Search NASA⌕ Search

SEARCH · Search NASA

Results for “circular nozzles”

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.

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

At least 127 records · Page 7
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

An interactive adaptive remeshing algorithm for the two-dimensional Euler equations

An interactive adaptive remeshing algorithm utilizing a frontal grid generator and a variety of time integration schemes for the two-dimensional Euler equations on unstructured meshes is presented. Several device dependent interactive graphics interfaces have been developed along with a device independent DI-3000 interface which can be employed on any computer that has the supporting software including the Cray-2 supercomputers Voyager and Navier. The time integration methods available include: an explicit four stage Runge-Kutta and a fully implicit LU decomposition. A cell-centered finite volume upwind scheme utilizing Roe's approximate Riemann solver is developed. To obtain higher order accurate results a monotone linear reconstruction procedure proposed by Barth is utilized. Results for flow over a transonic circular arc and flow through a supersonic nozzle are examined.

Slack, David C.↗
NASA NTRS1999

Application of the RNS3D Code to a Circular-Rectangular Transition Duct With and Without Inlet Swirl and Comparison with Experiments

Circular-rectangular transition ducts are used between engine exhausts and nozzles with rectangular cross sections that are designed for high performance aircraft. NASA Glenn Research Center has made experimental investigations of a series of circular-rectangular transition ducts to provide benchmark flow data for comparison with numerical calculations. These ducts are all designed with superellipse cross sections to facilitate grid generation. In response to this challenge, the three-dimensional RNS3D code has been applied to one of these transition ducts. This particular duct has a length-to-inlet diameter ratio of 1.5 and an exit-plane aspect ratio of 3.0. The inlet Mach number is 0.35. Two GRC experiments and the code were run for this duct without inlet swirl. One GRC experiment and the code were also run with inlet swirl. With no inlet swirl the code was successful in predicting pressures and secondary flow conditions, including a pair of counter-rotating vortices at both sidewalls of the exit plane. All these phenomena have been reported from the two GRC experiments. However, these vortices were suppressed in the one experiment when inlet swirl was used; whereas the RNS3D code still predicted them. The experiment was unable to provide data near the sidewalls, the very region where the vortices were predicted.

Cavicchi, Richard H.↗
NASA NTRS1978

Boundary-layer separation on isolated boattail nozzles

An angle of attack of 0 deg was investigated in the Langley 16 foot transonic tunnel at free-stream Mach numbers from 0.40 to 0.95 to study the phenomenon of separated flow on a series of circular-arc afterbodies. Both high-pressure air and solid circular cylinders with the cylinder diameter equal to the nozzle-exit diameter were used to simulate jet exhausts. The results indicate that boundary-layer separation is most extensive on steep boattails at high Mach numbers. The jet total-pressure ratio changes (jet total pressure to free-stream static pressure) affected the extent of separation very little; however, comparison of the separation data obtained by using the two jet-simulation techniques indicate that entrainment associated with the presence of a jet had a significant effect on the extent of separation. The solid-simulator separation data were also used to evaluate the predictions of eight separation criteria.

Abeyounis, W. K.↗
NASA NTRS1975

Suppression of nonlinear oscillations in combustors with partial length acoustic liners

An analytical model is formulated for a three-dimensional nonlinear stability problem in a rocket motor combustion chamber. The chamber is modeled as a right circular cylinder with a short (multi-orifice) nozzle, and an acoustic linear covering an arbitrary portion of the cylindrical periphery. The combustion is concentrated at the injector and the gas flow field is characterized by a mean Mach number. The unsteady combustion processes are formulated using the Crocco time lag model. The resulting equations are solved using a Green's function method combined with numerical evaluation techniques. The influence of acoustic liners on the nonlinear waveforms is predicted. Nonlinear stability limits and regions where triggering is possible are also predicted for both lined and unlined combustors in terms of the combustion parameters.

