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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 433 records · Page 24
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NASA NTRS2024

Wind Tunnel Test of A Scale Model of A Venus Probe to Determine Aeroacoustics Environment

The Zephyr probe for NASA’s DAVINCI project is being designed to fly to Venus and collect measurements to characterize the Venusian atmosphere while it descends towards the surface. Despite a slow descent speed as the probe approaches the surface, the very dense atmosphere means the vibro-acoustic environment that the probe must withstand could be very harsh. A wind tunnel test campaign was conducted to characterize the surface-pressure fluctuations (acoustics) on the external surface of the probe during the descent. The goal was to provide forcing functions for vibro-acoustic analysis of all instruments inside the probe. A 25% scale-model of the Zephyr probe was manufactured, instrumented with 18 microphones, and tested in two wind tunnels located in the Fluid Mechanics Lab at NASA Ames Research Center at velocities ranging from 13 m/s to 48 m/s. Trip strips were added to the model to ensure a turbulent boundary layer on the model. Microphones were located appropriately to provide spectra of pressure fluctuations in regions with different flow characteristics as well as for calculating various two-point statistics. The acoustic spectra measured in the wind tunnel tests were scaled up to flight-scale using the estimated flight trajectory information. Besides fluctuating surface-pressure measurements, hot-film velocity measurements as well as smoke-laser flow visualization were conducted to better understand the flow around the probe. Lastly, several different geometry configurations were tested to see whether any reduction in acoustic levels could be achieved, but none of the configurations that were tested offered significant improvements. The flow visualization showed 3 primary flow regimes on the model: an attached turbulent boundary layer on the front of the model, a region of separated flow upstream of the flared lip, and a large, separated wake at the rear of the model. The acoustic spectra generally show three different shapes corresponding to these flow regions, with the highest levels seen at the rear of the model, and directly in front of the drag plate. The scaled overall levels of pressure fluctuations were found to be high: in the range of 130dB to 155dB.

planetary probe↗
NASA NTRS1958

Investigation of the Static Lateral Stability and Aileron Characteristics of a 0.067-Scale Model of the Bell X-2 Airplane at Mach Numbers of 2.29, 2.78, 3.22, and 3.71

Tests were performed in the high. Mach number test section of the Langley Unitary Plan wind tunnel to determine the static lateral stability. and aileron characteristics of a 0.067-scale model of the Bell X-2 airplane at Mach numbers of 2.29, 2. 78, 3.22, and. 3.71. The results of this investigation indicated that the directional stability of the model was low with directional instability occurring at Mach numbers higher than 3.1 and. angles of attack higher than about 5.0 deg (equivalent lift coefficient of about 0.18). The yaw due to aileron deflection was adverse and, with 10 deg of differential aileron deflection, large enough to overbalance the available directional restoring moment at all angles of attack higher than about 5.0 deg (equivalent lift coefficient of about 0.21) and Mach numbers higher than 2. 5. The model also had positive effective dihedral for all test attitudes and. Mach numbers. A combination of the lateral-stability parameters with the aileron characteristics to form a lateral-stability criterion for a maneuver using ailerons alone indicated that the model has characteristics which would. give unstable aperiodic behavior (divergence) over a large part of the test Mach number and angle-of-attack range.

Fournier, Roger H.↗
NASA NTRS1987

Wind-tunnel free-flight investigation of a 0.15-scale model of the F-106B airplane with vortex flaps

An investigation to determine the effects of vortex flaps on the flight dynamic characteristics of the F-106B in the area of low-speed, high-angle-of-attack flight was undertaken on a 0.15-scale model of the airplane in the Langley 30- by 60-Foot Tunnel. Static force tests, dynamic forced-oscillation tests, as well as free-flight tests were conducted to obtain a data base on the flight characteristics of the F-106B airplane with vortex flaps. Vortex flap configurations tested included a full-span gothic flap, a full-span constant-chord flap, and a part-span gothic flap.

Yip, Long P.↗
NASA NTRS1947

An Investigation of the Aerodynamic Characteristics of an 0.08-Scale Model of the Chance Vought XF7U-1 Airplane in the Langley High-Speed 7- by 10-Foot Tunnel. Part I - Basic Longitudinal Stability Characteristics, TED No. NACA DE308: Basic Longitudinal Stability Characteristics, TED No. NACA DE308 - Part 1

The stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane have been investigated over a Mach number range from 0.40 to 0.91. Results of the basic longitudinal tests of the complete model with undeflected control surfaces are given in the present report with a very limited analysis of the results.

