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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 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Modeling of solid/porous wall boundary conditions for the validation of computational fluid dynamics codes

A computational study has been undertaken to investigate method of modeling solid and porous wall boundary conditions in computational fluid dynamics (CFD) codes. The procedure utilizes experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution is then imposed as a boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the boundary conditions at the wind tunnel walls using only sparse experimental measurements has been investigated. Position and refinement of experimental measurement locations required to describe porous wall boundary conditions has also been considered.

Beutner, Thomas J.↗
NASA NTRS1993

Determination of solid/porous wall boundary conditions from wind tunnel data for computational fluid dynamics codes

A computational and experimental study has been undertaken to investigate methods of modeling solid and porous wall boundary conditions in CFD codes. The procedure utilizes experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution is then imposed as a boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the boundary conditions at the wind tunnel walls using only sparse experimental measurements has been investigated. Verification of the approach using computational studies has been carried out using an incompressible flow solver. The current work demonstrates this technique for low speed flows and compares the result with experimental data obtained from a heavily instrumented variable porosity test section. Position and refinement of experimental measurements required to describe porous wall boundary conditions has also been considered for application to other porous wall wind tunnels.

Beutner, Thomas J.↗
NASA NTRS1986

Pretest Report for the Full Span Propulsive Wing/Canard Model Test in the NASA Langley 4 x 7 Meter Low Speed Wind Tunnel Second Series Test

A full span propulsive wing/canard model is to be tested in the NASA Langley Research Center (LaRC) 4 x 7 meter low speed wind tunnel. These tests are a continuation of the tests conducted in Feb. 1984, NASA test No.290, and are being conducted under NASA Contract NAS1-17171. The purpose of these tests is to obtain extensive lateral-directional data with a revised fuselage concept. The wings, canards, and vertical tail of this second test series model are the same as tested in the previous test period. The fuselage and internal flow path have been modified to better reflect an external configuration suitable for a fighter airplane. Internal ducting and structure were changed as required to provide test efficiency and blowing control. The model fuselage tested during the 1984 tests was fabricated with flat sides to provide multiple wing and canard placement variations. The locations of the wing and canard are important variables in configuration development. With the establishment of the desired relative placement of the lifting surfaces, a typically shaped fuselage has been fabricated for these tests. This report provides the information necessary for the second series tests of the propulsive wing/canard model. The discussion in this report is limited to that affected by the model changes and to the second series test program. The pretest report information for test 290 which is valid for the second series test was published in Rockwell report NR 83H-79. This report is presented as Appendix 1 and the modified fuselage stress report is presented as Appendix 2 to this pretest report.

Stewart, V. R.↗
NASA NTRS1989

Penetrative convective flows induced by internal heating and mantle compressibility

Penetrative convective flows induced in a spherical shell by combined effects of internal heating and mantle compressibility are investigated using mathematical and numerical formulations for compressible spherical shell convection. Isothermal stress-free boundary conditions applied at the top and the bottom of the shell are solved using a time-dependent finite difference code in a temperature, vorticity, stream function formulation for Rayleigh numbers ranging from the critical Rc up to 2000 Rc. Results indicate that compressibility, together with internal heating, could be a mechanism capable of generating spontaneously layered convection and local melting in the mantle and that non-Boussinesq effects must be considered in interpretations of geophysical phenomena.

Machetel, Philippe↗
NASA NTRS1971

Detailed investigation of flowfields within large scale hypersonic inlet models

Analytical and experimental investigations were conducted to determine the characteristics of the internal flows in model passages representative of hypersonic inlets and also sufficiently large for meaningful data to be obtained. Three large-scale inlet models, each having a different compression ratio, were designed to provide high performance and approximately uniform static-pressure distributions at the throat stations. A wedge forebody was used to simulate the flowfield conditions at the entrance of the internal passages, thus removing the actual vehicle forebody from consideration in the design of the wind-tunnel models. Tests were conducted in a 3.5 foot hypersonic wind tunnel at a nominal test Mach number of 7.4 and freestream unit Reynolds number of 2,700,000 per foot. From flowfield survey data the inlet entrance, the entering inviscid and viscous flow conditions were determined prior to the analysis of the data obtained in the internal passages. Detailed flowfield survey data were obtained near the centerlines of the internal passages to define the boundary-layer development on the internal surfaces and the internal shock-wave configuration. Finally, flowfield data were measured across the throats of the inlet models to evaluate the internal performance of the internal passages. These data and additional results from surface instrumentation and flow visualization studies were utilized to determine the internal flowfield patterns and the inlet performance.

