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Results for “Freestream Velocity”

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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

Hypersonic Second-Mode Instability Response to Shaped Roughness

An experimental campaign was conducted on a 7-degree half-angle cone in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to examine the influence of arrays of regularly spaced roughness elements on instability growth and transition. The primary element shape was a pair of elliptical planform ramps that were inclined at equal and opposite angles with respect to the local streamwise direction. The element shapes were designed to induce transient growth disturbances that would lead to sustained azimuthal modulation of the boundary layer flow while limiting the nearfield disturbances to avoid an immediate, i.e., effective tripping of the boundary layer. The bulk of the run matrix consisted of testing different element height sat free stream unit Reynolds numbers ranging from 9.8 to 13.1 million per meter. Other element shapes previously designed for tripping hypersonic boundary layers were also implemented. The model was instrumented with surface mounted Kulite ® and PCB ® pressure transducers and thermocouples. Spectra from the PCBs ® indicated clear suppression of the second-mode instability; however, neither the PCB ® spectra nor the heat transfer data presented strong evidence for delayed turbulent flow. Complementary stability computations likewise demonstrated second-mode reduction, particularly just downstream of the roughness, but also revealed a rise in first mode (streak-instability) amplitudes from the baseline that was likely responsible for the earlier transition observed for taller roughness cases.

Hypersonics↗
NASA NTRSDate not supplied

Wake Velocimetry of a Sphere-Cone Model in a Mach 10 Air Freestream using FLEET

Velocity measurements in the wake behind a sphere-cone-shaped vehicle wind tunnel model in a Mach 10 hypersonic air flow using femtosecond laser electronic excitation tagging (FLEET) are reported in this work. The FLEET measurements were performed at 1 kHz using a femtosecond laser centered at 800 nm and an ultrahigh-speed intensified camera system. The results reported here are for a single freestream condition corresponding to approximately Re ∞ /L = 3.6∙10 6 m -1 . The performance of the velocimetry instrument was evaluated in the quiescent test section at conditions relevant to the wake measurements. For velocity measurements in the wake, the FLEET measurement volume was translated to several radial locations from the model centerline in the vertical direction and at a streamwise location corresponding to approximately 12 mm from the payload. Reverse flow with mean velocities ranging from -12 m/s to -48 m/s were observed in the recirculating region of the wake, whereas hypersonic velocities in the range of 1079 m/s to 1183 m/s were observed in the free shear layer. A velocity uncertainty methodology is also outlined and applied for the reported measurements. The velocity data reported in this work is expected to play a significant role in the validation of computational codes modeling the hard-to-predict separated hypersonic wake.

FLEET↗
NASA NTRSDate not supplied

Wake Velocimetry of a Sphere-Cone Model in a Mach 10 Air Freestream using FLEET

Velocity measurements in the wake behind a sphere-cone-shaped vehicle wind tunnel model in a Mach 10 hypersonic air flow using femtosecond laser electronic excitation tagging (FLEET) are reported in this work. The FLEET measurements were performed at 1 kHz using a femtosecond laser centered at 800 nm and an ultrahigh-speed intensified camera system. The results reported here are for a single freestream condition corresponding to approximately Re ∞ /L = 3.6∙10 6 m -1 . The performance of the velocimetry instrument was evaluated in the quiescent test section at conditions relevant to the wake measurements. For velocity measurements in the wake, the FLEET measurement volume was translated to several radial locations from the model centerline in the vertical direction and at a streamwise location corresponding to approximately 12 mm from the payload. Reverse flow with mean velocities ranging from -12 m/s to -48 m/s were observed in the recirculating region of the wake, whereas hypersonic velocities in the range of 1079 m/s to 1183 m/s were observed in the free shear layer. A velocity uncertainty methodology is also outlined and applied for the reported measurements. The velocity data reported in this work is expected to play a significant role in the validation of computational codes modeling the hard-to-predict separated hypersonic wake.

