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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Nonideal isentropic gas flow through converging-diverging nozzles

A method for treating nonideal gas flows through converging-diverging nozzles is described. The method incorporates the Redlich-Kwong equation of state. The Runge-Kutta method is used to obtain a solution. Numerical results were obtained for methane gas. Typical plots of pressure, temperature, and area ratios as functions of Mach number are given. From the plots, it can be seen that there exists a range of reservoir conditions that require the gas to be treated as nonideal if an accurate solution is to be obtained.

Bober, W.↗
NASA NTRS1943

Study of the Two-Dimensional Flow Through a Converging-Diverging Nozzle

A study of a flow through a straight converging-diverging nozzle of simple design has been made preliminary to studies of other supersonic flows. The diverging part of the nozzle was designed by the Prandtl-Busemann method to give a uniform pressure at its exit of 0.298 times the initial total head, that is, to give a Mach number of 1.436. Schlieren photographs of the flow and pressure-distribution measurements along the diverging part of the nozzle were made. A comparison of the theory with these measurements is presented.

Kantrowitz, Arthur↗
NASA NTRS1995

Numerical simulations of small amplitude acoustic wave propagation in a converging-diverging nozzle

Numerical simulations of the small amplitude acoustic wave propagation in a converging-diverging nozzle proposed in Category 5 are performed and presented here. The quasi 1-D unsteady flow equations in conservative form are discretized by the DRP scheme with the artificial damping developed by Tam and Webb. Characteristic boundary conditions are used at both the inlet and outlet of the nozzle. The effect of different numerical implementations of the subsonic inflow boundary conditions on the convergence of the solution to the steady state is studied. In the case of a subsonic outflow in which a shock is formed in the nozzle, the interaction between acoustics and the shock is also investigated.

Dong, Thomas Z.↗
NASA NTRS2002

Noise, Turbulence, and Thrust of Subsonic Free Jets from Lobed Nozzles

A study of noise benefit, vis-a-vis thrust penalty, and its correlation to turbulence intensities was conducted for free jets issuing from lobed nozzles. Four convergent nozzles with constant exit area were used in the experiments. Three of these were of rectangular lobed configuration having six, ten and fourteen lobes; the fourth was a circular nozzle. Increasing the number of lobes resulted in a progressive reduction in the turbulence intensities as well as in the overall radiated noise. The noise reduction was pronounced at the low frequency end of the spectrum. However, there was an increase in the high frequency noise that rendered the overall benefit less attractive when compared on a scaled-up A-weighted basis. A reduction in noise was accompanied by a commensurate reduction in the turbulent kinetic energy in the flow field. As expected, increasing the number of lobes involved progressive reduction in the thrust coefficient. Among the cases studied, the six-lobed nozzle had the optimum reduction in turbulence and noise with the least thrust penalty.

Zaman, K. B. M. Q.↗
NASA NTRS1991

Divergence thrust loss calculations for convergent-divergent nozzles: Extensions to the classical case

The analytical derivations of the non-axial thrust divergence losses for convergent-divergent nozzles are described as well as how these calculations are embodied in the Navy/NASA engine computer program. The convergent-divergent geometries considered are simple classic axisymmetric nozzles, two dimensional rectangular nozzles, and axisymmetric and two dimensional plug nozzles. A simple, traditional, inviscid mathematical approach is used to deduce the influence of the ineffectual non-axial thrust as a function of the nozzle exit divergence angle.

