SEARCH · Engineering Papers

Results for “SONIC NOZZLE”

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 109 records · Page 6

All sources NASA NTRS DOE OSTI

NASA NTRS2026

Thermochemically-Closed Sonic-Flow Inversion for Enthalpy and Temperature in Multispecies Arc-Jet Flows

A thermochemically-closed sonic-flow inversion framework (TSIF) is developed to infer bulk enthalpy and total temperature upstream of a choked nozzle in arc-jet flows. The formulation recasts a pressure-rise total enthalpy quantification technique as an inverse problem in characteristic-velocity c * space using measured mass flow rate, upstream total pressure, gas composition, and nozzle throat geometry as inputs. Unlike calorimetric energy-balance approaches or optical diagnostics, the method relies primarily on routinely measured facility quantities combined with explicit thermochemical closure. Thermochemical states are obtained using NASA’s open-source Chemical Equilibrium with Applications (CEA) code, enabling construction of a chemistry-consistent relation between characteristic velocity, total enthalpy, and total temperature under equilibrium or frozen assumptions. A discharge coefficient is self-calibrated using cold-flow (arc-off) operation data and applied to hot-flow (arc-on) measurements, enabling upstream losses to be accounted for without empirical correlations. The framework is applied to air, N 2 , and CO 2 –N 2 arc-jet flows and demonstrates expected trends for the inferred thermochemical states as function of arc power, specific energy input, mass-flow, heater configuration, and test gas. In the air limit, under equilibrium assumptions, the method recovers the classical high-enthalpy asymptotic correlation of Winovich with a mean residual of 4.4%, demonstrating compatibility with established sonic-flow scaling, while extending applicability to arbitrary multi-species mixtures and non-equilibrium chemistry. The framework provides a mixture-flexible methodology for determining bulk thermochemical states in modern arc-jet environments using routine facility pressure, mass-flow, gas-composition, and nozzle-geometry information together with a cold-flow calibration.

inviscid theory↗

NASA NTRS1986

Acoustic properties associated with rectangular geometry supersonic nozzles

Acoustic property experiments have been conducted to ascertain the behavior of rectangular geometry supersonic nozzles whose throat aspect ratios vary over a 2.0-7.6 range, and whose three partial sidewall geometries range from full to 75-percent cutback. The tests employed unheated air at static conditions for nozzle Mach numbers of 1.35-1.66. It is found that sonic fatigue failures are possible at certain partial sidewall geometries and high nozzle aspect ratios. Unlike axisymmetric supersonic nozzles, shock noise dominates both the rear and forward arc for throat aspect ratio cases greater than 5.6. Jet screech frequency was adequately predicted with a simple vortex sheel model.

Seiner, J. M.↗

NASA NTRS1959

Near-Field and Far-Field Noise Surveys of Solid-Fuel Rocket Engines for a Range of Nozzle Exit Pressures

Measurements of near- and far-field noise pressures are presented for a 1,500-pound-thrust engine and for several 5,000-pound-thrust engines for which the nozzle exit pressure was changed systematically in order to study its effects on the noise level and spectra. Near-field surveys indicated that the highest noise pressure occurred at about 20 exit diameters downstream if the nozzle near the transition from super-sonic to subsonic flow. The acoustical power radiated from all engines averaged about 0.5 percent of the mechanical power of the exhaust stream, the least noise being radiated by the nozzle having an exit pressure less than atmospheric. The rocket engines of these tests radiate more power per cycle at the lower frequencies than arte reported for subsonic jets in other related studies.

Mayes, William H.↗

NASA NTRSDate not supplied

Subscale Motor Test Bed Development to Simulate Solid Rocket Motor Internal Environments for Material Characterization

