model in the test section of the 8- by 6-Foot Supersonic Wind Tunnel. The 8- by 6 was not only the laboratory’s first large supersonic wind tunnel, but it was also the NACA’s first facility capable of testing an operating engine at supersonic speeds. The 8- by 6-foot tunnel has been used to study engine inlets, fuel injectors, flameholders, exit nozzles, and controls on ramjet and turbojet propulsion systems. The 8-foot wide and 6-foot tall test section consisted of 1-inch thick steel plates with hatches on the floor and ceiling to facilitate the installation of the test article. The two windows seen on the right wall allowed photographic equipment to be set up. The test section was modified in 1956 to accommodate transonic research. NACA engineers drilled 4,700 holes into the test section walls to reduce transonic pressure disturbances and shock waves. NACA Lewis undertook an extensive research program on ramjets in the 1940s using several of its facilities. Ramjets provide a very simple source of propulsion. They are basically a tube which ingests high speed air, ignites it, and then expels the heated air at a significantly higher velocity. Ramjets are extremely efficient and powerful but can only operate at high speeds. Therefore, they require a booster rocket or aircraft drop to accelerate them to high speeds before they can operate.

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer

2014-01-01

This is part 2 of a two part document. Part 1 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression." A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind tunnel model of a joined wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind tunnel model was mated to a new, two degree of freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at10 percent static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free flying wind tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

The F2 wind tunnel at Fauga-Mauzac

NASA Technical Reports Server (NTRS)

Afchain, D.; Broussaud, P.; Frugier, M.; Rancarani, G.

1984-01-01

Details on the French subsonic wind-tunnel F2 that becomes operational on July 1983 are presented. Some of the requirements were: (1) installation of models on any wall of the facility, (2) good observation points due to transparent walls, (3) smooth flow, (4) a laser velocimeter, and (5) easy access and handling. The characteristics include a nonpressurized return circuit, dimensions of 5 x 1.4 x 1.8 m, maximum velocity of 100 m/s and a variable speed fan of 683 kW.

Modeling cumulus clouds in a two-phase wind tunnel

NASA Astrophysics Data System (ADS)

Bordás, R.; Thévenin, D.

2009-04-01

Experiments in wind-tunnels concerning meteorological flows are not very frequent in the literature. However, they are indispensable for a well-controlled and accurate investigation of turbulence-droplet interactions at the micro-scale. Of course it is impossible to reproduce perfectly the turbulent properties of clouds in a comparatively small wind-tunnel. The enormous length scales that are predominant in nature (integral length scale of typically 100 meters) lead to very high Reynolds numbers, roughly 107 calculated with the cloud dimensions or 104 as Taylor Reynolds number Reλ. Nevertheless, it is not necessary to reproduce exactly the whole turbulence spectrum to investigate the issue of rain formation in cumulus clouds. Only those scales and turbulence properties should be reproduced in the wind tunnel, which are physically important for the droplet population. In this work the key properties of cumulus clouds will be identified and implemented in a two-phase wind tunnel, allowing reproducible and accurate measurements. These properties are in particular the droplet number density, the turbulent kinetic energy and its dissipation rate. It is demonstrated by means of non-intrusive optical measurement techniques that the flow velocity, droplet number density, and key turbulence properties have been matched and are in the right order of magnitude. In this manner wind-tunnel investigations become possible and deliver realistic information concerning the interaction between droplets and turbulence in cumulus clouds.

Aeroelastic modeling of the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Heeg, Jennifer; Bennett, Robert M.

1991-01-01

The primary issues involved in the generation of linear, state-space equations of motion of a flexible wind tunnel model, the Active Flexible Wing (AFW), are discussed. The codes that were used and their inherent assumptions and limitations are also briefly discussed. The application of the CAP-TSD code to the AFW for determination of the model's transonic flutter boundary is included as well.

Control Surface Interaction Effects of the Active Aeroelastic Wing Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Heeg, Jennifer

2006-01-01

This paper presents results from testing the Active Aeroelastic Wing wind tunnel model in NASA Langley s Transonic Dynamics Tunnel. The wind tunnel test provided an opportunity to study aeroelastic system behavior under combined control surface deflections, testing for control surface interaction effects. Control surface interactions were observed in both static control surface actuation testing and dynamic control surface oscillation testing. The primary method of evaluating interactions was examination of the goodness of the linear superposition assumptions. Responses produced by independently actuating single control surfaces were combined and compared with those produced by simultaneously actuating and oscillating multiple control surfaces. Adjustments to the data were required to isolate the control surface influences. Using dynamic data, the task increases, as both the amplitude and phase have to be considered in the data corrections. The goodness of static linear superposition was examined and analysis of variance was used to evaluate significant factors influencing that goodness. The dynamic data showed interaction effects in both the aerodynamic measurements and the structural measurements.

An electronic scanner of pressure for wind tunnel models

NASA Technical Reports Server (NTRS)

Kauffman, Ronald C.; Coe, Charles F.

1986-01-01

An electronic scanner of pressure (ESOP) has been developed by NASA Ames Research Center for installation in wind tunnel models. An ESOP system consists of up to 20 pressure modules (PMs), each with 48 pressure transducers and a heater, an analog-to-digital (A/D) converter module, a microprocessor, a data controller, a monitor unit, a control and processing unit, and a heater controller. The PMs and the A/D converter module are sized to be installed in the models tested in the Ames Aerodynamics Division wind tunnels. A unique feature of the pressure module is the lack of moving parts such as a pneumatic switch used in other systems for in situ calibrations. This paper describes the ESOP system and the results of the initial testing of the system. The initial results indicate the system meets the original design goal of 0.15 percent accuracy.

An evaluation in a modern wind tunnel of the transonic adaptive wall adjustment strategy developed by NPL in the 1940's

NASA Technical Reports Server (NTRS)

Lewis, M. C.

1988-01-01

The first documented wind tunnel employing a flexible walled test section for the purpose of eliminating wall interference was constructed in England by the National Physical Laboratory (NPL) during the late 1930's. The tunnel was transonic and designed for two-dimensional testing. In an attempt to eliminate the top and bottom wall interference effects on the model NPL developed a strategy to adjust two flexible walls to streamlined shapes. This report covers an evaluation of the NPL wall adjustment strategy in a modern wind tunnel, e.g., the Transonic Self-Streamlining Wind Tunnel (TSWT) at the University of Southampton, England. The evaluation took the form of performance comparisons with other modern strategies which have been developed for use in, and proven in, the TSWT.

Design and Characterization of the UTIAS Anechoic Wind Tunnel

NASA Astrophysics Data System (ADS)

Chow, Derrick H. F.

The anechoic open-jet wind tunnel facility at the University of Toronto Institute for Aerospace Studies was updated and characterized to meet the needs of current and future aeroacoustic experiments. The wind tunnel inlet was resurfaced and flow-conditioning screens were redesigned to improve the freestream turbulence intensity to below 0.4% in the test section. The circular nozzle was replaced with a square secondary contraction that increased the maximum test section velocity to 75 m/s and improved flow uniformity to over 99% across a usable cross-sectional area of 500 mm x 500 mm. Acoustic baffles were installed in front of the wind tunnel inlet and foam wedges were installed in the anechoic chamber. The overall background sound pressure levels in the chamber were improved by 8-18 db over the range of operational freestream velocities. The anechoic chamber cut-off frequency is 170 Hz and the reverberation time for a 60 dB sound power decay is 0.032 s.

Altitude Wind Tunnel at NASA Glenn Research Center: An Interactive History

NASA Technical Reports Server (NTRS)

2008-01-01

When constructed in the Early 1940s, the Altitude Wind Tunnel (AWT) at NASA Glenn Research Center was the nation's only wind tunnel capable of studying full scale engines under realistic flight conditions. It played a significant role in the development of the first U.S. jet engines as well as technologies such as the afterburner and variable-area nozzle. In the late 1950s, the tunnels interior components were removed so that hardware for Project Mercury could be tested in altitude conditions. In 1961, a portion of the tunnel was converted into one of the country's first large vacuum tanks and renamed the Space Power Chamber (SPC). SPC was used extensively throughout the 1960s for the Centaur rocket program. This multimedia piece allows one to interactively learn about the Altitude Wind Tunnel facility. and the research performed there. The piece contains: (1) A chronological history of the AWT from its construction during World War II and the testing of early jet engines, through the Mercury and Centaur programs of the 1960s and up to the final use of the building for the Microwave Systems laboratory. (2) Photographic surveys of the facility in it wind tunnel, vacuum tank and final configurations. (3) Browsable gallery of over 200 captioned photographs and video clips.(4) A nine minute documentary of the AWT produced by NASA in 1961 (5) Links to over 70 reports and publications related to AWT research and the history of the NACA.

Rocket Plume Scaling for Orion Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Brauckmann, Gregory J.; Greathouse, James S.; White, Molly E.

2011-01-01

A wind tunnel test program was undertaken to assess the jet interaction effects caused by the various solid rocket motors used on the Orion Launch Abort Vehicle (LAV). These interactions of the external flowfield and the various rocket plumes can cause localized aerodynamic disturbances yielding significant and highly non-linear control amplifications and attenuations. This paper discusses the scaling methodologies used to model the flight plumes in the wind tunnel using cold air as the simulant gas. Comparisons of predicted flight, predicted wind tunnel, and measured wind tunnel forces-and-moments and plume flowfields are made to assess the effectiveness of the selected scaling methodologies.

Wind tunnel model surface gauge for measuring roughness

NASA Technical Reports Server (NTRS)

Vorburger, T. V.; Gilsinn, D. E.; Teague, E. C.; Giauque, C. H. W.; Scire, F. E.; Cao, L. X.

1987-01-01

The optical inspection of surface roughness research has proceeded along two different lines. First, research into a quantitative understanding of light scattering from metal surfaces and into the appropriate models to describe the surfaces themselves. Second, the development of a practical instrument for the measurement of rms roughness of high performance wind tunnel models with smooth finishes. The research is summarized, with emphasis on the second avenue of research.

Experiences with a high-blockage model tested in the NASA Ames 12-foot pressure wind tunnel

NASA Technical Reports Server (NTRS)

Coder, D. W.

1984-01-01

Representation of the flow around full-scale ships was sought in the subsonic wind tunnels in order to a Hain Reynolds numbers as high as possible. As part of the quest to attain the largest possible Reynolds number, large models with high blockage are used which result in significant wall interference effects. Some experiences with such a high blockage model tested in the NASA Ames 12-foot pressure wind tunnel are summarized. The main results of the experiment relating to wind tunnel wall interference effects are also presented.

Computed and Experimental Flutter/LCO Onset for the Boeing Truss-Braced Wing Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Scott, Robert C.; Funk, Christie J.; Allen, Timothy J.; Sexton, Bradley W.

2014-01-01

This paper presents high fidelity Navier-Stokes simulations of the Boeing Subsonic Ultra Green Aircraft Research truss-braced wing wind-tunnel model and compares the results to linear MSC. Nastran flutter analysis and preliminary data from a recent wind-tunnel test of that model at the NASA Langley Research Center Transonic Dynamics Tunnel. The simulated conditions under consideration are zero angle of attack, so that structural nonlinearity can be neglected. It is found that, for Mach number greater than 0.78, the linear flutter analysis predicts flutter onset dynamic pressure below the wind-tunnel test and that predicted by the Navier-Stokes analysis. Furthermore, the wind-tunnel test revealed that the majority of the high structural dynamics cases were wing limit cycle oscillation (LCO) rather than flutter. Most Navier-Stokes simulated cases were also LCO rather than hard flutter. There is dip in the wind-tunnel test flutter/LCO onset in the Mach 0.76-0.80 range. Conditions tested above that Mach number exhibited no aeroelastic instability at the dynamic pressures reached in the tunnel. The linear flutter analyses do not show a flutter/LCO dip. The Navier-Stokes simulations also do not reveal a dip; however, the flutter/LCO onset is at a significantly higher dynamic pressure at Mach 0.90 than at lower Mach numbers. The Navier-Stokes simulations indicate a mild LCO onset at Mach 0.82, then a more rapidly growing instability at Mach 0.86 and 0.90. Finally, the modeling issues and their solution related to the use of a beam and pod finite element model to generate the Navier-Stokes structure mode shapes are discussed.

Slotted-wall research with disk and parachute models in a low-speed wind tunnel

DOE Office of Scientific and Technical Information (OSTI.GOV)

Macha, J.M.; Buffington, R.J.; Henfling, J.L.

1990-01-01

An experimental investigation of slotted-wall blockage interference has been conducted using disk and parachute models in a low speed wind tunnel. Test section open area ratio, model geometric blockage ratio, and model location along the length of the test section were systematically varied. Resulting drag coefficients were compared to each other and to interference-free measurements obtained in a much larger wind tunnel where the geometric blockage ratio was less than 0.0025. 9 refs., 10 figs.

Data Fusion in Wind Tunnel Testing; Combined Pressure Paint and Model Deformation Measurements (Invited)

NASA Technical Reports Server (NTRS)

Bell, James H.; Burner, Alpheus W.

2004-01-01

As the benefit-to-cost ratio of advanced optical techniques for wind tunnel measurements such as Video Model Deformation (VMD), Pressure-Sensitive Paint (PSP), and others increases, these techniques are being used more and more often in large-scale production type facilities. Further benefits might be achieved if multiple optical techniques could be deployed in a wind tunnel test simultaneously. The present study discusses the problems and benefits of combining VMD and PSP systems. The desirable attributes of useful optical techniques for wind tunnels, including the ability to accommodate the myriad optical techniques available today, are discussed. The VMD and PSP techniques are briefly reviewed. Commonalties and differences between the two techniques are discussed. Recent wind tunnel experiences and problems when combining PSP and VMD are presented, as are suggestions for future developments in combined PSP and deformation measurements.

Screens Would Protect Wind-Tunnel Fan Blades

NASA Technical Reports Server (NTRS)

Farmer, Moses G.

1992-01-01

Butterfly screen installed in wind tunnel between test section and fan blades to prevent debris from reaching fan blades if model structure fails. Protective screens deployed manually or automatically. Concept beneficial anywhere wind tunnels employed. Also useful in areas outside of aerospace industry, such as in airflow design of automobiles and other vehicles.

Wind Tunnel Testing of Powered Lift, All-Wing STOL Model

NASA Technical Reports Server (NTRS)

Collins, Scott W.; Westra, Bryan W.; Lin, John C.; Jones, Gregory S.; Zeune, Cal H.

2008-01-01

Short take-off and landing (STOL) systems can offer significant capabilities to warfighters and, for civil operators thriving on maximizing efficiencies they can improve airspace use while containing noise within airport environments. In order to provide data for next generation systems, a wind tunnel test of an all-wing cruise efficient, short take-off and landing (CE STOL) configuration was conducted in the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) 14- by 22-foot Subsonic Wind Tunnel. The test s purpose was to mature the aerodynamic aspects of an integrated powered lift system within an advanced mobility configuration capable of CE STOL. The full-span model made use of steady flap blowing and a lifting centerbody to achieve high lift coefficients. The test occurred during April through June of 2007 and included objectives for advancing the state-of-the-art of powered lift testing through gathering force and moment data, on-body pressure data, and off-body flow field measurements during automatically controlled blowing conditions. Data were obtained for variations in model configuration, angles of attack and sideslip, blowing coefficient, and height above ground. The database produced by this effort is being used to advance design techniques and computational tools for developing systems with integrated powered lift technologies.

Assessment of Scaled Rotors for Wind Tunnel Experiments.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Maniaci, David Charles; Kelley, Christopher Lee; Chiu, Phillip

2015-07-01

Rotor design and analysis work has been performed to support the conceptualization of a wind tunnel test focused on studying wake dynamics. This wind tunnel test would serve as part of a larger model validation campaign that is part of the Department of Energy Wind and Water Power Program’s Atmosphere to electrons (A2e) initiative. The first phase of this effort was directed towards designing a functionally scaled rotor based on the same design process and target full-scale turbine used for new rotors for the DOE/SNL SWiFT site. The second phase focused on assessing the capabilities of an already available rotor,more » the G1, designed and built by researchers at the Technical University of München.« less

Cryogenic wind tunnels: Unique capabilities for the aerodynamicist

NASA Technical Reports Server (NTRS)

Hall, R. M.

1976-01-01

The cryogenic wind-tunnel concept as a practical means for improving ground simulation of transonic flight conditions. The Langley 1/3-meter transonic cryogenic tunnel is operational, and the design of a cryogenic National Transonic Facility is undertaken. A review of some of the unique capabilities of cryogenic wind tunnels is presented. In particular, the advantages of having independent control of tunnel Mach number, total pressure, and total temperature are highlighted. This separate control over the three tunnel parameters will open new frontiers in Mach number, Reynolds number, aeroelastic, and model-tunnel interaction studies.

System Dynamic Analysis of a Wind Tunnel Model with Applications to Improve Aerodynamic Data Quality

NASA Technical Reports Server (NTRS)

Buehrle, Ralph David

1997-01-01

The research investigates the effect of wind tunnel model system dynamics on measured aerodynamic data. During wind tunnel tests designed to obtain lift and drag data, the required aerodynamic measurements are the steady-state balance forces and moments, pressures, and model attitude. However, the wind tunnel model system can be subjected to unsteady aerodynamic and inertial loads which result in oscillatory translations and angular rotations. The steady-state force balance and inertial model attitude measurements are obtained by filtering and averaging data taken during conditions of high model vibrations. The main goals of this research are to characterize the effects of model system dynamics on the measured steady-state aerodynamic data and develop a correction technique to compensate for dynamically induced errors. Equations of motion are formulated for the dynamic response of the model system subjected to arbitrary aerodynamic and inertial inputs. The resulting modal model is examined to study the effects of the model system dynamic response on the aerodynamic data. In particular, the equations of motion are used to describe the effect of dynamics on the inertial model attitude, or angle of attack, measurement system that is used routinely at the NASA Langley Research Center and other wind tunnel facilities throughout the world. This activity was prompted by the inertial model attitude sensor response observed during high levels of model vibration while testing in the National Transonic Facility at the NASA Langley Research Center. The inertial attitude sensor cannot distinguish between the gravitational acceleration and centrifugal accelerations associated with wind tunnel model system vibration, which results in a model attitude measurement bias error. Bias errors over an order of magnitude greater than the required device accuracy were found in the inertial model attitude measurements during dynamic testing of two model systems. Based on a theoretical modal

Modifications to the 4x7 meter tunnel for acoustic research: Engineering feasibility study

NASA Technical Reports Server (NTRS)

1986-01-01

The NASA-Langley Research Center 4 x 7 Meter Low Speed Wind Tunnel is currently being used for low speed aerodynamics, V/STOL aerodynamics and, to a limited extent, rotorcraft noise research. The deficiencies of this wind tunnel for both aerodynamics and aeroacoustics research have been recognized for some time. Modifications to the wind tunnel are being made to improve the test section flow quality and to update the model cart systems. A further modification of the 4 x 7 Meter Wind Tunnel to permit rotorcraft model acoustics research has been proposed. As a precursor to the design of the proposed modifications, NASA is conducted both in-house and contracted studies to define the acoustic environment within the wind tunnel and to provide recommendations or the reduction of the wind tunnel background noise to a level acceptable to acoustics researchers. One of these studies by an acoustics consultant, has produced the primary reference documents that define the wind tunnel noise sources and outline recommended solutions.

Numerical investigation of air flow in a supersonic wind tunnel

NASA Astrophysics Data System (ADS)

Drozdov, S. M.; Rtishcheva, A. S.

2017-11-01

In the framework of TsAGI’s supersonic wind tunnel modernization program aimed at improving flow quality and extending the range of test regimes it was required to design and numerically validate a new test section and a set of shaped nozzles: two flat nozzles with flow Mach number at nozzle exit M=4 and M=5 and two axisymmetric nozzles with M=5 and M=6. Geometric configuration of the nozzles, the test section (an Eiffel chamber) and the diffuser was chosen according to the results of preliminary calculations of two-dimensional air flow in the wind tunnel circuit. The most important part of the work are three-dimensional flow simulation results obtained using ANSYS Fluent software. The following flow properties were investigated: Mach number, total and static pressure, total and static temperature and turbulent viscosity ratio distribution, heat flux density at wind tunnel walls (for high-temperature flow regimes). It is demonstrated that flow perturbations emerging from the junction of the nozzle with the test section and spreading down the test section behind the boundaries of characteristic rhomb’s reverse wedge are nearly impossible to eliminate. Therefore, in order to perform tests under most uniform flow conditions, the model’s center of rotation and optical window axis should be placed as close to the center of the characteristic rhomb as possible. The obtained results became part of scientific and technical basis of supersonic wind tunnel design process and were applied to a generalized class of similar wind tunnels.

Semi-span wind tunnel testing without conventional peniche

NASA Astrophysics Data System (ADS)

Skinner, S. N.; Zare-Behtash, H.

2017-12-01

Low-speed wind tunnel tests of a flexible wing semi-span model have been implemented in the 9× 7 ft de Havilland wind tunnel at the University of Glasgow. The main objective of this investigation is to quantify the effect of removing the traditional peniche boundary layer spacer utilised in this type of testing. Removal of the peniche results in a stand-off gap between the wind tunnel wall and the model's symmetry plane. This offers the advantage of preventing the development of a horseshoe vortex in front of the model, at the peniche/wall juncture. The formation of the horseshoe vortex is known to influence the flow structures around the entire model and thus alters the model's aerodynamic behaviours. To determine the influence of the stand-off gap, several gap heights have been tested for a range of angles of attack at Re=1.5× 10^6, based on the wing mean aerodynamic chord (MAC). Force platform data have been used to evaluate aerodynamic coefficients, and how they vary with stand-off heights. Stereoscopic Particle Imaging Velocimetry (sPIV) was used to examine the interaction between the tunnel boundary layer and model's respective stand-off gap. In addition, clay and tuft surface visualisation enhanced the understanding of how local flow structures over the length of the fuselage vary with stand-off height and angle of attack. The presented results show that a stand-off gap of four-to-five times the displacement thickness of the tunnel wall boundary layer is capable of achieving a flow field around the model fuselage that is representative of what would be expected for an equivalent full-span model in free-air—this cannot be achieved with the application of a peniche.

Heating requirements and nonadiabatic surface effects for a model in the NTF cryogenic wind tunnel

NASA Technical Reports Server (NTRS)

Macha, J. M.; Landrum, D. B.; Pare, L. A., III; Johnson, C. B.

1988-01-01

A theoretical study has been made of the severity of nonadiabatic surface conditions arising from internal heat sources within a model in a cryogenic wind tunnel. Local surface heating is recognized as having an effect on the development of the boundary layer, which can introduce changes in the flow about the model and affect the wind tunnel data. The geometry was based on the NTF Pathfinder I wind tunnel model. A finite element heat transfer computer code was developed and used to compute the steady state temperature distribution within the body of the model, from which the surface temperature distribution was extracted. Particular three dimensional characteristics of the model were represented with various axisymmetric approximations of the geometry. This analysis identified regions on the surface of the model susceptible to surface heating and the magnitude of the respective surface temperatures. It was found that severe surface heating may occur in particular instances, but could be alleviated with adequate insulating material. The heat flux through the surface of the model was integrated to determine the net heat required to maintain the instrumentation cavity at the prescribed temperature. The influence of the nonadiabatic condition on boundary layer properties and on the validity of the wind tunnel simulation was also investigated.

