SEARCH · Engineering Papers

Results for “Flow Conditions”

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.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 613 records · Page 34

All sources NASA NTRS DOE OSTI

NASA NTRS1993

Monte Carlo and experimental studies of nozzle flow in a low-power hydrogen arcjet

The flow in the nozzle of a small hydrogen arcjet intended for low-thrust propulsion is studied without arc ignition using numerical and experimental techniques. The flow conditions in the thruster indicate that low-density nonequilibrium effects will significantly influence the flow. Therefore, the numerical analysis is undertaken using a Monte Carlo approach. The experimental studies employ laser Raman scattering. Comparisons of the measured and computed results are made for total number density, rotational temperature, and for the number density of the first rotational level. The numerical results are found to be quite sensitive to the rotational relaxation rate, and a strong degree of thermal nonequilibrium is observed at the exit plane of the nozzle. Successful comparisons between experiment and analysis permit estimation of the rotational relaxation rate for hydrogen.

Boyd, Iain D.↗

NASA NTRS1995

Eliminating Flow Separation and Reducing Viscous Drag Through Boundary Layer Analysis and Manipulation

As both computers and flow-analyzing equations have increased in sophistication, Computational Fluid Dynamics (CFD) has evolved into a fixture for advanced aircraft design. While CFD codes have improved in accuracy and efficiency, their ability to encompass viscous effects is lacking in certain areas. For example, current CFD codes cannot accurately predict or correct for the increased drag due to these viscous effects at some flow conditions. However, by analyzing an airfoil's turbulent boundary layer, one can predict not only flow separation via the shape factor parameter, but also viscous drag via the momentum thickness. Various codes have been written which can calculate turbulent boundary layer parameters. The goal of my research is to develop procedures for modifying an airfoil (via its local pressure distribution) to eliminate boundary layer separation and/or to reduce viscous drag. The modifications to the local pressure distribution necessary to achieve these objectives will be determined using a direct-iterative method installed into a turbulent boundary layer analyzer. Furthermore, the modifications should preserve the basic characteristics of the original airfoil.

Oser, Matthew D.↗

NASA NTRS2002

Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade

A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.

Lepicovsky, J.↗

NASA NTRS2008

Balanced Flow Meters without Moving Parts

Balanced flow meters are recent additions to an established class of simple, rugged flow meters that contain no moving parts in contact with flow and are based on measurement of pressure drops across objects placed in flow paths. These flow meters are highly accurate, minimally intrusive, easily manufacturable, and reliable. A balanced flow meter can be easily mounted in a flow path by bolting it between conventional pipe flanges. A balanced flow meter can be used to measure the flow of any of a variety of liquids or gases, provided that it has been properly calibrated. Relative to the standard orifice-plate flow meter, the balanced flow meter introduces less turbulence and two times less permanent pressure loss and is therefore capable of offering 10 times greater accuracy and repeatability with less dissipation of energy. A secondary benefit of the reduction of turbulence is the reduction of vibration and up to 15 times less acoustic noise generation. Both the balanced flow meter and the standard orifice-plate flow meter are basically disks that contain holes and are instrumented with pressure transducers on their upstream and downstream faces. The most obvious difference between them is that the standard orifice plate contains a single, central hole while the balanced flow meter contains multiple holes. The term 'balanced' signifies that in designing the meter, the sizes and locations of the holes are determined in an optimization procedure that involves balancing of numerous factors, including volumetric flow, mass flow, dynamic pressure, kinetic energy, all in an effort to minimize such undesired effects as turbulence, pressure loss, dissipation of kinetic energy, and non-repeatability and nonlinearity of response over the anticipated range of flow conditions. Due to proper balancing of these factors, recent testing demonstrated that the balanced flow-meter performance was similar to a Venturi tube in both accuracy and pressure recovery, but featured reduced cost and pipe-length requirements.

