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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 37 records · Page 2
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

The effect of test-cell pressure on resistojet nozzle flow

Previous experimental work has shown that measured resistojet thrust decreases from that obtained at hard vacuum conditions as the test-cell pressure rises above 0.001 torr. Thrust losses have been observed for both cold and heated flow conditions, and the most significant losses have been experienced using thrusters with low Reynolds number flow and high area ratio nozzles. In order to further investigate nozzle flow characteristics, a pressure probe having four degrees of freedom has been used to obtain stagnation pressure surveys across the nozzle exit planes of four resistojets. The surveys show a change in the ratio of the supersonic core to the viscous boundary layer flow areas as the test-cell pressure increases. The surveys are also used for detecting whether an oblique shock is present in an overexpanded nozzle flow. Thruster temperature measurements and nozzle exit plane pressure surveys indicate that thrust losses are the combined result of convection heat losses and nozzle flow momentum effects.

Penko, Paul F.↗
NASA NTRS1992

Qualitative spectroscopic study of magnetic nozzle flow

The physics of the magnetic nozzle flow for a 100-kW-level quasi-steady MPD thruster was studied by photographic spectroscopy focusing on the plasma model in the flow and the acceleration mechanism. Spectroscopic visualization for the flow-species analysis indicates that the plasma-exhaust flow dominated by NII species were confined by the magnetic nozzle effect to collimate the flow for the better thruster performance. Inside the nozzle, the plasma flow was found to be in nonhomogeneous collisional-radiative condition. There appears to be a substantial flow acceleration from the magnetic nozzle inlet to the outlet with slight expansion. This suggests that the flow resembles that of constant area supersonic duct flow with cooling.

Umeki, T.↗
NASA NTRS2004

Experimental, Theoretical, and Computational Investigation of Separated Nozzle Flows

A detailed experimental, theoretical, and computational study of separated nozzle flows has been conducted. Experimental testing was performed at the NASA Langley 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a sub-scale, non-axisymmetric, two-dimensional, convergent- divergent nozzle. In addition, two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and algebraic Reynolds stress modeling. For reference, experimental and computational results were compared with theoretical predictions based on one-dimensional gas dynamics and an approximate integral momentum boundary layer method. Experimental results from this study indicate that off-design overexpanded nozzle flow was dominated by shock induced boundary layer separation, which was divided into two distinct flow regimes; three- dimensional separation with partial reattachment, and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results indicate that with controlled separation, the entire overexpanded range of nozzle performance would be within 10% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical exhaust system, this may allow a fixed geometry nozzle to cover an entire flight envelope. The computational simulation was in excellent agreement with experimental data over most of the test range, and did a good job of modeling internal flow and thrust performance. An exception occurred at low nozzle pressure ratios, where the two-dimensional computational model was inconsistent with the three-dimensional separation observed in the experiment. In general, the computation captured the physics of the shock boundary layer interaction and shock induced boundary layer separation in the nozzle, though there were some differences in shock structure compared to experiment. Though minor, these differences could be important for studies involving flow control or thrust vectoring of separated nozzles. Combined with other observations, this indicates that more detailed, three-dimensional computational modeling needs to be conducted to more realistically simulate shock-separated nozzle flows.

Hunter, Craig A.↗
NASA NTRS1972

Splitting supersonic nozzle flow into separate jets by overexpansion into a multilobed divergent nozzle

Air flowing from a convergent nozzle at pressure ratios greater than 2.5 has been split into eight separate jets by overexpansion of the flow into a divergent, eight-lobed passage. The splitting of the flow is accompanied by a decrease in the nozzle axial centerline Mach number. This in part is due to the radial inflow of secondary air between the lobes toward the nozzle centerline. Each of the smaller jets is partially split after it leaves the end of the divergent lobed section of the nozzle, thus creating a velocity profile having 16 peaks. At a pressure ratio of 3.5 the flow decelerates to Mach 1 in three convergent nozzle throat diameters. Convergent nozzle flow normally requires 12 diameters to reach Mach 1. The nozzle has a sound attenuation of 12 decibels with a thrust loss of 9 percent for the best configuration tested.

Huff, R. G.↗
DOE OSTI2023

Computational study of the ECN spray C via one-way coupling of internal nozzle flow and ensuing spray

Here, in this study, computational fluid dynamics simulations of the Engine Combustion Network Spray C injector were performed to investigate the effect of the internal nozzle flow on the development of the ensuing spray in automotive fuel injectors. A set of simulation best practices was first developed to achieve an accurate representation of the internal nozzle flow. Detailed, spatiotemporally defined information was extracted at the injector’s orifice exit and stored in maps that were subsequently employed to initialize a Lagrangian spray by means of a static, one-way coupling (OWC) approach. To highlight the impact of the in-nozzle flow on the spray, the OWC approach was compared with the conventional rate-of-injection (ROI) approach, for which the spray parcels are commonly initialized using a blob injection model. The results indicated that, using the ROI approach, considerable evaporation was observed within the plume in the near-nozzle region. On the other hand, the OWC approach showed evidence of evaporation farther downstream and at the plume periphery. The different behavior observed with the two approaches was mostly due to substantial differences in fuel mass distribution. The liquid fuel concentration predicted by the OWC approach matched well with X-ray tomography measurements of projected fuel mass per unit volume, in terms of both magnitude and distribution. In contrast, an unphysical distribution of fuel mass characterized by an extremely low liquid concentration was predicted by the ROI approach. In the far-field, the effect of the in-nozzle flow was found to be weaker as the spray penetrated and interacted with the ambient gas. In this region, the OWC and ROI approaches provided similar predictions of vapor penetration, spray morphology, and mixing-related characteristics.

