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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 NTRS2001

The Tribological Properties of Several Silahydrocarbons for Use in Space Mechanisms

Silahydrocarbons are members of a relatively new class of liquid lubricants with great potential for use in space mechanisms. They are unimolecular species consisting of silicon, carbon, and hydrogen. They possess unique wear, viscosity, and volatility properties while retaining the ability to solubilize conventional additives. The tribological properties of several members of this class, including tri, tetra- and penta-compounds, are presented. These properties include: viscosity-temperature (ASTM D446), viscosity-pressure coefficient, vapor pressure, volatility, lubricant lifetimes, traction, reciprocating and four ball wear rates and bearing performance. Lubricant lifetimes were determined using a vacuum ball bearing simulator, the spiral orbit tribometer (SOT). Wear was measured using a Cameron Plint reciprocating tribometer and wear rates with a vacuum four ball tribometer. Conventional viscometry was used for viscosity-temperature measurements and a Knudsen cell for vapor pressure. Vacuum Thermogravimetric Analysis (TGA) was also used for volatility measurements. Pressure viscosity coefficients (a values) were estimated from EHL (elastohydrodynamic lubrication) film thickness measurements. Traction coefficients were measured with a twin disk traction rig. Bearing tests were performed in a vacuum bearing test facility. These properties are compared to existing state-of-the-art space lubricants.

Jones, W. R., Jr.↗
NASA NTRS1959

Flow Characteristics About Two Thin Wings of Low Aspect Ratio Determined from Surface Pressure Measurements Obtained in Flight at Mach Numbers from 0.73 to 1.90

Surface pressure measurements were obtained at three chordwise stations on the wings of the X-3 and X-lE airplanes at Mach numbers from 0.73 to 1.13 for the X-3, and from 0.82 to 1.90 for the X-IE. Leading-edge separation is present on the X-3 wing at a Mach number of about 0.73 and an angle of attack of about 6 deg. However., when the Mach number is increased to 0.88, the trailing-edge separation dominates the pressure distribution and no leading-edge separation is visible although it is anticipated at the higher angles of attack shown. Conversely, the X-lE wing shows no indication of leading-edge separation within the scope of this investigation, but an overexpansion immediately behind the leading edge is present at a Mach number of approximately 0.82. Two separate normal shocks are present on the X-3 wing at a Mach number of about 0.88 and at a low angle of attack as an effect of wing geometry. These shocks merge to form a single shock when the angle of attack is increased to about 6 deg. At supersonic speeds the upper-surface expansion on the X-lE wing is limited by the approach of the pressure coefficients to the pressure coefficient for a vacuum.

Taillon, Norman V.↗
NASA NTRS1986

Three-dimensional boundary layer on a compressor rotor blade at peak pressure rise coefficient

A comprehensive study of the three-dimensional turbulent boundary layer on a compressor rotor blade at peak pressure rise coefficient is reported in this paper. The measurements were carried out at various chordwise and radial locations on a compressor rotor blade using a rotating miniature 'V' configuration hot-wire probe. The data are compared with the measurement at the design condition. Substantial changes in the blade boundary layer characteristics are observed, especially in the outer sixteen percent of the blade span. The increased chordwise pressure gradient and the leakage flow at the peak pressure coefficient have a cumulative effect in increasing the boundary layer growth on the suction surface. The leakage flow has a beneficial effect on the pressure surface. The momentum and boundary layer thicknesses increase substantially from those at the design condition, especially near the outer radii of the suction surface.