Kelley, Anthony R.↗

NASA NTRS2020

Tuning Neural Network Models for Improved Prediction of Boundary Layer Transition

Boundary layer transition can strongly impact flight vehicle performance as it influences surface skin friction and aerodynamic heating, making accurate transition prediction a key to designing next generation aircraft. Artificial neural networks (ANNs) have shown promise toward predicting laminar-turbulent transition based on linear stability correlations. The computational efficiency of ANNs and the substantially reduced user involvement in relation to direct computations based on the linear stability theory (LST) makes them an attractive methodology for integrating the LST based correlations in computational fluid dynamics codes. Tollmien-Schlichting (TS) waves correspond to the dominant transition mechanism in 2D or weakly 3D subsonic boundary layers, such as those encountered in general aviation applications. Improvements to neural network model accuracy in predicting the amplification rates of TS instability waves have been investigated by leveraging recent machine learning developments in conjunction with surrogate optimization techniques and via suitable augmentation of the data used to train the networks. The optimized models trained on the modified dataset reduced the average transition location errors on different airfoils at several flow conditions by 51% of the original manually-tuned network’s errors on the same flow cases. The actual transition locations were derived from the Langley Stability and Transition Analysis Code (LASTRAC).

Machine Learning↗

NASA NTRS2005

Closed Loop Active Flow Separation Detection and Control in a Multistage Compressor

Active closed loop flow control was successfully demonstrated on a full annulus of stator vanes in a low speed axial compressor. Two independent methods of detecting separated flow conditions on the vane suction surface were developed. The first technique detects changes in static pressure along the vane suction surface, while the second method monitors variation in the potential field of the downstream rotor. Both methods may feasibly be used in future engines employing embedded flow control technology. In response to the detection of separated conditions, injection along the suction surface of each vane was used. Injected mass flow on the suction surface of stator vanes is known to reduce separation and the resulting limitation on static pressure rise due to lowered diffusion in the vane passage. A control algorithm was developed which provided a proportional response of the injected mass flow to the degree of separation, thereby minimizing the performance penalty on the compressor system.

Bright, Michelle M.↗

NASA NTRS1984

Investigation of the external flow analysis for density measurements at high altitude

The results of analysis performed on the external flow around the shuttle orbiter nose regions at the Shuttle Upper Atmosphere Mass Spectrometer (SUMS) inlet orifice are presented. The purpose of the analysis is to quantitatively characterize the flow conditions to facilitate SUMS flight data reduction and subsequent determination of orbiter aerodynamic force coefficients in the hypersonic rarefied flow regime. Experimental determination of aerodynamic force coefficients requires accurate simultaneous measurement of forces (or acceleration) and dynamic pressure along with independent knowledge of density and velocity. The SUMS provides independent measurement of dynamic pressure; however, it does so indirectly and requires knowledge of the relationship between measured orifice conditions and the dynamic pressure which can only be determined on the basis of molecule or theory for a winged configuration. Monte Carlo direct simulation computer codes were developed for both the flow field solution at the orifice and for the internal orifice flow. These codes were used to study issues associated with geometric modeling of the orbiter nose geometry and the modeling of intermolecular collisions including rotational energy exchange and a preliminary analysis of vibrational excitation and dissociation effects. Data obtained from preliminary simulation runs are presented.

Bienkowski, G. K.↗

NASA NTRS1993

Experimental investigation of turbine disk cavity aerodynamics and heat transfer

An experimental investigation of turbine disk cavity aerodynamics and heat transfer was conducted to provide an experimental data base that can guide the aerodynamic and thermal design of turbine disks and blade attachments for flow conditions and geometries simulating those of the space shuttle main engine (SSME) turbopump drive turbines. Experiments were conducted to define the nature of the aerodynamics and heat transfer of the flow within the disk cavities and blade attachments of a large scale model simulating the SSME turbopump drive turbines. These experiments include flow between the main gas path and the disk cavities, flow within the disk cavities, and leakage flows through the blade attachments and labyrinth seals. Air was used to simulate the combustion products in the gas path. Air and carbon dioxide were used to simulate the coolants injected at three locations in the disk cavities. Trace amounts of carbon dioxide were used to determine the source of the gas at selected locations on the rotors, the cavity walls, and the interstage seal. The measurements on the rotor and stationary walls in the forward and aft cavities showed that the coolant effectiveness was 90 percent or greater when the coolant flow rate was greater than the local free disk entrainment flow rate and when room temperature air was used as both coolant and gas path fluid. When a coolant-to-gas-path density ratio of 1.51 was used in the aft cavity, the coolant effectiveness on the rotor was also 90 percent or greater at the aforementioned condition. However, the coolant concentration on the stationary wall was 60 to 80 percent at the aforementioned condition indicating a more rapid mixing of the coolant and flow through the rotor shank passages. This increased mixing rate was attributed to the destabilizing effects of the adverse density gradients.