Kemp, William B., Jr.↗
NASA NTRS1959

An Investigation of a 0.05-Scale Model of the XSM-64A Navaho Missile and Booster: Force Study at Mach Numbers from 1.77 to 3.51 - Part 1

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic loads and the static longitudinal and lateral stability of a 0.05-scale model of the XSM-64A Navaho missile and booster and its various components. Tests were conducted through a Mach number range of 1.77 to 3.51 with a corresponding Reynolds number range of 2.4 x 10(exp 6) to 2.9 x 10(exp 6). Results are presented for an angle-of-attack range of -8 deg to 4 deg for the missile-booster combination and -10 deg to 10 deg for the missile-alone configuration. Tests for both configurations were conducted through an angle-of-sideslip range of -8 deg to 8 deg. Also presented are some effects on the model characteristics of the deflection of various components including canard, tip aileron, vertical stabilizer, speed brakes, and booster pitch and yaw thrust chambers. The various components on which loads were measured include the wing, tip aileron, rudder, booster, booster separating surface, booster fin, and booster yaw and pitch thrust chambers. These data are presented without analysis.

Church, James D.↗
NASA NTRS1954

Preliminary Investigation of the Static Longitudinal and Lateral Stability Characteristics of a 0.05-Scale Model of the Convair F2Y-1 Airplane at High Subsonic Speeds

At the request of the Bureau of Aeronautics, Department of the Navy, National Advisory Committee for Aeronautics has conducted a preliminary investigation at high subsonic speeds of the static longitudinal and lateral stability characteristics of a 0.05-scale model of the Convair F2Y-1 water-based fighter airplane. The tests covered a Mach number range from 0.5 to 0.94 and corresponding Reynolds numbers, based on the wing mean aerodynamic chord, from 3.3 x 10(exp 6) to 4.3 x 10(exp 6). The maximum angle-of-attack range (obtained at the lower Mach numbers) was from -2 degrees to 25 degrees. Sideslip angles from -4 degrees to 12 degrees also were investigated. The investigation included effects of various arrangements of wing fences and of rocket packages.

Spreemann, Kenneth P.↗
NASA NTRS2015

Space Launch System Scale Model Acoustic Test Ignition Overpressure Testing

The overpressure phenomenon is a transient fluid dynamic event occurring during rocket propulsion system ignition. This phenomenon results from fluid compression of the accelerating plume gas, subsequent rarefaction, and subsequent propagation from the exhaust trench and duct holes. The high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and surrounding structure. Commonly known as ignition overpressure (IOP), this is an important design-to environment for the Space Launch System (SLS) that NASA is currently developing. Subscale testing is useful in validating and verifying the IOP environment. This was one of the objectives of the Scale Model Acoustic Test, conducted at Marshall Space Flight Center. The test data quantifies the effectiveness of the SLS IOP suppression system and improves the analytical models used to predict the SLS IOP environments. The reduction and analysis of the data gathered during the SMAT IOP test series requires identification and characterization of multiple dynamic events and scaling of the event waveforms to provide the most accurate comparisons to determine the effectiveness of the IOP suppression systems. The identification and characterization of the overpressure events, the waveform scaling, the computation of the IOP suppression system knockdown factors, and preliminary comparisons to the analytical models are discussed.

Nance, Donald↗
NASA NTRS2015

Space Launch System Scale Model Acoustic Test Ignition Overpressure Testing

The overpressure phenomenon is a transient fluid dynamic event occurring during rocket propulsion system ignition. This phenomenon results from fluid compression of the accelerating plume gas, subsequent rarefaction, and subsequent propagation from the exhaust trench and duct holes. The high-amplitude unsteady fluid-dynamic perturbations can adversely affect the vehicle and surrounding structure. Commonly known as ignition overpressure (IOP), this is an important design-to environment for the Space Launch System (SLS) that NASA is currently developing. Subscale testing is useful in validating and verifying the IOP environment. This was one of the objectives of the Scale Model Acoustic Test (SMAT), conducted at Marshall Space Flight Center (MSFC). The test data quantifies the effectiveness of the SLS IOP suppression system and improves the analytical models used to predict the SLS IOP environments. The reduction and analysis of the data gathered during the SMAT IOP test series requires identification and characterization of multiple dynamic events and scaling of the event waveforms to provide the most accurate comparisons to determine the effectiveness of the IOP suppression systems. The identification and characterization of the overpressure events, the waveform scaling, the computation of the IOP suppression system knockdown factors, and preliminary comparisons to the analytical models are discussed.