FLEET↗
NASA NTRS1990

Detailed flow-field measurements over a 75 deg swept delta wing

Results from an experimental investigation documenting the flowfield over a 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0, and 1.5 million based on the root chord were conducted with both a Pitot pressure probe and a 5-hole pressure probe; and 3-component laser velocimeter surveys were conducted at a Reynolds number of 1.0 million. The Pitot pressure surveys were obtained at 5 chordwise stations, the 5-hole probe surveys were obtained at 3 chordwise stations and the laser velocimeter surveys were obtained at one station. The results confirm the classical roll up of the flow into a pair of primary vortices over the delta wing. The velocity measurements indicate that Reynolds number has little effect on the global structure of the flowfield for the Reynolds number range investigated. Measurements of the non-dimensional axial velocity in the core of the vortex indicate a jet like flow with values greater than twice freestream. Comparisons between velocity measurements from the 5-hole pressure probe and the laser velocimeter indicate that the pressure probe does a reasonable job of measuring the flowfield quantities where the velocity gradients in the flowfield are low.

Kjelgaard, Scott O.↗
NASA NTRS2000

Effects of Hole Length, Supply Plenum Geometry, and Freestream Turbulence on Film Cooling Performance

Experimental measurements are presented in this report to document the sensitivity of film cooling performance to the hole length and coolant delivery plenum geometry. Measurements with hot-wire anemometry detail velocity, local turbulence, and spectral distributions over the exit plane of film cooling holes and downstream of injection in the coolant-freestream interaction zone. Measurements of discharge coefficients and adiabatic effectiveness are also provided. Coolant is supplied to the film cooling holes by means of a large, open plenum and through plenums which force the coolant to approach the holes either co-current or counter-current to the freestream. A single row of film cooling holes with 35 degree-inclined streamwise at two coolant-to-freestream velocity ratios, 0.5 and 1.0, is investigated. The coolant-to-freestream density ratio is maintained in the range 0.96 to 1.0. Measurements were taken under high-freestream (FSTI = 12%) and low-freestream turbulence intensity (FSTI = 0.5%) conditions. The results document the effects of the hole L/D, coolant supply plenum geometry, velocity ratio, and FSTI. In general, hole L/D and the supply plenum geometry play influential roles in the film cooling performance. Hole L/D effects, however, are more pronounced. Film cooling performance is also dependent upon the velocity ratio and FSTI.

Burd, Steven W.↗
NASA NTRS1988

The mean flow structure around and within a turbulent junction or horseshoe vortex. I - The upstream and surrounding three-dimensional boundary layer

An experimental study of the mean flow structure upstream, around, and in a turbulent junction or horseshoe vortex is presented for the case of an incompressible subsonic flow. Mean velocity field and turbulent kinetic energy field measurements (obtained with hot film anemometry) and local wall shear stress measurements are used to monitor the upstream and surrounding three-dimensional turbulent boundary layer-like flow away from separation. Freestream or edge velocity and floor static pressure results are reported.

Menna, J. D.↗
NASA NTRS1985

Effects of non-uniform velocity profiles on dual jets in a crossflow

The interaction between engine exhaust jets and the freestream affects the aerodynamic and stability characteristics of VTOL aircraft during the transition from hover to forward flight. This interaction is often modeled as a simple uniform jet issuing from a flat plate into a subsonic crossflow. The distribution of pressures induced by the jet on the surface of the plate can be used to predict the lift loss and pitching moment for a full-scale aircraft. The uniform jet model has limitations because an actual turbofan engine generates a rather nonuniform exit velocity profile. The purpose of this work is to study the effect of a nonuniform velocity profile on the surface pressure distribution. The mutual interaction of dual jets also is investigated in side-by-side and tandem configurations. Detailed pressure distributions are presented for two jet-to-freestream velocity ratios of 2.2 and 4.0. One important finding is that a nonuniform jet with a high velocity periphery and a low velocity core has a higher effective velocity ratio than a uniform jet with the same mass flow.

Moore, C. L.↗
NASA NTRS2011

Coupled Fluids-Radiation Analysis of a High-Mass Mars Entry Vehicle

The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQAIR code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a 15 m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the MSL-based vehicle.

