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At least 901 records · Page 50
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

Wind tunnel↗
NASA NTRS2014

Space Launch System Ascent Static Aerodynamic Database Development

This paper describes the wind tunnel testing work and data analysis required to characterize the static aerodynamic environment of NASA's Space Launch System (SLS) ascent portion of flight. Scaled models of the SLS have been tested in transonic and supersonic wind tunnels to gather the high fidelity data that is used to build aerodynamic databases. A detailed description of the wind tunnel test that was conducted to produce the latest version of the database is presented, and a representative set of aerodynamic data is shown. The wind tunnel data quality remains very high, however some concerns with wall interference effects through transonic Mach numbers are also discussed. Post-processing and analysis of the wind tunnel dataset are crucial for the development of a formal ascent aerodynamics database.

Pinier, Jeremy T.↗
NASA NTRS1989

Flutter testing of modern aircraft

Structural dynamicists avoid aircraft structural flutter through the use of flutter analyses, wind tunnel tests, ground-vibration tests, and flight flutter testing. FEM and unsteady aerodynamics models are often employed in analyses whose results' accuracies are verified by wind tunnel test results. Ground-vibratiuon testing is used to ascertain an airframe's resonant modes of vibration and their associated frequencies and damping rates; these data are then compared to the FEM analysis results. Flutter wind tunnel testing serves the same purpose for unsteady aerodynamic analysis as ground vibration testing does for the vibration analysis. Finally, flight flutter testing ensures that no flutter is present at any point in the envelope.

Johnson, Jeff↗
NASA NTRSDate not supplied

Overview of Mars Sample Return – Earth Entry System Woven Roughness Heating Augmentation Test in NASA Langley’s Mach 6 Wind Tunnel

The Mars Sample Return Mission (MSR) is a planned NASA flagship mission in which a sample retrieval lander (SRL) with a rover will be flown to Mars to obtain sample tubes on the surface that were dropped by the Mars 2020 rover [1]. After obtaining the sam-ples, the rover will return and ascend back to Martian orbit onboard the Mars Ascent Vehicle (MAV). Upon return to Earth orbit, the samples will perform Entry, Descent, and Landing (EDL) with the Earth Entry Sys-tem (EES) architecture, and land in Utah. The EES vehicle will utilize a HEEET-variant as its TPS, which will be the first time a woven TPS will be used on a flagship NASA mission [2]. This TPS offers a unique challenge for Computational Fluid Dynamics (CFD) modeling of the aerothermal envi-ronment of the vehicle, as woven roughness heating augmentation has not been extensively investigated experimentally. As a result, in order to validate com-putational models for woven roughness heating aug-mentation, a wind tunnel test campaign at NASA Langley Research Center’s Mach 6 wind tunnel was performed in April of 2023. This test campaign consisted of over a hundred runs with Reynolds numbers spanning from 1-7 mil-lion 1/ft and with six separate wind tunnel models used. A second campaign with a suite of new models will be conducted in Summer 2023 as well as a cam-paign with a flat plate model, both of which are of great interest to the MSR-EES project. The data obtained from this test are extremely vital for the MSR mission, as they will validate CFD roughness heating models which will be directly used to design the TPS of the EES portion of MSR and characterize the heating environment that the entry ve-hicle will experience. Further extensions of the MSR-EES test campaign will continue to provide validation data for developing more effective computational tools.

Jonathan Cheatwood↗
NASA NTRS1975

Test data from solid propellant plume aerodynamics test program in Ames 6 x 6 foot supersonic wind tunnel (shuttle test FA7) (Ames test 033-66)

The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solid propellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.

Hair, L. M.↗
NASA NTRSDate not supplied

LA-8 Computational Analysis and Validation Studies Using FlightStream

With the emerging market of Urban Air Mobility and the associated unique new flight vehicles with the capability of vertical takeoff and landing, design tools need to be calibrated to assist in the development of these new flight vehicle concepts. Distributed electric propulsion (DEP) covers these vehicles with high-energy flow from the array of motors, which leads to extensive propulsion/airframe interaction effects. With limited experimental data of vehicles using DEP and vehicle operation across a large range of angles of attack, including post-stall conditions, there is a lack of validated design tools to advance new vehicle concepts that have blown wings and operate in multiple flight modes. Although some CFD tools can handle the integration of multiple propulsors and unique geometries, the computational time with a full vehicle is extraordinary. FlightStream, a surface vorticity flow solver that can run in a fraction of the time of other CFD methods, has been evaluated using the Langley Aerodrome No. 8 (LA-8) vehicle as a test case. The LA-8 vehicle is a DEP tandem tilt-wing research vehicle that was designed, built, and tested at NASA Langley Research Center. The LA-8 design features both high-performance and high-technical risks aspects. The vehicle has gone through extensive design of experiments wind tunnel testing in the NASA Langley 12-Foot Low-Speed Wind Tunnel in order to capture all modes of operation. In addition, the LA-8 vehicle has undergone wind tunnel testing with and without propellers, which contributes to the validation of FlightStream for aerodynamic performance in blown and unblown wing cases. This study shows that FlightStream can produce trends and magnitudes of the lift and drag coefficients similar to the data collected from the wind tunnel testing of LA-8. The test cases were both blown and unblown wing cases and with and without high lift control surface deflections.

