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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

Prediction of Wind Tunnel Effects on the Installed F/A-18A Inlet Flow Field at High Angles-of-attack

NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18A, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Wind tunnel tests were a major component of the Lewis program. Since the available wind tunnel was small (9 x 15 ft) as compared to the scale of the model of the F/A-18A (19.78 percent), there were questions about the capability to obtain useful inlet performance data. The blockage effects were expected to be very large. This report represents the results of an analysis to determine how the wind tunnel walls effect inlet performance at several angles of attack. The predictions for the external particle traces along the fuselage indicate the influence of the wind tunnel side wall under the model is greater at 30 deg angle of attack than at 50 deg angle of attack on the under Leading Edge Extension (LEX) vortex trajectory. The side wall above the model appears to have negligible influence on the under LEX vortex. This may be due to the LEX acting as 'shield' to the upper wall effects. As expected, the wind tunnel has a significant influence on the external forces. The lift and drag coefficients increase significantly for the wind tunnel model as compared to free stream conditions. The wind tunnel had a small effect on the inlet recovery and on inlet total pressure distortion patterns. The predicted recoveries for the wind tunnel model are within one percentage point of the model recoveries in free stream conditions.

Smith, Crawford F.↗
NASA NTRS1961

The Development of an 8-inch by 8-inch Slotted Tunnel for Mach Numbers up to 1.28

An 8-inch by 8-inch transonic tunnel model with test section slotted on two opposite walls was constructed in which particular emphasis -was given to the development of slot geometry, slot-flow reentry section, and short-diffuser configurations for good test-region flow and minimum total-pressure losses. Center-line static pressures through the test section, wall static pressures through the other parts of the tunnel, and total-pressure distributions at the inlet and exit stations of the diffuser were measured- With a slot length equal to two tunnel heights and 1/14 open-area-ratio slotted walls) a test region one tunnel height in length was obtained in which the deviation from the mean Mach number was less than +/- 0.01 up to Mach number 1.15. With 1/7 open-area-ratio slotted walls, a test region 0.84 tunnel heights in length with deviation less than +/- O.01 was obtained up to Mach number 1.26. Increasing the tunnel diffuser angle from 6.4 to 10 deg. increased pressure loss through the tunnel at Mach number 1.20 from 15 percent to 20 percent of the total pressure. The use of other diffusers with equivalent angles of 10 deg. but contoured so that the initial diffusion angle was less than 10 deg. and the final angle was 200 reduced the losses to as low as 16 percent. A method for changing the test-section Mach number rapidly by controlling the flow through a bypass line from the tunnel settling chamber to the slot-flow plenum chamber of the test section was very effective. The test-section Mach number was reduced approximately 5 percent in 1/8 second by bleeding into the test section a flow of air equal to 2 percent of the mainstream flow and 30 percent in 1/4 second with bleed flow equal to 10 percent of the mainstream flow. The rate of reduction was largely determined by the opening rate of the bleed-flow-control valve.

Little, B. H., Jr.↗
NASA NTRS2006

Dynamic Tunnel Usability Study: Format Recommendations for Synthetic Vision System Primary Flight Displays

A usability study evaluating dynamic tunnel concepts has been completed under the Aviation Safety and Security Program, Synthetic Vision Systems Project. The usability study was conducted in the Visual Imaging Simulator for Transport Aircraft Systems (VISTAS) III simulator in the form of questionnaires and pilot-in-the-loop simulation sessions. Twelve commercial pilots participated in the study to determine their preferences via paired comparisons and subjective rankings regarding the color, line thickness and sensitivity of the dynamic tunnel. The results of the study showed that color was not significant in pilot preference paired comparisons or in pilot rankings. Line thickness was significant for both pilot preference paired comparisons and in pilot rankings. The preferred line/halo thickness combination was a line width of 3 pixels and a halo of 4 pixels. Finally, pilots were asked their preference for the current dynamic tunnel compared to a less sensitive dynamic tunnel. The current dynamic tunnel constantly gives feedback to the pilot with regard to path error while the less sensitive tunnel only changes as the path error approaches the edges of the tunnel. The tunnel sensitivity comparison results were not statistically significant.