Wind-tunnel investigation of a flush airdata system at Mach numbers from 0.7 to 1.4

NASA Technical Reports Server (NTRS)

Larson, Terry J.; Moes, Timothy R.; Siemers, Paul M., III

1990-01-01

Flush pressure orifices installed on the nose section of a 1/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wing-tunnel tests were conducted in the 11- by 11-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which were published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2 deg increments from -4 deg to 20 deg. Sideslip angles were varied in 4 deg increments from -8 deg to 8 deg. Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

Aeroservoelastic Wind-Tunnel Tests of a Free-Flying, Joined-Wing SensorCraft Model for Gust Load Alleviation

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Castelluccio, Mark A.; Coulson, David A.; Heeg, Jennifer

2011-01-01

A team comprised of the Air Force Research Laboratory (AFRL), Boeing, and the NASA Langley Research Center conducted three aeroservoelastic wind-tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, exible vehicles. In the first of these three tests, a full-span, aeroelastically scaled, wind-tunnel model of a joined-wing SensorCraft vehicle was mounted to a force balance to acquire a basic aerodynamic data set. In the second and third tests, the same wind-tunnel model was mated to a new, two-degree-of-freedom, beam mount. This mount allowed the full-span model to translate vertically and pitch. Trimmed flight at -10% static margin and gust load alleviation were successfully demonstrated. The rigid body degrees of freedom required that the model be own in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free-flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort. The balance and free ying wind-tunnel tests will be summarized. The design of the trim and gust load alleviation control laws along with the associated results will also be discussed.

Two-dimensional wind tunnel

NASA Technical Reports Server (NTRS)

1982-01-01

Information on the Japanese National Aerospace Laboratory two dimensional transonic wind tunnel, completed at the end of 1979 is presented. Its construction is discussed in detail, and the wind tunnel structure, operation, test results, and future plans are presented.

Engineering and fabrication cost considerations for cryogenic wind tunnel models

NASA Technical Reports Server (NTRS)

Boykin, R. M., Jr.; Davenport, J. B., Jr.

1983-01-01

Design and fabrication cost drivers for cryogenic transonic wind tunnel models are defined. The major cost factors for wind tunnel models are model complexity, tolerances, surface finishes, materials, material validation, and model inspection. The cryogenic temperatures require the use of materials with relatively high fracture toughness but at the same time high strength. Some of these materials are very difficult to machine, requiring extensive machine hours which can add significantly to the manufacturing costs. Some additional engineering costs are incurred to certify the materials through mechanical tests and nondestructive evaluation techniques, which are not normally required with conventional models. When instrumentation such as accelerometers and electronically scanned pressure modules is required, temperature control of these devices needs to be incorporated into the design, which requires added effort. Additional thermal analyses and subsystem tests may be necessary, which also adds to the design costs. The largest driver to the design costs is potentially the additional static and dynamic analyses required to insure structural integrity of the model and support system.

40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

Code of Federal Regulations, 2012 CFR

2012-07-01

... Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

Code of Federal Regulations, 2013 CFR

2013-07-01

... Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

Code of Federal Regulations, 2014 CFR

2014-07-01

... Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test Primary Partical Mean Size a (µm) Full Wind Tunnel Test 2 km/hr 24 km/hr Inlet...

Measurement of unsteady loading and power output variability in a micro wind farm model in a wind tunnel

NASA Astrophysics Data System (ADS)

Bossuyt, Juliaan; Howland, Michael F.; Meneveau, Charles; Meyers, Johan

2017-01-01

Unsteady loading and spatiotemporal characteristics of power output are measured in a wind tunnel experiment of a microscale wind farm model with 100 porous disk models. The model wind farm is placed in a scaled turbulent boundary layer, and six different layouts, varied from aligned to staggered, are considered. The measurements are done by making use of a specially designed small-scale porous disk model, instrumented with strain gages. The frequency response of the measurements goes up to the natural frequency of the model, which corresponds to a reduced frequency of 0.6 when normalized by the diameter and the mean hub height velocity. The equivalent range of timescales, scaled to field-scale values, is 15 s and longer. The accuracy and limitations of the acquisition technique are documented and verified with hot-wire measurements. The spatiotemporal measurement capabilities of the experimental setup are used to study the cross-correlation in the power output of various porous disk models of wind turbines. A significant correlation is confirmed between streamwise aligned models, while staggered models show an anti-correlation.

A Photogrammetric System for Model Attitude Measurement in Hypersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Jones, Thomas W.; Lunsford, Charles B.

2007-01-01

A series of wind tunnel tests have been conducted to evaluate a multi-camera videogrammetric system designed to measure model attitude in hypersonic facilities. The technique utilizes processed video data and photogrammetric principles for point tracking to compute model position including pitch, roll and yaw. A discussion of the constraints encountered during the design, and a review of the measurement results obtained from the NASA Langley Research Center (LaRC) 31-Inch Mach 10 tunnel are presented.

Scale Model of Agena/Mariner-C in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1964-02-21

Researcher Bobby Sanders examines a 0.10-scale model of the Mariner-C shroud and Agena rocket in the 8- by 6-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Mariner-C and Mariner-D were identical spacecraft designed by the Jet Propulsion Laboratory to flyby Mars and photograph the Martian surface. The two Mariner spacecraft were launched by Atlas-Agena-D rockets. Lewis had taken over management of the Agena Program in October 1962. Lewis researchers investigated two different types of shrouds for the Mariner missions—an over-the-nose design and a backup pyrotechnic design. The new shroud was wider in diameter than the Agena rocket, so there was concern that this disparity might create air flow instability that could damage the shroud or destroy the vehicle. The tests in the 8- by 6 tunnel simulated launch speeds from Mach 0.56 to 1.96. Afterwards the Agena-Mariner-C model was studied in the 10- by 10-Foot Supersonic Wind Tunnel at speeds of Mach 2.0 to 3.5. Mariner-C was launched on November 4, 1964, but the payload shroud did not jettison properly and the spacecraft’s battery power did not function. The mission ended unsuccessfully two days later. Mariner-D was launched on November 28, 1964 and became the first successful mission to Mars. It was the first time a planet was photographed from space. Mariner-D’s 21 photographs revealed an inhospitable and barren landscape.

Full-Span Tiltrotor Aeroacoustic Model (TRAM) Overview and 40- by 80-Foot Wind Tunnel Test. [conducted in the 40- by 80-Foot Wind Tunnel at Ames Research Center

NASA Technical Reports Server (NTRS)

McCluer, Megan S.; Johnson, Jeffrey L.; Rutkowski, Michael (Technical Monitor)

2001-01-01

Most helicopter data trends cannot be extrapolated to tiltrotors because blade geometry and aerodynamic behavior, as well as rotor and fuselage interactions, are significantly different for tiltrotors. A tiltrotor model has been developed to investigate the aeromechanics of tiltrotors, to develop a comprehensive database for validating tiltrotor analyses, and to provide a research platform for supporting future tiltrotor designs. The Full-Span Tiltrotor Aeroacoustic Model (FS TRAM) is a dual-rotor, powered aircraft model with extensive instrumentation for measurement of structural and aerodynamic loads. This paper will present the Full-Span TRAM test capabilities and the first set of data obtained during a 40- by 80-Foot Wind Tunnel test conducted in late 2000 at NASA Ames Research Center. The Full-Span TRAM is a quarter-scale representation of the V-22 Osprey aircraft, and a heavily instrumented NASA and U.S. Army wind tunnel test stand. Rotor structural loads are monitored and recorded for safety-of-flight and for information on blade loads and dynamics. Left and right rotor balance and fuselage balance loads are monitored for safety-of-flight and for measurement of vehicle and rotor aerodynamic performance. Static pressure taps on the left wing are used to determine rotor/wing interactional effects and rotor blade dynamic pressures measure blade airloads. All of these measurement capabilities make the FS TRAM test stand a unique and valuable asset for validation of computational codes and to aid in future tiltrotor designs. The Full-Span TRAM was tested in the NASA Ames Research Center 40- by 80-Foot Wind Tunnel from October through December 2000. Rotor and vehicle performance measurements were acquired in addition to wing pressures, rotor acoustics, and Laser Light Sheet (LLS) flow visualization data. Hover, forward flight, and airframe (rotors off) aerodynamic runs were performed. Helicopter-mode data were acquired during angle of attack and thrust sweeps for

Anomalous Shocks on the Measured Near-Field Pressure Signatures of Low-Boom Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Mack, Robert J.

2006-01-01

Unexpected shocks on wind-tunnel-measured pressure signatures prompted questions about design methods, pressure signature measurement techniques, and the quality of measurements in the flow fields near lifting models. Some of these unexpected shocks were the result of component integration methods. Others were attributed to the three-dimension nature of the flow around a lifting model, to inaccuracies in the prediction of the area-ruled lift, or to wing-tip stall effects. This report discusses the low-boom model wind-tunnel data where these unexpected shocks were initially observed, the physics of the lifting wing/body model's flow field, the wind-tunnel data used to evaluate the applicability of methods for calculating equivalent areas due to lift, the performance of lift prediction codes, and tip stall effects so that the cause of these shocks could be determined.

Wind Tunnel and Hover Performance Test Results for Multicopter UAS Vehicles

NASA Technical Reports Server (NTRS)

Russell, Carl R.; Jung, Jaewoo; Willink, Gina; Glasner, Brett

2016-01-01

There is currently a lack of published data for the performance of multicopter unmanned aircraft system (UAS) vehicles, such as quadcopters and octocopters, often referred to collectively as drones. With the rapidly increasing popularity of multicopter UAS, there is interest in better characterizing the performance of this type of aircraft. By studying the performance of currently available vehicles, it will be possible to develop models for vehicles at this scale that can accurately predict performance and model trajectories. This paper describes a wind tunnel test that was recently performed in the U.S. Army's 7- by 10-ft Wind Tunnel at NASA Ames Research Center. During this wind tunnel entry, five multicopter UAS vehicles were tested to determine forces and moments as well as electrical power as a function of wind speed, rotor speed, and vehicle attitude. The test is described here in detail, and a selection of the key results from the test is presented.

Space Shuttle wind tunnel testing program

NASA Technical Reports Server (NTRS)

Whitnah, A. M.; Hillje, E. R.

1984-01-01

A major phase of the Space Shuttle Vehicle (SSV) Development Program was the acquisition of data through the space shuttle wind tunnel testing program. It became obvious that the large number of configuration/environment combinations would necessitate an extremely large wind tunnel testing program. To make the most efficient use of available test facilities and to assist the prime contractor for orbiter design and space shuttle vehicle integration, a unique management plan was devised for the design and development phase. The space shuttle program is reviewed together with the evolutional development of the shuttle configuration. The wind tunnel testing rationale and the associated test program management plan and its overall results is reviewed. Information is given for the various facilities and models used within this program. A unique posttest documentation procedure and a summary of the types of test per disciplines, per facility, and per model are presented with detailed listing of the posttest documentation.

Computational Modeling of the Ames 11-Ft Transonic Wind Tunnel in Conjunction with IofNEWT

NASA Technical Reports Server (NTRS)

Djomehri, M. Jahed; Buning, Pieter G.; Erickson, Larry L.; George, Michael W. (Technical Monitor)

1995-01-01

Technical advances in Computational Fluid Dynamics have now made it possible to simulate complex three-dimensional internal flows about models of various size placed in a Transonic Wind Tunnel. TWT wall interference effects have been a source of error in predicting flight data from actual wind tunnel measured data. An advantage of such internal CFD calculations is to directly compare numerical results with the actual tunnel data for code assessment and tunnel flow analysis. A CFD capability has recently been devised for flow analysis of the NASA/Ames 11-Ft TWT facility. The primary objectives of this work are to provide a CFD tool to study the NASA/Ames 11-Ft TWT flow characteristics, to understand the slotted wall interference effects, and to validate CFD codes. A secondary objective is to integrate the internal flowfield calculations with the Pressure Sensitive Paint data, a surface pressure distribution capability in Ames' production wind tunnels. The effort has been part of the Ames IofNEWT, Integration of Numerical and Experimental Wind Tunnels project, which is aimed at providing further analytical tools for industrial application. We used the NASA/Ames OVERFLOW code to solve the thin-layer Navier-Stokes equations. Viscosity effects near the model are captured by Baldwin-Lomax or Baldwin-Barth turbulence models. The solver was modified to model the flow behavior in the vicinity of the tunnel longitudinal slotted walls. A suitable porous type wall boundary condition was coded to account for the cross-flow through the test section. Viscous flow equations were solved in generalized coordinates with a three-factor implicit central difference scheme in conjunction with the Chimera grid procedure. The internal flow field about the model and the tunnel walls were descretized by the Chimera overset grid system. This approach allows the application of efficient grid generation codes about individual components of the configuration; separate minor grids were developed

Lockheed XFV-1 model in the 40x80 foot wind tunnel at NASA Ames Research Center

NASA Image and Video Library

1952-05-16

Wide shot of 40x 80 wind tunnel settling chamber with Lockheed XFV-1 model. Project engineer Mark Kelly (not shown). Remote controlled model flown in the settling chamber of the 40x80 wind tunnel. Electric motors in the model, controlled the counter-rotating propellers to test vertical takeoff. Test no. 71

Wind Tunnel Model and Test to Evaluate the Effectiveness of a Passive Gust Alleviation Device for a Flying Wing Aircraft

DTIC Science & Technology

2016-10-04

model of 1.24 m with the PGAD and control surface 3 1.2. Design and manufacture of the gust generator (frame, blades , actuation and control system...Chapter 3, a gust generator with two rotating blades was designed and manufactured to induce a transverse turbulence for wind tunnel test. A CFD...velocity at 8C (eight times of blade chord length) achieved 1.3%. In Chapter 4, the wind tunnel test of the scaled wing model is presented, including the

A flying superconducting magnet and cryostat for magnetic suspension of wind-tunnel models

NASA Technical Reports Server (NTRS)

Britcher, C.; Goodyer, M. J.; Scurlock, R. G.; Wu, Y. Y.

1984-01-01

The engineering practicality of a persistent high-field superconducting solenoid cryostat as a magnetic suspension and balance system (MSBS) for wind-tunnel testing of aircraft and missile models is examined. The test apparatus is a simple solenoid of filamentary NbTi superconductor with a cupronickel matrix. The apparatus, with a length-to-diameter ratio of 6 to 1 and a radius of 32 mm, used a 0.25 mm wire with a critical current of 27 A in an external field of 6 T. The total heat inleak of 150 mW was achieved. Helium boiloff rates were tested over a range of operating conditions, including pitch attitudes from 10 deg nose down to 90 deg nose up; the rate was estimated as low, but the aerodynamic acceptability of venting gaseous helium has not been determined. It is shown that the effectiveness of the concept increases with increasing scale, and performance in excess of that of conventional ferromagnets is achievable with reduction in size and costs, and with aptness to transonic wind-tunnel testing. Detailed specifications and schematics are included.

Full-scale S-76 rotor performance and loads at low speeds in the NASA Ames 80- by 120-Foot Wind Tunnel. Vol. 1

NASA Technical Reports Server (NTRS)

Shinoda, Patrick M.

1996-01-01

A full-scale helicopter rotor test was conducted in the NASA Ames 80- by 120-Foot Wind Tunnel with a four-bladed S-76 rotor system. Rotor performance and loads data were obtained over a wide range of rotor shaft angles-of-attack and thrust conditions at tunnel speeds ranging from 0 to 100 kt. The primary objectives of this test were (1) to acquire forward flight rotor performance and loads data for comparison with analytical results; (2) to acquire S-76 forward flight rotor performance data in the 80- by 120-Foot Wind Tunnel to compare with existing full-scale 40- by 80-Foot Wind Tunnel test data that were acquired in 1977; (3) to evaluate the acoustic capability of the 80- by 120- Foot Wind Tunnel for acquiring blade vortex interaction (BVI) noise in the low speed range and compare BVI noise with in-flight test data; and (4) to evaluate the capability of the 80- by 120-Foot Wind Tunnel test section as a hover facility. The secondary objectives were (1) to evaluate rotor inflow and wake effects (variations in tunnel speed, shaft angle, and thrust condition) on wind tunnel test section wall and floor pressures; (2) to establish the criteria for the definition of flow breakdown (condition where wall corrections are no longer valid) for this size rotor and wind tunnel cross-sectional area; and (3) to evaluate the wide-field shadowgraph technique for visualizing full-scale rotor wakes. This data base of rotor performance and loads can be used for analytical and experimental comparison studies for full-scale, four-bladed, fully articulated rotor systems. Rotor performance and structural loads data are presented in this report.

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

NASA Technical Reports Server (NTRS)

Kmak, Frank J.; Hudgins, M.; Hergert, D.; George, Michael W. (Technical Monitor)

2001-01-01

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.

Smart wing wind tunnel test results

NASA Astrophysics Data System (ADS)

Scherer, Lewis B.; Martin, Christopher A.; Appa, Kari; Kudva, Jayanth N.; West, Mark N.

1997-05-01

The use of smart materials technologies can provide unique capabilities in improving aircraft aerodynamic performance. Northrop Grumman built and tested a 16% scale semi-span wind tunnel model of the F/A-18 E/F for the on-going DARPA/WL Smart Materials and Structures-Smart Wing Program. Aerodynamic performance gains to be validated included increase in the lift to drag ratio, increased pitching moment (Cm), increased rolling moment (Cl) and improved pressure distribution. These performance gains were obtained using hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist via a SMA torque tube and are compared to a conventional wind tunnel model with hinged control surfaces. This paper presents an overview of the results from the first wind tunnel test performed at the NASA Langley's 16 ft Transonic Dynamic Tunnel. Among the benefits demonstrated are 8 - 12% increase in rolling moment due to wing twist, a 10 - 15% increase in rolling moment due to contoured aileron, and approximately 8% increase in lift due to contoured flap, and improved pressure distribution due to trailing edge control surface contouring.

Convair XF-102 Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1953-08-21

A .10-scale model of Convair’s XF-102 in the 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory for jet exit studies. The XF-102 was a prototype of the F-102 Delta Dagger. The F-102 served as an interceptor against long range bombers from the Soviet Union. The aircraft was powered by a Pratt and Whitney J57 turbojet. The first prototype crashed two weeks after is first flight on October 24, 1953, just months after this photograph. Engineers then incorporated the fixed-wing design to reduce drag at supersonic speeds. The production model F-102 became the first delta-wing supersonic aircraft in operation. The 8- by 6-Foot Supersonic Wind Tunnel is used to study propulsion systems, including inlets and exit nozzles, combustion fuel injectors, flame holders, exit nozzles, and controls on ramjet and turbojet engines. Flexible sidewalls alter the tunnel’s nozzle shape to vary the Mach number during operation. A seven-stage axial compressor, driven by three electric motors that yield a total of 87,000 horsepower, generates air speeds from Mach 0.36 to 2.0.

Wind Tunnel Testing of a 120th Scale Large Civil Tilt-Rotor Model in Airplane and Helicopter Modes

NASA Technical Reports Server (NTRS)

Theodore, Colin R.; Willink, Gina C.; Russell, Carl R.; Amy, Alexander R.; Pete, Ashley E.

2014-01-01

In April 2012 and October 2013, NASA and the U.S. Army jointly conducted a wind tunnel test program examining two notional large tilt rotor designs: NASA's Large Civil Tilt Rotor and the Army's High Efficiency Tilt Rotor. The approximately 6%-scale airframe models (unpowered) were tested without rotors in the U.S. Army 7- by 10-foot wind tunnel at NASA Ames Research Center. Measurements of all six forces and moments acting on the airframe were taken using the wind tunnel scale system. In addition to force and moment measurements, flow visualization using tufts, infrared thermography and oil flow were used to identify flow trajectories, boundary layer transition and areas of flow separation. The purpose of this test was to collect data for the validation of computational fluid dynamics tools, for the development of flight dynamics simulation models, and to validate performance predictions made during conceptual design. This paper focuses on the results for the Large Civil Tilt Rotor model in an airplane mode configuration up to 200 knots of wind tunnel speed. Results are presented with the full airframe model with various wing tip and nacelle configurations, and for a wing-only case also with various wing tip and nacelle configurations. Key results show that the addition of a wing extension outboard of the nacelles produces a significant increase in the lift-to-drag ratio, and interestingly decreases the drag compared to the case where the wing extension is not present. The drag decrease is likely due to complex aerodynamic interactions between the nacelle and wing extension that results in a significant drag benefit.

Dynamic response of NASA Rotor Test Apparatus and Sikorsky S-76 hub mounted in the 80- by 120-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Peterson, Randall L.; Hoque, Muhammed S.

1994-01-01

A shake test was conducted in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center, using the NASA Ames Rotor Test Apparatus (RTA) and the Sikorsky S-76 rotor hub. The primary objective of this shake test was to determine the modal properties of the RTA, the S-76 rotor hub, and the model support system installed in the wind tunnel. Random excitation was applied at the rotor hub, and vibration responses were measured using accelerometers mounted at various critical locations on the model and the model support system. Transfer functions were computed using the load cell data and the accelerometer responses. The transfer function data were used to compute the system modal parameters with the aid of modal analysis software.

Plasma Wind Tunnel Testing of Electron Transpiration Cooling Concept

DTIC Science & Technology

2017-02-28

AFRL-AFOSR-UK-TR-2017-0012 Plasma Wind Tunnel Testing of Electron Transpiration Cooling Concept Olivier Chazot INSTITUT VON KARMAN DE DYNAMIQUE DES...28-02-2017 2. REPORT TYPE Final 3. DATES COVERED (From - To) 01 Dec 2015 to 30 Nov 2016 4. TITLE AND SUBTITLE Plasma Wind Tunnel Testing of Electron ...Aeronautics and Aerospace Department B-1640 Rhode Saint Genèse Belgium Internal Ref: ARR 1605 February 2017 Plasma Wind Tunnel Testing of Electron

Wind Tunnel Investigation of Ground Wind Loads for Ares Launch Vehicle

NASA Technical Reports Server (NTRS)

Keller, Donald F.; Ivanco, Thomas G.

2010-01-01

A three year program was conducted at the NASA Langley Research Center (LaRC) Aeroelasticity Branch (AB) and Transonic Dynamics Tunnel (TDT) with the primary objective to acquire scaled steady and dynamic ground-wind loads (GWL) wind-tunnel data for rollout, on-pad stay, and on-pad launch configurations for the Ares I-X Flight Test Vehicle (FTV). The experimental effort was conducted to obtain an understanding of the coupling of aerodynamic and structural characteristics that can result in large sustained wind-induced oscillations (WIO) on such a tall and slender launch vehicle and to generate a unique database for development and evaluation of analytical methods for predicting steady and dynamic GWL, especially those caused by vortex shedding, and resulting in significant WIO. This paper summarizes the wind-tunnel test program that employed two dynamically-aeroelastically scaled GWL models based on the Ares I-X Flight Test Vehicle. The first model tested, the GWL Checkout Model (CM), was a relatively simple model with a secondary objective of restoration and development of processes and methods for design, fabrication, testing, and data analysis of a representative ground wind loads model. In addition, parametric variations in surface roughness, Reynolds number, and protuberances (on/off) were investigated to determine effects on GWL characteristics. The second windtunnel model, the Ares I-X GWL Model, was significantly more complex and representative of the Ares I-X FTV and included the addition of simplified rigid geometrically-scaled models of the Kennedy Space Center (KSC) Mobile Launch Platform (MLP) and Launch Complex 39B primary structures. Steady and dynamic base bending moment as well as model response and steady and unsteady pressure data was acquired during the testing of both models. During wind-tunnel testing of each model, flow conditions (speed and azimuth) where significant WIO occurred, were identified and thoroughly investigated. Scaled data from

Real-Gas Flow Properties for NASA Langley Research Center Aerothermodynamic Facilities Complex Wind Tunnels

NASA Technical Reports Server (NTRS)

Hollis, Brian R.

1996-01-01

A computational algorithm has been developed which can be employed to determine the flow properties of an arbitrary real (virial) gas in a wind tunnel. A multiple-coefficient virial gas equation of state and the assumption of isentropic flow are used to model the gas and to compute flow properties throughout the wind tunnel. This algorithm has been used to calculate flow properties for the wind tunnels of the Aerothermodynamics Facilities Complex at the NASA Langley Research Center, in which air, CF4. He, and N2 are employed as test gases. The algorithm is detailed in this paper and sample results are presented for each of the Aerothermodynamic Facilities Complex wind tunnels.

40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

Code of Federal Regulations, 2011 CFR

2011-07-01

... 40 Protection of Environment 5 2011-07-01 2011-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test...

40 CFR Table F-2 to Subpart F of... - Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and...

Code of Federal Regulations, 2010 CFR

2010-07-01

... 40 Protection of Environment 5 2010-07-01 2010-07-01 false Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test, and Static Chamber Test F Table F-2 to Subpart F... Part 53—Particle Sizes and Wind Speeds for Full Wind Tunnel Test, Wind Tunnel Inlet Aspiration Test...