Lakshminarayana, B.↗
NASA NTRS2006

Room-temperature Broadening and Pressure-shift Coefficients in the nu(exp 2) Band of CH3D-O2: Measurements and Semi-classical Calculations

We report measured Lorentz O2-broadening and O2-induced pressure-shift coefficients of CH3D in the nu(exp 2) fundamental band. Using a multispectrum fitting technique we have analyzed 11 laboratory absorption spectra recorded at 0.011 cm(exp 1) resolution using the McMath-Pierce Fourier transform spectrometer, Kitt Peak, Arizona. Two absorption cells with path lengths of 10.2 and 25 cm were used to record the spectra. The total sample pressures ranged from 0.98 to 339.85 Torr with CH3D volume mixing ratios of 0.012 in oxygen. We report measurements for O2 pressure-broadening coefficients of 320 nu(exp 2) transitions with quantum numbers as high as J0(sup w) = 17 and K = 14, where K(sup w) = K' is equivalent to K (for a parallel band). The measured O2-broadening coefficients range from 0.0153 to 0.0645 cm(exp -1) atm(exp -1) at 296 K. All the measured pressure-shifts are negative. The reported O2-induced pressure-shift coefficients vary from about -0.0017 to -0.0068 cm(exp -1) atm(exp -1). We have examined the dependence of the measured broadening and shift parameters on the J(sup W), and K quantum numbers and also developed empirical expressions to describe the broadening coefficients in terms of m (m = -J(sup W), J(sup W), and J(sup w) + 1 in the QP-, QQ-, and QR-branch, respectively) and K. On average, the empirical expressions reproduce the measured broadening coefficients to within 4.4%. The O2-broadening and pressure shift coefficients were calculated on the basis of a semiclassical model of interacting linear molecules performed by considering in addition to the electrostatic contributions the atom-atom Lennard-Jones potential. The theoretical results of the broadening coefficients are generally larger than the experimental data. Using for the trajectory model an isotropic Lennard-Jones potential derived from molecular parameters instead of the spherical average of the atom-atom model, a better agreement is obtained with these data, especially for |m| <= 12 values (11.3% for the first calculation and 8.1% for the second calculation). The O2-pressure shifts whose vibrational contribution are either derived from parameters fitted in the QQ-branch of selfinduced shifts of CH3D or those obtained from pressure shifts induced by Xe in the nu(sup 3) band of CH3D are in reasonable agreement with the scattered experimental data (17.0% for the first calculation and 18.7% for the second calculation).

Predoi-Cross, Adriana↗
NASA NTRS1986

Measurement of viscosity of gaseous mixtures at atmospheric pressure

Coefficients of viscosity of various types of gas mixtures, including simulated natural-gas samples, have been measured at atmospheric pressure and room temperature using a modified capillary tube method. Pressure drops across the straight capillary tube section of a thermal mass flowmeter were measured for small, well-defined, volume flow rates for the test gases and for standard air. In this configuration, the flowmeter provides the volumetric flow rates as well as a well-characterized capillary section for differential pressure measurements across it. The coefficients of viscosity of the test gases were calculated using the reported value of 185.6 micro P for the viscosity of air. The coefficients of viscosity for the test mixtures were also calculated using Wilke's approximation of the Chapman-Enskog (C-E) theory. The experimental and calculated values for binary mixtures are in agreement within the reported accuracy of Wilke's approximation of the C-E theory. However, the agreement for multicomponent mixtures is less satisfactory, possible because of the limitations of Wilkes's approximation of the classical dilute-gas state model.

Singh, J. J.↗
NASA NTRS1947

Altitude-Wind-Tunnel Investigation of the 19B-2, 19B-8 and 19XB-1 Jet- Propulsion Engines: Analysis of Compressor Performance - 4

Investigations were conducted in the Cleveland altitude wind tunnel to determine the performance and operational characteristics of the 19B-2, 19B-8, and 19XS-1 turbojet engines. One objective was to determine the effect of altitude, flight Mach number, and tail-pipe-nozzle area on the performance characteristics of the six-stage and ten-stage axial-flow compressors of the 19B-8 and 19XB-1 engines, respectively, The data were obtained over a range of simulated altitudes and flight Mach numbers. At each simulated flight condition the engine was run over its full operable range of speeds. Performance characteristics of the 19B-8 and 19XB-1 compressors for the range of operation obtainable in the turboJet-engine installation are presented. Compressor characteristics are presented as functions of air flow corrected to sea-level conditions, compressor Mach number, and compressor load coefficient. For the range of compressor operation investigated, changes in Reynolds number had no measurable effect on the relations among compressor Mach number, corrected air flow, compressor load coefficient, compressor pressure ratio, and compressor efficiency. The operating lines for the 19B-8 compressor lay on the low-air-flow side of the region of maximum compressor efficiency; the 19B-8 compressor operated at higher average pressure coefficients per stage and produced a lower over-all pressure ratio than did the 19XB-1 compressor.

