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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 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

The effects of finite rate chemical processes on high enthalpy nozzle performance - A comparison between SPARK and SEAGULL

The generalized chemistry version of the computer code SPARK is extended to include two higher-order numerical schemes, yielding fourth-order spatial accuracy for the inviscid terms. The new and old formulations are used to study the influences of finite rate chemical processes on nozzle performance. A determination is made of the computationally optimum reaction scheme for use in high-enthalpy nozzles. Finite rate calculations are compared with the frozen and equilibrium limits to assess the validity of each formulation. In addition, the finite rate SPARK results are compared with the constant ratio of specific heats (gamma) SEAGULL code, to determine its accuracy in variable gamma flow situations. Finally, the higher-order SPARK code is used to calculate nozzle flows having species stratification. Flame quenching occurs at low nozzle pressures, while for high pressures, significant burning continues in the nozzle.

Carpenter, M. H.↗
NASA NTRS1989

A combined Eulerian-Lagrangian two-phase flow analysis of SSME HPOTP nozzle plug trajectories. II - Results

An analysis of Space Shuttle Main Engine high-pressure oxygen turbopump nozzle plug trajectories has been performed, using a Lagrangian method to track nozzle plug particles expelled from a turbine through a high Reynolds number flow in a turnaround duct with turning vanes. Axisymmetric and parametric analyses reveal that if nozzle plugs exited the turbine they would probably impact the LOX heat exchanger with impact velocities which are significantly less than the penetration velocity. The finding that only slight to moderate damage will result from nozzle plug failure in flight is supported by the results of a hot-fire engine test with induced nozzle plug failures.

Mcconnaughey, P. K.↗
NASA NTRS1989

RSRM nozzle fixed housing cooldown test

Flight 5 aft segments with nozzles were exposed to -17 F temperatures while awaiting shipment to KSC in February, 1989. No records were found which show that any previous nozzles were exposed to air temperatures as low as those seen by the Flight 5 nozzles. Thermal analysis shows that the temperature of the fixed housing, and forward and aft exit cone components dropped as low as -10 F. Structural analysis of the nozzles at these low temperatures show the forward and aft exit cone adhesive bonds to have a positive margin of safety, based on a 2.0 safety factor. These analyses show the normal and shear stresses in the fixed housing bond as low values. However, the hoop and meridinal stresses were predicted to be in the 4000 psi range; the failure stress allowable of EA913NA adhesive at -7 F. If the bonds did break in directions perpendicular to the surfaces, called bond crazing, no normal bond strength would be lost. Testing was conducted in two phases, showing that no degradation to the adhesive bonds occurred while the Flight 5 nozzles were subjected to subzero temperatures. The results of these tests are documented. Phase 1 testing cooled a full-scale RSRM insulated fixed housing to -13 F, with extensive bondline inspections. Phase 2 testing cooled the witness panel adhesive tensile buttions to -13 F, with failure strengths recorded before, during, and after the cooldown.

Bolieau, D. J.↗
NASA NTRS1990

An experimental investigation of thrust vectoring two-dimensional convergent-divergent nozzles installed in a twin-engine fighter model at high angles of attack

An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine thrust vectoring capability of subscale 2-D convergent-divergent exhaust nozzles installed on a twin engine general research fighter model. Pitch thrust vectoring was accomplished by downward rotation of nozzle upper and lower flaps. The effects of nozzle sidewall cutback were studied for both unvectored and pitch vectored nozzles. A single cutback sidewall was employed for yaw thrust vectoring. This investigation was conducted at Mach numbers ranging from 0 to 1.20 and at angles of attack from -2 to 35 deg. High pressure air was used to simulate jet exhaust and provide values of nozzle pressure ratio up to 9.