Nance, Donald K.↗
NASA NTRS1998

Rotating Pressure and Temperature Measurements on Scale Model Fans Using Luminescent Paints

Pressure and temperature sensitive paint (PSP, TSP) technology is a rapidly developing measurement technique that is widely accepted in external aerodynamic wind tunnel measurements. Almost all quantitative full field measurements have been on stationary surfaces. Rotating pressure and temperature measurements were successfully acquired during scale-model fan tests in the NASA Lewis Research Center 9 foot by 15 foot low speed wind tunnel at speeds as high as 9500 RPM. The use of PSP and TSP on rotating surfaces presents additional challenges to data acquisition. This paper describes in detail the techniques used to acquire global rotating pressure and temperature measurements and overcome the problems associated with this type of measurement. These include illumination and detection problems associated with moving surfaces, temperature dependence of PSP and limited emitted light available from short duration illumination sources. Solutions to these problems as well as pressure and temperature results will be discussed.

Bencic, Timothy J.↗
NASA NTRS2002

A New Statistically based Autoconversion rate Parameterization for use in Large-Scale Models

The autoconversion rate is a key process for the formation of precipitation in warm clouds. In climate models, physical processes such as autoconversion rate, which are calculated from grid mean values, are biased, because they do not take subgrid variability into account. Recently, statistical cloud schemes have been introduced in large-scale models to account for partially cloud-covered grid boxes. However, these schemes do not include the in-cloud variability in their parameterizations. In this paper, a new statistically based autoconversion rate considering the in-cloud variability is introduced and tested in three cases using the Canadian Single Column Model (SCM) of the global climate model. The results show that the new autoconversion rate improves the model simulation, especially in terms of liquid water path in all three case studies.

Lin, Bing↗
NASA NTRS1979

Results of transonic tests in the NASA/MSFC 14-inch trisonic wind tunnel on a 0.004-scale model (74-OTS) space shuttle launch vehicle (FA25)

The primary objective of the test was to determine the aerodynamic increments due to the attach structure. Secondary objectives were to determine the effects of: (1) orbiter nose mold line changes; (2) wire bundle fairings on data measurements; and (3) flow angularity. The scale model was tested over the Mach range from 0.60 to 1.25 at angles of attack and sideslip from -8 to +8 deg. The total pressure was 22 psia for all runs. Six-component orbiter data were obtained from a balance in the orbiter which was sting supported. The external tank was attached to the solid rocket booster, each of which was sting supported. An alternate two sting/two balance arrangement was also tested with a single sting and balance in the external tank measuring combined ET/SRB aero data replacing the two stings in the SRB's. Two runs were also made at Mach number 4.96 with the two SRB's removed. The aerodynamic coefficients obtained are tabulated as a function of angle of attack or sideslip for each Mach number value.

Lundy, T. E.↗
NASA NTRS1976

An investigation of the aerodynamic characteristics of a 0.00548 scale model (model no. 486) of the space shuttle 146-inch diameter solid rocket booster at angels of attack from 113 deg to 180 deg in the AEDC PWT 4-foot transonic wind tunnel (SA16F)

An experimental investigation (SA16F) was conducted in the AEDC PWT 4T to determine the entry static stability of a 0.00548 scale space shuttle solid rocket booster (SRB). The primary objective was to improve the definition of the aerodynamic characteristics in the angle of attack range beyond 90 deg in the vicinity of the entry trim point. The SRB scale model consisted of the reentry configuration with all major protuberances. A simulated heat shield around the engine nozzle was also included. Data were obtained for a 60 deg side mounted sting and a straight nose mounted sting. The angle of attack range for the side mounted sting was 113 deg to 147 deg and for the nose mounted sting 152 deg to 187 deg. The Mach number range consisted of 0.4 to 1.2 at roll angles of 0 and 90 deg. The resulting 6-component aerodynamic force data was presented as the variation of coefficients with angle of attack for each Mach number and roll angle.