Wind-tunnel investigation of the thrust augmentor performance of a large-scale swept wing model. [in the Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Koenig, D. G.; Falarski, M. D.

1979-01-01

Tests were made in the Ames 40- by 80-foot wind tunnel to determine the forward speed effects on wing-mounted thrust augmentors. The large-scale model was powered by the compressor output of J-85 driven viper compressors. The flap settings used were 15 deg and 30 deg with 0 deg, 15 deg, and 30 deg aileron settings. The maximum duct pressure, and wind tunnel dynamic pressure were 66 cmHg (26 in Hg) and 1190 N/sq m (25 lb/sq ft), respectively. All tests were made at zero sideslip. Test results are presented without analysis.

Wall Interference in Two-Dimensional Wind Tunnels

NASA Technical Reports Server (NTRS)

Kemp, William B., Jr.

1986-01-01

Viscosity and tunnel-wall constraints introduced via boundary conditions. TWINTN4 computer program developed to implement method of posttest assessment of wall interference in two-dimensional wind tunnels. Offers two methods for combining sidewall boundary-layer effects with upper and lower wall interference. In sequential procedure, Sewall method used to define flow free of sidewall effects, then assessed for upper and lower wall effects. In unified procedure, wind-tunnel flow equations altered to incorporate effects from all four walls at once. Program written in FORTRAN IV for batch execution.

Wind tunnel test results of a 1/8-scale fan-in-wing model

NASA Technical Reports Server (NTRS)

Wilson, John C.; Gentry, Garl L.; Gorton, Susan A.

1996-01-01

A 1/8-scale model of a fan-in-wing concept considered for development by Grumman Aerospace Corporation for the U.S. Army was tested in the Langley 14- by 22-Foot Subsonic Tunnel. Hover testing, which included height above a pressure-instrumented ground plane, angle of pitch, and angle of roll for a range of fan thrust, was conducted in a model preparation area near the tunnel. The air loads and surface pressures on the model were measured for several configurations in the model preparation area and in the tunnel. The major hover configuration change was varying the angles of the vanes attached to the exit of the fans for producing propulsive force. As the model height above the ground was decreased, there was a significant variation of thrust-removed normal force with constant fan speed. The greatest variation was generally for the height-to-fan exit diameter ratio of less than 2.5; the variation was reduced by deflecting fan exit flow outboard with the vanes. In the tunnel angles of pitch and sideslip, height above the tunnel floor, and wind speed were varied for a range of fan thrust and different vane angle configurations. Other configuration features such as flap deflections and tail incidence were evaluated as well. Though the V-tail empennage provided an increase in static longitudinal stability, the total model configuration remained unstable.

Modal Correction Method For Dynamically Induced Errors In Wind-Tunnel Model Attitude Measurements

NASA Technical Reports Server (NTRS)

Buehrle, R. D.; Young, C. P., Jr.

1995-01-01

This paper describes a method for correcting the dynamically induced bias errors in wind tunnel model attitude measurements using measured modal properties of the model system. At NASA Langley Research Center, the predominant instrumentation used to measure model attitude is a servo-accelerometer device that senses the model attitude with respect to the local vertical. Under smooth wind tunnel operating conditions, this inertial device can measure the model attitude with an accuracy of 0.01 degree. During wind tunnel tests when the model is responding at high dynamic amplitudes, the inertial device also senses the centrifugal acceleration associated with model vibration. This centrifugal acceleration results in a bias error in the model attitude measurement. A study of the response of a cantilevered model system to a simulated dynamic environment shows significant bias error in the model attitude measurement can occur and is vibration mode and amplitude dependent. For each vibration mode contributing to the bias error, the error is estimated from the measured modal properties and tangential accelerations at the model attitude device. Linear superposition is used to combine the bias estimates for individual modes to determine the overall bias error as a function of time. The modal correction model predicts the bias error to a high degree of accuracy for the vibration modes characterized in the simulated dynamic environment.

Initial adaptation testing of the bidimensionally self-adapting wall of the French T2 wind tunnel, around a three-dimensional object

NASA Technical Reports Server (NTRS)

Archambaud, J. P.; Dor, J. B.; Mignosi, A.; Lamarche, L.

1986-01-01

The test series was carried out at ONERA/CERT at the T2 wind tunnel in September 1984. The objective of this series was to minimize wall interference through a bidimensional adaptation around the models, inducing tridimensional flows. For this, three different models were used, measuring either the pressures or the forces and moment of pitch (balance). The adaptation was derived from a correction computation in the compressible axisymmetric tridimensional.

Airloads Correlation of the UH-60A Rotor Inside the 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Chang, I-Chung; Norman, Thomas R.; Romander, Ethan A.

2013-01-01

The presented research validates the capability of a loosely-coupled computational fluid dynamics (CFD) and comprehensive rotorcraft analysis (CRA) code to calculate the flowfield around a rotor and test stand mounted inside a wind tunnel. The CFD/CRA predictions for the full-scale UH-60A Airloads Rotor inside the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center are compared with the latest measured airloads and performance data. The studied conditions include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed up to and including stall. For the speed sweep, wind tunnel modeling becomes important at advance ratios greater than 0.37 and test stand modeling becomes increasingly important as the advance ratio increases. For the thrust sweep, both the wind tunnel and test stand modeling become important as the rotor approaches stall. Despite the beneficial effects of modeling the wind tunnel and test stand, the new models do not completely resolve the current airload discrepancies between prediction and experiment.

Development of a distributed-parameter mathematical model for simulation of cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Tripp, J. S.

1983-01-01

A one-dimensional distributed-parameter dynamic model of a cryogenic wind tunnel was developed which accounts for internal and external heat transfer, viscous momentum losses, and slotted-test-section dynamics. Boundary conditions imposed by liquid-nitrogen injection, gas venting, and the tunnel fan were included. A time-dependent numerical solution to the resultant set of partial differential equations was obtained on a CDC CYBER 203 vector-processing digital computer at a usable computational rate. Preliminary computational studies were performed by using parameters of the Langley 0.3-Meter Transonic Cryogenic Tunnel. Studies were performed by using parameters from the National Transonic Facility (NTF). The NTF wind-tunnel model was used in the design of control loops for Mach number, total temperature, and total pressure and for determining interactions between the control loops. It was employed in the application of optimal linear-regulator theory and eigenvalue-placement techniques to develop Mach number control laws.

RSRA sixth scale wind tunnel test. [of scale model of Sikorsky Whirlwind Helicopter

NASA Technical Reports Server (NTRS)

Flemming, R.; Ruddell, A.

1974-01-01

The sixth scale model of the Sikorsky/NASA/Army rotor systems research aircraft was tested in an 18-foot section of a large subsonic wind tunnel for the purpose of obtaining basic data in the areas of performance, stability, and body surface loads. The model was mounted in the tunnel on the struts arranged in tandem. Basic testing was limited to forward flight with angles of yaw from -20 to +20 degrees and angles of attack from -20 to +25 degrees. Tunnel test speeds were varied up to 172 knots (q = 96 psf). Test data were monitored through a high speed static data acquisition system, linked to a PDP-6 computer. This system provided immediate records of angle of attack, angle of yaw, six component force and moment data, and static and total pressure information. The wind tunnel model was constructed of aluminum structural members with aluminum, fiberglass, and wood skins. Tabulated force and moment data, flow visualization photographs, tabulated surface pressure data are presented for the basic helicopter and compound configurations. Limited discussions of the results of the test are included.

Glide back booster wind tunnel model testing

NASA Astrophysics Data System (ADS)

Pricop, M. V.; Cojocaru, M. G.; Stoica, C. I.; Niculescu, M. L.; Neculaescu, A. M.; Persinaru, A. G.; Boscoianu, M.

2017-07-01

Affordable space access requires partial or ideally full launch vehicle reuse, which is in line with clean environment requirement. Although the idea is old, the practical use is difficult, requiring very large technology investment for qualification. Rocket gliders like Space Shuttle have been successfullyoperated but the price and correspondingly the energy footprint were found not sustainable. For medium launchers, finally there is a very promising platform as Falcon 9. For very small launchers the situation is more complex, because the performance index (payload to start mass) is already small, versus medium and heavy launchers. For partial reusable micro launchers this index is even smaller. However the challenge has to be taken because it is likely that in a multiyear effort, technology is going to enable the performance recovery to make such a system economically and environmentally feasible. The current paper is devoted to a small unitary glide back booster which is foreseen to be assembled in a number of possible configurations. Although the level of analysis is not deep, the solution is analyzed from the aerodynamic point of view. A wind tunnel model is designed, with an active canard, to enablea more efficient wind tunnel campaign, as a national level premiere.

Qualification of the T2 wind tunnel in cryogenic operation. A: Thermal field, preliminary study of a schematic model

NASA Technical Reports Server (NTRS)

Dor, J. B.; Mignosi, A.; Plazanet, M.

1984-01-01

The T2 wind tunnel is described. The process of generating a cyrogenic gust using the example of a test made at very low temperature is presented. Detailed results of tests on temperatures for flow in the settling chamber, the interior walls of the system, and the metal casing are given. The transverse temperature distribution in the settling chamber and working section, and of the thermal gradients in the walls, are given as a function of the temperature level of the test.

Wind tunnel measurements for dispersion modelling of vehicle wakes

NASA Astrophysics Data System (ADS)

Carpentieri, Matteo; Kumar, Prashant; Robins, Alan

2012-12-01

Wind tunnel measurements downwind of reduced scale car models have been made to study the wake regions in detail, test the usefulness of existing vehicle wake models, and draw key information needed for dispersion modelling in vehicle wakes. The experiments simulated a car moving in still air. This is achieved by (i) the experimental characterisation of the flow, turbulence and concentration fields in both the near and far wake regions, (ii) the preliminary assessment of existing wake models using the experimental database, and (iii) the comparison of previous field measurements in the wake of a real diesel car with the wind tunnel measurements. The experiments highlighted very large gradients of velocities and concentrations existing, in particular, in the near-wake. Of course, the measured fields are strongly dependent on the geometry of the modelled vehicle and a generalisation for other vehicles may prove to be difficult. The methodology applied in the present study, although improvable, could constitute a first step towards the development of mathematical parameterisations. Experimental results were also compared with the estimates from two wake models. It was found that they can adequately describe the far-wake of a vehicle in terms of velocities, but a better characterisation in terms of turbulence and pollutant dispersion is needed. Parameterised models able to predict velocity and concentrations with fine enough details at the near-wake scale do not exist.

Comparison of results of an obstacle resolving microscale model with wind tunnel data

NASA Astrophysics Data System (ADS)

Grawe, David; Schlünzen, K. Heinke; Pascheke, Frauke

2013-11-01

The microscale transport and stream model MITRAS has been improved and a new technique has been implemented to improve numerical stability for complex obstacle configurations. Results of the updated version have been compared with wind tunnel data using an evaluation method that has been established for simple obstacle configurations. MITRAS is a part of the M-SYS model system for the assessment of ambient air quality. A comparison of model results for the flow field against quality ensured wind tunnel data has been carried out for both idealised and realistic test cases. Results of the comparison show a very good agreement of the wind field for most test cases and identify areas of possible improvement of the model. The evaluated MITRAS results can be used as input data for the M-SYS microscale chemistry model MICTM. This paper describes how such a comparison can be carried out for simple as well as realistic obstacle configurations and what difficulties arise.

Development of an apparatus to measure thermophysical properties of wind tunnel heat transfer models

NASA Technical Reports Server (NTRS)

Romanowski, R. F.; Steinberg, I. H.

1974-01-01

The apparatus and technique for measuring the thermophysical properties of models used with the phase-change paint method for obtaining wind tunnel heat transfer data are described. The method allows rapid measurement of the combined properties in a transient manner similar to an actual wind tunnel test. An effective value of the thermophysical properties can be determined which accounts for changes in thermal properties with temperature or with depth into the model surface. The apparatus was successfully tested at various heating rates between 19,000 and 124,000 watts per square meter.

Measurement Techniques for Flow Diagnostic in ITAM Impulse Wind Tunnels

DTIC Science & Technology

2010-04-01

time of wind - tunnel operation, so that oscillations caused by initial shock loads could decay and a comparatively long time period with constant flow...Flow Diagnostic in ITAM Impulse Wind Tunnels 7 - 4 RTO-EN-AVT-186 A strain-gauge pressure probe is an elastic element (membrane) in a sealed...Diagnostic in ITAM Impulse Wind Tunnels RTO-EN-AVT-186 7 - 5 probes are individually calibrated. Piezoelectric pressure gauges are based

Ares I Aerodynamic Testing at the Boeing Polysonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Pinier, Jeremy T.; Niskey, Charles J.; Hanke, Jeremy L.; Tomek, William G.

2011-01-01

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.

Real-time simulator for helicopter rotor wind-tunnel operations

NASA Technical Reports Server (NTRS)

Talbot, P. D.; Peterson, R. L.; Graham, D. R.

1986-01-01

This paper describes the elements and operation of a simulator that is being used to train operators of the Rotor Test Apparatus (RTA) in the large-scale 40- by 80-Foot Wind Tunnel at Ames Research Center. The simulator, named TUTOR (for Tunnel Utilization Trainer with Operating Rotor) duplicates the controls of the rotor and its dynamic behavior, as well as the wind-tunnel controls. The simulation software uses a preexisting blade-element model of a four-bladed rotor with flapping and lead-lag degrees of freedom. Equations were developed for all hardware and controls of the RTA and of the wind tunnel that are normally required to perform a wind-tunnel test of a helicopter rotor. The simulator hardware consists of consoles designed to have the same appearance and functions as those in the control room of the 40- by 80-Foot Wind Tunnel, allowing input from three operators who normally establish the required operating conditions during a test run. Normal operating procedures can be practiced, as well as simulated emergencies such as rotor power failure.

A new electronic scanner of pressure designed for installation in wind-tunnel models

NASA Technical Reports Server (NTRS)

Coe, C. T.; Parra, G. T.; Kauffman, R. C.

1981-01-01

A new electronic scanner of pressure (ESOP) has been developed by NASA Ames Research Center for installation in wind-tunnel models. An ESOP system includes up to 20 pressure modules, each with 48 pressure transducers, an A/D converter, a microprocessor, a data controller, a monitor unit, and a heater controller. The system is sized so that the pressure modules and A/D converter module can be installed within an average-size model tested in the Ames Aerodynamics Division wind tunnels. This paper describes the ESOP system, emphasizing the main element of the system - the pressure module. The measured performance of the overall system is also presented.

Surface flow visualization of separated flows on the forebody of an F-18 aircraft and wind-tunnel model

NASA Technical Reports Server (NTRS)

Fisher, David F.; Richwine, David M.; Banks, Daniel W.

1988-01-01

A method of in-flight surface flow visualization similar to wind-tunnel-model oil flows is described for cases where photo-chase planes or onboard photography are not practical. This method, used on an F-18 aircraft in flight at high angles of attack, clearly showed surface flow streamlines in the fuselage forebody. Vortex separation and reattachment lines were identified with this method and documented using postflight photography. Surface flow angles measured at the 90 and 270 degrees meridians show excellent agreement with the wind tunnel data for a pointed tangent ogive with an aspect ratio of 3.5. The separation and reattachment line locations were qualitatively similar to the F-18 wind-tunnel-model oil flows but neither the laminar separation bubble nor the boundary-layer transition on the wind tunnel model were evident in the flight surface flows. The separation and reattachment line locations were in fair agreement with the wind tunnel data for the 3.5 ogive. The elliptical forebody shape of the F-18 caused the primary separation lines to move toward the leeward meridian. Little effect of angle of attack on the separation locations was noted for the range reported.

Computational Results for the KTH-NASA Wind-Tunnel Model Used for Acquisition of Transonic Nonlinear Aeroelastic Data

NASA Technical Reports Server (NTRS)

Silva, Walter A.; Chwalowski, Pawel; Wieseman, Carol D.; Eller, David; Ringertz, Ulf

2017-01-01

A status report is provided on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the aeroelastic analyses of a full-span fighter configuration wind-tunnel model. This wind-tunnel model was tested in the Transonic Dynamics Tunnel (TDT) in the summer of 2016. Large amounts of data were acquired including steady/unsteady pressures, accelerations, strains, and measured dynamic deformations. The aeroelastic analyses presented include linear aeroelastic analyses, CFD steady analyses, and analyses using CFD-based reduced-order models (ROMs).

Aerodynamic performance of a low-speed wind tunnel.

PubMed

Frechen, F-B; Frey, M; Wett, M; Löser, C

2004-01-01

The determination of the odour mass flow emitted from a source is a very important step and forms the basis for all subsequent considerations and calculations. Wastewater treatment plants, as well as waste treatment facilities, consist of different kinds of odour sources. Unfortunately, most of the sources are passive sources, where no outward air flow-rate can be measured, but where odorants are obviously emitted. Thus, a type of sampling is required that allows to measure the emitted odour flow-rate (OFR). To achieve this, different methods are in use worldwide. Besides indirect methods, such as micrometeorological atmospheric dispersion models, which have not been used in Germany (in other countries due to different problems, direct methods are also used). Direct measurements include hood methods, commonly divided into static flux chambers, dynamic flux chambers and wind tunnels. The wind tunnel that we have been operating in principle since 1983 is different from all subsequent presented wind tunnels, in that we operate it at a considerably lower wind speed than the others. To describe the behaviour of this wind tunnel, measurement of the flow pattern in this low-speed tunnel are under way, and some initial results are presented here.

Advanced Turboprop Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1979-08-21

NASA Lewis Research Center researcher, John S. Sarafini, uses a laser doppler velocimeter to analyze a Hamilton Standard SR-2 turboprop design in the 8- by 6-Foot foot Supersonic Wind Tunnel. Lewis researchers were analyzing a series of eight-bladed propellers in their wind tunnels to determine their operating characteristics at speeds up to Mach 0.8. The program, which became the Advanced Turboprop (ATP), was part of a NASA-wide Aircraft Energy Efficiency Program undertaken to reduce aircraft fuel costs by 50 percent. The ATP concept was different from the turboprops in use in the 1950s. The modern versions had at least eight blades and were swept back for better performance. Bell Laboratories developed the laser doppler velocimeter technology in the 1960s to measure velocity of transparent fluid flows or vibration motion on reflective surfaces. Lewis researchers modified the device to measure the flow field of turboprop configurations in the transonic speed region. The modifications were necessary to overcome the turboprop’s vibration and noise levels. The laser beam was split into two beams which were crossed at a specific point. This permits researchers to measure two velocity components simultaneously. This data measures speeds both ahead and behind the propeller blades. Researchers could use this information as they sought to advance flow fields and to verify computer modeling codes.

Development and Evaluation of a Dynamic, 3-Degree-of-Freedom (DOF) Wind Tunnel Model

DTIC Science & Technology

2016-11-01

DATE (DD-MM-YYYY) November 2016 2. REPORT TYPE Contractor Report 3. DATES COVERED (From - To) November 2014–October 2015 4. TITLE AND SUBTITLE...eliminates any gimbal-induced flow disturbances. All 3 axes can be locked independently and are instrumented with high -resolution feedback. Experiments... high -speed video frames of the subsonic model in the transonic wind tunnel, Run 63

Mach 4 free-jet tunnel starting experiments for a hypersonic research engine model causing high blockage

NASA Technical Reports Server (NTRS)

Carson, G. T., Jr.; Midden, R. E.

1976-01-01

Tests of a full scale hypersonic research engine (HRE) were conducted in the hypersonic tunnel facility at Mach numbers of 5, 6, and 7. Since the HRE would cause a rather high blockage (48.83 percent of the nozzle area), subscale tests were conducted in various available small wind tunnels prior to the full scale tests to study the effects of model blockage on tunnel starting. The results of the Mach 4 subscale tests which utilized a model system at 0.0952 scale which simulated the HRE in the test section of the tunnel are presented. A satisfactory tunnel starting could not be achieved by varying the free jet length or diffuser size nor by inserting the model into the test stream after tunnel starting. However, the installation of a shroud around the HRE model allowed the tunnel to start with the model preset in the tunnel at a tunnel stagnation pressure to atmospheric exit pressure ratio of 13.4. The simulation of the discharge of instrumentation cooling water and the addition of test hardware at the aft end of the HRE model did not have a significant effect on the tunnel starting.

Development of a quiet supersonic wind tunnel with a cryogenic adaptive nozzle

NASA Technical Reports Server (NTRS)

Wolf, Stephen W. D.

1995-01-01

Low-disturbance or 'quiet' wind tunnels are now considered an essential part of meaningful boundary layer transition research. Advances in Supersonic Laminar Flow Control (SLFC) technology for swept wings depends on a better understanding of the receptivity of the transition phenomena to attachment-line contamination and cross-flows. This need has provided the impetus for building the Laminar Flow Supersonic Wind Tunnel (LFSWT) at NASA-Ames, as part of the NASA High Speed Research Program (HSRP). The LFSWT was designed to provide NASA with an unequaled capability for transition research at low supersonic Mach numbers (<2.5). The following are the objectives in support of the new Fluid Mechanic Laboratory (FML) quiet supersonic wind tunnel: (I) Develop a unique injector drive system using the existing FML indraft compressor; (2) Develop an FML instrumentation capability for quiet supersonic wind tunnel evaluation and transition studies at NASA-Ames; (3) Determine the State of the Art in quiet supersonic wind tunnel design; (4) Build and commission the LFSWT; (5) Make detailed flow quality measurements in the LFSWT; (6) Perform tests of swept wing models in the LFSWT in support of the NASA HSR program; and (7) Provide documentation of research progress.

7. VIEW WEST OF SCALE ROOM IN FULLSCALE WIND TUNNEL; ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

7. VIEW WEST OF SCALE ROOM IN FULL-SCALE WIND TUNNEL; SCALES ARE USED TO MEASURE FORCES ACTING ON MODEL AIRCRAFT SUSPENDED ABOVE. - NASA Langley Research Center, Full-Scale Wind Tunnel, 224 Hunting Avenue, Hampton, Hampton, VA

Study of optical techniques for the Ames unitary wind tunnel, part 7

NASA Technical Reports Server (NTRS)

Lee, George

1993-01-01

A summary of optical techniques for the Ames Unitary Plan wind tunnels are discussed. Six optical techniques were studied: Schlieren, light sheet and laser vapor screen, angle of attack, model deformation, infrared imagery, and digital image processing. The study includes surveys and reviews of wind tunnel optical techniques, some conceptual designs, and recommendations for use of optical methods in the Ames Unitary Plan wind tunnels. Particular emphasis was placed on searching for systems developed for wind tunnel use and on commercial systems which could be readily adapted for wind tunnels. This final report is to summarize the major results and recommendations.

Analysis of heat-transfer measurements from 2 AEDC wind tunnels on the Shuttle external tank

NASA Technical Reports Server (NTRS)

Nutt, K. W.

1984-01-01

Previous aerodynamic heating tests have been conducted in the AEDC/VKF Supersonic Wind Tunnel (A) to aid in defining the design thermal environment for the space shuttle external tank. The quality of these data has been under discussion because of the effects of low tunnel enthalpy and slow model injection rates. Recently the AEDC/VKF Hypersonic Wind Tunnel (C) has been modified to provide a Mach 4 capability that has significantly higher tunnel enthalpy with more rapid model injection rates. Tests were conducted in Tunnel C at Mach 4 to obtain data on the external tank for comparison with Tunnel A results. Data were obtained on a 0.0175 scale model of the Space Shuttle Integrated Vehicle at Re/ft = 4 x 10 to the 6th power with the tunnel stagnation temperature varying from 740 to 1440 R. Model attitude varied from an angle of attack of -5 to 5 deg and an angle of sideslip of -3 to 3 deg. One set of data was obtained in Tunnel C at Re/ft = 6.9 x 10 to the 6th for comparison with flight data. Data comparisons between the two tunnels for numerous regions on the external tank are given.

Digital control of wind tunnel magnetic suspension and balance systems

NASA Technical Reports Server (NTRS)

Britcher, Colin P.; Goodyer, Michael J.; Eskins, Jonathan; Parker, David; Halford, Robert J.