Arthur, Jarvis J., III↗
NASA NTRS2009

Compartment Venting Analyses of Ares I First Stage Systems Tunnel

Compartment venting analyses have been performed for the Ares I first stage systems tunnel using both the lumped parameter method and the three-dimensional (31)) transient computational fluid dynamics (CFD) approach. The main objective of venting analyses is to predict the magnitudes of differential pressures across the skin so the integrity of solid walls can be evaluated and properly designed. The lumped parameter method assumes the gas pressure and temperature inside the systems tunnel are spatially uniform, which is questionable since the tunnel is about 1,700 in. long and 4 in. wide. Therefore, 31) transient CFD simulations using the commercial CFD code FLUENT are performed in order to examine the gas pressure and temperature variations inside the tunnel. It was found that the uniform pressure and temperature assumptions inside the systems tunnel are valid during ascent. During reentry, the uniform pressure assumption is also reasonable but the uniform temperature assumption is not valid. Predicted pressure and temperature inside the systems tunnel using CFD are also compared with those from the lumped parameter method using the NASA code CHCHVENT. In general, the average pressure and temperature inside the systems tunnel from CFD are between the burst and crush results from CHCHVENT during both ascent and reentry. The skin differential pressure and pressure inside the systems tunnel relative to freestream pressure from CHCHVENT as well as velocity vectors and streamlines are also discussed in detail.

Wang, Qunzhen↗
NASA NTRS2011

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

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11-ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure and surface flow visualization data were obtained in both facilities but only the force and moment data are presented herein. Nacelle/pylon, tail effects and tunnel to tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon gave an increase in drag, decrease in lift and a less nose down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, all of the data shown fall within the 2-sigma limits for repeatability. The tunnel to tunnel differences are negligible for lift and pitching moment, while the drag shows a difference of less than ten counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

Rivers, S. M.↗
NASA NTRS2014

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

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.

Aeroelasticity↗
NASA NTRS1995

Fundamental and subharmonic excitation for an oscillator with several tunneling diodes in series

Connecting several tunneling diodes in series shows promise as a method for increasing the output power of these devices as millimeter-wave oscillators. However, due to the negative differential resistance (NDR) region in the dc I-V curve of a single tunneling diode, a circuit using several devices connected in series, and biased simultaneously in the NDR region, is dc unstable. Because of this instability, an oscillator with several tunneling diodes in series has a demanding excitation condition. Excitation using an externally applied RF signal is one approach to solving this problem. This is experimentally demonstrated using an RF source, both with frequency close to as well as with frequency considerably lower than the oscillation frequency. Excitation by an RF (radio frequency) source with a frequency as low as one sixth of the oscillation frequency was demonstrated in a proof-of-principle experiment at 2 GHz, for an oscillator with two tunnel diodes connected in series. Strong harmonics of the oscillation signal were generated as a result of the highly nonlinear dc I-V curve of the tunnel diode and a large signal oscillator design. Third harmonic output power comparable to that of the fundamental was observed in one oscillator circuit. If submillimeter wave resonant-tunneling diodes (RTD's) are used instead of tunnel diodes, this harmonic output may be useful for generating signals at frequencies well into the terahertz range.