Daniels, W. A.↗

NASA NTRS2004

Time-Average Measurement of Velocity, Density, Temperature, and Turbulence Using Molecular Rayleigh Scattering

Measurement of time-averaged velocity, density, temperature, and turbulence in gas flows using a nonintrusive, point-wise measurement technique based on molecular Rayleigh scattering is discussed. Subsonic and supersonic flows in a 25.4-mm diameter free jet facility were studied. The developed instrumentation utilizes a Fabry-Perot interferometer to spectrally resolve molecularly scattered light from a laser beam passed through a gas flow. The spectrum of the scattered light contains information about velocity, density, and temperature of the gas. The technique uses a slow scan, low noise 16-bit depth CCD camera to record images of the fringes formed by Rayleigh scattered light passing through the interferometer. A kinetic theory model of the Rayleigh scattered light is used in a nonlinear least squares fitting routine to estimate the unknown parameters from the fringe images. The ability to extract turbulence information from the fringe image data proved to be a challenge since the fringe is broadened by not only turbulence, but also thermal fluctuations and aperture effects from collecting light over a range of scattering angles. Figure 1 illustrates broadening of a Rayleigh spectrum typical of flow conditions observed in this work due to aperture effects and turbulence for a scattering angle, chi(sub s), of 90 degrees, f/3.67 collection optics, mean flow velocity, u(sub k), of 300 m/s, and turbulent velocity fluctuations, sigma (sub uk), of 55 m/s. The greatest difficulty in processing the image data was decoupling the thermal and turbulence broadening in the spectrum. To aid in this endeavor, it was necessary to seed the ambient air with smoke and dust particulates; taking advantage of the turbulence broadening in the Mie scattering component of the spectrum of the collected light (not shown in the figure). The primary jet flow was not seeded due to the difficulty of the task. For measurement points lacking particles, velocity, density, and temperature information could reliably be recovered, however the turbulence estimates contained significant uncertainty. Resulting flow parameter estimates are presented for surveys of Mach 0.6, 0.95, and 1.4 jet flows. Velocity, density, and temperature were determined with accuracies of 5 m/s, 1.5%, and 1%, respectively, in flows with no particles present, and with accuracies of 5 m/s, 1-4%, and 2% in flows with particles. Comparison with hotwire data for the Mach 0.6 condition demonstrated turbulence estimates with accuracies of about 5 m/s outside the jet core where Mie scattering from dust/smoke particulates aided in the estimation of turbulence. Turbulence estimates could not be recovered with any significant accuracy for measurement points where no particles were present.

Mielke, Amy F.↗

NASA NTRS1973

Overall and blade-element performance of a multiple-circular-arc bladed transonic compressor rotor with tip speed of 1375 feet per second

The design and experimental performance of a 20-inch-diameter multiple-circular-arc bladed axial-flow transonic compressor rotor is presented. Radial surveys of the flow conditions were made. At design speed the peak efficiency was 0.882 and occurred at a weight flow of 64.0 pounds per second. At this point the total-pressure and total-temperature ratios were 1.79 and 1.205, respectively. The stall margin at design speed was 8 percent based on weight flows and total-pressure ratios at experimental peak efficiency and near stall. The measured stall margin was 20 percent at design weight flow and speed.