Berton, Jeffrey J.↗
NASA NTRSDate not supplied

Study of a Plug Nozzle for Supersonic Aircraft Concepts

Nozzles with external plugs are candidates for the exhaust system of supersonic aircraft concepts in an ongoing research effort at NASA. As part of that effort a fundamental study is being conducted at NASA Glenn Research Center (GRC). Experimental and limited numerical simulation results from this study, for a plug nozzle with a convergent nozzle (cowl),have been reported in recent publications[1,2]. Conventional method-of-characteristics (MoC) design for the plug requires that the flow is choked at the nozzle exit. This is why a sharp convergence of the cowl is invoked in such a design in most previous studies [3-5], leading to the earlier choice of the convergent cowl. Such a nozzle, however, is not suitable for supersonic flights because of large boat tail drag suffered by its outer surface. The boat tail drag consideration dictates that the cowl be more or less cylindrical in shape. While various other shapes have been studied in the past [6], essentially a cylindrical shape is adopted in the larger research program at NASA. The cylindrical shape has certain advantages. With the plug crown located somewhat inside the nozzle, the flow experiences convergence and divergence as with a C-D nozzle. By translating the plug, different throat-to-exit area ratio could be achieved to obtain fully expanded condition at different values of NPR. This offers a relatively easy active control strategy for the nozzle geometry throughout the flight regime from landing and takeoff (LTO) to cruise conditions. A plug nozzle model mimicking the geometry being considered in the NASA program, i.e., having a cylindrical exit, have been fabricated recently for continued fundamental studies. Some of the plugs used with the earlier convergent cowl [1] are readily adopted to this model. In addition, a few other plugs have been obtained following a numerical optimization study [7]. The various configurations are to be investigated for thrust, noise and flow fields, with complementing numerical simulations. Thrust data for a variety of configurations as well as limited noise data are in hand. Further explorations including schlieren visualization are to be conducted in the near future. The proposed paper is to summarize all these results. In the following, a few key results obtained so far are described.

jets↗
NASA NTRS1980

Investigation of convergent-divergent nozzles applicable to reduced-power supersonic cruise aircraft

An investigation was conducted of isolated convergent-divergent nozzles to determine the effect of several design parameters on nozzle performance. Tests were conducted using high pressure air for propulsion simulation at Mach numbers from 0.60 to 2.86 at an angle of attack of 0 deg and at nozzle pressure ratios from jet off to 46.0. Three power settings (dry, partial afterburning, and maximum afterburning), three nozzle lengths, and nozzle expansion ratios from 1.22 to 2.24 were investigated. In addition, the effects of nozzle throat radius and a cusp in the external boattail geometry were studied. The results of this study indicate that, for nozzles operating near design conditions, increasing nozzle length increases nozzle thrust-minus-drag performance. Nozzle throat radius and an external boattail cusp had negligible effects on nozzle drag or internal performance.

Berrier, B. L.↗
NASA NTRS1980

The effect of throat contouring on two-dimensional converging-diverging nozzles at static conditions

An experiment was conducted at static conditions to determine the internal performance effects of nozzle throat contouring, the result of increasing the circular-arc throat radius. Five nonaxisymmetric converging-diverging nozzles were tested at nozzle pressure ratios up to 9.0. Data are presented as internal thrust ratios, discharge coefficients, and static-pressure distributions. Comparisons of internal performance data for the five nozzles show that throat contouring increases the value of discharge coefficient but has no significant effect on internal thrust ratio except in cases of internal flow separation. To illustrate the use of the two dimensional converging-diverging (2-D C-D) nozzle data base, a two dimensional inviscid theory was applied to the five configurations. The generally good agreement of data with theoretical results indicates that two-dimensional inviscid theory can be successfully applied to the prediction of 2-D C-D nozzle internal flow.

Mason, M. L.↗
NASA NTRS1983

Investigation of installation effects on twin-engine convergent-divergent nozzles

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine installation effects on convergent-divergent nozzles applicable to twin-engine reduced-power supersonic cruise aircraft. Tests were conducted at Mach numbers from 0.50 to 1.20, angles of attack from -5 deg to 9 deg, and at nozzle pressure ratios from jet off (1.0) to 8.0. The effects of empennage arrangement, nozzle length, and afterbody closure on total and component drag coefficients were investigated.

Bare, E. A.↗
NASA NTRS1987

Static internal performance of a two-dimensional convergent-divergent nozzle with thrust vectoring

A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.

Bare, E. Ann↗
NASA NTRS1958

Experimental Behavior of Pentaborane-Air Combustion Products During Expansion in a Convergent Divergent Nozzle

In order to evaluate the post combustion behavior of boric oxide, pentaborane-air mixtures, burned to completion at a combustor pressure of 3 atmospheres, were expanded through a 7.1-inch-long convergent-divergent nozzle having a 4-inch-diameter throat and an exit-to-throat area ratio of 1.68. The experimentally determined thrust performance was in good agreement with the ideal equilibrium performance at stagnation temperatures of 3300 deg R and lower. The boric oxide vapor at the combustor exit required about 400 F deg supercooling before any condensed phase was observed. For a given thrust, fuel consumption was as much as 20 percent greater than predicted from vapor-pressure data for combustor outlet temperatures i n the vicinity of 3600 deg R. A similar result could be expected in full-scale engines, since the test combustor provided an unusually long dwell time and a highly turbulent environment. During the expansion process, the vapor (when present) did not condense to the extent predicted for an equilibrium expansion process. Moreover, condensation was observed only i n the form of small, abrupt phase changes i n the subsonic flow near the throat. Friction, due to liquid boric oxide films on the nozzle surfaces, was negligible when the surface temperature was above 800 F.

