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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Aeroacoustic data for high Reynolds number supersonic axisymmetric jets

Both aerodynamic and near field acoustic behavior of several unheated axisymmetric shock free and shock containing high speed jet plumes are reported. The exit Mach number range for these data is from 0.9 to 2.5. The aerodynamic measurements include both mean and turbulence quantities for a shock free jet plume produced by a convergent divergent nozzle designed to have an exit Mach number of 2. The near field acoustic measurements presented include narrow band spectra, directivity and contour plots of select one third octave band data, and near field microphone correlations from a linear array. Shock noise results are also included as obtained by operating an underexpanded convergent nozzle at the design point of two supersonic exist Mach number convergent divergent nozzles.

Seiner, J. M.↗
NASA NTRS1985

Design optimization for a space based, reusable orbit transfer vehicle

Future NASA and DOD missions will benefit from high performance, reusable orbit transfer vehicles. With the advent of a space station, advanced engine technology, and various new vehicle concepts, reusable orbit transfer vehicles that provide significant economic benefits and mission capability improvements will be realized. Engine and vehicle design criteria previously have lacked definition with regard to issues such as space basing and servicing, man-rating and reliability, performance, mission flexibility, and life cycle cost for a reusable vehicle. The design study described here has resulted in the definition of a reusable orbit transfer vehicle concept and subsequent recommendations for the design criteria of an advanced LO2/LH2 engine. These design criteria include number of engines per vehicle, nozzle design, etc. The major characteristics of the vehicle preliminary design include low lift to drag aerocapture capability, a main propulsion system failure criteria of fail operational/fail safe, and either two main engines with a high performance attitude control system for back-up or three main engines with which to meet this failure criteria. In addition, a maintenance approach has been established for the advanced vehicle concept.

Redd, L.↗
NASA NTRS1949

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Special Nozzles: Efficiency with 20 Degrees-Inlet-Angle Rotor Blades - II

A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzle designs in combination with special rotor blades having a 20 inlet angle. The performance is presented in terms of blade, rotor, and brake efficiency as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. The blade efficiency with the nozzle having circular pas- sages (K) was equal to or higher than that with the nozzle having rectangular passages (J) for all pressure ratios and speeds investigated. The maximum blade efficiency of 0.571 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.296. The difference in blade efficiency was negligible at a pressure ratio of 8 at the low speeds; the maxim difference was 0.040 at a pressure ratio of 20 and a blade-jet speed ratio of 0.260.

Schum, Harold J.↗
NASA NTRS1991

Comparison of a quasi-3D analysis and experimental performance for three compact radial turbines

An experimental aerodynamic evaluation of three compact radial turbine builds was performed. Two rotors which were 40 to 50 percent shorter in axial length than conventional state of the art radial rotors were tested. A single nozzle design was used. One rotor was tested with the nozzle at two stagger angle settings. A second rotor was tested with the nozzle in only the closed down setting. Experimental results were compared to predict results from a quasi-3D inviscid and boundary layer analysis, called Meridl/Tsonic/Blayer (MTSB). This analysis was used to predict turbine performance. It has previously been calibrated only for axial, not radial, turbomachinery. The predicted and measured efficiencies were compared at the design point for the three turbines. At the design points the analysis overpredicted the efficiency by less than 1.7 points. Comparisons were also made at off-design operating points. The results of these comparisons showed the importance of an accurate clearance model for efficiency predictions and also that there are deficiencies in the incidence loss model used.

Simonyi, P. S.↗
NASA NTRS1991

Comparison of a quasi-3D analysis and experimental performance for three compact radial turbines

An experimental aerodynamic evaluation of three compact radial turbine builds was performed. Two rotors which were 40-50 percent shorter in axial length than conventional state-of-the-art radial rotors were tested. A single nozzle design was used. One rotor was tested with the nozzle at two stagger angle settings. A second rotor was tested with the nozzle in only the closed down setting. Experimental results were compared to predicted results from a quasi-3D inviscid and boundary layer analysis, called MTSB (Meridl/Tsonic/Blayer). This analysis was used to predict turbine performance. It has previously been calibrated only for axial, not radial, turbomachinery. The predicted and measured efficiencies were compared at the design point for the three turbines. At the design points the analysis overpredicted the efficiency by less than 1.7 points. Comparisons were also made at off-design operating points. The results of these comparisons showed the importance of an accurate clearance model for efficiency predictions and also that there are deficiencies in the incidence loss model used.