SUBHARMONIC EXCITATION↗
NASA NTRS2016

Pre-Test Assessment of the Use Envelope of the Normal Force of a Wind Tunnel Strain-Gage Balance

The relationship between the aerodynamic lift force generated by a wind tunnel model, the model weight, and the measured normal force of a strain-gage balance is investigated to better understand the expected use envelope of the normal force during a wind tunnel test. First, the fundamental relationship between normal force, model weight, lift curve slope, model reference area, dynamic pressure, and angle of attack is derived. Then, based on this fundamental relationship, the use envelope of a balance is examined for four typical wind tunnel test cases. The first case looks at the use envelope of the normal force during the test of a light wind tunnel model at high subsonic Mach numbers. The second case examines the use envelope of the normal force during the test of a heavy wind tunnel model in an atmospheric low-speed facility. The third case reviews the use envelope of the normal force during the test of a floor-mounted semi-span model. The fourth case discusses the normal force characteristics during the test of a rotated full-span model. The wind tunnel model's lift-to-weight ratio is introduced as a new parameter that may be used for a quick pre-test assessment of the use envelope of the normal force of a balance. The parameter is derived as a function of the lift coefficient, the dimensionless dynamic pressure, and the dimensionless model weight. Lower and upper bounds of the use envelope of a balance are defined using the model's lift-to-weight ratio. Finally, data from a pressurized wind tunnel is used to illustrate both application and interpretation of the model's lift-to-weight ratio.

strain-gage balance↗
NASA NTRS2015

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

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-ε 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 variable. Power differences between free field and wind tunnel cases were found from -7 % to 0 % in the 80- by 120-Foot Wind Tunnel test section and -1.6 % to 4.8 % in the 40- by 80-Foot Wind Tunnel, depending on the TTR orientation, tunnel velocity and blade setting. The TTR will be used in 2016 to test the Bell 609 rotor in a similar fashion to the research in this report.

CFD↗
NASA NTRS2016

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

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-ε 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 for the rotor to compare the power as a unique variable. Power differences between free field and wind tunnel cases were found from -7 to 0 percent in the 80- by 120-Foot Wind Tunnel and -1.6 to 4.8 percent in the 40- by 80-Foot Wind Tunnel, depending on the TTR orientation, tunnel velocity, and blade setting. The TTR will be used in 2016 to test the Bell 609 rotor in a similar fashion to the research in this report.

CFD↗
NASA NTRSDate not supplied

Computational Fluid Dynamics Simulations of the Transonic Dynamics Tunnel Airstream Oscillator System

This paper presents a Computational Fluid Dynamics (CFD) model of the flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT) with the Airstream Oscillator System (AOS) in operation. The TDT is a continuous-flow, closed circuit wind tunnel with a 16- by 16-foot slotted test section with cropped corners. The tunnel was originally built as the 19-ft Pressure Tunnel in 1938, but it was converted to the current transonic tunnel in the 1950s, with capabilities to use either air or heavy gas (R-134a) as the test medium. The TDT was also fitted with an AOS which can be used to create a gust field in the test section. To date, no computational analyses of the TDT involving the AOS have been performed. In this study, experimental data acquired of an oscillating airstream in the tunnel will be used to calibrate the computational analyses. A significant motivation of this work is to attempt to compare computational gust velocities to those recorded in the TDT literature. In addition to a validation of this model with experimental data, it may also be possible to supplement the experimental data with computational data. The experimental data is rather sparse and at only a few Mach numbers. Computational results may be able to expand the AOS data set. Another motivation for this work is that the Integrated Adaptive Wing Technology Maturation (IAWTM) semi-span model of the high-aspect-ratio CRM with 10 trailing edge control surfaces is currently being fabricated and will be delivered to the NASA Langley Transonic Dynamics Tunnel (TDT) for testing in late 2020. Among tests to be conducted, gust load alleviation (GLA) will be demonstrated at transonic conditions using the AOS.