33 ADVANCED PROPULSION SYSTEMS↗
NASA NTRS2002

Measurements of Turbulent Flow Field in Separate Flow Nozzles with Enhanced Mixing Devices - Test Report

As part of the Advanced Subsonic Technology Program, a series of experiments was conducted at NASA Glenn Research Center on the effect of mixing enhancement devices on the aeroacoustic performance of separate flow nozzles. Initial acoustic evaluations of the devices showed that they reduced jet noise significantly, while creating very little thrust loss. The explanation for the improvement required that turbulence measurements, namely single point mean and RMS statistics and two-point spatial correlations, be made to determine the change in the turbulence caused by the mixing enhancement devices that lead to the noise reduction. These measurements were made in the summer of 2000 in a test program called Separate Nozzle Flow Test 2000 (SFNT2K) supported by the Aeropropulsion Research Program at NASA Glenn Research Center. Given the hot high-speed flows representative of a contemporary bypass ratio 5 turbofan engine, unsteady flow field measurements required the use of an optical measurement method. To achieve the spatial correlations, the Particle Image Velocimetry technique was employed, acquiring high-density velocity maps of the flows from which the required statistics could be derived. This was the first successful use of this technique for such flows, and shows the utility of this technique for future experimental programs. The extensive statistics obtained were likewise unique and give great insight into the turbulence which produces noise and how the turbulence can be modified to reduce jet noise.

Bridges, James↗
NASA NTRS1993

DSMC simulation of low Reynolds number nozzle flows

A numerical analysis of low Reynolds number nozzle flows was performed to investigate the loss mechanisms involved and to determine the nozzle wall contour that minimizes these losses. DSMC was used to simulate flows through three different nozzle configurations at two different stagnation chamber temperatures so that the heat transfer losses could be separated from the wall contour effects on performance. A trumpet-shaped nozzle had 5 percent higher efficiency than a conical nozzle and a 3 percent higher efficiency than a bell-shaped nozzle with the unheated flow. With heated flow both the trumpet and bell-shaped nozzles had a 6.5 percent higher efficiency than the conical nozzle. The conical nozzle had the highest discharge coefficient of the three configurations, 0.92, and the trumpet-shaped nozzle had the lowest, 0.82. The discharge coefficient of each nozzle was unaffected by the change in stagnation temperature; however the increase in stagnation temperature increased the heat transfer and viscous losses in the boundary layer. These results suggest that the trumpet-shaped wall contour performed most efficiently except near the throat region, where it incurred large viscous losses. However, the bell-shaped nozzle may increase its overall performance with an increase in stagnation temperature.

Zelesnik, D.↗
NASA NTRS1991

Penn State axial flow turbine facility: Performance and nozzle flow field

The objective is to gain a thorough understanding of the flow field in a turbine stage including three-dimensional inviscid and viscid effects, unsteady flow field, rotor-stator interaction effects, unsteady blade pressures, shear stress, and velocity field in rotor passages. The performance of the turbine facility at the design condition is measured and compared with the design distribution. The data on the nozzle vane static pressure and wake characteristics are presented and interpreted. The wakes are found to be highly three-dimensional, with substantial radial inward velocity at most spanwise locations.

Lakshminarayana, B.↗
NASA NTRS1976

Analysis of nitrogen condensation in an expanding nozzle flow

Condensation of nitrogen flow in an expanding nozzle flow is analyzed using one-dimensional gas dynamic equations and the equations for nucleation and droplet growth. Effects of variations in the Tolman constant and the mass accommodation factor are discussed as well as the effect of foreign nuclei. Comparisons are made with experimental data obtained from a small, contoured nozzle.

Wang, F. C.↗
NASA NTRS2011

Jet Engine Exhaust Nozzle Flow Effector

A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

Turner, Travis L.↗
NASA NTRS2014

Jet Engine Exhaust Nozzle Flow Effector

A jet engine exhaust nozzle flow effector is a chevron formed with a radius of curvature with surfaces of the flow effector being defined and opposing one another. At least one shape memory alloy (SMA) member is embedded in the chevron closer to one of the chevron's opposing surfaces and substantially spanning from at least a portion of the chevron's root to the chevron's tip.

Turner, Travis L.↗
NASA NTRS1991

Low thrust viscous nozzle flow fields prediction

A Navier-Stokes code was developed for low thrust viscous nozzle flow field prediction. An implicit finite volume in an arbitrary curvilinear coordinate system lower-upper (LU) scheme is used to solve the governing Navier-Stokes equations and species transportation equations. Sample calculations of carbon dioxide nozzle flow are presented to verify the validity and efficiency of this code. The computer results are in reasonable agreement with the experimental data.

Liaw, G. S.↗