Simonyi, P. S.↗
NASA NTRS1954

Thrust characteristics of a series of convergent-divergent exhaust nozzles at subsonic and supersonic flight speeds

An experimental investigation of a series of four convergent-divergent exhaust nozzles was conducted in the Lewis 8-by-6 foot supersonic wind tunnel at Mach numbers of 0.1, 0.6, 1.6, and 2.0 over a range of nozzle pressure ratios. The thrust characteristics of these nozzles were determined by a pressure-integration technique. From a thrust standpoint, a nozzle designed to give uniform parallel flow at the exit had no advantage over the simple geometric design with conical convergent and divergent sections. The rapid-divergent nozzles might be competitive with the more gradual-divergent nozzles since the relatively short length of these nozzles would be advantageous from a weight standpoint and might result in smaller thrust losses due to friction. The thrusts, with friction losses neglected, were predicted satisfactorily by one-dimensional theory for the nozzles with relatively gradual divergence. The thrusts of the rapid-divergent designs were several percentages below the theoretical values at the design pressure ratio or above, while at low pressure ratios there was a considerable effect of free-stream Mach number, with thrusts considerably above theoretical values at subsonic speeds and somewhat above theoretical values at supersonic speeds. This Mach numb effect appeared to be related to the variation of the model base pressure with free-stream Mach number.

Fradenburgh, Evan A↗
NASA NTRS2008

Design of a Mach-3 Nozzle for TBCC Testing in the NASA LaRC 8-ft High Temperature Tunnel

A new nozzle is being constructed for the NASA Langley Research Center 8-Foot High Temperature Tunnel. The axisymmetric nozzle was designed with a Mach-3 exit flow for testing Turbine-Based Combined-Cycle engines at a Mach number in the vicinity of the transition from turbojet to ramjet operation. The nozzle contour was designed using the NASA Langley IMOCND computer program which solves the potential equation using the classical method of characteristics. To include viscous effects, the design procedure iterated the MOC contour generation with CFD Navier-Stokes calculations, adjusting MOC input parameters until target nozzle-exit conditions were achieved in the Navier-Stokes calculations. The design process was complicated by a requirement to use the final 29.5 inches of an existing 54.5-inch exit-diameter Mach-5 nozzle contour. This was accomplished by generating a Mach-3 contour that matched the radius of the Mach-5 contour at the match point and using a 3rd order polynomial to create a smooth transition between the two contours. During the final evaluation of the design it was realized that the throat diameter is more than half that of the upstream mixing chamber. This led to the concern that large vortical structures generated in the mixer would persist downstream, affecting nozzle-exit flow. This concern was addressed by analyzing the results of three-dimensional, viscous, numerical simulations of the entire flowfield, from the exit of the facility combustor to the nozzle exit. An analysis of the solution indicated that large scale structures do not pass through the throat and that both the total temperature and species (CO2) are well mixed in the mixer, providing uniform flow to the nozzle and subsequently the test cabin.

Gaffney, Richard L., Jr.↗
NASA NTRS1979

Pulsed plasma thruster contamination studies

The exhaust plume of the one millipound pulsed plasma thruster has a measurable backflow upstream of the nozzle exit plane which may deposit on and degrade the performance of exposed spacecraft surfaces. High speed photographs and Faraday cup measurements suggest that this backflow is predominantly an electrically neutral, relatively low energy vapor. Articulated collimator quartz crystal microbalance measurements of this backflow were made for a thruster with a radically modified nozzle and a flat plate backflow shield, to determine the backflow sensitivity to nozzle design changes. The results are compared with the original nozzle backflow and show a measurable reduction in the backflow directly upstream of the shield.