Computational Fluid Dynamics↗
NASA NTRSDate not supplied

Mars Sample Return, Sample Retrieval Lander, Reaction Control System Jet Interaction Supersonic Wind Tunnel Test Overview with CFD Predictions

NASA's Mars Sample Return campaign will be launching several missions over the next decade that will work together to return rock samples from Mars back to Earth. The Sample Retrieval Lander (SRL) will deliver the Mars Ascent Vehicle and fetch rover to the surface of Mars in 2006. Rock samples collected by the Mars 2020 Perseverance rover, landing in early 2021, will be loaded on the the ascent vehicle to be launched into Mars orbit for retrieval by yet another spacecraft. The Sample Retrieval Lander will be a blunt entry capsule similar to past Mars entry vehicles like Mars Science Laboratory and Viking. The vehicle will fly a guided entry, using a small lift vector produced by a non-zero trim angle of attack to eliminate downrange and crossrange position errors at the point of parachute deploy. This energy and heading management is achieved with a reaction control system (RCS) that directs the bank angle of the vehicle and also minimizes unwanted capsule dynamics. The reaction control system and control design is based on the Mars Science Laboratory and Mars 2020 RCS systems. However, due to packaging constraints, the backshell of this new entry vehicle has a different geometry than those earlier designs. To certify the RCS system for flight the project must characterize the jet plume interactions with the capsule backshell that could impair or significantly augment the RCS control authority. This characterization will be done through a combination of computational fluid dynamics (CFD) analysis and wind tunnel test. Two candidate arrangements of the RCS jets have been identified for the SRL vehicle and are currently under evaluation before final selection. The aero/RCS plume interactions of these candidate configurations have been measured in a supersonic wind tunnel test in NASA Langley's Unitary Plan Wind Tunnel. The test was conducted in the fall of 2020 and data is currently being reduced. An overview of the candidate RCS configurations are presented here with an overview of the wind tunnel model design, jet scaling and scaled nozzle design, and the test matrix. Preliminary CFD runs are presented with an assessment of the predicted plumes and their interaction with the wake flow of the vehicle. The predicted effects of the model sting is provided as well. This high fidelity wind tunnel test is being conducted much earlier in the SRL project than would normally be done. The test was funded as part of a CFD evaluation task funded by NASA's Aerosciences Evaluation and Test Capabilities Project. The objective of the evaluation task was to compare the ability of CFD to predict complex flows with data that can be measured in the Langley Unitary Plan Wind Tunnel. RCS Jet interactions were selected as a type of complex flow that is important to NASA missions. In addition to providing useful data to the SRL project, there was added emphasis on quantifying the accuracy of the CFD predictions and wind tunnel test data. An overview of the uncertainty quantification methodologies for computational and experimental portions of this test is presented.

blunt body↗
NASA NTRS2022

Interaction of a Tunnel-like Acoustic Disturbance Field with a Blunt Cone Boundary Layer at Mach 8

The existing measurements of laminar-to-turbulent transition over circular cones in conventional (i.e., “noisy”) hypersonic wind tunnels have established that the transition location moves downstream when the nose radius is increased from zero. However, this initially downstream movement slows down and ultimately reverses beyond a critical value of the nose radius, and may be related to external forcing in the form of freestream disturbances and/or surface roughness. To understand the effects of freestream acoustic disturbances on transition reversal over a blunt body, hypersonic boundary-layer receptivity to broadband freestream acoustic disturbances from the nozzle wall of a digital conventional wind tunnel is investigated by both direct numerical simulations (DNS) and modal and nonmodal stability analysis. A Mach 8 flow over a 7 deg half-angle cone with a nose radius of 𝑅 𝑛 = 5.2mm and freestream Reynolds number of 12.2 × 10 6 m -1 is considered. The results show that the broadband tunnel noise in the free stream of a convectional hypersonic wind tunnel (i.e., outside of the nozzle-wall turbulent boundary layer) can be well represented by an acoustic model with an ansatz of slow acoustic waves. With successful calibration of the model parameters against the precursor tunnel DNS, such an acoustic ansatz can successfully reproduce both the frequency-wavenumber spectra and the temporal evolution of the broadband tunnel noise radiated from the nozzle wall. Additionally, the DNS of the Mach 8 blunt cone with tunnel-like acoustic input above the bow shock showed that the spectra of wall-pressure and heat-transfer fluctuations recovers the signature of the axisymmetric waves predicted by the nonmodal analysis. Furthermore, the azimuthal wavenumber and frequency spectrum of the temperature fluctuations as a function of the wall-normal distance show higher amplitudes for three-dimensional waves above the boundary-layer edge. The numerical schlieren contours show the inclined structures commonly observed in blunt cone experiments, demonstrating that they correspond to three-dimensional structures due to freestream disturbances in the presence of an entropy layer.