Steven Curtis Geuther↗
NASA NTRS1983

Aerodynamic analysis of the loft anomaly observed on orbital flight tests of the space shuttle

The loft anomaly observed during the launch phase of the orbital flight tests (OFT) of the Space Shuttle is described and evidence that the loft anomaly resulted from previously unobserved aerodynamic phenomena is presented. The anomaly was that the altitude at staging was higher than anticipated. The anticipated altitude profile was predicted on wind tunnel test results that did not accurately simulate the flow between the orbiter vehicle and external tank and did not adequately simulate the engine plumes and thus the base pressures. An analogy is used to relate the flow between the orbiter and external tank to the flow in a two dimensional channel. Plume simulation is identified as a major goal during wind tunnel testing, and a wind tunnel test that was conducted to provide the best possible representation of the plume effect on the channel flow field is described.

Surber, T. E.↗
NASA NTRS1975

Flight Flutter Testing of the P6M

On the P6M the shake behavior, i.e., the response to random excitation at subcritical speeds of lowly damped airplane modes, is as important as the actual flutter speed. The approach is to first study the problem by means of analyses and wind-tunnel tests. These predictions are compared with flight test data obtained by spectral analysis of tape recordings of the airplane vibration responses to random aerodynamic turbulence. A similar spectrum analysis approach was used in high speed wind-tunnel tests. A resonance excitation technique was developed for low speed wind-tunnel testing, and well defined V-g curves were obtained. The effect of various parameters on both shake and flutter of T-tails with and without dihedral were studied. Preliminary flight tests yielded good correlation; they also yielded interesting information concerning a low frequency transonic snaking mode, and excitation by shed vortices.

Kachadourian, G.↗
NASA NTRS2013

Aerodynamic Stability and Performance of Next-Generation Parachutes for Mars Descent

The Low Density Supersonic Decelerator Project is developing a next-generation supersonic parachute for use on future Mars missions. In order to determine the new parachute configuration, a wind tunnel test was conducted at the National Full-scale Aerodynamics Complex 80- by 120-foot Wind Tunnel at the NASA Ames Research Center. The goal of the wind tunnel test was to quantitatively determine the aerodynamic stability and performance of various canopy configurations in order to help select the design to be flown on the Supersonic Flight Dynamics tests. Parachute configurations included the diskgap- band, ringsail, and ringsail-variant designs referred to as a disksail and starsail. During the wind tunnel test, digital cameras captured synchronized image streams of the parachute from three directions. Stereo hotogrammetric processing was performed on the image data to track the position of the vent of the canopy throughout each run. The position data were processed to determine the geometric angular history of the parachute, which were then used to calculate the total angle of attack and its derivatives at each instant in time. Static and dynamic moment coefficients were extracted from these data using a parameter estimation method involving the one-dimensional equation of motion for a rotation of parachute. The coefficients were calculated over all of the available canopy states to reconstruct moment coefficient curves as a function of total angle of attack. From the stability curves, useful metrics such as the trim total angle of attack and pitch stiffness at the trim angle could be determined. These stability metrics were assessed in the context of the parachute's drag load and geometric porosity. While there was generally an inverse relationship between the drag load and the stability of the canopy, the data showed that it was possible to obtain similar stability properties as the disk-gap-band with slightly higher drag loads by appropriately tailoring the geometric porosity distribution.

Low Density Supersonic Decelerator (LDSD)↗
NASA NTRS1982

Evaluation of wind tunnel performance testings of an advanced 45 deg swept 8-bladed propeller at Mach numbers from 0.45 to 0.85

The increased emphasis of fuel conservation in the world and the rapid increase in the cost of jet fuel has stimulated a series of studies of both conventional and unconventional propulsion systems for commercial aircraft. The results of these studies indicate that a fuel saving of 15 to 30 percent may be realized by the use of an advanced high-speed turboprop (Prop-Fan) compared to aircraft equipped with high bypass turbofan engines of equivalent technology. The Prop-Fan propulsion system is being investigated as part of the NASA Aircraft Energy Efficient Program. This effort includes the wind tunnel testing of a series of 8 and 10-blade Prop-Fan models incorporate swept blades. Test results indicate efficiency levels near the goal of 80 percent at Mach 0.8 cruise and an altitude of 10.67 km (35,000 ft). Each successive swept model has shown improved efficiency relative to the straight blade model. The fourth model, with 45 deg swept blades reported herein, shows a net efficiency of 78.2 at the design point with a power loading of 301 kW/sq meter and a tip speed of 243.8 m/sec (800 ft/sec.).