Rudolph, L. K.↗
NASA NTRS1932

Coefficients of discharge of fuel-injection nozzles for compression-ignition engines

This report presents the results of an investigation to determine the coefficients of discharge of nozzles with small, round orifices of the size used with high-speed compression-ignition engines. The injection pressures and chamber back pressures employed were comparable to those existing in compression-ignition engines during injection. The construction of the nozzles was varied to determine the effect of the nozzle design on the coefficient. Tests were also made with nozzles assembled in an automatic injection valve, both with a plain and with a helically grooved stem. It was found that a smooth passage before the orifice is requisite for high flow efficiency. A beveled leading edge before the orifice gave a higher coefficient of discharge than a rounded edge. The results with the nozzles assembled in an automatic injection valve having a plain stem duplicated those with the nozzles assembled at the end of a straight tube of constant diameter. Lower coefficients were obtained with the nozzles assembled in an injection valve having a helically grooved stem. When the coefficients of nozzles of any one geometrical shape were plotted against values of corresponding Reynold's numbers for the orifice diameters and rates of flow tested, it was found that experimental points were distributed along a single curve.

Gelalles, A G↗
NASA NTRS2003

Rayleigh Scattering Diagnostic Used to Measure Velocity and Density Fluctuation Spectra

A new, molecular Rayleigh-scattering-based flow diagnostic developed at the NASA Glenn Research Center has been used for the first time to measure the power spectrum of both gas density and radial velocity components in the plumes of high-speed jets. The objective of the work is to develop an unseeded, nonintrusive dynamic measurement technique for studying turbulent flows in NASA test facilities. This technique provides aerothermodynamic data not previously obtainable. It is particularly important for supersonic flows, where hot wire and pitot probes are difficult to use and disturb the flow under study. The effort is part of the nonintrusive instrumentation development program supporting propulsion research at the NASA Glenn Research Center. In particular, this work is measuring fluctuations in flow velocity, density, and temperature for jet noise studies. These data are valuable to researchers studying the correlation of flow fluctuations with far-field noise. One of the main objectives in jet noise research is to identify noise sources in the jet and to determine their contribution to noise generation. The technique is based on analyzing light scattered from molecules within the jet using a Fabry-Perot interferometer operating in a static imaging mode. The PC-based data acquisition system can simultaneously sample velocity and density data at rates to about 100 kHz and can handle up to 10 million data records. We used this system to interrogate three different jet nozzle designs in a Glenn free-jet facility. Each nozzle had a 25.4-mm exit diameter. One was convergent, used for subsonic flow measurements and to produce a screeching underexpanded jet with a fully expanded Mach number of 1.42. The other nozzles (Mach 1.4 and 1.8) were convergent-divergent types. The radial component of velocity and gas density were simultaneously measured in this work.

Seasholtz, Richard G.↗
NASA NTRS1990

Advanced Mach 3.5 Axisymmetric Quiet Nozzle

To advance boundary-layer stability and transition research and to ultimately provide reliable predictions of transition for supersonic flight vehicles, a wind tunnel is required with very low stream disturbance levels comparable to free flight conditions. Experimental and theoretical research to develop a low-disturbance supersonic wind tunnel has achieved a breakthrough. A new concept for nozzle design is presented which promises a large increase in the length of the quiet test core. The Advanced Mach 3.5 Axisymmetric Quiet Nozzle is the first prototype built to prove the new design concept. Experimental results from this new nozzle on the extent of laminar wall boundary layers are compared with data from other nozzles and with theoretical predictions based on linear stability theory. The Reynolds numbers based on the measured length of the quiet test core for this new nozzle are in excellent agreement with the theoretical predictions. The effect of surface finish on the nozzle performance is also discussed.