1987-01-01

Digital controllers are being developed for wind tunnel magnetic suspension and balance systems, which in turn permit wind tunnel testing of aircraft models free from support interference. Hardware and software features of two existing digital control systems are reviewed. Some aspects of model position sensing and system calibration are also discussed.

Cone-Probe Rake Design and Calibration for Supersonic Wind Tunnel Models

NASA Technical Reports Server (NTRS)

Won, Mark J.

1999-01-01

A series of experimental investigations were conducted at the NASA Langley Unitary Plan Wind Tunnel (UPWT) to calibrate cone-probe rakes designed to measure the flow field on 1-2% scale, high-speed wind tunnel models from Mach 2.15 to 2.4. The rakes were developed from a previous design that exhibited unfavorable measurement characteristics caused by a high probe spatial density and flow blockage from the rake body. Calibration parameters included Mach number, total pressure recovery, and flow angularity. Reference conditions were determined from a localized UPWT test section flow survey using a 10deg supersonic wedge probe. Test section Mach number and total pressure were determined using a novel iterative technique that accounted for boundary layer effects on the wedge surface. Cone-probe measurements were correlated to the surveyed flow conditions using analytical functions and recursive algorithms that resolved Mach number, pressure recovery, and flow angle to within +/-0.01, +/-1% and +/-0.1deg , respectively, for angles of attack and sideslip between +/-8deg. Uncertainty estimates indicated the overall cone-probe calibration accuracy was strongly influenced by the propagation of measurement error into the calculated results.

5. VIEW NORTH OF TEST SECTION IN FULLSCALE WIND TUNNEL ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

5. VIEW NORTH OF TEST SECTION IN FULL-SCALE WIND TUNNEL WITH FREE-FLIGHT MODEL OF A BOEING 737 SUSPENDED FROM A SAFETY CABLE. - NASA Langley Research Center, Full-Scale Wind Tunnel, 224 Hunting Avenue, Hampton, Hampton, VA

Simulating flow around scaled model of a hypersonic vehicle in wind tunnel

NASA Astrophysics Data System (ADS)

Markova, T. V.; Aksenov, A. A.; Zhluktov, S. V.; Savitsky, D. V.; Gavrilov, A. D.; Son, E. E.; Prokhorov, A. N.

2016-11-01

A prospective hypersonic HEXAFLY aircraft is considered in the given paper. In order to obtain the aerodynamic characteristics of a new construction design of the aircraft, experiments with a scaled model have been carried out in a wind tunnel under different conditions. The runs have been performed at different angles of attack with and without hydrogen combustion in the scaled propulsion engine. However, the measured physical quantities do not provide all the information about the flowfield. Numerical simulation can complete the experimental data as well as to reduce the number of wind tunnel experiments. Besides that, reliable CFD software can be used for calculations of the aerodynamic characteristics for any possible design of the full-scale aircraft under different operation conditions. The reliability of the numerical predictions must be confirmed in verification study of the software. The given work is aimed at numerical investigation of the flowfield around and inside the scaled model of the HEXAFLY-CIAM module under wind tunnel conditions. A cold run (without combustion) was selected for this study. The calculations are performed in the FlowVision CFD software. The flow characteristics are compared against the available experimental data. The carried out verification study confirms the capability of the FlowVision CFD software to calculate the flows discussed.

Development of an Intelligent Videogrammetric Wind Tunnel Measurement System

NASA Technical Reports Server (NTRS)

Graves, Sharon S.; Burner, Alpheus W.

2004-01-01

A videogrammetric technique developed at NASA Langley Research Center has been used at five NASA facilities at the Langley and Ames Research Centers for deformation measurements on a number of sting mounted and semispan models. These include high-speed research and transport models tested over a wide range of aerodynamic conditions including subsonic, transonic, and supersonic regimes. The technique, based on digital photogrammetry, has been used to measure model attitude, deformation, and sting bending. In addition, the technique has been used to study model injection rate effects and to calibrate and validate methods for predicting static aeroelastic deformations of wind tunnel models. An effort is currently underway to develop an intelligent videogrammetric measurement system that will be both useful and usable in large production wind tunnels while providing accurate data in a robust and timely manner. Designed to encode a higher degree of knowledge through computer vision, the system features advanced pattern recognition techniques to improve automated location and identification of targets placed on the wind tunnel model to be used for aerodynamic measurements such as attitude and deformation. This paper will describe the development and strategy of the new intelligent system that was used in a recent test at a large transonic wind tunnel.

Flutter suppression digital control law design and testing for the AFW wind tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1994-01-01

The design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a sting mounted fixed-in-roll aeroelastic wind-tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory, and it also involved control law order reduction, a gain root-locus study, and use of previous experimental results. A 23 percent increase in the open-loop flutter dynamic pressure was demonstrated during the wind-tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Flutter suppression digital control law design and testing for the AFW wind-tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

Design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a string mounted fixed-in-roll aeroelastic wind tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory and involved control law order reduction, a gain root-locus study, and the use of previous experimental results. A 23 percent increase in open-loop flutter dynamic pressure was demonstrated during the wind tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Flutter suppression digital control law design and testing for the AFW wind tunnel model

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

1992-01-01

Design of a control law for simultaneously suppressing the symmetric and antisymmetric flutter modes of a sting mounted fixed-in-roll aeroelastic wind tunnel model is described. The flutter suppression control law was designed using linear quadratic Gaussian theory, and involved control law order reduction, a gain root-locus study and use of previous experimental results. A 23 percent increase in the open-loop flutter dynamic pressure was demonstrated during the wind tunnel test. Rapid roll maneuvers at 11 percent above the symmetric flutter boundary were also performed when the model was in a free-to-roll configuration.

Wind tunnel technology for the development of future commercial aircraft

NASA Technical Reports Server (NTRS)

Szodruch, J.

1986-01-01

Requirements for new technologies in the area of civil aircraft design are mainly related to the high cost involved in the purchase of modern, fuel saving aircraft. A second important factor is the long term rise in the price of fuel. The demonstration of the benefits of new technologies, as far as these are related to aerodynamics, will,for the foreseeable future, still be based on wind tunnel measurements. Theoretical computation methods are very successfully used in design work, wing optimization, and an estimation of the Reynolds number effect. However, wind tunnel tests are still needed to verify the feasibility of the considered concepts. Along with other costs, the cost for the wind tunnel tests needed for the development of an aircraft is steadily increasing. The present investigation is concerned with the effect of numerical aerodynamics and civil aircraft technology on the development of wind tunnels. Attention is given to the requirements for the wind tunnel, investigative methods, measurement technology, models, and the relation between wind tunnel experiments and theoretical methods.

Using wind tunnels to predict bird mortality in wind farms: the case of griffon vultures.

PubMed

de Lucas, Manuela; Ferrer, Miguel; Janss, Guyonne F E

2012-01-01

Wind farms have shown a spectacular growth during the last 15 years. Avian mortality through collision with moving rotor blades is well-known as one of the main adverse impacts of wind farms. In Spain, the griffon vulture incurs the highest mortality rates in wind farms. As far as we know, this study is the first attempt to predict flight trajectories of birds in order to foresee potentially dangerous areas for wind farm development. We analyse topography and wind flows in relation to flight paths of griffon vultures, using a scaled model of the wind farm area in an aerodynamic wind tunnel, and test the difference between the observed flight paths of griffon vultures and the predominant wind flows. Different wind currents for each wind direction in the aerodynamic model were observed. Simulations of wind flows in a wind tunnel were compared with observed flight paths of griffon vultures. No statistical differences were detected between the observed flight trajectories of griffon vultures and the wind passages observed in our wind tunnel model. A significant correlation was found between dead vultures predicted proportion of vultures crossing those cells according to the aerodynamic model. Griffon vulture flight routes matched the predominant wind flows in the area (i.e. they followed the routes where less flight effort was needed). We suggest using these kinds of simulations to predict flight paths over complex terrains can inform the location of wind turbines and thereby reduce soaring bird mortality.

Fluorescence Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Alderfer, D. W.; Danehy, P. M.; Inma, J. A.; Berger, K. T.; Buck, G. M.; Schwartz, R J.

2007-01-01

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Most of the models did not survive repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2-inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various configurations were studied including different sting placements relative to the models, different model orientations and attachment angles, and different NO seeding methods. The angle of attack of the models was also varied and the location of the laser sheet was scanned to provide three-dimensional flowfield information. Virtual Diagnostics Interface technology, developed at NASA Langley, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images. Lessons learned and recommendations for future experiments are discussed.

NASA Glenn Wind Tunnel Model Systems Criteria

NASA Technical Reports Server (NTRS)

Soeder, Ronald H.; Roeder, James W.; Stark, David E.; Linne, Alan A.

2004-01-01

This report describes criteria for the design, analysis, quality assurance, and documentation of models that are to be tested in the wind tunnel facilities at the NASA Glenn Research Center. This report presents two methods for computing model allowable stresses on the basis of the yield stress or ultimate stress, and it defines project procedures to test models in the NASA Glenn aeropropulsion facilities. Both customer-furnished and in-house model systems are discussed. The functions of the facility personnel and customers are defined. The format for the pretest meetings, safety permit process, and model reviews are outlined. The format for the model systems report (a requirement for each model that is to be tested at NASA Glenn) is described, the engineers responsible for developing the model systems report are listed, and the timetable for its delivery to the project engineer is given.

The problem of dimensional instability in airfoil models for cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Wigley, D. A.

1982-01-01

The problem of dimensional instability in airfoil models for cryogenic wind tunnels is discussed in terms of the various mechanisms that can be responsible. The interrelationship between metallurgical structure and possible dimensional instability in cryogenic usage is discussed for those steel alloys of most interest for wind tunnel model construction at this time. Other basic mechanisms responsible for setting up residual stress systems are discussed, together with ways in which their magnitude may be reduced by various elevated or low temperature thermal cycles. A standard specimen configuration is proposed for use in experimental investigations into the effects of machining, heat treatment, and other variables that influence the dimensional stability of the materials of interest. A brief classification of various materials in terms of their metallurgical structure and susceptability to dimensional instability is presented.

Wind tunnel and ground static tests of a .094 scale powered model of a modified T-39 lift/cruise fan V/STOL research airplane

NASA Technical Reports Server (NTRS)

Hunt, D.; Clinglan, J.; Salemann, V.; Omar, E.

1977-01-01

Ground static and wind tunnel test of a scale model modified T-39 airplane are reported. The configuration in the nose and replacement of the existing nacelles with tilting lift/cruise fans. The model was powered with three 14 cm diameter tip driven turbopowered simulators. Forces and moments were measured by an internal strain guage balance. Engine simulator thrust and mass flow were measured by calibrated pressure and temperature instrumentation mounted downstream of the fans. The low speed handling qualities and general aerodynamic characteristics of the modified T-39 were defined. Test variables include thrust level and thrust balance, forward speed, model pitch and sideslip angle at forward speeds, model pitch, roll, and ground height during static tests, lift/cruise fan tilt angle, flap and aileron deflection angle, and horizonal stabilizer angle. The effects of removing the landing gear, the lift/cruise fans, and the tail surfaces were also investigated.

Implementation of a Particle Image Velocimetry System for Wind Tunnel Flowfield Measurements

DTIC Science & Technology

2014-12-01

Instrumentation Wind tunnel speed was measured by two pitot probes mounted on opposite tunnel walls upstream of the model and above the ground...board. The pitot probes were connected differentially to Scanivalve 1-psi transducers. A secondary measurement of wind tunnel speed was made with the...Manf. Model Range 1 Tunnel Vel (south pitot ) Transducer Scanivalve CR24D 1 psi 2 Tunnel Vel (north pitot ) Transducer Scanivalve CR24D 1 psi 3

Numerical Simulation of Selecting Model Scale of Cable in Wind Tunnel Test

NASA Astrophysics Data System (ADS)

Huang, Yifeng; Yang, Jixin

The numerical simulation method based on computational Fluid Dynamics (CFD) provides a possible alternative means of physical wind tunnel test. Firstly, the correctness of the numerical simulation method is validated by one certain example. In order to select the minimum length of the cable as to a certain diameter in the numerical wind tunnel tests, the numerical wind tunnel tests based on CFD are carried out on the cables with several different length-diameter ratios (L/D). The results show that, when the L/D reaches to 18, the drag coefficient is stable essentially.

An economical model for simulating droplet spectrum evolution in turbulent cloud chambers and wind tunnels

NASA Astrophysics Data System (ADS)

Krueger, Steven; Cantrell, W.; Niedermeier, D.; Shaw, R.; Stratmann, F.

2017-11-01

Although airborne instruments provide detailed information about the microphysical structure of clouds, the measurements provide only a few snapshots of each cloud. Deducing the droplet spectrum evolution from such measurements is next to impossible. We are using two alternative approaches: laboratory studies and numerical simulations. The former relies on a new turbulent cloud chamber (the Pi Chamber) at Michigan Technical University, as well as the first humid turbulent wind tunnel (LACIS-T) at the Leibniz Institute for Tropospheric Research. Both produce conditions for droplet growth (i.e., supersaturation) by mixing saturated vapor at different temperatures. The Pi Chamber produces turbulence by inducing Rayleigh-Bénard convection, while the wind tunnel generates turbulence with a grid. We are using the Explicit Mixing Parcel Model (EMPM) to numerically simulate droplet spectrum evolution in these flows. The EMPM explicitly links turbulent mixing and droplet spectrum evolution by representing a turbulent flow in a 1D domain with the linear eddy model. The EMPM can economically span scales from those of the smallest turbulent eddies to those of the largest. The EMPM grows or evaporates thousands of individual cloud droplets according to their local environments.

Altitude Wind Tunnel Drive Fan being Assembled

NASA Image and Video Library

1943-07-21

National Advisory Committee for Aeronautics (NACA) engineers assembled the Altitude Wind Tunnel’s (AWT) large wooden drive fan inside the hangar at the Aircraft Engine Research Laboratory. When it was built at the in the early 1940s the AWT was among the most complex test facilities ever designed. It was the first wind tunnel capable of operating full-scale engines under realistic flight conditions. This simulation included the reduction of air temperature, a decrease in air pressure, and the creation of an airstream velocity of up to 500 miles per hour. The AWT was constructed in 1942 and 1943. This photograph shows NACA engineers Lou Hermann and Jack Aust assembling the tunnel’s drive fan inside the hangar. The 12-bladed, 31-foot-diameter spruce wood fan would soon be installed inside the wind tunnel to create the high-speed airflow. This massive propeller was designed and constructed by the engine lab's design team at Langley Field. John Breisch, a Langley technician with several years of wind tunnel installation experience, arrived in Cleveland at the time of this photograph to supervise the fan assembly inside the hangar. He would return several weeks later to oversee the actual installation in the tunnel. The fan was driven at 410 revolutions per minute by an 18,000-horsepower General Electric induction motor that was located in the rear corner of the Exhauster Building. An extension shaft connected the motor to the fan. A bronze screen protected the fan against damage from failed engine parts sailing through the tunnel. Despite this screen the blades did become worn or cracked over time and had to be replaced. An entire new fan was installed in 1951.

The steady-state flow quality in a model of a non-return wind tunnel

NASA Technical Reports Server (NTRS)

Mort, K. W.; Eckert, W. T.; Kelly, M. W.

1972-01-01

The structural cost of non-return wind tunnels is significantly less than that of the more conventional closed-circuit wind tunnels. However, because of the effects of external winds, the flow quality of non-return wind tunnels is an area of concern at the low test speeds required for V/STOL testing. The flow quality required at these low speeds is discussed and alternatives to the traditional manner of specifying the flow quality requirements in terms of dynamic pressure and angularity are suggested. The development of a non-return wind tunnel configuration which has good flow quality at low as well as at high test speeds is described.

Reduction of Background Noise in the NASA Ames 40- by 80-Foot Wind Tunnel

NASA Technical Reports Server (NTRS)

Jaeger, Stephen M.; Allen, Christopher S.; Soderman, Paul T.; Olson, Larry E. (Technical Monitor)

1995-01-01

Background noise in both open-jet and closed wind tunnels adversely affects the signal-to-noise ratio of acoustic measurements. To measure the noise of increasingly quieter aircraft models, the background noise will have to be reduced by physical means or through signal processing. In a closed wind tunnel, such as the NASA Ames 40- by 80- Foot Wind Tunnel, the principle background noise sources can be classified as: (1) fan drive noise; (2) microphone self-noise; (3) aerodynamically induced noise from test-dependent hardware such as model struts and junctions; and (4) noise from the test section walls and vane set. This paper describes the steps taken to minimize the influence of each of these background noise sources in the 40 x 80.

Enhancements to AERMOD's building downwash algorithms based on wind-tunnel and Embedded-LES modeling

NASA Astrophysics Data System (ADS)

Monbureau, E. M.; Heist, D. K.; Perry, S. G.; Brouwer, L. H.; Foroutan, H.; Tang, W.

2018-04-01

Knowing the fate of effluent from an industrial stack is important for assessing its impact on human health. AERMOD is one of several Gaussian plume models containing algorithms to evaluate the effect of buildings on the movement of the effluent from a stack. The goal of this study is to improve AERMOD's ability to accurately model important and complex building downwash scenarios by incorporating knowledge gained from a recently completed series of wind tunnel studies and complementary large eddy simulations of flow and dispersion around simple structures for a variety of building dimensions, stack locations, stack heights, and wind angles. This study presents three modifications to the building downwash algorithm in AERMOD that improve the physical basis and internal consistency of the model, and one modification to AERMOD's building pre-processor to better represent elongated buildings in oblique winds. These modifications are demonstrated to improve the ability of AERMOD to model observed ground-level concentrations in the vicinity of a building for the variety of conditions examined in the wind tunnel and numerical studies.

Low-speed wind tunnel test results of the Canard Rotor/Wing concept

NASA Technical Reports Server (NTRS)

Bass, Steven M.; Thompson, Thomas L.; Rutherford, John W.; Swanson, Stephen

1993-01-01

The Canard Rotor/Wing (CRW), a high-speed rotorcraft concept, was tested at the National Aeronautics and Space Administration (NASA) Ames Research Center's 40- by 80-Foot Wind Tunnel in Mountain View, California. The 1/5-scale model was tested to identify certain low-speed, fixed-wing, aerodynamic characteristics of the configuration and investigate the effectiveness of two empennages, an H-Tail and a T-Tail. The paper addresses the principal test objectives and the results achieved in the wind tunnel test. These are summarized as: i) drag build-up and differences between the H-Tail and T-Tail configuration, ii) longitudinal stability of the H-Tail and T-Tail configurations in the conversion and cruise modes, iii) control derivatives for the canard and elevator in the conversion and cruise modes, iv) aerodynamic characteristics of varying the rotor/wing azimuth position, and v) canard and tail lift/trim capability for conversion conditions.

System Identification and Uncertainty Quantification Using Orthogonal Excitations and the Semi-span Super Sonic Transport (S4T) Model

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Wieseman, Carol D.

2012-01-01

Orthogonal harmonic multisine excitations were utilized in a wind tunnel test and in simulation of the SemiSpan Supersonic Transport model to assess aeroservoelastic characteristics. Fundamental issues associated with analyzing sinusoidal signals were examined, including spectral leakage, excitation truncation, and uncertainties on frequency response functions and mean-square coherence. Simulation allowed for evaluation of these issues relative to a truth model, while wind tunnel data introduced real-world implementation issues.

Using Wind Tunnels to Predict Bird Mortality in Wind Farms: The Case of Griffon Vultures

PubMed Central

de Lucas, Manuela; Ferrer, Miguel; Janss, Guyonne F. E.

2012-01-01

Background Wind farms have shown a spectacular growth during the last 15 years. Avian mortality through collision with moving rotor blades is well-known as one of the main adverse impacts of wind farms. In Spain, the griffon vulture incurs the highest mortality rates in wind farms. Methodology/Principal Findings As far as we know, this study is the first attempt to predict flight trajectories of birds in order to foresee potentially dangerous areas for wind farm development. We analyse topography and wind flows in relation to flight paths of griffon vultures, using a scaled model of the wind farm area in an aerodynamic wind tunnel, and test the difference between the observed flight paths of griffon vultures and the predominant wind flows. Different wind currents for each wind direction in the aerodynamic model were observed. Simulations of wind flows in a wind tunnel were compared with observed flight paths of griffon vultures. No statistical differences were detected between the observed flight trajectories of griffon vultures and the wind passages observed in our wind tunnel model. A significant correlation was found between dead vultures predicted proportion of vultures crossing those cells according to the aerodynamic model. Conclusions Griffon vulture flight routes matched the predominant wind flows in the area (i.e. they followed the routes where less flight effort was needed). We suggest using these kinds of simulations to predict flight paths over complex terrains can inform the location of wind turbines and thereby reduce soaring bird mortality. PMID:23152764

Overview of the 1989 Wind Tunnel Calibration Workshop

NASA Technical Reports Server (NTRS)

Henderson, Arthur, Jr.; Mckinney, L. Wayne

1993-01-01

An overview of the 1989 Wind Tunnel Calibration Workshop held at NASA LaRC in Hampton, VA on 19-20 Apr. 1989 is presented. The purpose of the Workshop was to explore wind tunnel calibration requirements as they relate to test quality and data accuracy, with the ultimate goal of developing wind tunnel calibration requirements for the major NASA wind tunnels at ARC, LaRC, and LeRC. The two sessions addressed the following topics: (1) what constitutes a properly calibrated wind tunnel; and (2) the status of calibration of NASA's major wind tunnels. The most significant contributions to the stated goals are highlighted, and the consensus of the Workshop's conclusions and recommendations regarding formulation and implementation of that goal are presented.

Flight and wind-tunnel correlation of boundary-layer transition on the AEDC transition cone

NASA Technical Reports Server (NTRS)

Fisher, D. L.; Dougherty, N. S., Jr.

1982-01-01

Transition and fluctuating surface pressure data were acquired on a 10 deg included angle cone, using the same instrumentation and technique over a wide range of Mach and Reynolds numbers in 23 wind tunnels and in flight. Transition was detected with a traversing pitot-pressure probe in contact with the surface. The surface pressure fluctuations were measured with microphones set flush in the cone surface. Good correlation of end of transition Reynolds number RE(T) was obtained between data from the lower disturbance wind tunnels and flight up to a boundary layer edge Mach number, M(e) = 1.2. Above M(e) = 1.2, however, this correlation deteriorates, with the flight Re(T) being 25 to 30% higher than the wind tunnel Re(T) at M(e) = 1.6. The end of transition Reynolds number correlated within + or - 20% with the surface pressure fluctuations, according to the equation used. Broad peaks in the power spectral density distributions indicated that Tollmien-Schlichting waves were the probable cause of transition in flight and in some of the wind tunnels.

Wind tunnel test of a variable-diameter tiltrotor (VDTR) model

NASA Technical Reports Server (NTRS)

Matuska, David; Dale, Allen; Lorber, Peter

1994-01-01

This report documents the results from a wind tunnel test of a 1/6th scale Variable Diameter Tiltrotor (VDTR). This test was a joint effort of NASA Ames and Sikorsky Aircraft. The objective was to evaluate the aeroelastic and performance characteristics of the VDTR in conversion, hover, and cruise. The rotor diameter and nacelle angle of the model were remotely changed to represent tiltrotor operating conditions. Data is presented showing the propulsive force required in conversion, blade loads, angle of attack stability and simulated gust response, and hover and cruise performance. This test represents the first wind tunnel test of a variable diameter rotor applied to a tiltrotor concept. The results confirm some of the potential advantages of the VDTR and establish the variable diameter rotor a viable candidate for an advanced tiltrotor. This wind tunnel test successfully demonstrated the feasibility of the Variable Diameter rotor for tilt rotor aircraft. A wide range of test points were taken in hover, conversion, and cruise modes. The concept was shown to have a number of advantages over conventional tiltrotors such as reduced hover downwash with lower disk loading and significantly reduced longitudinal gust response in cruise. In the conversion regime, a high propulsive force was demonstrated for sustained flight with acceptable blade loads. The VDTR demonstrated excellent gust response capabilities. The horizontal gust response correlated well with predictions revealing only half the response to turbulence of the conventional civil tiltrotor.