Palmer, Grant↗
NASA NTRS2011

Coupled Fluids-Radiation Analysis of a High-Mass Mars Vehicle

The NEQAIR line-by-line radiation code has been incorporated into the DPLR Navier-Stokes flow solver such that the NEQAIR subroutines are now callable functions of DPLR. The coupled DPLR-NEQAIR code was applied to compute the convective and radiative heating rates over high-mass Mars entry vehicles. Two vehicle geometries were considered - a 15 m diameter 70-degree sphere cone configuration and a slender, mid-L/D vehicle with a diameter of 5 m called an Ellipsled. The entry masses ranged from 100 to 165 metric tons. Solutions were generated for entry velocities ranging from 6.5 to 9.1 km/s. The coupled fluids-radiation solutions were performed at the peak heating location along trajectories generated by the Traj trajectory analysis code. The impact of fluids-radiation coupling is a function of the level of radiative heating and the freestream density and velocity. For the high-mass Mars vehicles examined in this study, coupling effects were greatest for entry velocities above 8.5 km/s where the surface radiative heating was reduced by up 17%. Generally speaking, the Ellipsled geometry experiences a lower peak radiative heating rate but a higher peak turbulent convective heating rate than the MSL-based vehicle.

Palmer, Grant↗
NASA NTRS2018

Application of STARFLEET Velocimetry in the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel

Selective two-photon absorptive resonance femtosecond laser electronic excitation tagging (STARFLEET) velocimetry is demonstrated for the first time in a NASA Langley wind tunnel with high repetition-rate and single-shot imaging. Experiments performed in the 0.3-meter Transonic Cryogenic Tunnel (TCT) allowed for testing at 300 K over a range of pressures (124 to 517 kPa) and Mach numbers (0.2-0.8) for freestream conditions and flow behind a cylindrical model. Measurement precision and accuracy are determined for the current set of experiments, as are signal intensity and lifetime. Precisions of 3-5 m/s (based on one standard deviation) were typical in the experiment; precisions better than 2% of the mean velocity were obtained for some of the highest velocity conditions. Agreement within a mean error of 3 m/s between STARFLEET freestream velocity measurements and facility DAS readings is demonstrated. STARFLEET is also shown to return spatially-resolved velocity profiles, though some binning of the signal is required.

STARFLEET↗
NASA NTRS1978

Flow and temperature fields following injection of a jet normal to a cross stream

The interaction of a jet entering into a freestream normal to the main flow direction has been studied with particular attention directed to visualization of the large-scale flow interactions and to measurement of the film-cooling performance. Large eddies are apparent downstream of the entering jet even at moderate blowing rate (defined as the ratio of the mass velocity of the jet to the mass velocity of the freestream). At higher blowing rate, there is only intermittent contact between the mass from the jet and the downstream wall. The film cooling downstream from a single normal jet yields a lower centerline effectiveness compared to an inclined jet through a greater lateral spreading. The greater spreading provides a more uniform effectiveness across the span of the downstream wall, in particular at large blowing rate.

Goldstein, R. J.↗
NASA NTRS1985

A consistent spatial differencing scheme for the transonic full-potential equation in three dimensions

A full-potential steady transonic wing flow solver has been modified so that freestream density and residual are captured in regions of constant velocity. This numerically precise freestream consistency is obtained by slightly altering the differencing scheme without affecting the implicit solution algorithm. The changes chiefly affect the fifteen metrics per grid point, which are computed once and stored. With this new method, the outer boundary condition is captured accurately, and the smoothness of the solution is especially improved near regions of grid discontinuity.

Thomas, S. D.↗
NASA NTRS1988

Experimental Study of Bypass Transition in a Boundary Layer

A detailed investigation to compare the boundary layer transition process in a low intensity disturbance environment to that in an environment in which the disturbances are initially non-linear in amplitude was conducted using a flat plate model. The transition mechanism based on linear growth of Tollmien Schlichting (T-S) waves was associated with a freestream turbulence level of 0.3 percent; however, for a freestream turbulence intensity of 0.65 percent and higher, the bypass transition mechanism prevailed. The results of detailed measurements acquired to study and compare the two transition mechanisms indicate that there exists a critical value for the peak rms of the velocity fluctuations within the boundary layer of approximately 3 to 3.5 percent of the freestream velocity. Once the unsteadiness within the boundary layer reached this critical value, turbulent bursting initiated, regardless of the transition mechanism. The two point correlations and simultaneous time traces within the transition region illustrate the features of a turbulent burst and its effect on the surrounding flowfield.

