TPSAS-NF1676L-18952-DND
Summary of near real time modeling techniques applied to flight tests and wind tunnel tests.
Jay Brandon↗
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Results for “wind tunnel testing”
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Technology Review of Wind-Tunnel Angle Measurement
Angle measurements are fundamental measurements in wind-tunnel testing, and are integral for most, if not all, wind-tunnel tests. Aerodynamic tests generally rely on angle or attitude measurement for research or testing objectives with demands of high accuracy and repeatable measurements. This paper reviews historical and recent developments in wind-tunnel angle measurement systems. The review covers sensor technologies including on-board and o-body measurement solutions. On-board sensor solutions, such as accelerometers, gyroscopes, and electrolytic bubbles, are compared with o-body solutions, such as laser interferometers and photogrammetry. Benefits, use cases, and limitations for each of the technologies are discussed to help guide wind-tunnel user selection, and provide direction for further research in sensor technologies which may provide enhanced measurement capability.
Toro, Kenneth G.↗
Design, develop, fabricate, assemble and checkout and advanced active damper system
An active electromagnetic damper system was developed for and used in a 5.5% aeroelastic model of Saturn 1B/SA206 launch configuration in wind tunnel tests to determine ground wind-induced loads. Detailed descriptions are given of the damper development program: approach, design analyses, fabrication, checkout, system calibration, wind tunnel test results, damper assembly and operating instructions, and trouble-shooting procedures. Conclusions and recommendations to improve the damper system for future applications are also presented.
Chang, C. S.↗
A forward speed effects study on jet noise from several suppressor nozzles in the NASA/Ames 40- by 80-foot wind tunnel
A test program was conducted in a 40 by 80 foot wind tunnel to evaluate the effect of relative velocity on the jet noise signature of a conical ejector, auxiliary inlet ejector, 32 spokes and 104 tube nozzle with and without an acoustically treated shroud. The freestream velocities in the wind tunnel were varied from 0 to 103.6 m/sec (300 ft/sec) for exhaust jet velocities of 259.1 m/sec (850 ft/sec) to 609.6 m/sec (2000 ft/sec). Reverberation corrections for the wind tunnel were developed and the procedure is explained. In conjunction with wind tunnel testing the nozzles were also evaluated on an outdoor test stand. The wind tunnel microphone arrays were duplicated during the outdoor testing. The data were then extrapolated for comparisons with data measured using a microphone array placed on a 30.5 meter (100 ft) arc. Using these data as a basis, farfield to nearfield arguments are presented with regards to the data measured in the wind tunnel. Finally, comparisons are presented between predictions made using existing methods and the measured data.
Beulke, M. R.↗
Parachute Aerodynamics From Video Data
A new data analysis technique for the identification of static and dynamic aerodynamic stability coefficients from wind tunnel test video data is presented. This new technique was applied to video data obtained during a parachute wind tunnel test program conducted in support of the Mars Exploration Rover Mission. Total angle-of-attack data obtained from video images were used to determine the static pitching moment curve of the parachute. During the original wind tunnel test program the static pitching moment curve had been determined by forcing the parachute to a specific total angle-of -attack and measuring the forces generated. It is shown with the new technique that this parachute, when free to rotate, trims at an angle-of-attack two degrees lower than was measured during the forced-angle tests. An attempt was also made to extract pitch damping information from the video data. Results suggest that the parachute is dynamically unstable at the static trim point and tends to become dynamically stable away from the trim point. These trends are in agreement with limit-cycle-like behavior observed in the video. However, the chaotic motion of the parachute produced results with large uncertainty bands.
Schoenenberger, Mark↗
Preliminary results of buffet tests in a cryogenic wind tunnel
Buffet tests of two wings with different leading-edge sweep show that it is feasible to use the standards wing root bending moment technique in a cryogenic wing tunnel. The results for the 65 deg sweep delta wing indicate the importance of matching the reduced frequency parameter in model tests for planforms which are sensitive to reduced frequency parameter if quantitative buffet measurements are required. The unique ability of a pressurized cryogenic wind tunnel to separate the effects of Reynolds number and of aeroelastic distortion by variations in the tunnel stagnation temperature and pressure was demonstrated.
Boyden, R. P.↗
Experimental Study of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Aeroshell with Axisymmetric Surface Deflection Patterns
A wind tunnel test program was conducted to obtain aeroheating environment data on Hypersonic Inflatable Aerodynamic Decelerator aeroshells with flexible thermal protection systems. Data were obtained on a set of rigid wind tunnel models with surface deflection patterns of various heights that simulated a range of potential in-flight aeroshell deformations. Wind tunnel testing was conducted at Mach 6 at unit Reynolds numbers from 2.1 × 10(exp 6)/ft to 8.3 × 10(exp 6)/ft and angles of attack from 0 deg to 18 deg. Boundary-layer transition onset and global surface heating distribution measurements were performed using phosphor thermography and flow field images were obtained through schlieren photography. Surface deflections were found to both promote early transition of the boundary layer and to augment heating levels for both laminar and turbulent flows. A complimentary computational flow field study was also performed to provide heating predictions for comparison with the measurements as well as boundary layer flow field properties for use in correlating the data. Correlations of the wind tunnel data were developed to predict deflection effects on boundary layer transition and surface heating and were applied to both the wind tunnel test conditions and to the trajectory of NASA's successful IRVE-3 flight test. In general, the correlations produced at least qualitative agreement with the wind tunnel data, although the heating levels were underpredicted for some of the larger surface deflections. For the flight conditions, the correlations suggested that peak heating levels on the leeward side conical flank of the IRVE-3 vehicle may have exceeded those at nose for times late in the trajectory after the peak heating time point. However, the flight estimates were based on a conservative assumption of surface deflection magnitude (i.e., larger) than likely was produced in flight.