Seebaugh, W. R.↗
NASA NTRS1989

Analysis of three-dimensional viscous flow in a supersonic throughflow fan

A 3-D Navier-Stokes code was developed for analysis of turbomachinery blade rows and other internal flows. The Navier-Stokes equations are written in a Cartesian coordinate system rotating about the x-axis, and then mapped to a general body-fitted coordinate system. Streamwise viscous terms are neglected using the thin layer assumption, and turbulence effects are modelled using the Baldwin-Lomax turbulence model. The equations are discretized using finite differences on stacked C-type grids and are solved using a multistage Runge-Kutta algorithm with a spatially varying time step and implicit residual smoothing. Calculations were made of the flow around a supersonic throughflow fan blade. The fan was designed as a key component in a supersonic cruise engine. The 3-D calculations were done on a 129x29x33 grid and took 50 minutes of cpu time. Comparisons with the quasi-3-D results show minor differences in loading due to 3-D effects. Particle traces show nearly 2-D flows near the pressure surface, but large secondary flows within the suction surface boundary layer. The horseshoe vortex ahead of the leading edge is clearly seen.

Chima, Rodrick V.↗
NASA NTRS1993

A method for the modelling of porous and solid wind tunnel walls in computational fluid dynamics codes

Porous wall wind tunnels have been used for several decades and have proven effective in reducing wall interference effects in both low speed and transonic testing. They allow for testing through Mach 1, reduce blockage effects and reduce shock wave reflections in the test section. Their usefulness in developing computational fluid dynamics (CFD) codes has been limited, however, by the difficulties associated with modelling the effect of a porous wall in CFD codes. Previous approaches to modelling porous wall effects have depended either upon a simplified linear boundary condition, which has proven inadequate, or upon detailed measurements of the normal velocity near the wall, which require extensive wind tunnel time. The current work was initiated in an effort to find a simple, accurate method of modelling a porous wall boundary condition in CFD codes. The development of such a method would allow data from porous wall wind tunnels to be used more readily in validating CFD codes. This would be beneficial when transonic validations are desired, or when large models are used to achieve high Reynolds numbers in testing. A computational and experimental study was undertaken to investigate a new method of modelling solid and porous wall boundary conditions in CFD codes. The method utilized experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution was then imposed as a pressure boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the effect of porosity changes on the wall was investigated. Also, the effectiveness of this method when only sparse experimental measurements were available has been investigated. The current work demonstrated this approach for low speed flows and compared the results with experimental data obtained from a heavily instrumented variable porosity test section. The approach developed was simple, computationally inexpensive, and did not require extensive or intrusive measurements of the boundary conditions during the wind tunnel test. It may be applied to both solid and porous wall wind tunnel tests.

Beutner, Thomas John↗
NASA NTRS1994

Determination of solid/porous wall boundary conditions from wind tunnel data for computational fluid dynamics codes

A computational and experimental study has been undertaken to investigate methods of modelling solid and porous wall boundary conditions in computational fluid dynamics (CFD) codes. The procedure utilizes experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution is then imposed as a pressure boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the boundary conditions at the wind tunnel walls using only sparse experimental measurements has been investigated. Verification of the approach using computational studies has been carried out using an incompressible flow solver. The current work demonstrates this technique for low speed flows and compares the result with experimental data obtained from a heavily instrumented variable porosity test section. Position and refinement of experimental measurements required to describe porous wall boundary conditions have also been considered for application to other porous wall wind tunnels. The approach developed is simple, computationally inexpensive, and does not require extensive or intrusive measurements. It may be applied to both solid and porous wall wind tunnel tests. Some consideration is given to the extension of this method to three dimensions.