Capone, Francis J.↗
NASA NTRS1990

Static investigation of a two-dimensional convergent-divergent exhaust nozzle with multiaxis thrust-vectoring capability

An investigation was conducted in the Static Test Facility of the NASA Langley 16-Foot Transonic Tunnel to determine the internal performance of two-dimensional convergent-divergent nozzles designed to have simultaneous pitch and yaw thrust vectoring capability. This concept utilized divergent flap rotation of thrust vectoring in the pitch plane and deflection of flat yaw flaps hinged at the end of the sidewalls for yaw thrust vectoring. The hinge location of the yaw flaps was varied at four positions from the nozzle exit plane to the throat plane. The yaw flaps were designed to contain the flow laterally independent of power setting. In order to eliminate any physical interference between the yaw flap deflected into the exhaust stream and the divergent flaps, the downstream corners of both upper and lower divergent flaps were cut off to allow for up to 30 deg of yaw flap deflection. The impact of varying the nozzle pitch vector angle, throat area, yaw flap hinge location, yaw flap length, and yaw flap deflection angle on nozzle internal performance characteristics, was studied. High-pressure air was used to simulate jet exhaust at nozzle pressure ratios up to 7.0. Static results indicate that configurations with the yaw flap hinge located upstream of the exit plane provide relatively high levels of thrust vectoring efficiency without causing large losses in resultant thrust ratio. Therefore, these configurations represent a viable concept for providing simultaneous pitch and yaw thrust vectoring.

Taylor, John G.↗
NASA NTRS1990

Experimental and analytical study of close-coupled ventral nozzles for ASTOVL aircraft

Flow in a generic ventral nozzle system was studied experimentally and analytically with a block version of the PARC3D computational fluid dynamics program (a full Navier-Stokes equation solver) in order to evaluate the program's ability to predict system performance and internal flow patterns. For the experimental work a one-third-size model tailpipe with a single large rectangular ventral nozzle mounted normal to the tailpipe axis was tested with unheated air at steady-state pressure ratios up to 4.0. The end of the tailpipe was closed to simulate a blocked exhaust nozzle. Measurements showed about 5 1/2 percent flow-turning loss, reasonable nozzle performance coefficients, and a significant aftward axial component of thrust due to flow turning loss, reasonable nozzle performance coefficients, and a significant aftward axial component of thrust due to flow turning more than 90 deg. Flow behavior into and through the ventral duct is discussed and illustrated with paint streak flow visualization photographs. For the analytical work the same ventral system configuration was modeled with two computational grids to evaluate the effect of grid density. Both grids gave good results. The finer-grid solution produced more detailed flow patterns and predicted performance parameters, such as thrust and discharge coefficient, within 1 percent of the measured values. PARC3D flow visualization images are shown for comparison with the paint streak photographs. Modeling and computational issues encountered in the analytical work are discussed.

Mcardle, Jack G.↗
NASA NTRS1990

Nature of convection-stabilized dc arcs in dual-flow nozzle geometry. I - The cold flow field and dc arc characteristics. II - Optical diagnostics and theory

Steady-state low-current air arcs in a dual-flow nozzle system are studied experimentally. The cold flow field with no arc is investigated using a 12.7-mm diameter dual-flow nozzle in a steady-flow facility. Mach number and mass flux distributions are determined for various nozzle-pressure ratios and nozzle-gap spacing. It is found that the shock waves in the converging-diverging nozzles result in a decrease in overal resistance by about 15 percent. Also, Schlieren and differential interferometry techniques are used to visualize the density gradients within the arc plasma and thermal mantle. Both optical techniques reveal a laminar arc structure for a reservoir pressure of 1 atm at various current levels. Experimentally determined axial static pressure and cold-flow mass flux rate distributions and a channel-flow model with constant arc temperatre are used to solve the energy integral for the arc radius as a function of axial distance. The arc electric field strength, voltage, resistance, and power are determined with Ohm's law and the total heat transfer is related to arc power.

Serbetci, Ilter↗
NASA NTRS1990

External nozzle flap dynamic load measurements on F-15 S/MTD model

Dynamic pressure loads were obtained on 1/12 scale models of the F-15B production aircraft and the F-15 S/MTD experimental aircraft with rectangular nozzles and canards. Flight Mach numbers from 0.51 to 1.20 were studied for aircraft angles of attack from 0 to 10 deg and nozzle pressure ratios from 1.00 to 5.09. The results show that dynamic levels are lower in the internozzle region of twin rectangular nozzles than are levels found with twin axisymmetric nozzles. At other locations, the levels associated with both geometries are of the same order of magnitude when normalized by aircraft dynamic Q. At Mach number of 0.51, the loads spectrum is dominated by plume shock noise processes for both geometries. Above Mach 0.51, this mechanism is associated with either vortex bursting from a forward location or turbulent boundary layer separation over the nozzle external flaps. At supersonic speeds both geometries show significantly decreased load levels.