Comparison of propeller cruise noise data taken in the NASA Lewis 8- by 6-foot wind tunnel with other tunnel and flight data

NASA Technical Reports Server (NTRS)

Dittmar, James H.

1989-01-01

The noise of advanced high speed propeller models measured in the NASA 8- by 6-foot wind tunnel has been compared with model propeller noise measured in another tunnel and with full-scale propeller noise measured in flight. Good agreement was obtained for the noise of a model counterrotation propeller tested in the 8- by 6-foot wind tunnel and in the acoustically treated test section of the Boeing Transonic Wind Tunnel. This good agreement indicates the relative validity of taking cruise noise data on a plate in the 8- by 6-foot wind tunnel compared with the free-field method in the Boeing tunnel. Good agreement was also obtained for both single rotation and counter-rotation model noise comparisons with full-scale propeller noise in flight. The good scale model to full-scale comparisons indicate both the validity of the 8- by 6-foot wind tunnel data and the ability to scale to full size. Boundary layer refraction on the plate provides a limitation to the measurement of forward arc noise in the 8- by 6-foot wind tunnel at the higher harmonics of the blade passing tone. The use of a validated boundary layer refraction model to adjust the data could remove this limitation.

Comparison of propeller cruise noise data taken in the NASA Lewis 8- by 6-foot wind tunnel with other tunnel and flight data

NASA Technical Reports Server (NTRS)

Dittmar, James

1989-01-01

The noise of advanced high speed propeller models measured in the NASA 8- by 6-foot wind tunnel has been compared with model propeller noise measured in another tunnel and with full-scale propeller noise measured in flight. Good agreement was obtained for the noise of a model counterrotation propeller tested in the 8- by 6-foot wind tunnel and in the acoustically treated test section of the Boeing Transonic Wind Tunnel. This good agreement indicates the relative validity of taking cruise noise data on a plate in the 8- by 6-foot wind tunnel compared with the free-field method in the Boeing tunnel. Good agreement was also obtained for both single rotation and counter-rotation model noise comparisons with full-scale propeller noise in flight. The good scale model to full-scale comparisons indicate both the validity of the 8- by 6-foot wind tunnel data and the ability to scale to full size. Boundary layer refraction on the plate provides a limitation to the measurement of forward arc noise in the 8- by 6-foot wind tunnel at the higher harmonics of the blade passing tone. The sue of a validated boundary layer refraction model to adjust the data could remove this limitation.

Study of the integration of wind tunnel and computational methods for aerodynamic configurations

NASA Technical Reports Server (NTRS)

Browne, Lindsey E.; Ashby, Dale L.

1989-01-01

A study was conducted to determine the effectiveness of using a low-order panel code to estimate wind tunnel wall corrections. The corrections were found by two computations. The first computation included the test model and the surrounding wind tunnel walls, while in the second computation the wind tunnel walls were removed. The difference between the force and moment coefficients obtained by comparing these two cases allowed the determination of the wall corrections. The technique was verified by matching the test-section, wall-pressure signature from a wind tunnel test with the signature predicted by the panel code. To prove the viability of the technique, two cases were considered. The first was a two-dimensional high-lift wing with a flap that was tested in the 7- by 10-foot wind tunnel at NASA Ames Research Center. The second was a 1/32-scale model of the F/A-18 aircraft which was tested in the low-speed wind tunnel at San Diego State University. The panel code used was PMARC (Panel Method Ames Research Center). Results of this study indicate that the proposed wind tunnel wall correction method is comparable to other methods and that it also inherently includes the corrections due to model blockage and wing lift.

NACA Technician Cleans a Ramjet in 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1950-04-21

A technician at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory cleans the pitot tube on a 16-inch diameter ramjet in the 8- by 6-Foot Supersonic Wind Tunnel. Pitot tubes are a measurement device used to determine the flow velocity at a specific location in the air stream, not the average velocity of the entire wind stream. NACA Lewis was in the midst of a multi-year program to determine the feasibility of ramjets and design improvements that could be employed for all models. The advantage of the ramjet was its ability to process large volumes of combustion air, resulting in the burning of fuel at the optimal stoichiometric temperatures. This was not possible with turbojets. The higher the Mach number, the more efficient the ramjet operated. The 8- by 6 Supersonic Wind Tunnel had been in operation for just over one year when this photograph was taken. The facility was the NACA’s largest supersonic tunnel and the only facility capable of running an engine at supersonic speeds. The 8- by 6 tunnel was also equipped with a Schlieren camera system that captured the air flow gradient as it passes over the test setup. The ramjet tests in the 8- by 6 tunnel complemented the NACA Lewis investigations using aircraft, the Altitude Wind Tunnel and smaller supersonic tunnels. Researchers studied the ramjet’s performance at different speeds and varying angles -of -attack.

Wind tunnel measurements of wake structure and wind farm power for actuator disk model wind turbines in yaw

NASA Astrophysics Data System (ADS)

Howland, Michael; Bossuyt, Juliaan; Kang, Justin; Meyers, Johan; Meneveau, Charles

2016-11-01

Reducing wake losses in wind farms by deflecting the wakes through turbine yawing has been shown to be a feasible wind farm control approach. In this work, the deflection and morphology of wakes behind a wind turbine operating in yawed conditions are studied using wind tunnel experiments of a wind turbine modeled as a porous disk in a uniform inflow. First, by measuring velocity distributions at various downstream positions and comparing with prior studies, we confirm that the nonrotating wind turbine model in yaw generates realistic wake deflections. Second, we characterize the wake shape and make observations of what is termed a "curled wake," displaying significant spanwise asymmetry. Through the use of a 100 porous disk micro-wind farm, total wind farm power output is studied for a variety of yaw configurations. Strain gages on the tower of the porous disk models are used to measure the thrust force as a substitute for turbine power. The frequency response of these measurements goes up to the natural frequency of the model and allows studying the spatiotemporal characteristics of the power output under the effects of yawing. This work has been funded by the National Science Foundation (Grants CBET-113380 and IIA-1243482, the WINDINSPIRE project). JB and JM are supported by ERC (ActiveWindFarms, Grant No. 306471).

Measurements of temperature and pressure fluctuations in the T prime 2 cryogenic wind tunnel

NASA Technical Reports Server (NTRS)

Blanchard, A.; Dor, J. B.; Breil, J. F.

1980-01-01

Cold wire measurement of temperature fluctuations were made in a DERAT T'2 induction powered cryogenic wind tunnel for 2 types of liquid nitrogen injectors. Thermal turbulence measured in the tranquilization chamber depends to a great extent on the injector used; for fine spray of nitrogen drops, this level of turbulence seemed completely acceptable. Fluctuations in static pressure taken from the walls of the vein by Kulite sensors showed that there was no increase in aerodynamic noise during cryogenic gusts.

Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

NASA Technical Reports Server (NTRS)

Koning, Witold J. F.

2016-01-01

Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tiltrotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity Unsteady Reynolds Averaged Navier-Stokes (URANS) solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade-element model (BEM) with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt, and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation, and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall, interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A "quasi linear trim" was used to trim the thrust

Wind Tunnel Interference Effects on Tilt Rotor Testing Using Computational Fluid Dynamics

NASA Technical Reports Server (NTRS)

Koning, Witold J. F.

2015-01-01

Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity URANS solver is used with an incompressible flow model and a realizable k-e turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade element model with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at NASA Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A 'quasi linear trim' was used to trim the thrust for the rotor to compare the power as a unique

Preliminary Tests in the NACA Free-Spinning Wind Tunnel

NASA Technical Reports Server (NTRS)

Zimmerman, C H

1937-01-01

Typical models and the testing technique used in the NACA free-spinning wind tunnel are described in detail. The results of tests on two models afford a comparison between the spinning characteristics of scale models in the tunnel and of the airplanes that they represent.

Experimental Results from the Active Aeroelastic Wing Wind Tunnel Test Program

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Spain, Charles V.; Florance, James R.; Wieseman, Carol D.; Ivanco, Thomas G.; DeMoss, Joshua; Silva, Walter A.; Panetta, Andrew; Lively, Peter; Tumwa, Vic

2005-01-01

The Active Aeroelastic Wing (AAW) program is a cooperative effort among NASA, the Air Force Research Laboratory and the Boeing Company, encompassing flight testing, wind tunnel testing and analyses. The objective of the AAW program is to investigate the improvements that can be realized by exploiting aeroelastic characteristics, rather than viewing them as a detriment to vehicle performance and stability. To meet this objective, a wind tunnel model was crafted to duplicate the static aeroelastic behavior of the AAW flight vehicle. The model was tested in the NASA Langley Transonic Dynamics Tunnel in July and August 2004. The wind tunnel investigation served the program goal in three ways. First, the wind tunnel provided a benchmark for comparison with the flight vehicle and various levels of theoretical analyses. Second, it provided detailed insight highlighting the effects of individual parameters upon the aeroelastic response of the AAW vehicle. This parameter identification can then be used for future aeroelastic vehicle design guidance. Third, it provided data to validate scaling laws and their applicability with respect to statically scaled aeroelastic models.

A wind tunnel study on the effects of complex topography on wind turbine performance

NASA Astrophysics Data System (ADS)

Howard, Kevin; Hu, Stephen; Chamorro, Leonardo; Guala, Michele

2012-11-01

A set of wind tunnel experiments were conducted to study the response of a wind turbine under flow conditions typically observed at the wind farm scale, in complex terrain. A scale model wind turbine was placed in a fully developed turbulent boundary layer flow obtained in the SAFL Wind Tunnel. Experiments focused on the performance of a turbine model, under the effects induced by a second upwind turbine or a by three-dimensional, sinusoidal hill, peaking at the turbine hub height. High frequency measurements of fluctuating streamwise and wall normal velocities were obtained with a X-wire anemometer simultaneously with the rotor angular velocity and the turbine(s) voltage output. Velocity measurements in the wake of the first turbine and of the hill were used to determine the inflow conditions for the downwind test turbine. Turbine performance was inferred by the mean and fluctuating voltage statistics. Specific experiments were devoted to relate the mean voltage to the mean hub velocity, and the fluctuating voltage to the unsteadiness in the rotor kinematics induced by the perturbed (hill or turbine) or unperturbed (boundary layer) large scales of the incoming turbulent flow. Results show that the voltage signal can be used to assess turbine performance in complex flows.

Analytical Models for Rotor Test Module, Strut, and Balance Frame Dynamics in the 40 by 80 Ft Wind Tunnel

NASA Technical Reports Server (NTRS)

Johnson, W.

1976-01-01

A mathematical model is developed for the dynamics of a wind tunnel support system consisting of a balance frame, struts, and an aircraft or test module. Data are given for several rotor test modules in the Ames 40 by 80 ft wind tunnel. A model for ground resonance calculations is also described.

Application of interactive computer graphics in wind-tunnel dynamic model testing

NASA Technical Reports Server (NTRS)

Doggett, R. V., Jr.; Hammond, C. E.

1975-01-01

The computer-controlled data-acquisition system recently installed for use with a transonic dynamics tunnel was described. This includes a discussion of the hardware/software features of the system. A subcritical response damping technique, called the combined randomdec/moving-block method, for use in windtunnel-model flutter testing, that has been implemented on the data-acquisition system, is described in some detail. Some results using the method are presented and the importance of using interactive graphics in applying the technique in near real time during wind-tunnel test operations is discussed.

A survey of the three-dimensional high Reynolds number transonic wind tunnel

NASA Technical Reports Server (NTRS)

Takashima, K.; Sawada, H.; Aoki, T.

1982-01-01

The facilities for aerodynamic testing of airplane models at transonic speeds and high Reynolds numbers are surveyed. The need for high Reynolds number testing is reviewed, using some experimental results. Some approaches to high Reynolds number testing such as the cryogenic wind tunnel, the induction driven wind tunnel, the Ludwieg tube, the Evans clean tunnel and the hydraulic driven wind tunnel are described. The level of development of high Reynolds number testing facilities in Japan is discussed.

Comparison of Flight Measured, Predicted and Wind Tunnel Measured Winglet Characteristics on a KC-135 Aircraft

NASA Technical Reports Server (NTRS)

Dodson, R. O., Jr.

1982-01-01

One of the objectives of the KC-135 Winglet Flight Research and Demonstration Program was to obtain experimental flight test data to verify the theoretical and wind tunnel winglet aerodynamic performance prediction methods. Good agreement between analytic, wind tunnel and flight test performance was obtained when the known differences between the tests and analyses were accounted for. The flight test measured fuel mileage improvements for a 0.78 Mach number was 3.1 percent at 8 x 10(5) pounds W/delta and 5.5 percent at 1.05 x 10(6) pounds W/delta. Correcting the flight measured data for surface pressure differences between wind tunnel and flight resulted in a fuel mileage improvement of 4.4 percent at 8 x 10(5) pounds W/delta and 7.2 percent at 1.05 x 10(6) pounds W/delta. The performance improvement obtained was within the wind tunnel test data obtained from two different wind tunnel models. The buffet boundary data obtained for the baseline configuration was in good agreement with previous established data. Buffet data for the 15 deg cant/-4 deg incidence configuration showed a slight improvement, while the 15 deg cant/-2 deg incidence and 0 deg cant/-4 deg incidence data showed a slight deterioration.

Evaluation of Simultaneous Multisine Excitation of the Joined Wing SensorCraft Aeroelastic Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Heeg, Jennifer; Morelli, Eugene A.

2011-01-01

Multiple mutually orthogonal signals comprise excitation data sets for aeroservoelastic system identification. A multisine signal is a sum of harmonic sinusoid components. A set of these signals is made orthogonal by distribution of the frequency content such that each signal contains unique frequencies. This research extends the range of application of an excitation method developed for stability and control flight testing to aeroservoelastic modeling from wind tunnel testing. Wind tunnel data for the Joined Wing SensorCraft model validates this method, demonstrating that these signals applied simultaneously reproduce the frequency response estimates achieved from one-at-a-time excitation.

Advances in Projection Moire Interferometry Development for Large Wind Tunnel Applications

NASA Technical Reports Server (NTRS)

Fleming, Gary A.; Soto, Hector L.; South, Bruce W.; Bartram, Scott M.

1999-01-01

An instrument development program aimed at using Projection Moire Interferometry (PMI) for acquiring model deformation measurements in large wind tunnels was begun at NASA Langley Research Center in 1996. Various improvements to the initial prototype PMI systems have been made throughout this development effort. This paper documents several of the most significant improvements to the optical hardware and image processing software, and addresses system implementation issues for large wind tunnel applications. The improvements have increased both measurement accuracy and instrument efficiency, promoting the routine use of PMI for model deformation measurements in production wind tunnel tests.

Rolls Royce Avon RA-14 Engine in the Altitude Wind Tunnel

NASA Image and Video Library

1956-03-21

A Rolls Royce Avon RA-14 engine was tested in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics’ (NACA) Lewis Flight Propulsion Laboratory. The Avon RA-14 engine was a 16-stage axial-flow compressor turbojet capable of producing 9,500 pounds of thrust. The Avon replaced Rolls Royce’s successful Nene engine in 1950 and remained in service until 1974. It was one of several British engines studied in the tunnel during the 1950s. The Altitude Wind Tunnel went through a series of modifications in 1951 to increase its capabilities. An annex was attached to the Exhauster Building to house three new Ingersoll-Rand compressors. The wooden blades on the tunnel’s 31-foot diameter fan were replaced, a pump house and exhaust cooler were constructed underneath the tunnel, and two new cells were added to the cooling tower. The modified wind tunnel continued to analyze jet engines in the 1950s, although the engines, like the RA-14 seen here, were much more powerful than those studied several years before. Lewis researchers studied the RA-14 turbojet engine in the Altitude Wind Tunnel for 11 months in 1956. The engine was mounted on a stand capable of gauging engine thrust, and the tunnel’s air was ducted to the engine through a venturi and bellmouth inlet, seen in this photograph. The initial studies established the engine’s performance characteristics with a fixed-area nozzle and its acceleration characteristics. The researchers also used the tunnel to investigate windmilling of the compressor blades, restarting at high altitudes, and the engine’s performance limits at altitude.

LACIS-T - A humid wind tunnel for investigating the Interactions between Cloud Microphysics and Turbulence

NASA Astrophysics Data System (ADS)

Voigtländer, Jens; Niedermeier, Dennis; Siebert, Holger; Shaw, Raymond; Schumacher, Jörg; Stratmann, Frank

2017-04-01

To improve the fundamental and quantitative understanding of the interactions between cloud microphysical and turbulent processes, the Leibniz Institute for Tropospheric Research (TROPOS) has built up a new humid wind tunnel (LACIS-T). LACIS-T allows for the investigation of cloud microphysical processes, such as cloud droplet activation and freezing, under-well defined thermodynamic and turbulent flow conditions. It therewith allows for the straight forward continuation, extension, and completion of the cloud microphysics related investigations carried out at the Leipzig Aerosol Cloud Interaction Simulator (LACIS) under laminar flow conditions. Characterization of the wind tunnel with respect to flow, thermodynamics, and droplet microphysics is carried out with probes mounted inside (pitot tube and hot-wire anemometer for mean velocity and fluctuations, Pt100 sensor for mean temperature, cold-wire sensor for temperature fluctuations is in progress, as well as a dew-point mirror for mean water vapor concentration, a Lyman-alpha sensor for water vapor fluctuations is in progress) the measurement section, and from outside with optical detection methods (a laser light sheet is available for cloud droplet visualization, a digital holography system for detection of cloud droplet size distributions will be installed for tests in February 2017), respectively. Computational fluid dynamics (CFD) simulations have been carried out for defining suitable experimental conditions and assisting the interpretation of the experimental data. In this work, LACIS-T, its fundamental operating principle, and first preliminary results from ongoing characterization efforts will be presented.

Experimental investigation of a newly designed supersonic wind tunnel

NASA Astrophysics Data System (ADS)

Wu, J.; Radespiel, R.

2015-06-01

The flow characteristics of the tandem nozzle supersonic wind tunnel at the Institute of Fluid Mechanics, Technische Universität Braunschweig, a are investigated. Conventional measurement techniques were utilized. The flow development is examined by pressure sensors installed at various streamwise positions. The temperature is measured in the storage tube and the settling chamber. The influence of flow treatment in the settling chamber on the flow quality is also studied. The flow quality of test section is evaluated by a 6-probe Pitot rake. The pressure fluctuations in the test section are studied by a sharp cone model. Eventually, good agreement between the measurements and numerical simulation of the tunnel design is achieved.

Aeroservoelastic wind-tunnel investigations using the Active Flexible Wing Model: Status and recent accomplishments

NASA Technical Reports Server (NTRS)

Noll, Thomas E.; Perry, Boyd, III; Tiffany, Sherwood H.; Cole, Stanley R.; Buttrill, Carey S.; Adams, William M., Jr.; Houck, Jacob A.; Srinathkumar, S.; Mukhopadhyay, Vivek; Pototzky, Anthony S.

1989-01-01

The status of the joint NASA/Rockwell Active Flexible Wing Wind-Tunnel Test Program is described. The objectives are to develop and validate the analysis, design, and test methodologies required to apply multifunction active control technology for improving aircraft performance and stability. Major tasks include designing digital multi-input/multi-output flutter-suppression and rolling-maneuver-load alleviation concepts for a flexible full-span wind-tunnel model, obtaining an experimental data base for the basic model and each control concept and providing comparisons between experimental and analytical results to validate the methodologies. The opportunity is provided to improve real-time simulation techniques and to gain practical experience with digital control law implementation procedures.

Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel (Postprint)

DTIC Science & Technology

2010-05-01

SCRAMJET WIND TUNNEL (POSTPRINT) 5a. CONTRACT NUMBER FA8650-10-D-5226-0002 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 62102F 6. AUTHOR(S...prototype scramjet engine as a wind tunnel . A sample holder was designed using combustion fluid dynamics results as inputs into structural models. The...Z39-18 Development of a Test to Evaluate Aerothermal Response of Materials to Hypersonic Flow Using a Scramjet Wind Tunnel Triplicane A

Advancing Test Capabilities at NASA Wind Tunnels

NASA Technical Reports Server (NTRS)

Bell, James

2015-01-01

NASA maintains twelve major wind tunnels at three field centers capable of providing flows at 0.1 M 10 and unit Reynolds numbers up to 45106m. The maintenance and enhancement of these facilities is handled through a unified management structure under NASAs Aeronautics and Evaluation and Test Capability (AETC) project. The AETC facilities are; the 11x11 transonic and 9x7 supersonic wind tunnels at NASA Ames; the 10x10 and 8x6 supersonic wind tunnels, 9x15 low speed tunnel, Icing Research Tunnel, and Propulsion Simulator Laboratory, all at NASA Glenn; and the National Transonic Facility, Transonic Dynamics Tunnel, LAL aerothermodynamics laboratory, 8 High Temperature Tunnel, and 14x22 low speed tunnel, all at NASA Langley. This presentation describes the primary AETC facilities and their current capabilities, as well as improvements which are planned over the next five years. These improvements fall into three categories. The first are operations and maintenance improvements designed to increase the efficiency and reliability of the wind tunnels. These include new (possibly composite) fan blades at several facilities, new temperature control systems, and new and much more capable facility data systems. The second category of improvements are facility capability advancements. These include significant improvements to optical access in wind tunnel test sections at Ames, improvements to test section acoustics at Glenn and Langley, the development of a Supercooled Large Droplet capability for icing research, and the development of an icing capability for large engine testing. The final category of improvements consists of test technology enhancements which provide value across multiple facilities. These include projects to increase balance accuracy, provide NIST-traceable calibration characterization for wind tunnels, and to advance optical instruments for Computational Fluid Dynamics (CFD) validation. Taken as a whole, these individual projects provide significant

Design and optimization of resistance wire electric heater for hypersonic wind tunnel

NASA Astrophysics Data System (ADS)

Rehman, Khurram; Malik, Afzaal M.; Khan, I. J.; Hassan, Jehangir

2012-06-01

The range of flow velocities of high speed wind tunnels varies from Mach 1.0 to hypersonic order. In order to achieve such high speed flows, a high expansion nozzle is employed in the converging-diverging section of wind tunnel nozzle. The air for flow is compressed and stored in pressure vessels at temperatures close to ambient conditions. The stored air is dried and has minimum amount of moisture level. However, when this air is expanded rapidly, its temperature drops significantly and liquefaction conditions can be encountered. Air at near room temperature will liquefy due to expansion cooling at a flow velocity of more than Mach 4.0 in a wind tunnel test section. Such liquefaction may not only be hazardous to the model under test and wind tunnel structure; it may also affect the test results. In order to avoid liquefaction of air, a pre-heater is employed in between the pressure vessel and the converging-diverging section of a wind tunnel. A number of techniques are being used for heating the flow in high speed wind tunnels. Some of these include the electric arc heating, pebble bed electric heating, pebble bed natural gas fired heater, hydrogen burner heater, and the laser heater mechanisms. The most common are the pebble bed storage type heaters, which are inefficient, contaminating and time consuming. A well designed electrically heating system can be efficient, clean and simple in operation, for accelerating the wind tunnel flow up to Mach 10. This paper presents CFD analysis of electric preheater for different configurations to optimize its design. This analysis has been done using ANSYS 12.1 FLUENT package while geometry and meshing was done in GAMBIT.

Ramjet Testing in the NACA's Altitude Wind Tunnel

NASA Image and Video Library

1946-02-21

A 20-inch diameter ramjet installed in the Altitude Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory. The Altitude Wind Tunnel was used in the 1940s to study early ramjet configurations. Ramjets provide a very simple source of propulsion. They are basically a tube which takes in high-velocity air, ignites it, and then expels the expanded airflow at a significantly higher velocity for thrust. Ramjets are extremely efficient and powerful but can only operate at high speeds. Therefore a turbojet or rocket was needed to launch the vehicle. This NACA-designed 20-inch diameter ramjet was installed in the Altitude Wind Tunnel in May 1945. The ramjet was mounted under a section of wing in the 20-foot diameter test section with conditioned airflow ducted directly to the engine. The mechanic in this photograph was installing instrumentation devices that led to the control room. NACA researchers investigated the ramjet’s overall performance at simulated altitudes up to 47,000 feet. Thrust measurements from these runs were studied in conjunction with drag data obtained during small-scale studies in the laboratory’s small supersonic tunnels. An afterburner was attached to the ramjet during the portions of the test program. The researchers found that an increase in altitude caused a reduction in the engine’s horsepower. They also determined the optimal configurations for the flameholders, which provided the engine’s ignition source.