Beutner, Thomas J.↗
NASA NTRS1996

Development of a Driver Code for the WICS Project

Wall Interference Correction System (WICS) is a computational technique to evaluate the wind tunnel wall interference corrections for blockage, Mach number, dynamic pressure and angle of attack. The objective is to predict the effects of the bounding walls on wind tunnel measurements [1]. The procedure attempts to compute these corrections in real time so that the test engineer can implement these by adjusting wind tunnel controls for angle of attack and dynamic pressure while the test is in progress in the 12 ft. pressure wind tunnel at the Ames Research Center. The technique utilizes an ideal flow solver PMARC (Panel Method developed at the Ames Research Center) to compute influence coefficients in an internal flow mode by representing a paneled model of the tunnel for fullspan and semispan tests.

Ghosh, Amitabha↗
NASA NTRS2003

Compressor Performance Enhanced by Active Flow Control Over Stator Vanes

The application of active flow control technology to enhance turbomachinery system performance is being investigated at the NASA Glenn Research Center through experimental studies. Active flow control involves the use of sensors and actuators embedded within engine components to dynamically alter the internal flow path during off nominal operation in order to optimize engine performance and maintain stable operation. Modern compressors are already highly optimized components that must be designed to accommodate a broad range of operating conditions in a safe and efficient manner. Since overall engine performance is driven by compressor performance, advances in compressor technology that reduce weight and parts count, reduce fuel consumption, and lower maintenance costs will have a significant impact on the cost of aircraft ownership. Active flow control holds the promise of delivering such technology advances.

Culley, Dennis E.↗
NASA NTRS2014

Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Clevleand, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSLthe first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal icing would occur. The ice crystal icing event, an uncommanded reduction in thrust, was able to be turned on and off by manipulating cloud TWC. A flight test point where no ice crystal icing event occurred was also duplicated in PSL. Physics based computational tools were successfully used to predict tunnel settings to induce ice buildup along the low pressure compression system flow path for several test points at incrementally lower altitudes, demonstrating that development of ice crystal icing scaling laws is potentially feasible. Analysis of PSL test data showed that uncommanded reduction in thrust occurs during ice crystal cloud on operation prior to fan speed reduction. This supports previous findings that the reduction of thrust for this test article is due to ice buildup leading to a restricted airflow from either physical or aerodynamic blockage in the engine core flow path.

Ice Crystal Icing↗
NASA NTRS2014

Validation Ice Crystal Icing Engine Test in the Propulsion Systems Laboratory at NASA Glenn Research Center

The Propulsion Systems Laboratory (PSL) is an existing altitude simulation jet engine test facility located at NASA Glenn Research Center in Cleveland, OH. It was modified in 2012 with the integration of an ice crystal cloud generation system. This paper documents the inaugural ice crystal cloud test in PSL--the first ever full scale, high altitude ice crystal cloud turbofan engine test to be conducted in a ground based facility. The test article was a Lycoming ALF502-R5 high bypass turbofan engine, serial number LF01. The objectives of the test were to validate the PSL ice crystal cloud calibration and engine testing methodologies by demonstrating the capability to calibrate and duplicate known flight test events that occurred on the same LF01 engine and to generate engine data to support fundamental and computational research to investigate and better understand the physics of ice crystal icing in a turbofan engine environment while duplicating known revenue service events and conducting test points while varying facility and engine parameters. During PSL calibration testing it was discovered than heated probes installed through tunnel sidewalls experienced ice buildup aft of their location due to ice crystals impinging upon them, melting and running back. Filtered city water was used in the cloud generation nozzle system to provide ice crystal nucleation sites. This resulted in mineralization forming on flow path hardware that led to a chronic degradation of performance during the month long test. Lacking internal flow path cameras, the response of thermocouples along the flow path was interpreted as ice building up. Using this interpretation, a strong correlation between total water content (TWC) and a weaker correlation between median volumetric diameter (MVD) of the ice crystal cloud and the rate of ice buildup along the instrumented flow path was identified. For this test article the engine anti-ice system was required to be turned on before ice crystal icing would occur. The ice crystal icing event, an uncommanded reduction in thrust, was able to be turned on and off by manipulating cloud TWC. A flight test point where no ice crystal icing event occurred was also duplicated in PSL. Physics based computational tools were successfully used to predict tunnel settings to induce ice buildup along the low pressure compression system flow path for several test points at incrementally lower altitudes, demonstrating that development of ice crystal icing scaling laws is potentially feasible. Analysis of PSL test data showed that uncommanded reduction in thrust occurs during ice crystal cloud on operation prior to fan speed reduction. This supports previous findings that the reduction of thrust for this test article is due to ice buildup leading to a restricted airflow from either physical or aerodynamic blockage in the engine core flow path.