Seiner, John M.↗
NASA NTRS1991

The 3D computation of single-expansion-ramp and scramjet nozzles

A description of the computations for three-dimensional nonaxisymmetric nozzles and an analysis of the flowfields are presented. Two different types of nozzles are investigated for compressible flows at high Reynolds numbers. These are the single-expansion-ramp and scramjet nozzles. The computation for the single-expansion-ramp nozzle focuses on the condition of low pressure ratio, which requires the simulation for turbulent flow that is not needed at high pressure ratios. The flowfield contains the external quiescent air, and the internal regions of subsonic and low supersonic flows. The second type is the scramjet nozzle, which typically has a very large area ratio and is designed to operate at high speeds and pressure ratios. The freestream external flow has a Mach number of 6, and the internal flow leaving the combustion chamber is at a Mach number of 1.62. The flowfield is mostly supersonic except in the viscous region near walls. The computed results from both cases are compared with experimental data for the surface pressure distributions.

Lai, H. T.↗
NASA NTRS1990

Comparison of UNL laser imaging and sizing system and a phase Doppler system for analyzing sprays from a NASA nozzle

Research was conducted on characteristics of aerosol sprays using a P/DPA and a laser imaging/video processing system on a NASA MOD-1 air assist nozzle being evaluated for use in aircraft icing research. Benchmark tests were performed on monodispersed particles and on the NASA MOD-1 nozzle under identical lab operating conditions. The laser imaging/video processing system and the P/DPA showed agreement on a calibration tests in monodispersed aerosol sprays of + or - 2.6 micron with a standard deviation of + or - 2.6 micron. Benchmark tests were performed on the NASA MOD-1 nozzle on the centerline and radially at 0.5 inch increments to the outer edge of the spray plume at a distance 2 ft downstream from the exit nozzle. Comparative results at two operation conditions of the nozzle are presented for the two instruments. For the 1st case studied, the deviation in arithmetic mean diameters determined by the two instruments was in a range of 0.1 to 2.8 micron, and the deviation in Sauter mean diameters varied from 0 to 2.2 micron. Severe operating conditions in the 2nd case resulted in the arithmetic mean diameter deviating from 1.4 to 7.1 micron and the deviation in the Sauter mean diameters ranging from 0.4 to 6.7 micron.

Alexander, Dennis R.↗
NASA NTRS1991

Progress toward synergistic hypermixing nozzles

Mean flow measurements were obtained for air-to-air mixing downstream of swept and unswept ramp wall mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations studied are identical. The nozzles inject three parallel supersonic jets at a 15 deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at underexpanded static pressure condition (pressure ratio = 5). Surface oil flow visualization was used to study the near wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.

Davis, D. O.↗
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 NTRS1991

Goertler instability and hypersonic quiet nozzle design

A concept for nozzle design which incorporates slow expansion rates and a radial flow region is developed for use on a hypersonic (Mach 6) nozzle. It is shown that the boundary-layer suction slot upstream of the nozzle throat is necessary to remove upstream turbulent boundary layers and to initialize a new laminar boundary layer on the downstream nozzle wall. In effect, the laminar boundary-layer flow can be extended more effectively farther downstream by a slow expansion contour based on the present design concept than by a rapid expansion contour used in previous pilot quiet nozzles.

Chen, Fang-Jenq↗
NASA NTRS1991

Progress toward synergistic hypermixing nozzles

Mean flow measurements have been obtained for air-to-air mixing downstream of swept and unswept ramp wall-mounted hypermixing nozzle configurations. Aside from the sweep of the ramps, the two nozzle configurations investigated are identical. The nozzles inject three parallel supersonic jets at a 15-deg angle (relative to the wind tunnel wall) into a supersonic freestream. Mach number and volume fraction distributions in a transverse plane 11.1 nozzle heights downstream from the nozzle exit plane were measured. Data are presented for a freestream Mach number of three at a matched static pressure condition and also at an underexpanded static pressure condition (pressure ratio equal to 5). Surface oil flow visualization was used to investigate the near-wall flow behavior. The results indicate that the swept ramp injectors produce stronger and larger vortex pairs than the unswept ramp injectors. The increased interaction between the swept ramp model's larger vortex pairs yields better mixing characteristics for this model.