NASA Marshall Space Flight Center’s Materials and Processes Department, with support from the Propulsion Systems Department has developed a test bed for exposing ablative thermal protection materials to an environment similar to that of solid rocket motors (SRMs) by using a hybrid-propellant combustor. The Solid Fuel Torch (SFT) is a pressurized. small-scale hybrid rocket motor which utilizes an aluminized solid fuel and a gaseous oxygen (GOx) oxidizer to vary internal environments which can be tailored to specific program needs. The baseline design for the SFT was developed to support the Space Shuttle Program and was redesigned to support Space Launch System (SLS) Booster program requirements. There are four configurations of the SFT testbed: 1.) Insulation configuration with the option of a “dome” or “converging cone” insulation test section, 2.) Super Sonic Blast Tube configuration to simulate nozzle exit cone environments, 3.) Flow Straightened configuration which includes an additional instrumented length between the fuel grain and the insulation, and 4.) Constant Diameter Throat configuration which includes a cylindrical phenolic test section to test samples with different ply angles under the same environment simultaneously. Motor pressures for each configuration can range from 150 to 1500 psi with heat fluxes up to 1200 BTU/sq.ft/s. This paper discusses the development of each configuration and the results obtained during the initial test series.

C.L. Bower↗

NASA NTRS2014

An Interactive Method of Characteristics Java Applet to Design and Analyze Supersonic Aircraft Nozzles

The Method of Characteristics (MOC) is a classic technique for designing supersonic nozzles. An interactive computer program using MOC has been developed to allow engineers to design and analyze supersonic nozzle flow fields. The program calculates the internal flow for many classic designs, such as a supersonic wind tunnel nozzle, an ideal 2D or axisymmetric nozzle, or a variety of plug nozzles. The program also calculates the plume flow produced by the nozzle and the external flow leading to the nozzle exit. The program can be used to assess the interactions between the internal, external and plume flows. By proper design and operation of the nozzle, it may be possible to lessen the strength of the sonic boom produced at the rear of supersonic aircraft. The program can also calculate non-ideal nozzles, such as simple cone flows, to determine flow divergence and nonuniformities at the exit, and its effect on the plume shape. The computer program is written in Java and is provided as free-ware from the NASA Glenn central software server.

Analysis↗

NASA NTRS1972

Factors which influence the analysis and design of ejector nozzles

A theoretical analysis of the viscous interaction between the primary and secondary streams of ejector nozzles was developed. The analysis accounts for real sonic line effects and the streamwise variation in stream mixing and boundary layer within the ejector. The aspects of the analysis are explained and illustrated by applying the theory to a variety of ejector configurations including cylindrical shroud, contoured flap and plug nozzles. Extensive comparisons are made between theory and data to show the importance of various analytical assumptions and such design variables as diameter ratio, spacing ratio, total temperature ratio, and primary nozzle geometry.

Anderson, B. H.↗

NASA NTRS1973

Double expansion nozzles for shock tunnels and Ludwieg tubes

Double expansion nozzles use a shock system located between two sonic throats in order to exchange total pressure for flow cross-sectional area. They can be useful for gaskinetic and aerodynamic flows that otherwise cannot take advantage of available pressure capability. This paper reports experiments with a 7.5 by 10 cm, Mach 3 nozzle operated in both conventional and double expansion nozzle modes using 7.5 by 10 cm and 2.5 cm I.D. Ludwieg drive tubes, respectively. Pressure traces document a simple starting model for the double expansion case, while holographic interferometry has been used to establish and compare nozzle flow quality. Interferometry is also used to estimate turbulence levels through onset of fringe breakdown. The measurements generally show that, with careful design, double expansion nozzles can be used to extend facility performance without loss in flow quality and with acceptable loss in testing time.

Russell, D. A.↗

NASA NTRS1973

Transonic laminar boundary layers with surface curvature.

The effect of surface curvature (both longitudinal and transverse) and the associated pressure gradient across the flow is investigated analytically for a laminar boundary layer subjected to pressure gradients along the flow. Property variation which results from heat transfer and compressibility is taken into account. Numerical solutions of the boundary layer equations are obtained for locally similar sonic flow through the throat of a nozzle for a range of flow conditions and for various shaped nozzle surfaces with different amounts of wall cooling. A few solutions were also obtained for the analogous flow around the shoulder of a flat-faced body in a supersonic flow. The effect of various parameters that arise in the equations upon application of the Levy-Mangler transformation are investigated and discussed with respect to their influence on the velocity and total enthalpy profiles and the corresponding profile slopes at the surface to which the shear stress and heat transfer are related. An important finding is that at throat Reynolds numbers less than 100,000 the heat transfer parameter at a nozzle throat decreases as the throat radius of curvature decreases.