Suder, Kenneth L.↗
NASA NTRS1988

Summary of an experimental investigation on the ground vortex

The results of an experimental investigation into the position and characteristics of the ground vortex are summarized. A 48-inch wind tunnel was modified to create a testing environment suitable for the ground vortex study. Flow visualization was used to document the jet-crossflow interaction and a two-component Laser Doppler Velocimeter (LDV) was used to survey the flowfield in detail. Measurements of the ground vortex characteristics and location as a function of freestream-to-jet velocity ratio, jet height, pressure gradient and upstream boundary layer thickness were obtained.

Billet, Michael L.↗
NASA NTRS1991

Holographic interferometry at the NASA-Ames Hypervelocity Free-Flight Aerodynamic Facility

A laser holographic interferometer has been designed and developed to improve flow visualization for models in free-flight at high velocities and low freestream static pressures. Of special interest is the flow structure due to high-temperature effects in the post-shock region of models traveling at hypervelocity. A dual-plate technique with a double-pass, Michelson-type configuration has been incorporated in an existing hypervelocity ballistic range. Holograms of 5-deg blunt conical models traveling at 5 km/sec in air at a freestream static pressure of 0.08 atm have been made. Finite- and infinite-fringe interferograms were reconstructed from the holograms and provided qualitative flowfield results. In addition, a digital scanning technique was also used to produce a 'synthetic' infinite-fringe interferogram from the analysis of an experimental finite-fringe interferogram. A 'synthetic' infinite-fringe interferogram shows detail in the wake region not found in the experimental interferograms.

Tam, Tim C.↗
NASA NTRS1992

A systematic experimental and computational investigation of a class of contoured wall fuel injectors

The performance of contoured wall fuel injectors for scramjet engine applications is considered. These fuel injectors were aimed at augmenting mixing through axial vorticity production arising from interaction of the fuel/air interface with an oblique shock. The effects of incoming boundary layer height, injector spacing, and injectant to freestream pressure and velocity ratios are examined. Results from 3D flow field surveys and Navier-Stokes simulations are presented.

Waitz, Ian A.↗
NASA NTRS1993

Reduction in size and unsteadiness of a VTOL ground vortex by ground fences

A ground vortex, produced when a jet impinges on the ground in the presence of cross flow, is encountered by V/STOL aircraft hovering near the ground and is known to be hazardous to the aircraft. The objective of this research was to identify a ground-based technique by which both the mean size and fluctuation in size of the ground vortex could be reduced. A simple passive method was identified and examined in the laboratory. Specifically, one or two fine wire mesh screens (ground fences) bent in a horseshoe shape and located on the ground in front of the jet impingement point proved to be very effective. The ground fences work by decreasing the momentum of the upstream-traveling wall jet, effectively causing a higher freestream-to-jet velocity ratio (V(sub infinity)/V(sub j)) and thus, a ground vortex smaller in size and unsteadiness. At(V(sub infinity)/V(sub j)) = 0.15, the addition of a single ground fence resulted in a 70 percent reduction in mean size of the ground vortex. With two ground fences, the mean size decreased by about 85 percent. Fluctuations in size decreased nearly in proportion to the mean size, for both the single and double fence configurations. These results were consistent over a wide range of jet Reynolds number (10(exp 4) less than Re(sub jet) less than 10(exp 5)); further development and full-scale Reynolds number testing are required, however, to determine if this technique can be made practical for the case of actual VTOL aircraft.

Cimbala, John M.↗
NASA NTRS1993

Aerothermodynamic heating due to shock wave/laminar boundary-layer interactions in high-enthalpy hypersonic flow

The interaction between a swept shock wave and a laminar boundary layer was investigated experimentally in high-enthalpy hypersonic flow. The effect of high-temperature, real gas physics on the interaction was examined by conducting tests in air and helium. Heat transfer measurements were made on the surface of a flat plate and a shock-generating fin using thin-film resistance sensors for fin incidence angles of 0, 5, and 10 deg at Mach numbers of 6.9 in air and 7.2 in helium. The experiments were conducted in the NASA HYPULSE expansion tube, an impulse-type facility capable of generating high-enthalpy, high-velocity flow with freestream levels of dissociated species that are particularly low. The measurements indicate that the swept shock wave creates high local heat transfer levels in the interaction region, with the highest heating found in the strongest interaction. The maximum measured heating rates in the interaction are order of magnitude greater than laminar flat plate boundary layer heating levels at the same location.

Hackett, Charles M.↗