Automatic control of cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Balakrishna, S.

1989-01-01

Inadequate Reynolds number similarity in testing of scaled models affects the quality of aerodynamic data from wind tunnels. This is due to scale effects of boundary-layer shock wave interaction which is likely to be severe at transonic speeds. The idea of operation of wind tunnels using test gas cooled to cryogenic temperatures has yielded a quantrum jump in the ability to realize full scale Reynolds number flow similarity in small transonic tunnels. In such tunnels, the basic flow control problem consists of obtaining and maintaining the desired test section flow parameters. Mach number, Reynolds number, and dynamic pressure are the three flow parameters that are usually required to be kept constant during the period of model aerodynamic data acquisition. The series of activity involved in modeling, control law development, mechanization of the control laws on a microcomputer, and the performance of a globally stable automatic control system for the 0.3-m Transonic Cryogenic Tunnel (TCT) are discussed. A lumped multi-variable nonlinear dynamic model of the cryogenic tunnel, generation of a set of linear control laws for small perturbation, and nonlinear control strategy for large set point changes including tunnel trajectory control are described. The details of mechanization of the control laws on a 16 bit microcomputer system, the software features, operator interface, the display and safety are discussed. The controller is shown to provide globally stable and reliable temperature control to + or - 0.2 K, pressure to + or - 0.07 psi and Mach number to + or - 0.002 of the set point value. This performance is obtained both during large set point commands as for a tunnel cooldown, and during aerodynamic data acquisition with intrusive activity like geometrical changes in the test section such as angle of attack changes, drag rake movements, wall adaptation and sidewall boundary-layer removal. Feasibility of the use of an automatic Reynolds number control mode with

The Variable Density Wind Tunnel of the National Advisory Committee for Aeronautics

NASA Technical Reports Server (NTRS)

Munk, Max M; Miller, Elton W

1926-01-01

This report contains an exact description of the new wind tunnel of the National Advisory Committee for Aeronautics. This is the first american type wind tunnel. It differs from ordinary wind tunnels by its being surrounded by a strong steel shell, 35 feet long and 15 feet in diameter. A compressor system is provided to fill this shell - and hence the entire wind tunnel - with air compressed to a density up to 25 times the ordinary atmospheric density. It is demonstrated in the report that the increase of the air density makes up for a corresponding decrease in the scale of the model. Hence such american type wind tunnel is free from scale effect. The report is illustrated by many drawings and photographs. All construction details are described, and many dimensions given. The method of conducting tests is also described and some preliminary results given in the report. So far, the tests have confirmed the chief feature of this wind tunnel - absence of scale effect.

Wind-tunnel tests on a 3-dimensional fixed-geometry scramjet inlet at M = 2.30 to 4.60

NASA Technical Reports Server (NTRS)

Mueller, J. N.; Trexler, C. A.; Souders, S. W.

1977-01-01

Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.

Aeroelastic characteristics of a rapid prototype multi-material wind tunnel model of a mechanically deployable aerodynamic decelerator

NASA Astrophysics Data System (ADS)

Raskin, Boris

Scaled wind tunnel models are necessary for the development of aircraft and spacecraft to simulate aerodynamic behavior. This allows for testing multiple iterations of a design before more expensive full-scale aircraft and spacecraft are built. However, the cost of building wind tunnel models can still be high because they normally require costly subtractive manufacturing processes, such as machining, which can be time consuming and laborious due to the complex surfaces of aerodynamic models. Rapid prototyping, commonly known as 3D printing, can be utilized to save on wind tunnel model manufacturing costs. A rapid prototype multi-material wind tunnel model was manufactured for this thesis to investigate the possibility of using PolyJet 3D printing to create a model that exhibits aeroelastic behavior. The model is of NASA's Adaptable Deployable entry and Placement (ADEPT) aerodynamic decelerator, used to decelerate a spacecraft during reentry into a planet's atmosphere. It is a 60° cone with a spherically blunted nose that consists of a 12 flexible panels supported by a rigid structure of nose, ribs, and rim. The novel rapid prototype multi-material model was instrumented and tested in two flow conditions. Quantitative comparisons were made of the average forces and dynamic forces on the model, demonstrating that the model matched expected behavior for average drag, but not Strouhal number, indicating that there was no aeroelastic behavior in this particular case. It was also noted that the dynamic properties (e.g., resonant frequency) associated with the mounting scheme are very important and may dominate the measured dynamic response.

Aeroelastic Analysis of a Flexible Wing Wind Tunnel Model with Variable Camber Continuous Trailing Edge Flap Design

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia

2015-01-01

This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope

Heating requirements and nonadiabatic surface effects for a model in the NTF (National Transonic Facility) cryogenic wind tunnel

DOE Office of Scientific and Technical Information (OSTI.GOV)

Macha, J.M.; Landrum, D.B.; Pare, L.A. III

1988-01-01

A theoretical study has been made of the severity of nonadiabatic surface conditions arising from internal heat sources within a model in a cryogenic wind tunnel. Local surface heating is recognized as having an effect on the development of the boundary layer, which can introduce changes in the flow about the model and affect the wind tunnel data. The geometry was based on the NTF Pathfinder I wind tunnel model. A finite element heat transfer computer code was developed and used to compute the steady state temperature distribution within the body of the model, from which the surface temperature distributionmore » was extracted. Particular three dimensional characteristics of the model were represented with various axisymmetric approximations of the geometry. This analysis identified regions on the surface of the model susceptible to surface heating and the magnitude of the respective surface temperatures. It was found that severe surface heating may occur in particular instances, but could be alleviated with adequate insulating material. The heat flux through the surface of the model was integrated to determine the net heat required to maintain the instrumentation cavity at the prescribed temperature. The influence of the nonadiabatic condition on boundary layer properties and on the validity of the wind tunnel simulation was also investigated. 20 refs., 12 figs.« less

9x15 Low Speed Wind Tunnel Improvements Update

NASA Technical Reports Server (NTRS)

Stephens, David

2017-01-01

The 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) at NASA Glenn Research Center was built in 1969 in the return leg of the 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT). The 9x15 LSWT was designed for performance testing of VSTOL aircraft models, but with the addition of the current acoustic treatment in 1986 the tunnel been used principally for acoustic and performance testing of aircraft propulsion systems. The present document describes an anticipated acoustic upgrade to be completed in 2018.

Design and calibration of the carousel wind tunnel

NASA Technical Reports Server (NTRS)

Leach, R. N.; Greeley, R.; Iversen, J.; White, B.; Marshall, J. R.

1986-01-01

In the study of planetary aeolian processes the effect of gravity is not readily modeled. Gravity appears in the equations of particle motion along with interparticle forces but the two terms are not separable. A wind tunnel that would permit variable gravity would allow separation of the forces and aid greatly in understanding planetary aeolian processes. The design Carousel Wind Tunnel (CWT) allows for a long flow distance in a small sized tunnel since the test section is a continuo us circuit and allows for a variable pseudo gravity. A prototype design was built and calibrated to gain some understanding of the characteristics of the design and the results presented.

Design and calibration of the carousel wind tunnel

NASA Technical Reports Server (NTRS)

Leach, R. N.; Greeley, Ronald; Iversen, James D.; White, Bruce R.; Marshall, John R.

1987-01-01

In the study of planetary aeolian processes the effect of gravity is not readily modeled. Gravity appears in the equations of particle motion along with interparticle forces but the two terms are not separable. A wind tunnel that would permit variable gravity would allow separation of the forces and aid greatly in understanding planetary aeolian processes. The design of the Carousel Wind Tunnel (CWT) allows for a long flow distance in a small sized tunnel since the test section is a continuous circuit and allows for a variable pseudo-gravity. A prototype design was built and calibrated to gain some understanding of the characteristics of the design and the results presented.

A vegetation modeling concept for Building and Environmental Aerodynamics wind tunnel tests and its application in pollutant dispersion studies.

PubMed

Gromke, Christof

2011-01-01

A new vegetation modeling concept for Building and Environmental Aerodynamics wind tunnel investigations was developed. The modeling concept is based on fluid dynamical similarity aspects and allows the small-scale modeling of various kinds of vegetation, e.g. field crops, shrubs, hedges, single trees and forest stands. The applicability of the modeling concept was validated in wind tunnel pollutant dispersion studies. Avenue trees in urban street canyons were modeled and their implications on traffic pollutant dispersion were investigated. The dispersion experiments proved the modeling concept to be practicable for wind tunnel studies and suggested to provide reliable concentration results. Unfavorable effects of trees on pollutant dispersion and natural ventilation in street canyons were revealed. Increased traffic pollutant concentrations were found in comparison to the tree-free reference case. Copyright © 2010 Elsevier Ltd. All rights reserved.

A tilting wind tunnel for fire behavior studies

Treesearch

David R. Weise

1994-01-01

The combined effects of wind velocity and slope on wildland fire behavior can be studied in the laboratory using a tilting wind tunnel. The tilting wind tunnel requires a commercially available fan to induce wind and can be positioned to simulate heading and backing fires spreading up and down slope. The tunnel is portable and can be disassembled for transport using a...

A New Method of Testing in Wind Tunnels

NASA Technical Reports Server (NTRS)

Margoulis, W

1921-01-01

Now, in existing wind tunnels, using a horsepower of 100 to 300, the models are generally made to a 1/10 scale and the speed is appreciably lower than the speeds currently attained by airplanes. The Reynolds number realized is thus 15 to 25 times smaller than that reached by airplanes in free flight, while the ratio of speed to the velocity of sound is between a third and three quarters of the true ratio. The necessary increases in either the diameter of the wind tunnel or the velocity of the airstream are too costly. However, the author shows that it is possible to have wind tunnels in which the Reynolds number will be greater than that now obtained by airplanes, and in which the ratio of the velocity to the velocity of sound will also be greater than that realized in practice, by employing a gas other than air, at a pressure and temperature different from those of the surrounding atmosphere. The gas is carbonic acid, a gas having a low coefficient of viscosity, high density, and a low ratio of specific heat. The positive results of using carbonic acid in wind tunnel tests are given.

Jet-Boundary Corrections for Reflection-Plane Models in Rectangular Wind Tunnels

DTIC Science & Technology

1943-01-01

clock- wiso) located at a distance d above the tunnel center line and at distances equal to VIand —vI from the reflection wall. Tho single trailing vortex...neglected. The angle-.xkmnge is usually small, les than %O. M & fw-refleciicur plane models in 7-by 10-foof cIosedrectangular wind fu-meIs 04

Wind tunnel pressurization and recovery system

NASA Technical Reports Server (NTRS)

Pejack, Edwin R.; Meick, Joseph; Ahmad, Adnan; Lateh, Nordin; Sadeq, Omar

1988-01-01

The high density, low toxicity characteristics of refrigerant-12 (dichlorofluoromethane) make it an ideal gas for wind tunnel testing. Present limitations on R-12 emissions, set to slow the rate of ozone deterioration, pose a difficult problem in recovery and handling of large quantities of R-12. This preliminary design is a possible solution to the problem of R-12 handling in wind tunnel testing. The design incorporates cold temperature condensation with secondary purification of the R-12/air mixture by adsorption. Also discussed is the use of Freon-22 as a suitable refrigerant for the 12 foot wind tunnel.

Photogrammetry Applied to Wind Tunnel Testing

NASA Technical Reports Server (NTRS)

Liu, Tian-Shu; Cattafesta, L. N., III; Radeztsky, R. H.; Burner, A. W.

2000-01-01

In image-based measurements, quantitative image data must be mapped to three-dimensional object space. Analytical photogrammetric methods, which may be used to accomplish this task, are discussed from the viewpoint of experimental fluid dynamicists. The Direct Linear Transformation (DLT) for camera calibration, used in pressure sensitive paint, is summarized. An optimization method for camera calibration is developed that can be used to determine the camera calibration parameters, including those describing lens distortion, from a single image. Combined with the DLT method, this method allows a rapid and comprehensive in-situ camera calibration and therefore is particularly useful for quantitative flow visualization and other measurements such as model attitude and deformation in production wind tunnels. The paper also includes a brief description of typical photogrammetric applications to temperature- and pressure-sensitive paint measurements and model deformation measurements in wind tunnels.

Digital-flutter-suppression-system investigations for the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Mukhopadhyay, Vivek; Hoadley, Sherwood Tiffany; Cole, Stanley R.; Buttrill, Carey S.

1990-01-01

Active flutter suppression control laws were designed, implemented, and tested on an aeroelastically-scaled wind-tunnel model in the NASA Langley Transonic Dynamics Tunnel. One of the control laws was successful in stabilizing the model while the dynamic pressure was increased to 24 percent greater than the measured open-loop flutter boundary. Other accomplishments included the design, implementation, and successful operation of a one-of-a-kind digital controller, the design and use of two simulation methods to support the project, and the development and successful use of a methodology for online controller performance evaluation.

TWINTAN: A program for transonic wall interference assessment in two-dimensional wind tunnels

NASA Technical Reports Server (NTRS)

Kemp, W. B., Jr.

1980-01-01

A method for assessing the wall interference in transonic two dimensional wind tunnel test was developed and implemented in a computer program. The method involves three successive solutions of the transonic small disturbance potential equation to define the wind tunnel flow, the perturbation attriburable to the model, and the equivalent free air flow around the model. Input includes pressure distributions on the model and along the top and bottom tunnel walls which are used as boundary conditions for the wind tunnel flow. The wall induced perturbation fields is determined as the difference between the perturbation in the tunnel flow solution and the perturbation attributable to the model. The methodology used in the program is described and detailed descriptions of the computer program input and output are presented. Input and output for a sample case are given.

Case Studies for the Statistical Design of Experiments Applied to Powered Rotor Wind Tunnel Tests

NASA Technical Reports Server (NTRS)

Overmeyer, Austin D.; Tanner, Philip E.; Martin, Preston B.; Commo, Sean A.

2015-01-01

The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

Fluorescence Imaging and Streamline Visualization of Hypersonic Flow over Rapid Prototype Wind-Tunnel Models

NASA Technical Reports Server (NTRS)

Danehy, Paul M.; Alderfer, David W.; Inman, Jennifer A.; Berger, Karen T.; Buck, Gregory M.; Schwartz, Richard J.

2008-01-01

Reentry models for use in hypersonic wind tunnel tests were fabricated using a stereolithography apparatus. These models were produced in one day or less, which is a significant time savings compared to the manufacture of ceramic or metal models. The models were tested in the NASA Langley Research Center 31-Inch Mach 10 Air Tunnel. Only a few of the models survived repeated tests in the tunnel, and several failure modes of the models were identified. Planar laser-induced fluorescence (PLIF) of nitric oxide (NO) was used to visualize the flowfields in the wakes of these models. Pure NO was either seeded through tubes plumbed into the model or via a tube attached to the strut holding the model, which provided localized addition of NO into the model s wake through a porous metal cylinder attached to the end of the tube. Models included several 2- inch diameter Inflatable Reentry Vehicle Experiment (IRVE) models and 5-inch diameter Crew Exploration Vehicle (CEV) models. Various model configurations and NO seeding methods were used, including a new streamwise visualization method based on PLIF. Virtual Diagnostics Interface (ViDI) technology, developed at NASA Langley Research Center, was used to visualize the data sets in post processing. The use of calibration "dotcards" was investigated to correct for camera perspective and lens distortions in the PLIF images.

Aerodynamic characteristics of the modified 40- by 80-foot wind tunnel as measured in a 1/50th-scale model

NASA Technical Reports Server (NTRS)

Smith, Brian E.; Naumowicz, Tim

1987-01-01

The aerodynamic characteristics of the 40- by 80-Foot Wind Tunnel at Ames Research Center were measured by using a 1/50th-scale facility. The model was configured to closely simulate the features of the full-scale facility when it became operational in 1986. The items measured include the aerodynamic effects due to changes in the total-pressure-loss characteristics of the intake and exhaust openings of the air-exchange system, total-pressure distributions in the flow field at locations around the wind tunnel circuit, the locations of the maximum total-pressure contours, and the aerodynamic changes caused by the installation of the acoustic barrier in the southwest corner of the wind tunnel. The model tests reveal the changes in the aerodynamic performance of the 1986 version of the 40- by 80-Foot Wind Tunnel compared with the performance of the 1982 configuration.

Wind tunnel wall interference

NASA Technical Reports Server (NTRS)

Newman, Perry A.; Mineck, Raymond E.; Barnwell, Richard W.; Kemp, William B., Jr.

1986-01-01

About a decade ago, interest in alleviating wind tunnel wall interference was renewed by advances in computational aerodynamics, concepts of adaptive test section walls, and plans for high Reynolds number transonic test facilities. Selection of NASA Langley cryogenic concept for the National Transonic Facility (NTF) tended to focus the renewed wall interference efforts. A brief overview and current status of some Langley sponsored transonic wind tunnel wall interference research are presented. Included are continuing efforts in basic wall flow studies, wall interference assessment/correction procedures, and adaptive wall technology.

A century of wind tunnels since Eiffel

NASA Astrophysics Data System (ADS)

Chanetz, Bruno

2017-08-01

Fly higher, faster, preserve the life of test pilots and passengers, many challenges faced by man since the dawn of the twentieth century, with aviation pioneers. Contemporary of the first aerial exploits, wind tunnels, artificially recreating conditions encountered during the flight, have powerfully contributed to the progress of aeronautics. But the use of wind tunnels is not limited to aviation. The research for better performance, coupled with concern for energy saving, encourages manufacturers of ground vehicles to perform aerodynamic tests. Buildings and bridge structures are also concerned. This article deals principally with the wind tunnels built at ONERA during the last century. Somme wind tunnels outside ONERA, even outside France, are also evocated when their characteristics do not exist at ONERA.

Scaling between Wind Tunnels-Results Accuracy in Two-Dimensional Testing

NASA Astrophysics Data System (ADS)

Rasuo, Bosko

The establishment of exact two-dimensional flow conditions in wind tunnels is a very difficult problem. This has been evident for wind tunnels of all types and scales. In this paper, the principal factors that influence the accuracy of two-dimensional wind tunnel test results are analyzed. The influences of the Reynolds number, Mach number and wall interference with reference to solid and flow blockage (blockage of wake) as well as the influence of side-wall boundary layer control are analyzed. Interesting results are brought to light regarding the Reynolds number effects of the test model versus the Reynolds number effects of the facility in subsonic and transonic flow.

Qualification of the T2 wind tunnel in cryogenic operation. B: Flow fluctuations, particle detection and qualification

NASA Technical Reports Server (NTRS)

Dor, J. B.; Mignosi, A.; Plazanet, M.

1984-01-01

This report presents part of the tests for verification of the T2 transonic induction wind tunnel in cryogenic operation. The first part of the results presented concerns fluctuations in pressure and temperature at ambient temperature and in cryogenic regulation. The second part presents the condensation phenomena which could be observed in the cryogenic flow by means of an optical particle detection system in the test section.

Wind Tunnel Database Development using Modern Experiment Design and Multivariate Orthogonal Functions

NASA Technical Reports Server (NTRS)

Morelli, Eugene A.; DeLoach, Richard

2003-01-01

A wind tunnel experiment for characterizing the aerodynamic and propulsion forces and moments acting on a research model airplane is described. The model airplane called the Free-flying Airplane for Sub-scale Experimental Research (FASER), is a modified off-the-shelf radio-controlled model airplane, with 7 ft wingspan, a tractor propeller driven by an electric motor, and aerobatic capability. FASER was tested in the NASA Langley 12-foot Low-Speed Wind Tunnel, using a combination of traditional sweeps and modern experiment design. Power level was included as an independent variable in the wind tunnel test, to allow characterization of power effects on aerodynamic forces and moments. A modeling technique that employs multivariate orthogonal functions was used to develop accurate analytic models for the aerodynamic and propulsion force and moment coefficient dependencies from the wind tunnel data. Efficient methods for generating orthogonal modeling functions, expanding the orthogonal modeling functions in terms of ordinary polynomial functions, and analytical orthogonal blocking were developed and discussed. The resulting models comprise a set of smooth, differentiable functions for the non-dimensional aerodynamic force and moment coefficients in terms of ordinary polynomials in the independent variables, suitable for nonlinear aircraft simulation.

Supercritical tests of a self-optimizing, variable-Camber wind tunnel model

NASA Technical Reports Server (NTRS)

Levinsky, E. S.; Palko, R. L.

1979-01-01

A testing procedure was used in a 16-foot Transonic Propulsion Wind Tunnel which leads to optimum wing airfoil sections without stopping the tunnel for model changes. Being experimental, the optimum shapes obtained incorporate various three-dimensional and nonlinear viscous and transonic effects not included in analytical optimization methods. The method is a closed-loop, computer-controlled, interactive procedure and employs a Self-Optimizing Flexible Technology wing semispan model that conformally adapts the airfoil section at two spanwise control stations to maximize or minimize various prescribed merit functions subject to both equality and inequality constraints. The model, which employed twelve independent hydraulic actuator systems and flexible skins, was also used for conventional testing. Although six of seven optimizations attempted were at least partially convergent, further improvements in model skin smoothness and hydraulic reliability are required to make the technique fully operational.

The Design of Low-Turbulence Wind Tunnels

NASA Technical Reports Server (NTRS)

Dryden, Hugh L; Abbott, Ira H

1949-01-01

Within the past 10 years there have been placed in operation in the United States four low-turbulence wind tunnels of moderate cross-sectional area and speed, one at the National Bureau of Standards, two at the NACA Langley Laboratory, and one at the NACA Ames Laboratory. This paper reviews briefly the state of knowledge and those features which make possible the attainment of low turbulence in wind tunnels. Specific applications to two wind tunnels are described.

Detection of boundary-layer transitions in wind tunnels

NASA Technical Reports Server (NTRS)

Wood, W. R.; Somers, D. M.

1978-01-01

Accelerometer replaces stethoscope in technique for detection of laminar-to-turbulent boundary-layer transitions on wind-tunnel models. Technique allows measurements above or below atmospheric pressure because human operator is not required within tunnel. Data may be taken from accelerometer, and pressure transducer simultaneously, and delivered to systems for analysis.

Finite element analysis of high aspect ratio wind tunnel wing model: A parametric study

NASA Astrophysics Data System (ADS)

Rosly, N. A.; Harmin, M. Y.

2017-12-01

Procedure for designing the wind tunnel model of a high aspect ratio (HAR) wing containing geometric nonlinearities is described in this paper. The design process begins with identification of basic features of the HAR wing as well as its design constraints. This enables the design space to be narrowed down and consequently, brings ease of convergence towards the design solution. Parametric studies in terms of the spar thickness, the span length and the store diameter are performed using finite element analysis for both undeformed and deformed cases, which respectively demonstrate the linear and nonlinear conditions. Two main criteria are accounted for in the selection of the wing design: the static deflections due to gravitational loading should be within the allowable margin of the size of the wind tunnel test section and the flutter speed of the wing should be much below the maximum speed of the wind tunnel. The findings show that the wing experiences a stiffness hardening effect under the nonlinear static solution and the presence of the store enables significant reduction in linear flutter speed.

The Brothers Were Wright - An Abridged History of Wind Tunnel Testing at Ames Research Center

NASA Technical Reports Server (NTRS)

Buchholz, Steve

2017-01-01

The Wright Brothers used wind tunnel data to refine their design for the first successful airplane back in 1903. Today, wind tunnels are still in use all over the world gathering data to improve the design of cars, trucks, airplanes, missiles and spacecraft. Ames Research Center is home to many wind tunnels, including the Unitary Plan Wind Tunnel complex. Built in the early 1950s, it is one of the premiere transonic and supersonic testing facilities in the country. Every manned spacecraft has been tested in the wind tunnels at Ames. This is a testing history from past to present.