Back, L. H.↗

NASA NTRS1974

Computer program for calculating the flow field of supersonic ejector nozzles

An analytical procedure for computing the performance of supersonic ejector nozzles is presented. This procedure includes real sonic line effects and an interaction analysis for the mixing process between the two streams. The procedure is programmed in FORTRAN 4 and has operated successfully on IBM 7094, IBM 360, CDC 6600, and Univac 1108.

Anderson, B. H.↗

NASA NTRS2018

Computational and Experimental Study of Plume and Shock Interaction Effects on Sonic Boom in the NASA Ames 9x7 Supersonic Wind Tunnel

A wind tunnel test was performed in the NASA Ames 9x7 Supersonic Wind Tunnel focusing on the shock waves traveling through and interacting with an exhaust nozzle plume. This experimental study was conducted to develop and validate the CFD capability required to accurately include nozzle flow with impinging shock effects on near field and ground‐propagated sonic boom signatures. The model was made to be generic, and included a simple nozzle shape, two different aft decks, and a few generic horizontal tails. High pressure air was pumped through a nozzle at various nozzle pressure ratios (NPR) to represent the engine plume in flight. The three different aft body representations each created a different shock wave signature that passed through the plume. An aft deck configuration, where part of the aircraft shields the nozzle plume, was also tested. Retroreflective Background-Oriented Schlieren (RBOS) was used to obtain schlieren images of the flow field around the model and behind the model. This study compares wind tunnel data and numerical simulations conducted by the NASA Tetrahedral Unstructured Software System CFD code, USM3D.

Winski, Courtney S.↗

NASA NTRS1972

Factors which influence the analysis and design of ejector nozzles.

A theoretical analysis of the viscous interaction between the primary and secondary streams of ejector nozzles was developed. The analysis accounts for real sonic-line effects and the streamwise variation in stream mixing and boundary layer within the ejector. The aspects of the analysis are explained and illustrated by applying the theory to a variety of ejector configurations including cylindrical shroud, contoured flap and plug nozzles. Extensive comparisons are made between theory and data to show the importance of various analytical assumptions and such design variables as diameter ratio, spacing ratio, total temperature ratio, and primary nozzle geometry.

Anderson, B. H.↗

NASA NTRS1989

Sonic-point capturing

A prototype scheme that produces perfectly smooth transonic solutions to nozzle-flow problems is derived and tested. The basic upwind scheme is described as well as satisfying the entropy condition, treatment of the source term, and numerical verification. The analysis yielded a numerical flux function for use near a sonic point, which is based on a full model of a transonic expansion wave, and a matched treatment for the source term.

Van Leer, Bram↗

NASA NTRS1972

Engine noise technology

The characteristics of aircraft engine noise are discussed. Data are provided to show the noise produced by the following aircraft components: (1) fan noise, (2) noise suppressing structures, (3) sonic inlets, (4) jet mixing noise due to nozzle flow, and (5) thrust reversers. Charts are developed to show the sound pressure level and the frequencies for each type of noise source. The use of laminates and composite materials to dissipate acoustic power is examined.

Luidens, R. W.↗

NASA NTRS1998

Mixing of Supersonic Streams

The Strutjet approach to Rocket Based Combined Cycle (RBCC) propulsion depends upon fuel-rich flows from the rocket nozzles and turbine exhaust products mixing with the ingested air for successful operation in the ramjet and scramjet modes. A model of the Strutjet device has been built and is undergoing test to investigate the mixing of the streams as a function of distance from the Strutjet exit plane. Initial cold flow testing of the model is underway to determine both, the behavior of the ingested air in the duct and to validate the mixing diagnostics. During the tests, each of the two rocket nozzles ejected up to two pounds mass per second into the 13.6 square inch duct. The tests showed that the mass flow of the rockets was great enough to cause the entrained air to go sonic at the strut, which is the location of the rocket nozzles. More tests are necessary to determine whether the entrained air chokes due to the reduction in the area of the duct at the strut (a physical choke), or because of the addition of mass inside the duct at the nozzle exit (a Fabri choke). The initial tests of the mixing diagnostics are showing promise.