Ramsey, P. E.↗
NASA NTRS1978

Radar cross section measurements of a scale model of the space shuttle orbiter vehicle

A series of microwave measurements was conducted to determine the radar cross section of the Space Shuttle Orbiter vehicle at a frequency and at aspect angles applicable to re-entry radar acquisition and tracking. The measurements were performed in a microwave anechoic chamber using a 1/15th scale model and a frequency applicable to C-band tracking radars. The data were digitally recorded and processed to yield statistical descriptions useful for prediction of orbiter re-entry detection and tracking ranges.

Yates, W. T.↗
NASA NTRS1956

Longitudinal and Lateral Stability Characteristics of a 0.04-Scale Model of the Lockheed F-104A Airplane at Mach Numbers of 1.82 and 2.01. Coord. No. AF-245

An investigation has been made in the Langley 4 by 4-foot supersonic pressure tunnel at Mach numbers of 1.82 and 2.01 to determine the longitudinal and lateral static-stability characteristics of a 0.04-scale model of the Lockheed F-104A airplane. The effects of a modified vertical tail, several ventral-fin arrangements, and several external store arrangements were also determined. The tests were made at Reynolds numbers of 1.02 (exp 6) and 1.382 (exp 6), respectively, based on the wing mean geometric chord. The tests were made of combined angles of attack and sideslip that varied from about -4 deg. to about 20 deg.

Spearman, M. Leroy↗
NASA NTRS1948

Supplementary Investigation in the Langley 20-Foot Free-Spinning Tunnel of the Spin and Recovery Characteristics of a 0.057-Scale Model of the Chance Vought XFU-1 Airplane, Ted No. NACA DE 311

An investigation has been conducted in the Langley 20-foot free-spinning tunnel to determine the effects of decreasing the rudder deflection, of decreasing the rudder span, and of differential rudder movements on the spin and recovery characteristics of a 0.057-scale model of the Chance Vought XF7U-1 airplane. The results indicated that decreasing the rudder span or the rudder deflections, individually or jointly, did not seriously alter the spin or recovery characteristics of the model; and recovery by normal use of controls (full rapid rudder reversal followed l/2 to 1 turn later by movement of the stick forward of neutral) remained satisfactory. Linking the original rudders so that the inboard rudder moves from full with the spin to neutral while the outboard rudder moves from neutral to full against the spin will also result in satisfactory spin and recovery characteristics. Calculations of rudder-pedal forces for recovery showed that the expected forces would probably be within the capabilities of a pilot but that it would be advisable to install some type of boost in the control system to insure easy and rapid movement of the rudders.

Berman, Theodore↗
NASA NTRS2012

Evaluation of Cloud Microphysics Simulated using a Meso-Scale Model Coupled with a Spectral Bin Microphysical Scheme through Comparison with Observation Data by Ship-Borne Doppler and Space-Borne W-Band Radars

Equivalent radar reflectivity factors (Ze) measured by W-band radars are directly compared with the corresponding values calculated from a three-dimensional non-hydrostatic meso-scale model coupled with a spectral-bin-microphysical (SBM) scheme for cloud. Three case studies are the objects of this research: one targets a part of ship-borne observation using 95 GHz Doppler radar over the Pacific Ocean near Japan in May 2001; other two are aimed at two short segments of space-borne observation by the cloud profiling radar on CloudSat in November 2006. The numerical weather prediction (NWP) simulations reproduce general features of vertical structures of Ze and Doppler velocity. A main problem in the reproducibility is an overestimation of Ze in ice cloud layers. A frequency analysis shows a strong correlation between ice water contents (IWC) and Ze in the simulation; this characteristic is similar to those shown in prior on-site studies. From comparing with the empirical correlations by the prior studies, the simulated Ze is overestimated than the corresponding values in the studies at the same IWC. Whereas the comparison of Doppler velocities suggests that large-size snowflakes are necessary for producing large velocities under the freezing level and hence rules out the possibility that an overestimation of snow size causes the overestimation of Ze. Based on the results of several sensitivity tests, we conclude that the source of the overestimation is a bias in the microphysical calculation of Ze or an overestimation of IWC. To identify the source of the problems needs further validation research with other follow-up observations.

Iguchi, T.↗