Rohrbach, C.↗
NASA NTRS2015

Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel

Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.

Diffuser↗
NASA NTRS2018

Independent Effects of Reynolds and Mach Numbers 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 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%-scale semispan wing based upon the Common Research Model airplane configuration. The wind-tunnel testing was conducted at the ONERA F1 pressurized wind tunnel with Reynolds numbers of 1.6×10(exp 6) to 11.9×10(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, leadingedge 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 icedairfoils. 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.

Broeren, Andy P.↗
NASA NTRS1982

Results of buffet tests in a cryogenic wind tunnel

Buffet tests on two semispan wing models with different leading edge sweep show that it is feasibile to use the standard dynamic wing root bending moment technique in a cryogenic wind tunnel. One model was a slender 65 deg swept delta wing with sharp leading edges. The other model was an unswept wing of aspect ratio 1.5 with a British NPL 9510 airfoil section. The results for the 65 deg swept 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 static aeroelastic distortion by variations in the tunnel stagnation temperature and pressure were demonstrated.

Boyden, R. P.↗
NASA NTRS1984

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

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.

Kilgore, R. A.↗
NASA NTRS2025

Subsonic Ultra Green Aircraft Research: Phase V – Buffet Test Report

This test report summarizes work performed by the Boeing Subsonic Ultra-Green Aircraft Research (SUGAR) team in the Phase V Transonic Truss-Braced Wing contract task. The task was awarded in September 2020, and ended at the end of September 2022. As a part of the SUGAR Phase V contract task, Boeing developed a wind tunnel test plan, model requirements, then fabricated and conducted a dedicated transonic buffet wind tunnel test. The test focused on the investigation of buffet onset, and the ability to accurately predict buffet onset for the TTBW configuration. This report summarizes the results and post-test analysis of the 9% scale, semispan, Transonic Truss-Braced Wing (TTBW) transonic buffet model tested at the NASA Ames Research Center Unitary Plan Wind Tunnel (UPWT) 11-Foot Transonic Wind Tunnel (11-Ft TWT) facility located at Moffett Field, CA. It was tested from January 24, 2022, through February 18, 2022. The wind tunnel test was successful in capturing buffet onset at the model design dynamic pressure. Multiple configurations were tested starting in the wing-body-strut configuration and ending with the full configuration including wing-body-strut-nacelle/pylon, and flap hinge. The test also measured the effect of control surface deflections (ailerons and spoilers), as well as the effect of an outboard strut flap. Steady data collected during the test included standard force and moment measurements, as well as wing, strut, nacelle, and body (fuselage) static pressures. These data were collected in a standard pitch-pause mode. Unsteady data captured in the test, collected in continuous pitch mode, included wing and strut dynamic pressures, accelerations, and strains. According to test data, previous concerns regarding critical buffet onset on the strut (at low vehicle lift coefficient) and in the wing-strut channel have been mitigated. Buffet onset on the TTBW appears to follow conventional wing-induced buffet onset mechanisms. However, the critical station where buffet onset first occurs has moved inboard to the wing-strut juncture span location (55-60% span) from the typical ~70% span location. Test data obtained by unsteady pressure transducers provided a good match to expected buffet onset modes. An assessment of different tools and methodologies for predicting buffet onset was conducted. Reasonable agreement was found for buffet boundary predictions between most of the methods examined when using a forward trip location, with the greatest discrepancies from the pitching moment break and trailing edge pressure divergence methods. When an aft trip location was used the buffet boundary predictions diverged more significantly from unsteady pressure measurements. In addition, good agreement between the buffet onset trends were found when the Reynolds number, configuration, or spanload was changed.

strut↗
NASA NTRS1985

The space shuttle ascent vehicle aerodynamic challenges configuration design and data base development

The phase B Space Shuttle systems definition studies resulted in a generic configuration consisting of a delta wing orbiter, and two solid rocket boosters (SRB) attached to an external fuel tank (ET). The initial challenge facing the aerodynamic community was aerodynamically optimizing, within limits, this configuration. As the Shuttle program developed and the sensitivities of the vehicle to aerodynamics were better understood the requirements of the aerodynamic data base grew. Adequately characterizing the vehicle to support the various design studies exploded the size of the data base to proportions that created a data modeling/management challenge for the aerodynamicist. The ascent aerodynamic data base originated primarily from wind tunnel test results. The complexity of the configuration rendered conventional analytic methods of little use. Initial wind tunnel tests provided results which included undesirable effects from model support tructure, inadequate element proximity, and inadequate plume simulation. The challenge to improve the quality of test results by determining the extent of these undesirable effects and subsequently develop testing techniques to eliminate them was imposed on the aerodynamic community. The challenges to the ascent aerodynamics community documented are unique due to the aerodynamic complexity of the Shuttle launch. Never before was such a complex vehicle aerodynamically characterized. The challenges were met with innovative engineering analyses/methodology development and wind tunnel testing techniques.

Dill, C. C.↗