Boundary-layer transition↗
NASA NTRSDate not supplied

In Pursuit of CFD-based Wind Tunnel Calibrations

Computational fluid dynamic 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 a sufficient characterization of both the physical geometry of the wind tunnel, the thermodynamics of the tunnel, and flow conditions in the test section. The condition setting process among subsonic wind tunnels, closed or open throat, is generally the same, though they can differ in specific details of the thermodynamics. The derivation, application, and error estimation of condition setting and calibration is discussed. Computational flow solutions of the high-speed leg of the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel are used to evaluate the calibration process and discuss calibration curve sensitivities. Duplicating the characteristics of a particular wind tunnel is difficult at best. A calibration curve derived from the computational method is the most consistent method to use for tunnel condition setting.

wind tunnel↗
NASA NTRS1973

Aeroacoustic research in wind tunnels: A status report

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.

Bender, J.↗
NASA NTRS1974

The cryogenic wind tunnel for high Reynolds number testing

Experiments performed at the NASA Langley Research Center in a cryogenic low-speed continuous-flow tunnel and in a cryogenic transonic continuous-flow pressure tunnel have demonstrated the predicted changes in Reynolds number, drive power, and fan speed with temperature, while operating with nitrogen as the test gas. The experiments have also demonstrated that cooling to cryogenic temperatures by spraying liquid nitrogen directly into the tunnel circuit is practical and that tunnel temperature can be controlled within very close limits. Whereas most types of wind tunnel could operate with advantage at cryogenic temperatures, the continuous-flow fan-driven tunnel is particularly well suited to take full advantage of operating at these temperatures. A continuous-flow fan-driven cryogenic tunnel to satisfy current requirements for test Reynolds number can be constructed and operated using existing techniques. Both capital and operating costs appear acceptable.

Kilgore, R. A.↗
NASA NTRS1974

Some tunnel-wall effects on transonic flutter

Significant effects of wind-tunnel walls were observed on the transonic flutter boundaries of wall-mounted models during two flutter model research studies. In these studies, flutter experiments with cantilevered SST-type wing models were conducted in three different wind tunnels. The experimental results are compared to flutter boundaries calculated for the models in free air. The results indicate that transonic flutter boundaries can be affected by tunnel-wall interference, tunnel resonances, and shock-wave reflections, and that flutter model data accuracy is a function of model/tunnel size and tunnel wall porosity. However, models within the recommended size limits should give accurate results in transonic tunnels with normal ventilation. A flutter trend analysis for a two-dimensional wing-demonstrating tunnel wall and resonance effects on flutter are also presented.

Ruhlin, C. L.↗
NASA NTRS1976

Correlation of low speed wind tunnel and flight test data for V/STOL aircraft

The availability of wind tunnel test data for correlation purposes of the same V/STOL aircraft tested in flight is very limited. This is due in a large part to size limitations of wind tunnels and the number of wind tunnels available for testing of full-scale aircraft. Wind tunnel tests are described for two research aircraft - the XV-5B fan-in-wing aircraft and the YOV-10 RCF (rotating cylinder flap) aircraft - in the NASA Ames 40- by 80-foot wind tunnel. The tests were conducted specifically to provide for correlation between wind tunnel and in-flight aerodynamics and noise test data. Correlation between aerodynamic and noise data are presented and testing techniques that are related to the accuracy of the data, or that might affect the correlations, are discussed. The correlation of noise measurements made with a J-85 engine mounted on a F-106 aircraft during low altitude flyovers with the same J-85 engine mounted on a model and tested in the Ames 40- by 80-foot wind tunnel are also reported.

Cook, W. L.↗