Kovich, G.↗

NASA NTRS1975

Coaxial injector response to axial instability under simulated cold-flow

The frequency dependence of the admittances and response factors of various gaseous rocket injector configurations subject to axial instabilities under cold-flow conditions, have been measured using the modified impedance-tube technique. The tested configurations simulate the flows in a gaseous-fuel injector, gaseous-oxidizer injector and a coaxial injector with both fuel and oxidizer elements. Comparison of the measured response data with corresponding data predicted by the Feiler and Heidmann model indicates good agreement between the two sets of data.

Janardan, B. A.↗

NASA NTRS1975

Viscous interaction of flow redevelopment after flow reattachment with supersonic external streams

A flow model has been developed to study the flow development after reattachment with supersonic external streams. Special attention is given to the pressure difference across the viscous layer, and it is suggested that such a flow redevelopment can be treated as a relaxation of this pressure difference. Upon correlating the pressure difference with a slope parameter of the velocity profile, the system of equations governing the flow would produce a saddle point singularity corresponding to the fully rehabilitated asymptotic flow condition. A method of calculation for this flowfield, in conjunction with the matching of the upstream flow, has been derived and is discussed. Samples of calculations are also presented. Reasonably good agreement with experimental data has also been observed.

Chow, W. L.↗

NASA NTRS1991

Estimation of additive forces and moments for supersonic inlets

A technique for estimating the additive forces and moments associated with supersonic, external compression inlets as a function of mass flow ratio has been developed. The technique makes use of a low order supersonic paneling method for calculating minimum additive forces at maximum mass flow conditions. A linear relationship between the minimum additive forces and the maximum values for fully blocked flow is employed to obtain the additive forces at a specified mass flow ratio. The method is applicable to two-dimensional inlets at zero or nonzero angle of attack, and to axisymmetric inlets at zero angle of attack. Comparisons with limited available additive drag data indicate fair to good agreement.

Perkins, Stanley C., Jr.↗

NASA NTRS1992

Comment on: The present state and future direction of eddy viscosity models

Two turbulence models are compared under various flow conditions. Presented first is a k-omega model which is best suited for wall bounded flows. Presented second is a k-epsilon model which is best suited for free shear flow. Generalizations of the k-omega model are discussed, and various aerodynamic parameters such as skin friction, velocity distribution, and hypersonic cylinder-flare, are compared for the different models. All results are presented in viewgraph format.

Coakley, T. J.↗

NASA NTRS1992

Lip leakage flow simulation for the Gravity Probe B gas spinup

The lip leakage flow for the Gravity Probe B (GP-B) gas spinup system is investigated using a particle simulation code on the Connection Machine (PSiCM). Particle simulation is employed because the flow conditions are in the transition regime between continuum and free molecule where particle methods are of greatest use. The dominant flow is Couette in nature and the simulation is first validated through comparison to theoretical results for Couette flow in the transition regime. A GP-B type geometry is then simulated and results are presented for two conditions, those corresponding to near the inlet and near the outlet of the spinup channel. Comparison to experiment is not made because experimental data is not yet available.

Dagum, Leonardo↗

NASA NTRS1993

Wall interference and boundary simulation in a transonic wind tunnel with a discretely slotted test section

A computational simulation of a transonic wind tunnel test section with longitudinally slotted walls is developed and described herein. The nonlinear slot model includes dynamic pressure effects and a plenum pressure constraint, and each slot is treated individually. The solution is performed using a finite-difference method that solves an extended transonic small disturbance equation. The walls serve as the outer boundary conditions in the relaxation technique, and an interaction procedure is used at the slotted walls. Measured boundary pressures are not required to establish the wall conditions but are currently used to assess the accuracy of the simulation. This method can also calculate a free-air solution as well as solutions that employ the classical homogeneous wall conditions. The simulation is used to examine two commercial transport aircraft models at a supercritical Mach number for zero-lift and cruise conditions. Good agreement between measured and calculated wall pressures is obtained for the model geometries and flow conditions examined herein. Some localized disagreement is noted, which is attributed to improper simulation of viscous effects in the slots.