Davis, D. O.↗
NASA NTRS1991

The spatial characteristics of the B helical mode for a choked axisymmetric nozzle

An acoustic near-field study was performed for an axisymmetric conical nozzle operated at a fully expanded Mach number of 1.3. The acoustic measurements were performed in the nozzle exit plane using multiple sensors arranged around the periphery of the nozzle. The acquired data were simultaneously digitized. The acoustic spatial characteristics, describing the large scale structure associated with the preferred shear layer instability mode, were determined for the dominant B screech mode. The nozzle was fitted with a lip thickening device to determine the effect of this geometric variable on the spatial structure of the jet. For the thin-lipped configuration, the flapping structure of the B screech mode was found to precess in a time-dependent manner about the jet axis. An increasing of the nozzle exit lip thickness altered the spatial characteristics of this mode from a flapping to a time-dependent flapping or spinning. Using the auto-bicoherence spectrum, frequency dependencies were found to exist in the acoustic data of the thick-lipped configuration. These dependencies may be related to the nonlinearity of the jet and/or the time-dependent nature of the B mode spatial structure.

Ponton, Michael K.↗
NASA NTRS1992

Static performance of a cruciform nozzle with multiaxis thrust-vectoring and reverse-thrust capabilities

A multiaxis thrust vectoring nozzle designed to have equal flow turning capability in pitch and yaw was conceived and experimentally tested for internal, static performance. The cruciform-shaped convergent-divergent nozzle turned the flow for thrust vectoring by deflecting the divergent surfaces of the nozzle, called flaps. Methods for eliminating physical interference between pitch and yaw flaps at the larger multiaxis deflection angles was studied. These methods included restricting the pitch flaps from the path of the yaw flaps and shifting the flow path at the throat off the nozzle centerline to permit larger pitch-flap deflections without interfering with the operation of the yaw flaps. Two flap widths were tested at both dry and afterburning settings. Vertical and reverse thrust configurations at dry power were also tested. Comparison with two dimensional convergent-divergent nozzles showed lower but still competitive thrust performance and thrust vectoring capability.

Wing, David J.↗
NASA NTRS1992

CFD analysis of the STME nozzle flowfield

The Space Transportation Main Engine (STME) uses a gas generator cycle to cool the nozzle wall by a film-dump of the turbine exhaust. The ability to cool the skirt is a key concern in the design of the STME. CFD calculations were undertaken to predict the film cooling effectiveness and performance sensitivities for various configurations, operating points, and inlet conditions. The results presented here were obtained for the subscale nozzle. The computations were performed using REFLEQS. The computational analysis showed that a chemical equilibrium model was necessary to obtain correct predictions of the specific impulse. The frozen composition model underpredicts the ISP by about 6 percent. It was also observed that the coolant film was successful in maintaining the nozzle wall well below the stagnation temperature of the core flow. The effect of the coolant flow on the performance of the engine was found to be negligible. The computed heat fluxes at the wall were in good agreement with the empirical data obtained by Pratt and Whitney. Further test data from Pratt and Whitney are forthcoming for the subscale nozzle. Calculations will be performed to determine cooling efficiencies and nozzle performance over a range of conditions, and model predictions will be compared with experimental data. Information is given in viewgraph form.

Krishnan, Anantha↗
NASA NTRS1992

Forebody vortex flow control on the F-16C using tangential slot and jet nozzle blowing

The present experimental investigation of the effects of tangential slot and nozzle forebody blowing on the aerodynamic and stability characteristics of the F-16C configuration, for up-to-55 deg alpha and up-to-20 deg sideslip, was conducted on a 1/15th-scale model at freestream Mach = 0.4 and Re of 2.5 million/ft, for one tangential slot-nozzle location and two chordwise jet nozzle locations. Both the slot and the jet nozzles generated large yawing moments at very low blowing rates, over a wide range of alpha and sideslip angles. The centers of pressure of the side forces which generated the yawing moments due to slot and nozzle blowing were located in the region of the leading-edge extensions (LEXs); the moments may accordingly be due to the manipulation of the combined forebody and LEX flowfields.

Lemay, Scott P.↗