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At least 793 records · Page 44

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NASA NTRS1985

Reconstruction of a 3-dimensional transonic rotor flow field from holographic interferogram data

Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2-ft-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in seeral planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.

Yu, Y. H.↗

NASA NTRS1985

Reconstruction of a three-dimensional, transonic rotor flow field from holographic interferogram data

Holographic interferometry and computer-assisted tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2-ft-diam view field near the model rotor-blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting fringe-order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in several planes above the blade surface. The value from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three-dimensional transonic rotor flow studies.

Yu, Y. H.↗

NASA NTRS1977

Rolling moments in a trailing vortex flow field

Pressure distributions are presented which were measured on a wing in close proximity to a tip vortex of known structure generated by a larger, upstream semispan wing. Overall loads calculated by integration of these pressures are checked by independent measurements made with an identical model mounted on a force balance. Several conventional methods of wing analysis are used to predict the loads on the following wing. Strip theory is shown to give uniformly poor results for loading distribution, although predictions of overall lift and rolling moment are sometimes acceptable. Good results are obtained for overall coefficients and loading distribution by using linearized pressures in vortex-lattice theory in conjunction with a rectilinear vortex. The equivalent relation from reverse-flow theory that can be used to give economic predictions for overall loads is presented.

Mcmillan, O. J.↗

NASA NTRS1977

Rolling moments in a trailing vortex flow field

Mcmillan, O. J.↗

NASA NTRS1989

The measurement of boundary layers on a compressor blade in cascade. IV - Flow fields for incidence angles of -1.5 and -8.5 degrees

Measurements, made with laser Doppler velocimetry, about a double-circular-arc compressor blade in a cascade are presented for -1.5 and -8.5 degree incidence angles and a chord Reynolds number near 500,000. Comparisons between the results of the current study and those of an earlier work at a 5.0 degree indicence are made. It is found that, in spite of the relative sophistication of the measurement techniques, transition on the pressure surface at the -1.5 degree incidence is dominated by a separation 'bubble' too small to be detected by the laser Doppler velocimeter. The development of the boundary layers at -1.5 and 5.0 degrees are found to be similar. In contrast to the flow at these two incidence angles, the leading edge separation 'bubble' is on the pressure surface for the -8.5 degree incidence.

Zierke, W. C.↗

NASA NTRS1988

Application of advanced grid generation techniques for flow field computations about complex configurations

In the computation of flowfields about complex configurations, it is very difficult to construct a boundary-fitted coordinate system. An alternative approach is to use several grids at once, each of which is generated independently. This procedure is called the multiple grids or zonal grids approach; its applications are investigated. The method conservative providing conservation of fluxes at grid interfaces. The Euler equations are solved numerically on such grids for various configurations. The numerical scheme used is the finite-volume technique with a three-stage Runge-Kutta time integration. The code is vectorized and programmed to run on the CDC VPS-32 computer. Steady state solutions of the Euler equations are presented and discussed. The solutions include: low speed flow over a sphere, high speed flow over a slender body, supersonic flow through a duct, and supersonic internal/external flow interaction for an aircraft configuration at various angles of attack. The results demonstrate that the multiple grids approach along with the conservative interfacing is capable of computing the flows about the complex configurations where the use of a single grid system is not possible.

Kathong, Monchai↗

NASA NTRS1992

Effect of initial acceleration on the development of the flow field of an airfoil pitching at constant rate

We present results from a series of experiments where an airfoil is pitched at constant rate from 0 to 60 degrees angle of attack. It is well documented that the dynamic stall behavior of such an airfoil strongly depends on the nondimensional pitch rate K = dot-alpha C/(2U(sub infinity)), where C is the chord, dot-alpha the constant pitch rate, and U(sub infinity) the free stream speed. In reality, the actual motion of the airfoil deviates from the ideal ramp due to the finite acceleration and deceleration periods imposed by the damping of drive system and response characteristics of the airfoil. It is possible that the pitch rate alone may not suffice in describing the flow and that the details of the motion trajectory before achieving a desired constant pitch rate may also affect the processes involved in the dynamic stall phenomenon. The effects of acceleration and deceleration periods are investigated by systematically varing the acceleration magnitude and its duration through the initial acceleration phase to constant pitch rate. The magnitude and duration of deceleration needed to bring the airfoil motion to rest is similarly controlled.