Espander, W. R.↗
NASA NTRS1984

Real jet effects on dual jets in a crossflow

A 6-ft by 6-ft wind tunnel section was modification to accommodate the 7-ft wide NASA dual-jet flate model in an effort to determine the effects of nonuniform and/or noncircular jet exhaust profiles on the pressure field induced on a nearby surface. Tests completed yield surface pressure measurements for a 90 deg circular injector producing exit profiles representative of turbofan nozzles (such as the TF-34 nozzle). The measurements were obtained for both tandem and side-by-side jet configurations, jet spacing of S/D =2, and velocity ratios of R=2.2 and 4.0. Control tests at the same mass flow rate but with uniform exit velocity profiles were also conducted, for comparison purposes. Plots for 90 deg injection and R=2.2 show that the effects of exit velocity profile nonuniformity are quite significant.

Schetz, J. A.↗
NASA NTRS1962

Performance Evaluation of Fixed- and Variable-Area Rocket Exhaust Diffusers Using Single and Clustered Nozzles with and Without Gimbaling

An investigation of exhaust diffusers used for altitude simulation in testing rocket engines was conducted with model diffusers and gaseous nitrogen as the working fluid. This investigation was conducted to eval­uate the effects on performance of gimbaling clustered nozzles in a single fixed-area exhaust diffuser. A two-nozzle cluster was gimbaled in all attitudes in four different exhaust-diffuser configurations: (1) a straight circular tube, (2) a straight figure-eight tube, (3) a circular tube with a second throat, and (4) a figure-eight tube with a second throat. Gimbaling clustered nozzles had little effect on the operating pressure ratio for either straight-tube or second-throat diffusers, but caused a large increase in the starting-pressure-ratio requirements in some straight-tube exhaust diffusers. The performance of two or four clustered nozzles with no gimbaling was compared with single-nozzle per­formance on the basis of the ratio of diffuser area to nozzle-throat area. The performance was found to be dependent on the diffuser- to nozzle-throat-area ratio and independent of both the number of primary nozzles and the nozzle-area ratio. A separate investigation was also conducted to evaluate the per­formance improvement obtainable with a variable-area exhaust diffuser. This type of diffuser achieved a 26-percent reduction in starting pressure ratio, a 40-percent reduction in operating pressure ratio, and a 50-percent reduction in the overall diffuser length over the values attainable with a fixed-area second-throat diffuser.

ROCKET↗
NASA NTRS1974

Stability of partially lined combustors with distributed combustion

An integral method for theoretical evaluation of the stability of confined flows with mass sources is presented. This method is applied to evaluate the effect of partial length acoustic liners on liquid propellant combustors with distributed combustion sources. The combustor is modeled as a right circular cylinder terminated by a multiorifice (constant Mach number) nozzle. The spatial spread of combustion in the chamber is represented by an arbitrary number of pressure sensitive planar mass sources. On the cylindrical periphery of the chamber is an acoustic absorber of arbitrary length, position, and damping capability. Integral equations are developed from the partial differential equations describing the nonsteady flow and are solved iteratively. Results are given in terms of linear neutral stability plots using a pressure sensitive combustion response. Calculations indicate that increasing the axial spread of combustion increases the stability of the combustor and the effectiveness of the absorber. The most effective linear placement appears at the location of greatest combustion concentration.

Baer, M. R.↗
NASA NTRS1982

Theoretical evaluation of a V/STOL fighter model utilizing the PAN AIR code

The PAN AIR computer code was investigated as a tool for predicting closely coupled aerodynamic and propulsive flowfields of arbitrary configurations. The NASA/Ames V/STOL fighter model, a configuration of complex geometry, was analyzed with the PAN AIR code. A successful solution for this configuration was obtained when the nozzle exit was treated as an impermeable surface and no wakes were included around the nozzle exit. When separated flow was simulated from the end of the nacelle, requiring the use of wake networks emanating from the nozzle exit, a number of problems were encountered. A circular body nacelle model was used to investigate various techniques for simulating the exhaust plume in PAN AIR. Several approaches were tested and eliminated because they could not correctly simulate the interference effects. Only one plume modeling technique gave good results. A PAN AIR computation that used a plume shape and inflow velocities obtained from the Navier-Stokes solution for the plume produced results for the effects of power that compared well with experimental data.