Al-Saadi, Jassim A.↗

NASA NTRS1993

Semi-span model testing in the National Transonic Facility

A semi-span testing technique has been proposed for the NASA Langley Research Center's National Transonic Facility (NTF). Semi-span testing has several advantages including (1) larger model size, giving increased Reynolds number capability; (2) improved model fidelity, allowing ease of flap and slat positioning which ultimately improves data quality; and (3) reduced construction costs compared with a full-span model. In addition, the increased model size inherently allows for increased model strength, reducing aeroelastic effects at the high dynamic pressure levels necessary to simulate flight Reynolds numbers. The Energy Efficient Transport (EET) full-span model has been modified to become the EET semi-span model. The full-span EET model was tested extensively at both NASA LRC and NASA Ames Research Center. The available full-span data will be useful in validating the semi-span test strategy in the NTF. In spite of the advantages discussed above, the use of a semi-span model does introduce additional challenges which must be addressed in the testing procedure. To minimize the influence of the sidewall boundary layer on the flow over the semi-span model, the model must be off-set from the sidewall. The objective is to remove the semi-span model from the sidewall boundary layer by use of a stand-off geometry. When this is done however, the symmetry along the centerline of the full-span model is lost when the semi-span model is mounted on the wind tunnel sidewall. In addition, the large semi-span model will impose a significant pressure loading on the sidewall boundary layer, which may cause separation. Even under flow conditions where the sidewall boundary layer remains attached, the sidewall boundary layer may adversely effect the flow over the semi-span model. Also, the increased model size and sidewall mounting requires a modified wall correction strategy. With these issues in mind, the semi-span model has been well instrumented with surface pressure taps to obtain data on the expected complex flow field in the near wall region. This status report summarizes the progress to date on developing the semi-span geometry definition suitable for generating structured grids for the computational research. In addition, the progress on evaluating three state-of-the-art Navier-Stokes codes is presented.

Chokani, Ndaona↗

NASA NTRS2005

Experimental Investigation of a Morphing Nacelle Ducted Fan

The application of Circulation Control to the nacelle of a shrouded fan is proposed as a means to enhance off-design performance of the shrouded fan. Typically, a fixed geometry shroud is efficient at a single operating condition. Modifying circulation about the fixed geometry is proposed as a means to virtually morph the shroud without moving surfaces. This approach will enhance off-design-point performance with minimal complexity, weight, and cost. Termed the Morphing Nacelle, this concept provides an attractive propulsion option for Vertical Take-off and Landing (VTOL) aircraft, such conceptual Personal Air Vehicle (PAV) configurations proposed by NASA. An experimental proof of concept investigation of the Morphing Nacelle is detailed in this paper. A powered model shrouded fan model was constructed with Circulation Control (CC) devices integrated in the inlet and exit of the nacelle. Both CC devices consisted of an annular jet slot directing a jet sheet tangent to a curved surface, generally described as a Coanda surface. The model shroud was tailored for axial flight, with a diffusing inlet, but was operated off-design condition as a static lifting fan. Thrust stand experiments were conducted to determine if the CC devices could effectively improve off-design performance of the shrouded fan. Additional tests were conducted to explore the effectiveness of the CC devices a means to reduce peak static pressure on the ground below a lifting fan. Experimental results showed that off-design static thrust performance of the model was improved when the CC devices were employed under certain conditions. The exhaust CC device alone, while effective in diffusing the fan exhaust and improving weight flow into shroud inlet, tended to diminish performance of the fan with increased CC jet momentum. The inlet CC device was effective at reattaching a normally stalled inlet flow condition, proving an effective means of enhancing performance. A more dramatic improvement in static thrust was obtained when the inlet and exit CC devices were operated in unison, but only over a limited range of CC jet momentum. Operating the nacelle inlet and exit CC devices together proved very effective in reducing peak ground plane static pressure, while maintaining static thrust. The Morphing Nacelle concept proved effective at enhancing off-design performance of the model; however, additional investigation is necessary to generalize the results.

Kondor, Shayne A.↗