Koochesfahani, M. M.↗

NASA NTRS1992

Efficient computation of inviscid flow fields around complex configurations using a multiblock multigrid method

The applicability of a multigrid technique to block-structured, body-fitted meshes is examined focusing on three different strategies. In the first strategy data are exchanged between blocks in each stage of a five-stage Runge-Kutta time-stepping scheme which keeps a possible time lag between blocks to a minimum, but requires a large amount of I/O operations and storage. The second strategy is based on performing a complete Runge-Kutta cycle within a block before switching to the next. In the third strategy both a complete Runge-Kutta cycle and the residual evaluation for the restriction operator are done within a block, allowing a minimum of I/O and storage. The inviscid flow around a wing-body/engine-pylon configuration was computed on a mesh consisting of 11 computational blocks. It was found that both the first and the second strategies delivered converged results, but the third failed due to larger time lag between blocks.

Rossow, C.-C.↗

NASA NTRS1988

Navier Stokes computation of the flow field over delta wings with spanwise leading edge blowing

The concept of spanwise leading edge blowing, a means of controlling the position and strength of leading edge vortices, is analyzed by numerical solutions of the three-dimensional Thin-Layer Navier Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results agree favorably with experimental measurements. It is found that the use of spanwise leading edge blowing not only magnifies the size and strength of the leading edge vortices, but also moves the vortex cores outboard and upward. As a result, the increase in lift comes primarily from the greater nonlinear vortex lift. The presence of the leading edge jet stream displaces the flow outboard, thereby increasing the effective aspect ratio of the delta wing. However, blowing causes earlier vortex breakdown, thus decreasing the stall angle.

Yeh, David T.↗

NASA NTRS1993

Comparison Between Predicted and Experimentally Measured Flow Fields at the Exit of the SSME HPFTP Impeller

Validation of CFD codes is a critical first step in the process of developing CFD design capability. The MSFC Pump Technology Team has recognized the importance of validation and has thus funded several experimental programs designed to obtain CFD quality validation data. The first data set to become available is for the SSME High Pressure Fuel Turbopump Impeller. LDV Data was taken at the impeller inlet (to obtain a reliable inlet boundary condition) and three radial positions at the impeller discharge. Our CFD code, TASCflow, is used within the Propulsion and Commercial Pump industry as a tool for pump design. The objective of this work, therefore, is to further validate TASCflow for application in pump design. TASCflow was used to predict flow at the impeller discharge for flowrates of 80, 100 and 115 percent of design flow. Comparison to data has been made with encouraging results.

Bache, George↗

NASA NTRS2019

Wind Tunnel Flow Field Visualizations of the Space Launch System Vehicle Ascent

A wind tunnel test was conducted to characterize the aeroacoustic environment of several configurations of the Space Launch System during ascent. The test was conducted in the 11-by-11 foot transonic and 9-by-7 foot supersonic test sections at NASA Ames research center. Throughout this experiment data was collected from several types of instrumentation including: dynamic and steady-state pressure sensors, unsteady and steady pressure sensitive paint, time-resolved shadowgraph and infrared imaging. The following details results and analysis from the time-resolved shadowgraph and infrared imaging data systems. The time-resolved shadowgraph provided a qualitative measurement of the near-field turbulent fluctuations. These results helped provide context to the relative magnitude and frequency content of the fluid-structure-interaction driving the surface pressure phenomena characterized by the discrete pressure transducers and unsteady pressure sensitive paint. The infrared imaging was used to verify boundary layer trip effectiveness and provide temperature correction for the unsteady pressure sensitive paint.