Hollis, Brian R.↗
Independent Effects of Reynolds Number and Mach Number on the Aerodynamics of an Iced Swept Wing
Aerodynamic assessment of icing effects on swept wings is an important component of a larger effort to improve three-dimensional (3D) icing simulation capabilities. An understanding of ice-shape geometric fidelity and Reynolds and Mach number effects on the iced-wing aerodynamics is needed to guide the development and validation of ice-accretion simulation tools. To this end, wind tunnel testing was carried out for a 13.3-percent-scale semispan wing based upon the Common Research Model airplane configuration. The wind tunnel testing was conducted at the Office National d’Etudes et de Recherches Aérospatiales (ONERA) F1 pressurized wind tunnel with Reynolds numbers of 1.6 x 10(exp 6) to 11.9 x 10(exp 6 ) and Mach numbers of 0.09 to 0.34. Five different configurations were investigated using fully 3D, high-fidelity artificial ice shapes that maintain nearly all of the 3D ice-accretion features documented in prior icing wind tunnel tests. These large, leading-edge ice shapes were nominally based upon airplane holding in icing conditions scenarios. For three of these configurations, lower fidelity simulations were also built and tested. The results presented in this paper show that while Reynolds and Mach number effects are important for quantifying the clean-wing performance, there is very little to no effect for an iced wing with 3D, high-fidelity artificial ice shapes or 3D smooth ice shapes with grit roughness. These conclusions are consistent with the large volume of past research on iced airfoils. However, some differences were also noted for the associated stalling angle of the iced swept wing and for various lower fidelity versions of the leading-edge ice accretion. More research is planned to further investigate the key features of ice-accretion geometry that must be simulated in lower fidelity versions in order to capture the essential aerodynamics.
Aircraft icing↗
Wind Tunnel Performance Tests of a Full-Scale Proprotor on the Tiltrotor Test Rig
The Tiltrotor Test Rig (TTR) is a new, large-scale proprotor test system for the National Full-Scale Aerodynamics Complex (NFAC). The TTR was jointly developed by NASA, the U. S. Army and the U. S. Air Force. The first wind-tunnel entry was completed in November 2018 with a modern, 26-ft diameter proprotor. The primary purpose was to complete the development of the TTR, including systems integration with the NFAC. The TTR and rotor were tested up to 273 knots in axial flow, which is the highest airspeed ever achieved by a full-scale proprotor in any wind tunnel. The entry constitutes the first full-scale wind-tunnel test of a gimballed, coning-flexure rotor. Extensive conversion-mode data were also acquired, and hover/climb conditions were explored. Additional testing included aerodynamic tares, motor tests, thermal tests, modal vibration tests, and other checkout activities. This report summarizes the results of the test, including examples of the most significant data.
Tiltrotor Test Rig↗
Wind Tunnel Model Design for Sonic Boom Studies of Nozzle Jet with Shock Interactions
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.
Wind↗
The Modern Design of Experiments: A Technical and Marketing Framework
A new wind tunnel testing process under development at NASA Langley Research Center, called Modern Design of Experiments (MDOE), differs from conventional wind tunnel testing techniques on a number of levels. Chief among these is that MDOE focuses on the generation of adequate prediction models rather than high-volume data collection. Some cultural issues attached to this and other distinctions between MDOE and conventional wind tunnel testing are addressed in this paper.
DeLoach, R.↗
Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration
Archived wind tunnel test data are available for flyback booster or other alternate recoverable configuration as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle, including contractor data for an extensive variety of configurations with an array of wing and body planforms. The test data have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration. Basic components include booster, orbiter, and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configurations include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. The digital database consists of 220 files containing basic tunnel data. Database structure is documented in a series of reports which include configuration sketches for the various planforms tested. This is Volume 3 -- launch configurations.