Howell, G. A.↗
NASA NTRS1995

Auxiliary Control Mechanisms For RUM Actuators

RUM actuator proposed device mounted on gimbaled, pivoted, or freely floating instrument and used to make instrument oscillate about one or more axes. Includes one or more (typically, two) unbalanced masses (e.g., lump masses on arms) for each axis about which oscillation desired. Unbalanced masses driven to rotate at constant angular velocity to produce oscillating centrifugal forces, which, in turn, exerts cyclic torques that make instrument oscillate about desired axis. Used in both terrestrial and spacecraft applications: for example, to oscillate instruments as diverse as x-ray telescopes, magnetometers, accelerometers, and agricultural spraying nozzles in variety of scan patterns that include linear, circular, and raster. Weights, power demands, and costs reduced.

Polites, Michael E.↗
NASA NTRS1990

Experimental investigation of a single flush-mounted hypermixing nozzle

Reported herein are the results of an experimental wind tunnel investigation of a circular supersonic jet (Mj = 3.47) injected at a 10 degree angle into a supersonic freestream. Measurements were made for nominal freestream Mach numbers of 1.6, 2.0, 2.5 and 3.0. Three jet total pressures were run at each freestream Mach number, resulting in twelve separate operating conditions. The measurements indicate the presence of two pairs of contra-rotating vortices. One pair follows the jet trajectory and tends to split the jet into two streams. A smaller pair, rotating in an opposite sense, develops in the near wall region. Reported results include Mach number and volume fraction distributions in the cross plane, as well as jet penetration and mixing efficiency.

Davis, D. O.↗
NASA NTRS1989

Supersonic jet noise and the high speed civil transport

An evaluation is made of the comparative advantages of prospective SST engine noise-suppression systems, with a view to their effectiveness in meeting the federally-mandated community noise standards of FAR 36 Stage III. A noise-suppression system must be capable of removing at least 4 EPNdB of noise percent thrust loss at takeoff. While none of the suppressors presently discussed is capable of meeting this goal, the inverted velocity profile/annular convergent-divergent plug/acoustically-treated ejector suppressor combination of configurational elements appears to represent the most efficient noise-control apparatus. Noncircular cross-section nozzle geometries also furnish a general noise reduction advantage over circular ones.

Seiner, John M.↗
NASA NTRS1983

Remarks on the layout of the subsonic free jet wind tunnels

By means of a recently installed wind tunnel with a circular free jet, it is shown that requirements concerning the flow parameters (e.g., uniform velocity profile and uniform and low turbulence level in the nozzle exit) can be easily and at moderate expenses fulfilled without changing the settling chamber and the nozzle itself. The installations in the settling chamber are adjustable. The structure is not limited to settling chambers with a circular cross section.

Vagt, J. D.↗
NASA NTRS1977

Jet noise source modification due to forward flight

The effects of forward flight on the turbulence characteristics of a jet in a co-flowing stream have been determined for a 2.22 inch circular jet in a 36 inch free jet wind tunnel. The nozzle exit velocity was 400 ft/sec, and the tunnel velocity was set at 0, 40, 120 and 200 ft/sec. Measurements of flow properties including mean velocity, turbulence intensity and spectra, convection velocity, integral length scale, and convected integral time scale were carried out using two linearized hot wires. Results were used to predict changes due to flight in the jet acoustic sources. The noise reductions for a cold jet with a velocity of 1000 ft/sec, due to the change in acoustic sources in flight, agreed well at all angles with measured noise reductions.

Larson, R. S.↗
NASA NTRS1979

A simplified computer program for the prediction of the linear stability behavior of liquid propellant combustors

A program for predicting the linear stability of liquid propellant rocket engines is presented. The underlying model assumptions and analytical steps necessary for understanding the program and its input and output are also given. The rocket engine is modeled as a right circular cylinder with an injector with a concentrated combustion zone, a nozzle, finite mean flow, and an acoustic admittance, or the sensitive time lag theory. The resulting partial differential equations are combined into two governing integral equations by the use of the Green's function method. These equations are solved using a successive approximation technique for the small amplitude (linear) case. The computational method used as well as the various user options available are discussed. Finally, a flow diagram, sample input and output for a typical application and a complete program listing for program MODULE are presented.