Hawk, Clark W.↗

NASA NTRS1990

Improved Insert For Variable Mach Number

Nozzle insert lowers wind-tunnel mach number while maintaining excellent flow quality. Essential components of improved design include sonic first throat, expansion surface, variable boundary-layer bleed, insert itself, and existing, unchanged, contour of nozzle. Modification involves creation of secondary throat and critical addition of boundary-layer bleed path between insert and original tunnel wall. Enables quick change of mach number in existing facility at relatively low cost and requires no special contouring of existing nozzle. Represents simple and cost-effective method of altering mach number in any supersonic wind tunnel.

Puster, Richard L.↗

NASA NTRS2006

Design Enhancements of the Two-Dimensional, Dual Throat Fluidic Thrust Vectoring Nozzle Concept

A Dual Throat Nozzle fluidic thrust vectoring technique that achieves higher thrust-vectoring efficiencies than other fluidic techniques, without sacrificing thrust efficiency has been developed at NASA Langley Research Center. The nozzle concept was designed with the aid of the structured-grid, Reynolds-averaged Navier-Stokes computational fluidic dynamics code PAB3D. This new concept combines the thrust efficiency of sonic-plane skewing with increased thrust-vectoring efficiencies obtained by maximizing pressure differentials in a separated cavity located downstream of the nozzle throat. By injecting secondary flow asymmetrically at the upstream minimum area, a new aerodynamic minimum area is formed downstream of the geometric minimum and the sonic line is skewed, thus vectoring the exhaust flow. The nozzle was tested in the NASA Langley Research Center Jet Exit Test Facility. Internal nozzle performance characteristics were defined for nozzle pressure ratios up to 10, with a range of secondary injection flow rates up to 10 percent of the primary flow rate. Most of the data included in this paper shows the effect of secondary injection rate at a nozzle pressure ratio of 4. The effects of modifying cavity divergence angle, convergence angle and cavity shape on internal nozzle performance were investigated, as were effects of injection geometry, hole or slot. In agreement with computationally predicted data, experimental data verified that decreasing cavity divergence angle had a negative impact and increasing cavity convergence angle had a positive impact on thrust vector angle and thrust efficiency. A curved cavity apex provided improved thrust ratios at some injection rates. However, overall nozzle performance suffered with no secondary injection. Injection holes were more efficient than the injection slot over the range of injection rates, but the slot generated larger thrust vector angles for injection rates less than 4 percent of the primary flow rate.

Flamm, Jeffrey D.↗

NASA NTRS2016

Computational and Experimental Study of Supersonic Nozzle Flow and Aft-Deck Interactions

NASA has been conducting research into reducing sonic boom and changing FAA regulations to allow for supersonic commercial transport over land in the United States. This particular study looks at a plume passing through a shock generated from an aft deck on a nacelle; the aft deck is meant to represent the trailing edge of a wing. NASA Langley Research Center USM3D CFD code results are compared to the experimental data taken at the NASA Glenn Research Center 1-foot by 1-foot Supersonic Wind Tunnel. This study included examining two turbulence models along with different volume sourcing methods for grid generation. The results show that using the k-epsilon turbulence model within USM3D produced shock signatures that closely follow the experimental data at a variety of nozzle pressure ratio settings.

Bruce, Walter E., IV↗

NASA NTRS1989

Performance of an aero-space plane propulsion nozzle

An inviscid and viscous analysis is provided for an exposed half nozzle that is used with a scramjet for thrust generation. The analysis is based on the inviscid theory of a two-dimensional, minimum length nozzle with a curved inlet surface, where the flow may be sonic or supersonic. Inlet conditions are prescribed and the gas is assumed to be perfect. Viscous, and when appropriate inviscid, nondimensional parametric results are provided for the thrust, lift, heat transfer, pitching moment, and a variety of boundary-layer thicknesses. In addition to global results, wall distributions of pressure, heat transfer, etc., are provided. The analysis demonstrates that the nozzle produces a considerable lift force whose magnitude may exceed the thrust and a significant pitching moment. The thrust is quite sensitive to the inlet Mach number; it rapidly decreases as the inlet Mach number increases There is little loss in the thrust as the nozzle's downstream wall is truncated. The corresponding decrease in lift and the pitching moment is moderate.

Emanuel, George↗