Glynn, J. L.↗
Preliminary Steps in Developing Rapid Aero Modeling Technology
The Rapid Aero Modeling (RAM) approach is a method to efficiently and automatically obtain aerodynamic models during testing, significantly saving time and resources. Motivation for this technology results from demand for experimental efficiency and model fidelity that has increased with growing aircraft complexity and aerodynamic nonlinearities. These issues are typical in the responses presented by a class of vehicles categorized as Urban Air Mobility aircraft where many features from both airplane and rotorcraft are present. For UAM configurations, with typically many more factors than conventional aircraft, traditional test methods can lead to increased costs and missed interactions. RAM guides the test to obtain high-fidelity, statistically rigorous aircraft models, and the approach is applicable to computational, ground, or flight-test experiments. It combines concepts from Design of Experiment theory and Aircraft System Identification theory that allow the user the freedom to choose, in advance of the test, a specific level of fidelity typically, in terms of prediction error. RAM only collects data required to meet the user-specified fidelity and fidelity is only limited by the facility and test article capabilities. An initial wind tunnel test to support development of RAM was conducted to assess potential metrics, algorithms, and procedures. This paper presents results from initial tests for the development of RAM technology and highlights some of the unique features of RAM applied toa conventional configuration during a ground-based, static, wind-tunnel test.
Patrick C Murphy↗
Generation of Fullspan Leading-Edge 3D Ice Shapes for Swept-Wing Aerodynamic Testing
The deleterious effect of ice accretion on aircraft is often assessed through dry-air flight and wind tunnel testing with artificial ice shapes. This paper describes a method to create fullspan swept-wing artificial ice shapes from partial span ice segments acquired in the NASA Glenn Icing Reserch Tunnel for aerodynamic wind-tunnel testing. Full-scale ice accretion segments were laser scanned from the Inboard, Midspan, and Outboard wing station models of the 65% scale Common Research Model (CRM65) aircraft configuration. These were interpolated and extrapolated using a weighted averaging method to generate fullspan ice shapes from the root to the tip of the CRM65 wing. The results showed that this interpolation method was able to preserve many of the highly three dimensional features typically found on swept-wing ice accretions. The interpolated fullspan ice shapes were then scaled to fit the leading edge of a 8.9% scale version of the CRM65 wing for aerodynamic wind-tunnel testing. Reduced fidelity versions of the fullspan ice shapes were also created where most of the local three-dimensional features were removed. The fullspan artificial ice shapes and the reduced fidelity versions were manufactured using stereolithography.
Icing↗
Standardization Tests of NACA No. 1 Wind Tunnel
The tests described in this report were made in the 5-foot atmospheric wind tunnel of the National Advisory Committee for Aeronautics, at Langley Field. The primary objective of collecting data on the characteristics of this tunnel for comparison with those of others throughout the world, in order that, in the future, the results of tests made in all the principle laboratories may be interpreted, compared, and coordinated on a basis of scientifically established relationships, a process hitherto impossible due to the lack of comparable data. The work includes tests of a disk, spheres, cylinders, and airfoils, explorations of the test section for static pressure and velocity distribution, and determination of the variations of air flow direction throughout the operating range of the tunnel. (author)
Reid, Elliott G↗
Characterization of the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel Supersonic Test Section (2020 Test)
Recent customer requirements for improved Mach number stability for sonic boom signature measurements has piqued interest in the use of the solid-wall supersonic test section in the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel. To supply customers with understanding of the operating conditions and flow quality in this region of the tunnel, characterization tests were performed in the solid-wall supersonic test section in January 2020. The test entry covered the operating range of the facility for three-drive-motor operation. Calibration models were generated along with an improved understanding of operational considerations for this region of the tunnel. The data set acquired included centerline axial and planar flow-field distributions to quantify test section flow quality. The collected data provided an improved understanding of operational considerations for this region of the tunnel as well as allowed for the development of detailed calibration models for the solid-wall supersonic test section.
Supersonic↗
A new airfoil design concept
The present airfoil design concept is based on utilizing unconventional geometry characteristics near the airfoil trailing edge which include a finite trailing edge thickness, strongly divergent trailing edge upper and lower surfaces, and high surface curvature on the lower surface at or near the lower surface trailing edge. This paper presents computational analyses of airfoils and a wing utilizing the concept, airfoil validation wind tunnel test results of several configurations, and wing-validation wind tunnel test results for a complete wing design. In addition to validating the concept, the airfoil and wing testing provided additional detailed data to better understand the aerodynamic advantage of such an unconventional trailing edge configuration. It is demonstrated that the concept represents a significant step in airfoil technology beyond that achieved with the Supercritical Airfoil. This concept provides the aerodynamicist an additional degree of design freedom and flexibility previously unrecognized.
Henne, P. A.↗
A system for model access in tunnels with an unbreathable test medium
In many specialty wind tunnels, test gases other than ambient air are used to meet special testing requirements. A typical example is the use of freon as the test gas to achieve a realistic density ratio between gas and model for exploring flutter stability boundaries. Another example is the use of pressurized air to elevate the stream density and enhance Reynolds number or dynamic pressure simulation. Such specialty tunnels require a system of access o the model which will allow services and changes to the model without exposing personnel to the unnatural and perhaps hostile environment or requiring the venting and purging of the entire tunnel circuit. This paper will describe the plenum and model access systems for the forthcoming U.S. National Transonic facility where gaseous nitrogen (N2) at temperatures between 338 and 78 K and at pressures to 9 bars is used as the test medium. The operation at cold temperatures imposes some additional requirements which make the access systems more difficult to design and time consuming to operate than for conventional wind tunnels.
Howell, R. R.↗