Mitchell, C. E.↗
NASA NTRS1940

Standards for discharge measurement with standardized nozzles and orifices

The following standards give the standardized forms for two throttling devices, standard nozzles and standard orifices, and enable them to be used in circular pipes without calibration. The definition of the standards are applicable in principle to the calibration and use of nonstandardized throttling devices, such as the venturi tube. The standards are valid, likewise, as a basis for discharge measurements in the German acceptance standards.

Source record↗
NASA NTRS2020

NASA Exploration Systems & Habitation (X-Hab)Academic Innovation Challenge 2020: Microgravity Gas/Liquid Separator for the CO2 Revitalization System Final Report

This report outlines the efforts of UNT X-Hab Team on a project titled, Microgravity Gas/Liquid Separator for the CO2 Revitalization System. The project is executed as a senior design project during academic year 2019-2020. The X-Hab 2020 Academic Innovation Challenge has selected eleven senior design teams from colleges across the US to demonstrate working prototypes for exploration systems and habitation. UNT has been tasked with the creation of a gas-liquid separator for an air revitalization system. Air revitalization technology has been used to support spaceflight by removing CO2 from enclosed systems in order to maintain breathable air. Solid sorbents such as zeolites or lithium hydroxide have been used in the past for these systems but are difficult to handle in microgravity environments and require a large amount of energy. This challenge aims to demonstrate vortex phase separator (VPS) technology for removing H2O from a CO2 stream. However, VPS technology also has the capability of using liquid sorbents for removing CO2 from an air stream. Further research could be done on liquid sorbents, such as liquid amine, in use with VPS systems. This would potentially be an alternative technology in replacing use of solid sorbents in CO2 removal systems. In 2019, NASA proposed a design that uses a gas/liquid contactor to allow for efficient contact between the two fluid phase. This was integrated into an overall CO2 removal system. The subsystems for gas-liquid separation and storage in NASA’s previous models for CO2 removal system could be replaced with a VPS. Innovative vortex separator technology is expected to allow for high throughput flow and highly efficient CO2 removal compared to other gas-liquid separation technologies. VPS relies on centripetal driven buoyancy forces to form a gas-liquid vortex within a fixed, right circular cylinder. The gas stream enters the separator through a tangential nozzle and breaks into very small bubbles (<<1 mm) resulting in a very large contact surface area for interaction with the liquid stream via energy and mass exchange. VPS technology can handle mismatches in inlet and outlet flow rates and system volume changes through the range of liquid thickness held in the separator (i.e., buffering and accumulating capability), and requires low pressure differences (<5-10 in H2O in most cases) for operation. The X-Hab 2020 team leverages these characteristics to investigate VPS technology as an alternative CO2 removal technology.

Alyssa Sarvadi↗
NASA NTRS1999

Subsonic Jet Noise from Non-Axisymmetric and Tabbed Nozzles

Subsonic jet noise from non-axisymmetric and tabbed nozzles are investigated experimentally and theoretically. It is shown that the noise spectra of these jets are in good agreement with the similarity spectra found empirically earlier by Tam, Golebiowski and Seiner through a detailed analysis of supersonic jet noise data. Further, the radiated noise fields of the jets under study, including elliptic and large aspect ratio rectangular jets, are found to be quite axisymmetric and are practically the same as that of a circular jet with the same exit area. These experimental results strongly suggest that nozzle geometry modification into elliptic or rectangular shapes is not an effective method for jet noise suppression. A lobed nozzle, on the other hand, is found to significantly impact the noise field. Noise from large scale turbulent structures, radiating principally in the downstream direction, is effectively suppressed. Tabs also impact the noise field, primarily by shifting the spectral peak to a higher frequency. A jetlets model is developed to provide a basic understanding of the noise from tabbed jets. The model predicts that the noise spectrum from a jet with N tabs (N greater than or equal to 2) can be obtained from that of the original jet (no tab) by a simple frequency shift. The shifted frequency is obtained by multiplying the original frequency by N(exp 1/2). This result is in fairly good agreement with experimental data.

Tam, Christopher K. W.↗