Shadowgraph↗

NASA NTRS2004

PIV Investigations of the Flow Field in the Volute of a Rotary Blood Pump

A full-size acrylic model of a rotary blood pump was developed in order to utilize Particle Image Velocimetry (PIV) to make measurements of the fluid velocities and turbulent stresses throughout the device. The development of an understanding of the hemodynamics within the blood pump is critical to the development and validation of computational models. A blood analog solution, consisting of sodium iodide solution and glycerin, was developed to match physiological kinematic viscosity. The refractive indecies of the fluid, the pump casing and the impeller were matched to facilitate the use of PIV to make velocity measurements. Velocity measurements made in the volute exit/diffuser region are presented for pumps speeds of 3000-3850 rpm. At each speed data were obtained at a physiological pressure of 90 mmHg and at a maximum flow condition. Four hundred data pairs were used for each resultant mean velocity vector value, representing greater than an order of magnitude more data pairs than reported previously in the literature on similar devices and resulting in velocity uncertainty levels of approximately 2.9%.

Sankovic, John M.↗

NASA NTRS2004

PIV Investigations of the Flow Field in the Volute of a Rotary Blood Pump

A full-size acrylic model of a rotary blood pump was developed in order to utilize Particle Image Velocimetry (PIV) to make measurements of the fluid velocities and turbulent stresses throughout the device. The development of an understanding of the hemodynamics within the blood pump is critical to the development and validation of computational models. A blood analog solution, consisting of sodium iodide solution and glycerin, was developed to match physiological kinematic viscosity. The refractive indices of the fluid, the pump casing, and the impeller were matched to facilitate the use of PIV to make velocity measurements. Velocity measurements made in the volute exit/diffuser region are presented for pumps speeds of 3000-3850 rpm. At each speed data were obtained at a physiological pressure of 12 kPa and at a maximum flow condition. Four hundred data pairs were used for each resultant mean velocity vector value, representing greater than an order of magnitude more data pairs than reported previously in the literature on similar devices and resulting in velocity uncertainty levels of approximately 22.9%.

Sankovic, John M.↗

NASA NTRS1975

Optimum performance and potential flow field of hovering rotors

Rotor and propeller performance and induced potential flowfields were studied on the basis of a rotating actuator disk concept, with special emphasis on rotors hovering out of ground effect. A new theory for the optimum performance of rotors hovering OGE is developed and presented. An extended theory for the optimum performance of rotors and propellers in axial motion is also presented. Numerical results are presented for the optimum distributions of blade-bound circulation together with axial inflow and ultimate wake velocities for the hovering rotor over the range of thrust coefficient of interest in rotorcraft applications. Shapes of the stream tubes and of the velocities in the slipstream are obtained, using available methods, for optimum and off-optimum circulation distributions for rotors hovering in and out of ground effect. A number of explicit formulae useful in computing rotor and propeller induced flows are presented for stream functions and velocities due to distributions of circular vortices over axi-symmetric surfaces.

Wu, J. C.↗

NASA NTRS1994

Computational study of generic hypersonic vehicle flow fields

The geometric data of the generic hypersonic vehicle configuration included body definitions and preliminary grids for the forebody (nose cone excluded), midsection (propulsion system excluded), and afterbody sections. This data was to be augmented by the nose section geometry (blunt conical section mated with the noncircular cross section of the forebody initial plane) along with a grid and a detailed supersonic combustion ramjet (scramjet) geometry (inlet and combustor) which should be merged with the nozzle portion of the afterbody geometry. The solutions were to be obtained by using a Navier-Stokes (NS) code such as TUFF for the nose portion, a parabolized Navier-Stokes (PNS) solver such as the UPS and STUFF codes for the forebody, a NS solver with finite rate hydrogen-air chemistry capability such as TUFF and SPARK for the scramjet and a suitable solver (NS or PNS) for the afterbody and external nozzle flows. The numerical simulation of the hypersonic propulsion system for the generic hypersonic vehicle is the major focus of this entire work. Supersonic combustion ramjet is such a propulsion system, hence the main thrust of the present task has been to establish a solution procedure for the scramjet flow. The scramjet flow is compressible, turbulent, and reacting. The fuel used is hydrogen and the combustion process proceeds at a finite rate. As a result, the solution procedure must be capable of addressing such flows.