Aeroservoelastic Testing of Free Flying Wind Tunnel Models Part 1: A Sidewall Supported Semispan Model Tested for Gust Load Alleviation and Flutter Suppression

NASA Technical Reports Server (NTRS)

Scott, Robert C.; Vetter, Travis K.; Penning, Kevin B.; Coulson, David A.; Heeg, Jennifer.

2013-01-01

of a two part document. Part 2 is titled: "Aeroservoelastic Testing of Free Flying Wind Tunnel Models, Part 2: A Centerline Supported Fullspan Model Tested for Gust Load Alleviation." A team comprised of the Air Force Research Laboratory (AFRL), Northrop Grumman, Lockheed Martin, and the NASA Langley Research Center conducted three aeroservoelastic wind tunnel tests in the Transonic Dynamics Tunnel to demonstrate active control technologies relevant to large, flexible vehicles. In the first of these three tests, a semispan, aeroelastically scaled, wind tunnel model of a flying wing SensorCraft vehicle was mounted to a force balance to demonstrate gust load alleviation. In the second and third tests, the same wing was mated to a new, multi-degree of freedom, sidewall mount. This mount allowed the half-span model to translate vertically and pitch at the wing root, allowing better simulation of the full span vehicle's rigid body modes. Gust load alleviation (GLA) and Body freedom flutter (BFF) suppression were successfully demonstrated. The rigid body degrees-of-freedom required that the model be flown in the wind tunnel using an active control system. This risky mode of testing necessitated that a model arrestment system be integrated into the new mount. The safe and successful completion of these free flying tests required the development and integration of custom hardware and software. This paper describes the many systems, software, and procedures that were developed as part of this effort.

Second-Generation Large Civil Tiltrotor 7- by 10-Foot Wind Tunnel Test Data Report

NASA Technical Reports Server (NTRS)

Theodore, Colin R.; Russell, Carl R.; Willink, Gina C.; Pete, Ashley E.; Adibi, Sierra A.; Ewert, Adam; Theuns, Lieselotte; Beierle, Connor

2016-01-01

An approximately 6-percent scale model of the NASA Second-Generation Large Civil Tiltrotor (LCTR2) Aircraft was tested in the U.S. Army 7- by 10-Foot Wind Tunnel at NASA Ames Research Center January 4 to April 19, 2012, and September 18 to November 1, 2013. The full model was tested, along with modified versions in order to determine the effects of the wing tip extensions and nacelles; the wing was also tested separately in the various configurations. In both cases, the wing and nacelles used were adopted from the U.S. Army High Efficiency Tilt Rotor (HETR) aircraft, in order to limit the cost of the experiment. The full airframe was tested in high-speed cruise and low-speed hover flight conditions, while the wing was tested only in cruise conditions, with Reynolds numbers ranging from 0 to 1.4 million. In all cases, the external scale system of the wind tunnel was used to collect data. Both models were mounted to the scale using two support struts attached underneath the wing; the full airframe model also used a third strut attached at the tail. The collected data provides insight into the performance of the preliminary design of the LCTR2 and will be used for computational fluid dynamics (CFD) validation and the development of flight dynamics simulation models.

A Wind-Tunnel Investigation of Tilt-Rotor Gust Alleviation Systems

NASA Technical Reports Server (NTRS)

Ham, N. D.; Whitaker, H. P.

1978-01-01

The alleviation of the effects of gusts on tilt rotor aircraft by means of active control systems was investigated. The gust generator, the derivation of the equations of motion of the rotor wing combination, the correlation of these equations with the results of wind tunnel model tests, the use of the equations to design various gust alleviating active control systems, and the testing and evaluation of these control systems by means of wind tunnel model tests were developed.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

NASA Technical Reports Server (NTRS)

Rivers, Melissa; Quest, Juergen; Rudnik, Ralf

2015-01-01

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.

Mercury Capsule Retrorocket Test in the Altitude Wind Tunnel

NASA Image and Video Library

1960-09-21

A mechanic at the National Aeronautics and Space Administration (NASA) Lewis Research Center prepares the inverted base of a Mercury capsule for a test of its posigrade retrorockets inside the Altitude Wind Tunnel. In October 1959 NASA’s Space Task Group allocated several Project Mercury assignments to Lewis. The Altitude Wind Tunnel was modified to test the Atlas separation system, study the escape tower rocket plume, train astronauts to bring a spinning capsule under control, and calibrate the capsule’s retrorockets. The turning vanes, makeup air pipes, and cooling coils were removed from the wide western end of the tunnel to create a 51-foot diameter test chamber. The Mercury capsule had a six-rocket retro-package affixed to the bottom of the capsule. Three of these were posigrade rockets used to separate the capsule from the booster and three were retrograde rockets used to slow the capsule for reentry into the earth’s atmosphere. Performance of the retrorockets was vital since there was no backup system. Qualification tests of the retrorockets began in April 1960 on a retrograde thrust stand inside the southwest corner of the Altitude Wind Tunnel. These studies showed that a previous issue concerning the delayed ignition of the propellant had been resolved. Follow-up test runs verified reliability of the igniter’s attachment to the propellant. In addition, the capsule’s retrorockets were calibrated so they would not alter the capsule’s attitude when fired.

Highly cited articles in wind tunnel-related research: a bibliometric analysis.

PubMed

Mo, Ziwei; Fu, Hui-Zhen; Ho, Yuh-Shan

2018-06-01

Wind tunnels have been widely employed in aerodynamic research. To characterize the high impact research, a bibliometric analysis was conducted on highly cited articles related to wind tunnel based on the Science Citation Index Expanded (SCI-EXPANDED) database from 1900 to 2014. Articles with at least 100 citations from the Web of Science Core Collection were selected and analyzed in terms of publication years, authors, institutions, countries/territories, journals, Web of Science categories, and citation life cycles. The results show that a total of 77 highly cited articles in 37 journals were published between 1959 and 2008. Journal of Fluid Mechanics published the most of highly cited articles. The USA was the most productive country and most frequent partner of internationally collaboration. The prolific institutions were mainly located in the USA and UK. The authors who were both first author and corresponding author published 88% of the articles. The Y index was also deployed to evaluate the publication characteristics of authors. Moreover, the articles with high citations in both history and the latest year with their citation life cycles were examined to provide insights for high impact research. The highly cited articles were almost earliest wind tunnel experimental data and reports on their own research specialty, and thus attracted high citations. It was revealed that classic works of wind tunnel research was frequently occurred in 1990s but much less in 2000s, probably due to the development of numerical models of computational fluid dynamic (CFD) in recent decades.

Static and Wind Tunnel Aero-Performance Tests of NASA AST Separate Flow Nozzle Noise Reduction Configurations

NASA Technical Reports Server (NTRS)

Mikkelsen, Kevin L.; McDonald, Timothy J.; Saiyed, Naseem (Technical Monitor)

2001-01-01

This report presents the results of cold flow model tests to determine the static and wind tunnel performance of several NASA AST separate flow nozzle noise reduction configurations. The tests were conducted by Aero Systems Engineering, Inc., for NASA Glenn Research Center. The tests were performed in the Channels 14 and 6 static thrust stands and the Channel 10 transonic wind tunnel at the FluiDyne Aerodynamics Laboratory in Plymouth, Minnesota. Facility checkout tests were made using standard ASME long-radius metering nozzles. These tests demonstrated facility data accuracy at flow conditions similar to the model tests. Channel 14 static tests reported here consisted of 21 ASME nozzle facility checkout tests and 57 static model performance tests (including 22 at no charge). Fan nozzle pressure ratio varied from 1.4 to 2.0, and fan to core total pressure ratio varied from 1.0 to 1.19. Core to fan total temperature ratio was 1.0. Channel 10 wind tunnel tests consisted of 15 tests at Mach number 0.28 and 31 tests at Mach 0.8. The sting was checked out statically in Channel 6 before the wind tunnel tests. In the Channel 6 facility, 12 ASME nozzle data points were taken and 7 model data points were taken. In the wind tunnel, fan nozzle pressure ratio varied from 1.73 to 2.8, and fan to core total pressure ratio varied from 1.0 to 1.19. Core to fan total temperature ratio was 1.0. Test results include thrust coefficients, thrust vector angle, core and fan nozzle discharge coefficients, total pressure and temperature charging station profiles, and boat-tail static pressure distributions in the wind tunnel.

The Role of Hierarchy in Response Surface Modeling of Wind Tunnel Data

NASA Technical Reports Server (NTRS)

DeLoach, Richard

2010-01-01

This paper is intended as a tutorial introduction to certain aspects of response surface modeling, for the experimentalist who has started to explore these methods as a means of improving productivity and quality in wind tunnel testing and other aerospace applications. A brief review of the productivity advantages of response surface modeling in aerospace research is followed by a description of the advantages of a common coding scheme that scales and centers independent variables. The benefits of model term reduction are reviewed. A constraint on model term reduction with coded factors is described in some detail, which requires such models to be well-formulated, or hierarchical. Examples illustrate the consequences of ignoring this constraint. The implication for automated regression model reduction procedures is discussed, and some opinions formed from the author s experience are offered on coding, model reduction, and hierarchy.

Residual interference and wind tunnel wall adaption

NASA Technical Reports Server (NTRS)

Mokry, Miroslav

1989-01-01

Measured flow variables near the test section boundaries, used to guide adjustments of the walls in adaptive wind tunnels, can also be used to quantify the residual interference. Because of a finite number of wall control devices (jacks, plenum compartments), the finite test section length, and the approximation character of adaptation algorithms, the unconfined flow conditions are not expected to be precisely attained even in the fully adapted stage. The procedures for the evaluation of residual wall interference are essentially the same as those used for assessing the correction in conventional, non-adaptive wind tunnels. Depending upon the number of flow variables utilized, one can speak of one- or two-variable methods; in two dimensions also of Schwarz- or Cauchy-type methods. The one-variable methods use the measured static pressure and normal velocity at the test section boundary, but do not require any model representation. This is clearly of an advantage for adaptive wall test section, which are often relatively small with respect to the test model, and for the variety of complex flows commonly encountered in wind tunnel testing. For test sections with flexible walls the normal component of velocity is given by the shape of the wall, adjusted for the displacement effect of its boundary layer. For ventilated test section walls it has to be measured by the Calspan pipes, laser Doppler velocimetry, or other appropriate techniques. The interface discontinuity method, also described, is a genuine residual interference assessment technique. It is specific to adaptive wall wind tunnels, where the computation results for the fictitious flow in the exterior of the test section are provided.

A position and attitude vision measurement system for wind tunnel slender model

NASA Astrophysics Data System (ADS)

Cheng, Lei; Yang, Yinong; Xue, Bindang; Zhou, Fugen; Bai, Xiangzhi

2014-11-01

A position and attitude vision measurement system for drop test slender model in wind tunnel is designed and developed. The system used two high speed cameras, one is put to the side of the model and another is put to the position where the camera can look up the model. Simple symbols are set on the model. The main idea of the system is based on image matching technique between the 3D-digital model projection image and the image captured by the camera. At first, we evaluate the pitch angles, the roll angles and the position of the centroid of a model through recognizing symbols in the images captured by the side camera. And then, based on the evaluated attitude info, giving a series of yaw angles, a series of projection images of the 3D-digital model are obtained. Finally, these projection images are matched with the image which captured by the looking up camera, and the best match's projection images corresponds to the yaw angle is the very yaw angle of the model. Simulation experiments are conducted and the results show that the maximal error of attitude measurement is less than 0.05°, which can meet the demand of test in wind tunnel.

Test Data Report, Low-Speed Wind Tunnel Drag Test of a 2/5 Scale Lockheed AH-56 Cheyenne Door-Hinge Hub

DTIC Science & Technology

2016-07-01

the U.S. Army 7– by 10–foot Wind Tunnel located at NASA Ames Research Center in Moffett Field, CA. The purpose of the test was to quantify the drag...drag test of a non-rotating 2/5 scale Lockheed AH-56 Cheyenne main rotor hub in the U.S. Army 7– by 10–foot Wind Tunnel located at NASA Ames Research...the U.S. Army 7– by 10–foot wind tunnel at NASA Ames Research Center 5 2.3 Perspective view of the hub mounted with major dimensions and model

Atmospheric Probe Model: Construction and Wind Tunnel Tests

NASA Technical Reports Server (NTRS)

Vogel, Jerald M.

1998-01-01

The material contained in this document represents a summary of the results of a low speed wind tunnel test program to determine the performance of an atmospheric probe at low speed. The probe configuration tested consists of a 2/3 scale model constructed from a combination of hard maple wood and aluminum stock. The model design includes approximately 130 surface static pressure taps. Additional hardware incorporated in the baseline model provides a mechanism for simulating external and internal trailing edge split flaps for probe flow control. Test matrix parameters include probe side slip angle, external/internal split flap deflection angle, and trip strip applications. Test output database includes surface pressure distributions on both inner and outer annular wings and probe center line velocity distributions from forward probe to aft probe locations.

A method for the modelling of porous and solid wind tunnel walls in computational fluid dynamics codes

NASA Technical Reports Server (NTRS)

Beutner, Thomas John

1993-01-01

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

Optimized aerodynamic design process for subsonic transport wing fitted with winglets. [wind tunnel model

NASA Technical Reports Server (NTRS)

Kuhlman, J. M.

1979-01-01

The aerodynamic design of a wind-tunnel model of a wing representative of that of a subsonic jet transport aircraft, fitted with winglets, was performed using two recently developed optimal wing-design computer programs. Both potential flow codes use a vortex lattice representation of the near-field of the aerodynamic surfaces for determination of the required mean camber surfaces for minimum induced drag, and both codes use far-field induced drag minimization procedures to obtain the required spanloads. One code uses a discrete vortex wake model for this far-field drag computation, while the second uses a 2-D advanced panel wake model. Wing camber shapes for the two codes are very similar, but the resulting winglet camber shapes differ widely. Design techniques and considerations for these two wind-tunnel models are detailed, including a description of the necessary modifications of the design geometry to format it for use by a numerically controlled machine for the actual model construction.

Gemini Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1962-09-21

A researcher at the National Aeronautics and Space Administration (NASA) Lewis Research Center examines a small-scale model of the Gemini capsule in the 10- by 10-Foot Supersonic Wind Tunnel test section. Gemini was added to NASA’s manned space program after its predecessor, Mercury, and its antecedent, Apollo, were already established. Gemini was a transitional mission designed provide the astronauts with practice docking with other spacecraft and withstanding durations in space up to two weeks. The program was officially announced on December 7, 1961, but planning began in mid-1959. It was named Gemini after the zodiac twins because of the spacecraft’s two passenger capacity. The Gemini Program was the first program to start at the new Manned Spacecraft Center in Houston, now the Johnson Space Center. Unlike Mercury and Apollo, Lewis had very little involvement with the Gemini Program. This model was tested in the 10- by 10 tunnel for several weeks in September 1962. Lewis began managing the Agena second-stage rocket program shortly after this photograph was taken. Agenas were used to launch a variety of spacecraft and satellites in the 1960s. They were also used on several Gemini missions to provide targets for the astronauts to practice their rendezvous maneuvers. Gemini had two unmanned and ten manned flights in 1965 and 1966. These yielded the first spacewalks, long-duration space missions, first onboard computer, docking with a second spacecraft, and rendezvous maneuvers.

The Denis-gruson Six-component Wind-tunnel Balance

NASA Technical Reports Server (NTRS)

1935-01-01

The 6.C.1 balance is the first fully automatic balance assuring a continuous and simultaneous record of the aerodynamic characteristics of an airfoil in a wind tunnel. Because of the rapidity of the measurements a complete polar (six components) requires only about three minutes of wind, that is to say, of motive power, which is of interest for wind tunnels with high efficiency factors and may lead to the economical design of large size wind tunnels.

Materials and construction techniques for cryogenic wind tunnel facilities for instruction/research use

NASA Technical Reports Server (NTRS)

Morse, S. F.; Roper, A. T.

1975-01-01

The results of the cryogenic wind tunnel program conducted at NASA Langley Research Center are presented to provide a starting point for the design of an instructional/research wind tunnel facility. The advantages of the cryogenic concept are discussed, and operating envelopes for a representative facility are presented to indicate the range and mode of operation. Special attention is given to the design, construction and materials problems peculiar to cryogenic wind tunnels. The control system for operation of a cryogenic tunnel is considered, and a portion of a linearized mathematical model is developed for determining the tunnel dynamic characteristics.

A wind-tunnel investigation of wind-turbine wakes in different yawed and loading conditions

NASA Astrophysics Data System (ADS)

Bastankhah, Majid; Porté-Agel, Fernando

2015-04-01

Wind-turbine wakes have negative effects on wind-farm performance. They are associated with: (a) the velocity deficit, which reduces the generated power of downwind turbines; and (b) the turbulence level, which increases the fatigue loads on downwind turbines. Controlling the yaw angle of turbines can potentially improve the performance of wind farms by deflecting the wake away from downwind turbines. However, except for few studies, wakes of yawed turbines still suffer from the lack of systematic research. To fill this research gap, we performed wind-tunnel experiments in the recirculating boundary-layer wind tunnel at the WIRE Laboratory of EPFL to better understand the wakes of yawed turbines. High-resolution stereoscopic particle image-velocimetry (S-PIV) was used to measure three velocity components in a horizontal plane located downwind of a horizontal-axis, three-blade model turbine. A servo-controller was connected to the DC generator of the turbine, which allowed us to apply different loadings. The power and thrust coefficients of the turbine were also measured for each case. These power and thrust measurements together with the highly-resolved flow measurements enabled us to study different wake characteristics such as the energy entrainment from the outer flow into the wake, the wake deflection and the helicoidal tip vortices for yawed turbines.

Investigation of correlation between full-scale and fifth-scale wind tunnel tests of a Bell helicopter Textron Model 222

NASA Technical Reports Server (NTRS)

Squires, P. K.

1982-01-01

Reasons for lack of correlation between data from a fifth-scale wind tunnel test of the Bell Helicopter Textron Model 222 and a full-scale test of the model 222 prototype in the NASA Ames 40-by 80-foot tunnel were investigated. This investigation centered around a carefully designed fifth-scale wind tunnel test of an accurately contoured model of the Model 222 prototype mounted on a replica of the full-scale mounting system. The improvement in correlation for drag characteristics in pitch and yaw with the fifth-scale model mounted on the replica system is shown. Interference between the model and mounting system was identified as a significant effect and was concluded to be a primary cause of the lack of correlation in the earlier tests.

Ceramic and coating applications in the hostile environment of a high temperature hypersonic wind tunnel. [Langley 8-foot high temperature structures tunnel

NASA Technical Reports Server (NTRS)

Puster, R. L.; Karns, J. R.; Vasquez, P.; Kelliher, W. C.

1981-01-01

A Mach 7, blowdown wind tunnel was used to investigate aerothermal structural phenomena on large to full scale high speed vehicle components. The high energy test medium, which provided a true temperature simulation of hypersonic flow at 24 to 40 km altitude, was generated by the combustion of methane with air at high pressures. Since the wind tunnel, as well as the models, must be protected from thermally induced damage, ceramics and coatings were used extensively. Coatings were used both to protect various wind tunnel components and to improve the quality of the test stream. Planned modifications for the wind tunnel included more extensive use of ceramics in order to minimize the number of active cooling systems and thus minimize the inherent operational unreliability and cost that accompanies such systems. Use of nonintrusive data acquisition techniques, such as infrared radiometry, allowed more widespread use of ceramics for models to be tested in high energy wind tunnels.

Computational Design and Analysis of a Transonic Natural Laminar Flow Wing for a Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Lynde, Michelle N.; Campbell, Richard L.

2017-01-01

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.

Digital-flutter-suppression-system investigations for the active flexible wing wind-tunnel model

NASA Technical Reports Server (NTRS)

Perry, Boyd, III; Mukhopadhyay, Vivek; Hoadley, Sherwood T.; Cole, Stanley R.; Buttrill, Carey S.; Houck, Jacob A.

1990-01-01

Active flutter suppression control laws were designed, implemented, and tested on an aeroelastically-scaled wind tunnel model in the NASA Langley Transonic Dynamics Tunnel. One of the control laws was successful in stabilizing the model while the dynamic pressure was increased to 24 percent greater than the measured open-loop flutter boundary. Other accomplishments included the design, implementation, and successful operation of a one-of-a-kind digital controller, the design and use of two simulation methods to support the project, and the development and successful use of a methodology for on-line controller performance evaluation.

National Wind Tunnel Complex (NWTC)

NASA Technical Reports Server (NTRS)

1996-01-01

The National Wind Tunnel Complex (NWTC) Final Report summarizes the work carried out by a unique Government/Industry partnership during the period of June 1994 through May 1996. The objective of this partnership was to plan, design, build and activate 'world class' wind tunnel facilities for the development of future-generation commercial and military aircraft. The basis of this effort was a set of performance goals defined by the National Facilities Study (NFS) Task Group on Aeronautical Research and Development Facilities which established two critical measures of improved wind tunnel performance; namely, higher Reynolds number capability and greater productivity. Initial activities focused upon two high-performance tunnels (low-speed and transonic). This effort was later descoped to a single multipurpose tunnel. Beginning in June 1994, the NWTC Project Office defined specific performance requirements, planned site evaluation activities, performed a series of technical/cost trade studies, and completed preliminary engineering to support a proposed conceptual design. Due to budget uncertainties within the Federal government, the NWTC project office was directed to conduct an orderly closure following the Systems Design Review in March 1996. This report provides a top-level status of the project at that time. Additional details of all work performed have been archived and are available for future reference.

Estimation of Unsteady Aerodynamic Models from Dynamic Wind Tunnel Data

NASA Technical Reports Server (NTRS)

Murphy, Patrick; Klein, Vladislav

2011-01-01

Demanding aerodynamic modelling requirements for military and civilian aircraft have motivated researchers to improve computational and experimental techniques and to pursue closer collaboration in these areas. Model identification and validation techniques are key components for this research. This paper presents mathematical model structures and identification techniques that have been used successfully to model more general aerodynamic behaviours in single-degree-of-freedom dynamic testing. Model parameters, characterizing aerodynamic properties, are estimated using linear and nonlinear regression methods in both time and frequency domains. Steps in identification including model structure determination, parameter estimation, and model validation, are addressed in this paper with examples using data from one-degree-of-freedom dynamic wind tunnel and water tunnel experiments. These techniques offer a methodology for expanding the utility of computational methods in application to flight dynamics, stability, and control problems. Since flight test is not always an option for early model validation, time history comparisons are commonly made between computational and experimental results and model adequacy is inferred by corroborating results. An extension is offered to this conventional approach where more general model parameter estimates and their standard errors are compared.

Wind tunnel study of helical and straight-bladed vertical-axis wind turbine wakes

NASA Astrophysics Data System (ADS)

Bagheri, Maryam; Araya, Daniel

2017-11-01

It is hypothesized that blade curvature can serve as a passive means to control fluid entrainment and wake recovery in vertical-axis wind turbine (VAWT) arrays. We test this experimentally in a wind tunnel using two different VAWT configurations, one with straight blades and another with helical blades, keeping all other experimental parameters fixed. A small-scale, commercially available VAWT (15W max power) is used as the baseline wind tunnel model in each case. The commercial VAWT blades are replaced with either straight or helical blades that are 3D-printed extrusions of the same airfoil cross-section. Results from smoke flow visualization, three-component wake velocity measurements, and turbine power data are presented. These results give insight into the potential use of VAWTs with curved blades in utility-scale wind farms.

Comparison of the NASA Common Research Model European Transonic Wind Tunnel Test Data to NASA Test Data

NASA Technical Reports Server (NTRS)

Rivers, Melissa B.; Quest, Jurgen; Rudnik, Ralf

2015-01-01

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment, surface pressure and wing bending and twist data are presented herein.

Active Control of Wind-Tunnel Model Aeroelastic Response Using Neural Networks

NASA Technical Reports Server (NTRS)

Scott, Robert C.