Chen, Fang-Jenq↗
NASA NTRS1979

Wing aerodynamic loading caused by jet-induced lift associated with STOL-OTW configurations

Surface pressure distributions were obtained with model-scale STOL-OTW configurations using various nozzles designed to promote flow attachment to the wing/flap surface. The nozzle configurations included slot-types and both circular and slot nozzles with external flow deflectors. The wing aerodynamic loading caused by the jet-induced lift is presented in conventional terms of delta p/q as a function of chordwise surface distance in the nozzle centerline plane as well as outboard of the nozzle centerline. Nozzle roof/deflector angle, chordwise location of the nozzle, wing size, and flap deflection angle are included in the geometric variables affecting the wing loading.

Vonglahn, U.↗
NASA NTRS1981

Experimental and analytical investigation of axisymmetric supersonic cruise nozzle geometry at Mach numbers from 0.60 to 1.30

Quantitative pressure and force data for five axisymmetric boattail nozzle configurations were examined. These configurations simulate the variable-geometry feature of a single nozzle design operating over a range of engine operating conditions. Five nozzles were tested in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.60 to 1.30. The experimental data were also compared with theoretical predictions.

Carson, G. T., Jr.↗
NASA NTRS1985

Investigation of arcjet nozzle performance

An investigation is performed to examine the feasibility of improving arcjet performance through the use of contoured nozzle designs. The results of preliminary experiments performed on two different nozzle configurations, a 19 degree half-angle cone and a 'bell' shaped nozzle, are described. A unique experimental arcjet apparatus, which effectively tests only the change in nozzle contour on arcjet performance is used in this investigation. The preliminary experimental results indicate approximately an 8 percent improvement in thrust and specific impulse for operation at 15 kW with an ammonia flow rate of 0.20 g/s for a 'bell' shaped nozzle compared to a conical nozzle.

Brophy, J. R.↗
NASA NTRS2020

Experimental Validation of Nozzle Flow Simulations for Rotating Detonation Rocket Engines

Rotating detonation engines (RDEs) promise increased thermodynamic performance that may significantly enhance the capabilities of current rocket platforms. Little work has been conducted thus far to characterize the effect of nozzle design on high chamber pressure RDEs. Previous computational work was completed to understand the nozzle performance of the Purdue methane/oxygen rocket RDE. A new experimental study on a similar kerosene/oxygen RDE was conducted to validate the results of the computational study. Both a nozzleless geometry and several aerospike designs were hot-fire tested. New pressure instrumentation on the different nozzle surfaces were included to better understand the flow physics unique to RDE chamber exit conditions. Both the previous computational study and the new experimental results confirmed that for a nozzleless geometry, the RDE cycle enhances suction on the base region, an important result to determining RDE engine performance separate from nozzle effects. While the aerospike experiments showed agreement with computational results, delay of flow separation due to the RDE cycle could not be confirmed. Two aerospike geometries with different nozzle pressure ratios were experimentally evaluated. This study showed both the strength and the necessity for 3D transient computations to better understand the RDE flow field with nozzle geometries.

Alexis J Harroun↗
NASA NTRS1981

The design and development of a mixer compound exhaust system for a quiet clean general aviation turbofan

Lower engine fuel consumption has become a dominant factor in turbofan engine design due to rapidly increasing fuel costs. One engine component with a large impact on engine performance is the exhaust system. Previous exhaust system studies have demonstrated the significant exhaust system efficiency gains available through mixing of the core and bypass flows. Typically, a large, costly rig and engine program are required to develop and optimize these gains. The purpose of this paper is to present the results of the low-cost design system used for the quiet, clean general aviation turbofan mixer nozzle design and development. The scale model and full-scale engine test results confirm the predicted 3 to 5% reduction in cruise fuel consumption. This unique design system, which is based on integrating advanced three-dimensional viscous numerical methods with empirical optimization techniques, is summarized and detailed comparisons with test data are presented. The ability to accurately predict relative performance of mixer systems with substantially reduced development time and cost savings is demonstrated.

Blackmore, W. L.↗