PENTABORANE↗
NASA NTRS1979

Two-phase choked flow of cryogenic fluids in converging-diverging nozzles

Data are presented for the two phase choked flow of three cryogenic fluids - nitrogen, methane, and hydrogen - in four converging-diverging nozzles. The data cover a range of inlet stagnation conditions, all single phase, from well below to well above the thermodynamic critical conditions. In almost all cases the nozzle throat conditions were two phase. The results indicate that the choked flow rates were not very sensitive to nozzle geometry. However, the axial pressure profiles, especially the throat pressure and the point of vaporization, were very sensitive to both nozzle geometry and operating conditions. A modified Henry-Fauske model correlated all the choked flow rate data to within + or - 10 percent. Neither the equilibrium model nor the Henry-Fauske model predicted throat pressures well over the whole range of data. Above the thermodynamic critical temperature the homogeneous equilibrium model was preferred for both flow rate and pressure ratio. The data of the three fluids could be normalized by the principle of corresponding states.

Simoneau, R. J.↗
NASA NTRS1982

Investigation of installation effects of single-engine convergent-divergent nozzles

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine installation effects on single-engine convergent-divergent nozzles applicable to reduced-power supersonic cruise aircraft. Tests were conducted at Mach numbers from 0.50 to 1.20, at angles of attack from -3 degrees to 9 degrees, and at nozzle pressure ratios from 1.0 (jet off) to 8.0. The effects of empennage arrangement, nozzle length, a cusp fairing, and afterbody closure on total aft-end drag coefficient and component drag coefficients were investigated. Basic lift- and drag-coefficient data and external static-pressure distributions on the nozzle and afterbody are presented and discussed.

Burley, J. R., II↗
NASA NTRS2002

Numerical Investigation of 'Transonic Resonance' with a Convergent-Divergent Nozzle

At pressure ratios lower than the design value, convergent-divergent (C-D) nozzles often undergo a flow resonance accompanied by the emission of acoustic tones. The phenomenon, driven by the unsteady shock within the divergent section of the nozzle, has been studied experimentally by Zaman et al. In this paper, the space-time conservation element solution element (CE/SE) method is employed to numerically investigate the phenomenon. The computations are performed for a given nozzle geometry for several different pressure ratios. Sustained 'limit cycle' oscillations are encountered in all cases. The oscillation frequencies, their variation with pressure ratio including a 'stage jump', agree well with the experimental results. The unsteady flow data confirm that stage 1 of the resonance (fundamental) involves a one-quarter standing wave while stage 2 (third harmonic) involves a three-quarter standing wave within the divergent section of the nozzle. Details of the shock motion, and the flow and near acoustic field, are documented for one case each of stages 1 and 2.

Loh, Ching Y.↗
NASA NTRS1975

Pressure distribution in a converging-diverging nozzle during two-phase choked flow of subcooled nitrogen

Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane is at the divergence end of the constant area throat section. At high stagnation pressures the fluid stays liquid well into the constant area throat region; at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region, and this anomaly is related to the two-phase process. The fluid is metastably all liquid below the saturation pressure.

Simoneau, R. J.↗
NASA NTRS1975

Pressure distribution in a converging-diverging nozzle during two-phase choked flow of subcooled nitrogen

Choked flow rates and axial pressure distributions were measured for subcooled nitrogen in a converging-diverging nozzle with a constant area section in the throat region. Stagnation pressures ranged from slightly above saturation to twice the thermodynamic critical pressure. Stagnation temperatures ranged from 0.75 to 1.03 times the thermodynamic critical temperature. The choking plane appears to be at the divergence end of the constant area throat section. At high stagnation pressures the fluid appears to stay liquid well into the constant area throat region; however, at near saturation stagnation pressures it appears that vaporization occurs at or before the entrance to the constant area throat region. The throat-to-stagnation pressure ratio data exhibits an anomalous flat region. This anomaly appears to be fundamentally related to the two-phase process and not merely to the present specific nozzle geometry. The fluid appears to be metastably all liquid below the saturation pressure. The data are compared to various flow models. No model adequately describes the whole range of the experiment.

Simoneau, R. J.↗