2000-01-01

NASA Langley Research Center, Hampton, VA 23681 Under a joint research and development effort conducted by the National Aeronautics and Space Administration and The Boeing Company (formerly McDonnell Douglas) three neural-network based control systems were developed and tested. The control systems were experimentally evaluated using a transonic wind-tunnel model in the Langley Transonic Dynamics Tunnel. One system used a neural network to schedule flutter suppression control laws, another employed a neural network in a predictive control scheme, and the third employed a neural network in an inverse model control scheme. All three of these control schemes successfully suppressed flutter to or near the limits of the testing apparatus, and represent the first experimental applications of neural networks to flutter suppression. This paper will summarize the findings of this project.

Altitude Wind Tunnel at the NACA’s Aircraft Engine Research Laboratory

NASA Image and Video Library

1945-06-21

Two men on top of the Altitude Wind Tunnel (AWT) at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The tunnel was a massive rectangular structure, which for years provided one of the highest vantage points on the laboratory. The tunnel was 263 feet long on the north and south legs and 121 feet long on the east and west sides. The larger west end of the tunnel, seen here, was 51 feet in diameter. The east side of the tunnel was 31 feet in diameter at the southeast corner and 27 feet in diameter at the northeast. The throat section, which connected the northwest corner to the test section, narrowed sharply from 51 to 20 feet in diameter. The AWT’s altitude simulation required temperature and pressure fluctuations that made the design of the shell more difficult than other tunnels. The simultaneous decrease in both pressure and temperature inside the facility produced uneven stress loads, particularly on the support rings. The steel used in the primary tunnel structure was one inch thick to ensure that the shell did not collapse as the internal air pressure was dropped to simulate high altitudes. It was a massive amount of steel considering the World War II shortages. The shell was covered with several inches of fiberglass insulation to retain the refrigerated air and a thinner outer steel layer to protect the insulation against the weather. A unique system of rollers was used between the shell and its support piers. These rollers allowed for movement as the shell expanded or contracted during the altitude simulations. Certain sections would move as much as five inches during operation.

Fast Laser Holographic Interferometry For Wind Tunnels

NASA Technical Reports Server (NTRS)

Lee, George

1989-01-01

Proposed system makes holographic interferograms quickly in wind tunnels. Holograms reveal two-dimensional flows around airfoils and provide information on distributions of pressure, structures of wake and boundary layers, and density contours of flow fields. Holograms form quickly in thermoplastic plates in wind tunnel. Plates rigid and left in place so neither vibrations nor photgraphic-development process degrades accuracy of holograms. System processes and analyzes images quickly. Semiautomatic micro-computer-based desktop image-processing unit now undergoing development moves easily to wind tunnel, and its speed and memory adequate for flows about airfoils.

Evaluation of spray drift using low speed wind tunnel measurements and dispersion modeling

USDA-ARS?s Scientific Manuscript database

The objective of this work was to evaluate the EPA’s proposed Test Plan for the validation testing of pesticide spray drift reduction technologies (DRTs) for row and field crops, focusing on the evaluation of ground application systems using the low-speed wind tunnel protocols and processing the dat...

Some anomalies observed in wind-tunnel tests of a blunt body at transonic and supersonic speeds

NASA Technical Reports Server (NTRS)

Brooks, J. D.

1976-01-01

An investigation of anomalies observed in wind tunnel force tests of a blunt body configuration was conducted at Mach numbers from 0.20 to 1.35 in the Langley 8-foot transonic pressure tunnel and at Mach numbers of 1.50, 1,80, and 2.16 in the Langley Unitary Plan wind tunnel. At a Mach number of 1.35, large variations occurred in axial force coefficient at a given angle of attack. At transonic and low supersonic speeds, the total drag measured in the wind tunnel was much lower than that measured during earlier ballistic range tests. Accurate measurements of total drag for blunt bodies will require the use of models smaller than those tested thus far; however, it appears that accurate forebody drag results can be obtained by using relatively large models. Shock standoff distance is presented from experimental data over the Mach number range from 1.05 to 4.34. Theory accurately predicts the shock standoff distance at Mach numbers up to 1.75.

DARPA/AFRL/NASA Smart Wing Second Wind Tunnel Test Results

NASA Technical Reports Server (NTRS)

Scherer, L. B.; Martin, C. A.; West, M.; Florance, J. P.; Wieseman, C. D.; Burner, A. W.; Fleming, G. A.

2001-01-01

To quantify the benefits of smart materials and structures adaptive wing technology, Northrop Grumman Corp. (NGC) built and tested two 16% scale wind tunnel models (a conventional and a "smart" model) of a fighter/attack aircraft under the DARPA/AFRL/NASA Smart Materials and Structures Development - Smart Wing Phase 1. Performance gains quantified included increased pitching moment (C(sub M)), increased rolling moment (C(subl)) and improved pressure distribution. The benefits were obtained for hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist effected by SMA torque tube mechanisms, compared to conventional hinged control surfaces. This paper presents an overview of the results from the second wind tunnel test performed at the NASA Langley Research Center s (LaRC) 16ft Transonic Dynamic Tunnel (TDT) in June 1998. Successful results obtained were: 1) 5 degrees of spanwise twist and 8-12% increase in rolling moment utilizing a single SMA torque tube, 2) 12 degrees of deflection, and 10% increase in rolling moment due to hingeless, contoured aileron, and 3) demonstration of optical techniques for measuring spanwise twist and deflected shape.

Wind tunnel interference factors for high-lift wings in closed wind tunnels. Ph.D. Thesis - Princeton Univ.

NASA Technical Reports Server (NTRS)

Joppa, R. G.

1973-01-01

A problem associated with the wind tunnel testing of very slow flying aircraft is the correction of observed pitching moments to free air conditions. The most significant effects of such corrections are to be found at moderate downwash angles typical of the landing approach. The wind tunnel walls induce interference velocities at the tail different from those induced at the wing, and these induced velocities also alter the trajectory of the trailing vortex system. The relocated vortex system induces different velocities at the tail from those experienced in free air. The effect of the relocated vortex and the walls is to cause important changes in the measured pitching moments in the wind tunnel.

A surface flow visualisation technique for use in cryogenic wind tunnels

NASA Technical Reports Server (NTRS)

Kell, D. M.

1978-01-01

A method of surface flow visualization for use in cryogenic wind tunnels is described which requires injection of a cryogenic liquid onto the model while the tunnel is running. This necessitates the use of a substance that remains liquid over a large range of cryogenic wind tunnel operating temperatures. It is found that propane (C3H8) is a suitable substance. Experiments are conducted in a subsonic cryogenic wind tunnel to assess the practical application of liquid propane flow visualization. The propane is stored in a chamber cooled by liquid nitrogen and when required is pumped through pipes to a gallery inside the model and then out onto the surface through small holes. To color the liquid a suspension of pigment particles is used. Propane is supplied to the cooled chamber in gaseous form from a standard liquefied gas cylinder. The sequence of events is illustrated on a propane temperature-entropy diagram. The use of liquefied propane for flow visualization in a cryogenic tunnel operating at pressures up to 40 atm appears to be feasible. Illustrative examples are provided.

Analysis and design of quiet hypersonic wind tunnels

NASA Astrophysics Data System (ADS)

Naiman, Hadassah

The purpose of the present work is to integrate CFD into the design of quiet hypersonic wind tunnels and the analysis of their performance. Two specific problems are considered. The first problem is the automated design of the supersonic portion of a quiet hypersonic wind tunnel. Modern optimization software is combined with full Navier-Stokes simulations and PSE stability analysis to design a Mach 6 nozzle with maximum quiet test length. A response surface is constructed from a user-specified set of contour shapes and a genetic algorithm is used to find the "optimal contour", which is defined as the shortest nozzle with the maximum quiet test length. This is achieved by delaying transition along the nozzle wall. It is found that transition is triggered by Goertler waves, which can be suppressed by including a section of convex curvature along the contour. The optimal design has an unconventional shape described as compound curvature, which makes the contour appear slightly wavy. The second problem is the evaluation of a proposed modification of the test section in the Boeing/AFOSR Mach 6 Quiet Tunnel. The new design incorporates a section of increased diameter with the intention of enabling the tunnel to start in the presence of larger blunt models. Cone models with fixed base diameter (and hence fixed blockage ratio) are selected for this study. Cone half-angles from 15° to 75° are examined to ascertain the effect of ii the strength of the test model shock wave on the tunnel startup. The unsteady, laminar, compressible Navier-Stokes equations are solved. The resulting flowfields are analyzed to see what affect the shocks and shear layers have on the quiet test section flow. This study indicates that cone angles ≤20° allow the tunnel to start. Keywords. automated optimization, response surface, parabolized stability equations, compound curvature, laminar, wind tunnel, unstart, test section.

Model Deformation and Optical Angle of Attack Measurement System in the NASA Ames Unitary Plan Wind Tunnel

NASA Technical Reports Server (NTRS)

Kushner, Laura K.; Drain, Bethany A.; Schairer, Edward T.; Heineck, James T.; Bell, James H.

2017-01-01

Both AoA and MDM measurements can be made using an optical system that relies on photogrammetry. Optical measurements are being requested by customers in wind tunnels with increasing frequency due to their non-intrusive nature and recent hardware and software advances that allow measurements to become near real time. The NASA Ames Research Center Unitary Plan Wind Tunnel is currently developing a system based on photogrammetry to measure model deformation and model angle of attack. This paper describes the new system, its development, its use on recent tests and plans to further develop the system.

A new low-turbulence wind tunnel for animal and small vehicle flight experiments

NASA Astrophysics Data System (ADS)

Quinn, Daniel B.; Watts, Anthony; Nagle, Tony; Lentink, David

2017-03-01

Our understanding of animal flight benefits greatly from specialized wind tunnels designed for flying animals. Existing facilities can simulate laminar flow during straight, ascending and descending flight, as well as at different altitudes. However, the atmosphere in which animals fly is even more complex. Flow can be laminar and quiet at high altitudes but highly turbulent near the ground, and gusts can rapidly change wind speed. To study flight in both laminar and turbulent environments, a multi-purpose wind tunnel for studying animal and small vehicle flight was built at Stanford University. The tunnel is closed-circuit and can produce airspeeds up to 50 m s-1 in a rectangular test section that is 1.0 m wide, 0.82 m tall and 1.73 m long. Seamless honeycomb and screens in the airline together with a carefully designed contraction reduce centreline turbulence intensities to less than or equal to 0.030% at all operating speeds. A large diameter fan and specialized acoustic treatment allow the tunnel to operate at low noise levels of 76.4 dB at 20 m s-1. To simulate high turbulence, an active turbulence grid can increase turbulence intensities up to 45%. Finally, an open jet configuration enables stereo high-speed fluoroscopy for studying musculoskeletal control in turbulent flow.

Use of 3D Printing for Custom Wind Tunnel Fabrication

NASA Astrophysics Data System (ADS)

Gagorik, Paul; Bates, Zachary; Issakhanian, Emin

2016-11-01

Small-scale wind tunnels for the most part are fairly simple to produce with standard building equipment. However, the intricate bell housing and inlet shape of an Eiffel type wind tunnel, as well as the transition from diffuser to fan in a rectangular tunnel can present design and construction obstacles. With the help of 3D printing, these shapes can be custom designed in CAD models and printed in the lab at very low cost. The undergraduate team at Loyola Marymount University has built a custom benchtop tunnel for gas turbine film cooling experiments. 3D printing is combined with conventional construction methods to build the tunnel. 3D printing is also used to build the custom tunnel floor and interchangeable experimental pieces for various experimental shapes. This simple and low-cost tunnel is a custom solution for specific engineering experiments for gas turbine technology research.

The design of two sonic boom wind tunnel models from conceptual aircraft which cruise at Mach numbers of 2.0 and 3.0

NASA Technical Reports Server (NTRS)

Mack, Robert J.; Needleman, Kathy E.

1990-01-01

A method for designing wind tunnel models of conceptual, low-boom, supersonic cruise aircraft is presented. Also included is a review of the procedures used to design the conceptual low-boom aircraft. In the discussion, problems unique to, and encountered during, the design of both the conceptual aircraft and the wind tunnel models are outlined. The sensitivity of low-boom characteristics in the aircraft design to control the volume and lift equivalent area distributions was emphasized. Solutions to these problems are reported; especially the two which led to the design of the wind tunnel model support stings.

The application of cryogenics to high Reynolds number testing in wind tunnels. I - Evolution, theory, and advantages

NASA Technical Reports Server (NTRS)

Kilgore, R. A.; Dress, D. A.

1984-01-01

During the time which has passed since the construction of the first wind tunnel in 1870, wind tunnels have been developed to a high degree of sophistication. However, their development has consistently failed to keep pace with the demands placed on them. One of the more serious problems to be found with existing transonic wind tunnels is their inability to test subscale aircraft models at Reynolds numbers sufficiently near full-scale values to ensure the validity of using the wind tunnel data to predict flight characteristics. The Reynolds number capability of a wind tunnel may be increased by a number of different approaches. However, the best solution in terms of model, balance, and model support loads, as well as in terms of capital and operating cost appears to be related to the reduction of the temperature of the test gas to cryogenic temperatures. The present paper has the objective to review the evolution of the cryogenic wind tunnel concept and to describe its more important advantages.

Iterative adaption of the bidimensional wall of the French T2 wind tunnel around a C5 axisymmetrical model: Infinite variation of the Mach number at zero incidence and a test at increased incidence

NASA Technical Reports Server (NTRS)

Archambaud, J. P.; Dor, J. B.; Payry, M. J.; Lamarche, L.

1986-01-01

The top and bottom two-dimensional walls of the T2 wind tunnel are adapted through an iterative process. The adaptation calculation takes into account the flow three-dimensionally. This method makes it possible to start with any shape of walls. The tests were performed with a C5 axisymmetric model at ambient temperature. Comparisons are made with the results of a true three-dimensional adaptation.

Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a data base and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Data Base is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configuration types include booster and orbiter components in various stacked and tandem combinations.

Survey Of Wind Tunnels At Langley Research Center

NASA Technical Reports Server (NTRS)

Bower, Robert E.

1989-01-01

Report presented at AIAA 14th Aerodynamic Testing Conference on current capabilities and planned improvements at NASA Langley Research Center's major wind tunnels. Focuses on 14 major tunnels, 8 unique in world, 3 unique in country. Covers Langley Spin Tunnel. Includes new National Transonic Facility (NTF). Also surveys Langley Unitary Plan Wind Tunnel (UPWT). Addresses resurgence of inexpensive simple-to-operate research tunnels. Predicts no shortage of tools for aerospace researcher and engineer in next decade or two.

Infrared radiometer for measuring thermophysical properties of wind tunnel models

NASA Technical Reports Server (NTRS)

Corwin, R. R.; Moorman, S. L.; Becker, E. C.

1978-01-01

An infrared radiometer is described which was developed to measure temperature rises of wind tunnel models undergoing transient heating over a temperature range of -17.8 C to 260 C. This radiometer interfaces directly with a system which measures the effective thermophysical property square root of rho ck. It has an output temperature fluctuation of 0.26 C at low temperatures and 0.07 C at high temperatures, and the output frequency response of the radiometer is from dc to 400 hertz.

Modeling the effects of wind tunnel wall absorption on the acoustic radiation characteristics of propellers

NASA Technical Reports Server (NTRS)

Baumeister, K. J.; Eversman, W.

1986-01-01

Finite element theory is used to calculate the acoustic field of a propeller in a soft walled circular wind tunnel and to compare the radiation patterns to the same propeller in free space. Parametric solutions are present for a 'Gutin' propeller for a variety of flow Mach numbers, admittance values at the wall, microphone position locations, and propeller to duct radius ratios. Wind tunnel boundary layer is not included in this analysis. For wall admittance nearly equal to the characteristic value of free space, the free field and ducted propeller models agree in pressure level and directionality. In addition, the need for experimentally mapping the acoustic field is discussed.

Modeling the effects of wind tunnel wall absorption on the acoustic radiation characteristics of propellers

NASA Technical Reports Server (NTRS)

Baumeister, K. J.; Eversman, W.

1986-01-01

Finite element theory is used to calculate the acoustic field of a propeller in a soft walled circular wind tunnel and to compare the radiation patterns to the same propeller in free space. Parametric solutions are present for a "Gutin" propeller for a variety of flow Mach numbers, admittance values at the wall, microphone position locations, and propeller to duct radius ratios. Wind tunnel boundary layer is not included in this analysis. For wall admittance nearly equal to the characteristic value of free space, the free field and ducted propeller models agree in pressure level and directionality. In addition, the need for experimentally mapping the acoustic field is discussed.

Application of Neural Networks to Wind tunnel Data Response Surface Methods

NASA Technical Reports Server (NTRS)

Lo, Ching F.; Zhao, J. L.; DeLoach, Richard

2000-01-01

The integration of nonlinear neural network methods with conventional linear regression techniques is demonstrated for representative wind tunnel force balance data modeling. This work was motivated by a desire to formulate precision intervals for response surfaces produced by neural networks. Applications are demonstrated for representative wind tunnel data acquired at NASA Langley Research Center and the Arnold Engineering Development Center in Tullahoma, TN.

Quantification of wind flow in the European Mars Simulation Wind Tunnel Facility

NASA Astrophysics Data System (ADS)

Holstein-Rathlou, C.; Merrison, J. P.; Iversen, J. J.; Nornberg, P.

2012-04-01

We present the European Mars Simulation Wind Tunnel facility, a unique prototype facility capable of simulating a wide range of environmental conditions, such as those which can be found at the surface of Earth or Mars. The chamber complements several other large-scale simulation facilities at Aarhus University, Denmark. The facility consists of a 50 m3 environmental chamber capable of operating at low pressure (0.02 - 1000 mbar) and cryogenic temperatures (-130 °C up to +60 °C). This chamber houses a re-circulating wind tunnel capable of generating wind speeds up to 25 m/s and has a dust injection system that can produce suspended particulates (aerosols). It employs a unique LED based optical illumination system (solar simulator) and an advanced network based control system. Laser based optoelectronic instrumentation is used to quantify and monitor wind flow, dust suspension and deposition. This involves a commercial Laser Doppler Anemometer (LDA) and a Particle Dynamics Analysis receiver (PDA), which are small laser based instruments specifically designed for measuring wind speed and sizes of particles situated in a wind flow. Wind flow calibrations will be performed with the LDA system and presented. Pressure and temperature calibrations will follow in order to enable the facility to be used for the testing, development, calibration and comparison of e.g. meteorological sensors under a wide range of environmental conditions as well as multi-disciplinary scientific studies. The wind tunnel is accessible to international collaborators and space agencies for instrument testing, calibration and qualification. It has been financed by the European Space Agency (ESA) as well as the Aarhus University Science Faculty and the Villum Kann Rasmussen Foundation.

Planning Image-Based Measurements in Wind Tunnels by Virtual Imaging

NASA Technical Reports Server (NTRS)

Kushner, Laura Kathryn; Schairer, Edward T.

2011-01-01

Virtual imaging is routinely used at NASA Ames Research Center to plan the placement of cameras and light sources for image-based measurements in production wind tunnel tests. Virtual imaging allows users to quickly and comprehensively model a given test situation, well before the test occurs, in order to verify that all optical testing requirements will be met. It allows optimization of the placement of cameras and light sources and leads to faster set-up times, thereby decreasing tunnel occupancy costs. This paper describes how virtual imaging was used to plan optical measurements for three tests in production wind tunnels at NASA Ames.

New methodologies for calculation of flight parameters on reduced scale wings models in wind tunnel =

NASA Astrophysics Data System (ADS)

Ben Mosbah, Abdallah

In order to improve the qualities of wind tunnel tests, and the tools used to perform aerodynamic tests on aircraft wings in the wind tunnel, new methodologies were developed and tested on rigid and flexible wings models. A flexible wing concept is consists in replacing a portion (lower and/or upper) of the skin with another flexible portion whose shape can be changed using an actuation system installed inside of the wing. The main purpose of this concept is to improve the aerodynamic performance of the aircraft, and especially to reduce the fuel consumption of the airplane. Numerical and experimental analyses were conducted to develop and test the methodologies proposed in this thesis. To control the flow inside the test sections of the Price-Paidoussis wind tunnel of LARCASE, numerical and experimental analyses were performed. Computational fluid dynamics calculations have been made in order to obtain a database used to develop a new hybrid methodology for wind tunnel calibration. This approach allows controlling the flow in the test section of the Price-Paidoussis wind tunnel. For the fast determination of aerodynamic parameters, new hybrid methodologies were proposed. These methodologies were used to control flight parameters by the calculation of the drag, lift and pitching moment coefficients and by the calculation of the pressure distribution around an airfoil. These aerodynamic coefficients were calculated from the known airflow conditions such as angles of attack, the mach and the Reynolds numbers. In order to modify the shape of the wing skin, electric actuators were installed inside the wing to get the desired shape. These deformations provide optimal profiles according to different flight conditions in order to reduce the fuel consumption. A controller based on neural networks was implemented to obtain desired displacement actuators. A metaheuristic algorithm was used in hybridization with neural networks, and support vector machine approaches and their

Scale effects in wind tunnel modeling of an urban atmospheric boundary layer

NASA Astrophysics Data System (ADS)

Kozmar, Hrvoje

2010-03-01

Precise urban atmospheric boundary layer (ABL) wind tunnel simulations are essential for a wide variety of atmospheric studies in built-up environments including wind loading of structures and air pollutant dispersion. One of key issues in addressing these problems is a proper choice of simulation length scale. In this study, an urban ABL was reproduced in a boundary layer wind tunnel at different scales to study possible scale effects. Two full-depth simulations and one part-depth simulation were carried out using castellated barrier wall, vortex generators, and a fetch of roughness elements. Redesigned “Counihan” vortex generators were employed in the part-depth ABL simulation. A hot-wire anemometry system was used to measure mean velocity and velocity fluctuations. Experimental results are presented as mean velocity, turbulence intensity, Reynolds stress, integral length scale of turbulence, and power spectral density of velocity fluctuations. Results suggest that variations in length-scale factor do not influence the generated ABL models when using similarity criteria applied in this study. Part-depth ABL simulation compares well with two full-depth ABL simulations indicating the truncated vortex generators developed for this study can be successfully employed in urban ABL part-depth simulations.

Altitude Wind Tunnel Drive Motor Installation

NASA Image and Video Library

1943-07-21

Construction workers install the drive motor for the Altitude Wind Tunnel (AWT) in the Exhauster Building at the National Advisory Committee for Aeronautics (NACA) Aircraft Engine Research Laboratory. The AWT was capable of operating full-scale engines in air density, speed, and temperature similar to that found at high altitudes. The tunnel could produce wind speeds up to 500 miles per hour through a 20-foot-diameter test section at the standard operating altitude of 30,000 feet. The airflow was created by a large wooden fan near the tunnel’s southeast corner. This photograph shows the installation of the 18,000-horsepower drive motor inside the adjoining Exhauster Building in July 1943. The General Electric motor, whose support frame is seen in this photograph, connected to a drive shaft that extended from the building, through the tunnel shell, and into a 12-bladed, 31-foot-diameter spruce wood fan. Flexible couplings on the shaft allowed for the movement of the shell. The corner of the Exhauster Building was built around the motor after its installation. The General Electric induction motor could produce 10 to 410 revolutions per minute and create wind speeds up to 500 miles per hour, or Mach 0.63, at 30,000 feet. The AWT became operational in January 1944 and tested piston, turbojet and ramjet engines for nearly 20 years.

Laminar Flow Supersonic Wind Tunnel primary air injector

NASA Technical Reports Server (NTRS)

Smith, Brooke Edward

1993-01-01

This paper describes the requirements, design, and prototype testing of the flex-section and hinge seals for the Laminar Flow Supersonic Wind Tunnel Primary Injector. The supersonic atmospheric primary injector operates between Mach 1.8 and Mach 2.2 with mass-flow rates of 62 to 128 lbm/s providing the necessary pressure reduction to operate the tunnel in the desired Reynolds number (Re) range.

Neural network feedforward control of a closed-circuit wind tunnel

NASA Astrophysics Data System (ADS)

Sutcliffe, Peter