Narayan, Johnny R.↗

NASA NTRS2004

Prediction of STS-107 Hypervelocity Flow Fields about the Shuttle Orbiter with Various Wing Leading Edge Damage

Computations were performed for damaged configurations of the Shuttle Orbiter in support of the STS-107 Columbia accident investigation. Two configurations with missing wing leading-edge reinforced carbon-carbon (RCC) panels were evaluated at conditions just prior to the peak heating trajectory point. The initial configuration modeled the Orbiter with an approximate missing RCC panel 6 to determine whether this damage could result in anomalous temperatures measured during the STS-107 reentry. This missing RCC panel 6 computation was found to produce heating augmentation factors of 5 times the nominal heating rates on the side fuselage with lesser heat increases on the front of the OMS pod. This is consistent with the thermocouple and resistance temperature detector sensors from the STS-107 re-entry which observed off nominal high early in the re-entry trajectory. A second damaged configuration modeled the Orbiter with missing RCC panel 9 and included ingestion of the flow into the outboard RCC channel. This computation lowered the level (only 2 times nominal) and moved the location of the heating augmentation on the leeside fuselage relative to the missing RCC panel 6 configuration. The lesser heating augmentation for missing RCC panel 9 was confined near the wing fuselage juncture. Near nominal heating was predicted on the remainder of the side fuselage with some lower than nominal heating on the front surface of the OMS pod. These results for missing RCC panel 9 are consistent with data from the STS-107 re-entry where the heating augmentation was observed to move off the side fuselage and OMS pod sensors at later times in the trajectory. As this solution requires supersonic mass ingestion into the RCC channel, it is probably not an appropriate model prior to penetration of the flow through the spar into the wing structure. It may, however, be representative of the conditions at later times and could account for the movement of the heating signature on the side fuselage.

Pulsonetti, Maria V.↗

NASA NTRS1994

Propagation of experimental uncertainties from the tunnel to the body coordinate system in 3-D LDV flow field studies

An analysis of experimental laser Doppler velocimetry (LDV) data uncertainties that propagate from measurements in the tunnel coordinate system to results in the model system are provided. Calculations of uncertainties as functions of the variables that comprise the final result requires assessment of the contribution each variable makes. Such an analysis enables and necessitates the experimentalists to identify and address the contributing error sources in the experimental measurement system. This provides an opportunity to improve the quality of data derived from experimental systems. This is especially important in experiments where small changes in test conditions are expected to produce small, detectable changes in results. In addition, the need for high-quality experimental data for CFD method validation demands a thorough assessment of experimental uncertainty. Transforming from one Cartesian coordinate system to another by three sequential rotations, equations were developed to transform the variables initially obtained in the original coordinates into variables in the final coordinate system. Based on the transformation equations, propagation equations for errors in the experimentally-derived flow quantities were derived for a model at angle of attack. Experimental uncertainties were then propagated from the tunnel coordinate system into the model system.

Neuhart, Dan H.↗

NASA NTRSDate not supplied

Three-Dimensional Instantaneous Flow-field Reconstruction Using Planar Spectral Proper Orthogonal Decomposition

A novel method for three-dimensional flowfield reconstruction is presented that utilizes spectral proper orthogonal decomposition (SPOD). The time-resolved PIV data are acquired in multiple two-dimensional spanwise planes, with adequate streamwise spacing to resolve the out-of-plane wavelength. Knowledge of the out-of-plane phase speed is also required, and for the present case this value was obtained using instantaneous hot-wire measurements. The method is validated by comparing phase-averaged results acquired with a forcing signal to the SPOD reconstructed results using the same data. The method is also applied to data with no forcing applied, and the results are shown to be similar to the forced case. This method is particularly useful for cases in which there are multiple frequencies of interest, or cases in which the phenomena occur over a broad range of frequencies. This method allows reconstruction of the frequencies of interest without the need for a forcing signal, so the flow can be studied in its natural (undisturbed) state.

boundary-layer transition↗