Turbofan Engine↗
NASA NTRSDate not supplied

Isolator Shock Train Response to Dual-Mode Scramjet Throttling-Like Forcing

Typical dual-mode scramjet designs employ an isolator component situated between the inlet and combustor. The isolator fluidically buffers the inlet from the combustor via the formation of a shock train, an interconnected series of shock-boundary layer interactions. The isolator is required during ramjet-engine operation and at the lower edges of the scramjet operational envelope. In an operational engine, transient behavior imposes fluidic dynamics due to attitude adjustments, combustor throttling, aberrant resonances within the combustor, etc. To accommodate these transients, the shock train adjusts on both temporal and spatial scales. Sufficiently large downstream pressure increases in the combustor play a critical role during unstart, whereby strong compression waves propagate upstream in the isolator to ultimately force the shock train out the inlet, if it cannot stabilize to the pressure increase. This critically reduces the engine mass flow rate, potentially causing a catastrophic reduction in thrust and loss of the vehicle. A nearly instantaneous unstart can, in principle, be predicted by a ratio of the normal shock pressure jump in an inviscid isolator. However, lower-magnitude perturbations, below this bound, also induce unstarts. Understanding the transient behavior inherent in unsteady isolator operation is critical to reducing safety margin in isolator length and refining engine control algorithms to improve engine performance. The Isolator Dynamics Research Lab (IDRL) is a cold-flow, direct-connect isolator facility ideal for assessing the fundamental fluid dynamics of shock trains. The internal flow of a scramjet isolator is modeled using a 2D, symmetric Mach 2.5 nozzle followed by a 695 mm long, 25.4x50.8 mm rectangular isolator test section. A blockage cone on a high-speed ball screw actuator downstream of the isolator is pulsed to simulate the throttling behavior in an operational combustor. Wall flush high-speed static pressure transducers detect the streamwise shock train within the isolator. Time delays between the onset of the throttling and the response motion of the shock train are analyzed to assess the controlling scales and parameters. From this analysis, the response time of the isolator shock train system is found to depend primarily on the isolator back pressure time-rate-of-change but quickly saturates to a maximum propagation speed along the isolator. These results inform dual-mode scramjet engine designers of the critical parameters and limits of an unstart resistant engine.

scramjet↗
NASA NTRS1993

Assessment of three numerical methods for the computation of a low-density plume flow

Results from three numerical methods including one based on the Navier-Stokes equations, one based on kinetic theory using the DSMC method, and one based on the Boltzmann equation with a Krook-type collision term are compared to each other and to experimental data for a model problem of heated nitrogen flow in a conical nozzle expanding into a vacuum. The problem simulates flow in a resistojet, a low-thrust, electrothermal rocket. The continuum method is applied to both the internal flow and near-field plume. The DSMC and Boltzmann methods are applied primarily to the plume. Experimental measurements of Pitot pressure and flow angle, taken with an apparatus that duplicates the model nozzle flow, are used in the comparisons.

Penko, Paul F.↗
NASA NTRS1952

Axisymmetric supersonic flow in rotating impellers

General equations are developed for isentropic, frictionless, axisymmetric flow in rotating impellers with blade thickness taken into account and with blade forces eliminated in favor of the blade-surface function. It is shown that the total energy of the gas relative to the rotating coordinate system is dependent on the stream function only, and that if the flow upstream of the impeller is vortex-free, a velocity potential exists which is a function of only the radial and axial distances in the impeller. The characteristic equations for supersonic flow are developed and used to investigate flows in several configurations in order to ascertain the effect of variations of the boundary conditions on the internal flow and the work input. Conditions varied are prerotation of the gas, blade turning rate, gas velocity at the blade tips, blade thickness, and sweep of the leading edge.

Goldstein, Arthur W↗