Dietz, Robert O.↗
NASA NTRS1956

Ram-jet Performance

The ram jet is basically one of the most dimple types of aircraft engine. It consists only of an inlet diffuser, a combustion system, and an exit nozzle. A typical ram-jet configuration is shown in figure 128. The engine operates on the Brayton cycle, and ideal cycle efficiency depends only on the ratio of engine to ambient pressure. The increased, engine pressures are obtained by ram action alone, and for this reason the ram jet has zero thrust at zero speed. Therefore, ram-jet-powered aircraft must be boosted to flight speeds close to a Mach number of 1.0 before appreciable thrust is generated by the engine. Since pressure increases are obtained by ram action alone, combustor-inlet pressures and temperatures are controlled by the flight speed, the ambient atmospheric condition, and by the efficiency of the inlet diffuser. These pressures and temperatures, as functions of flight speed and altitude, are shown in figure 129 for the NACA standard atmosphere and for practical values of diffuser efficiency. It can be seen that very wide ranges of combustor-inlet temperatures and pressures may be encountered over the ranges of flight velocity and altitude at which ram jets may be operated. Combustor-inlet temperatures from 500 degrees to 1500 degrees R and inlet pressures from 5 to 100 pounds per square inch absolute represent the approximate ranges of interest in current combustor development work. Since the ram jet has no moving parts in the combustor outlet, higher exhaust-gas temperatures than those used in current turbojets are permissible. Therefore, fuel-air ratios equivalent to maximum rates of air specific impulse or heat release can be used, and, for hydrocarbon fuels, this weight ratio is about 0.070. Lower fuel-air ratios down to about 0.015 may also be required to permit efficient cruise operation. This fuel-air-ratio range of 0.015 to 0.070 used in ram jets can be compared with the fuel-air ratios up to 0.025 encountered in current turbojets. Ram-jet combustor-inlet velocities range from 150 to 400 feet per second. These high linear velocities combined with the relatively low pressure ratios obtainable in ram jets require that the pressure drop through the combustor be kept low to avoid excessive losses in cycle efficiency. It has been estimated that, for a long-range ram-jet engine, an increase in pressure loss of one dynamic head would require a compensating 1-percent increase in combustion efficiency. Therefore, combustor pressure-loss coefficients (pressure drop/impact pressure) of the order of 1 to 4 are found in most current engines. The operating conditions described impose major problems in the design of stable and efficient ram-jet combustion systems. This chapter presents a survey of ram-jet combustor research and, where possible, points out criteria that may be useful in the design of ram-jet combustion systems.

Cervenko, A. J.↗
NASA NTRS2015

CFD Simulations of the Space Launch System Ascent Aerodynamics and Booster Separation

This paper presents details of Computational Fluid Dynamic modeling of the Space Launch System during ascent. The primary focus of the paper is the flow simulation of the vehicle during ascent using the Overflow Navier-Stokes code. Computations of 739 first-stage flight conditions covering a range of Mach numbers, angles of attack, and roll angles were computed. The overset grid system contained 375 million grid points, and over 28 mil- lion CPU hours were used in the simulations. The simulations were run on the Pleiades supercomputer at the NASA Advanced Supercomputer Center at Ames Research Center. The data products from this work include integrated line-loads, surface pressure coefficients, venting pressures, and protuberance air-loads. Detailed comparisons were made of the aerodynamic performance predicted by Overflow and the wind-tunnel derived aero- dynamic database. A small number of the cases were run with two different turbulence models and with two differencing schemes. These results were used to quantify the sensitivity to the choice of the turbulence model and to the differencing scheme. The paper also introduces an effort to use the inviscid, unstructured Cartesian solver Cart3D to compute the aerodynamics during booster separation. Adaptive mesh refinement is being used to enable accurate simulations of sixteen booster-separation-motor plumes. The use of this tool is explored in preparation for building a booster-separation aerodynamic database.

Ascent Aerodynamics↗
NASA NTRS1998

Pressure Broadening and Pressure Shift Coefficients in the 2nu(sup 0, sub 2) and nu(sub 1) Bands of (16)O(13)C(18)O

In this study we report the first measurements of the pressure broadening and the pressure-induced line shift coefficients due to air and nitrogen for 93 individual rovibrational lines, P(46) to R(46) and P(50) to R(45), respectively, belonging to the 2nu(sup 0, sub 2) and nu(sub 1) bands of (16)O(13)C(18)O. The results were obtained by analyzing ten long path, high-resolution laboratory, absorption spectra using a multispectrum nonlinear least-squares technique. Broadening due to nitrogen was about 4% larger than by air. Similar to broadening coefficients, the pressure-shift coefficients were found to be transition dependent, but different for the P- and R-branch lines with the same rotational quantum number, J. Except for a few R-branch lines, the measured shift coefficients were negative. The shift coefficients varied from + 4 x 10(exp -3)/cm.atm at 296 K to -5.6 x 10(exp -3)/cm.atm at 296 K. Comparisons of the broadening and shift coefficients between the two bands, between the P and R branches, and between the two broadening gases are reported. No significant difference between shift coefficients of the two bands or the two broadening gases was observed. The results obtained were compared with those reported in the literature for the more abundant (12)C(16)O2 species.

Devi, V. Malathy↗
DOE OSTI2022

Prediction of heat transfer coefficient from pressure drop using machine learning techniques in smooth horizontal pipes

Limited experimental work has been done focusing specifically on the relationship between pressure drop and heat transfer in the transient flow regime in pipes. The aim of this study was therefore to use machine learning to find a relationship between pressure drop and heat transfer coefficients of developing and fully developed flow in smooth horizontal circular tubes in the laminar, transient, quasi-turbulent and turbulent. The data are collected from literature, coming from experimentation. Pressure drops and heat transfer measurements were taken simultaneously and the relationship between pressure drop, and heat transfer was determined using multiple machine learning techniques.

20 FOSSIL-FUELED POWER PLANTS↗
NASA NTRS1975

The effect of Reynolds number on boattail drag

An investigation has been conducted in the Langley pilot transonic cryogenic tunnel to determine the effects of varying Reynolds number on boattail drag at subsonic speeds. Six boattailed cone-cylinder nacelle models were tested with the jet exhaust simulated by a cylindrical sting. Reynolds number was varied from about 2.6 million to 132 million by changing model length and unit Reynolds number. Boattail pressure coefficient distributions show that increasing Reynolds number tends to make the pressure coefficients in the expansion region more negative and the pressure coefficients in the recompression region more positive. These two effects were compensating and as a result there was little or no effect of Reynolds number on the pressure drag of the isolated boattails.

Reubush, D. E.↗
NASA NTRS1978

Flight-measured pressure characteristics of aft-facing steps in high Reynolds number flow at Mach numbers of 2.20, 2.50, and 2.80 and comparison with other data

The YF-12 airplane was studied to determine the pressure characteristics associated with an aft-facing step in high Reynolds number flow for nominal Mach numbers of 2.20, 2.50, and 2.80. Base pressure coefficients were obtained for three step heights. The surface static pressures ahead of and behind the step were measured for the no-step condition and for each of the step heights. A boundary layer rake was used to determine the local boundary layer conditions. The Reynolds number based on the length of flow ahead of the step was approximately 10 to the 8th power and the ratios of momentum thickness to step height ranged from 0.2 to 1.0. Base pressure coefficients were compared with other available data at similar Mach numbers and at ratios of momentum thickness to step height near 1.0. In addition, the data were compared with base pressure coefficients calculated by a semiempirical prediction method. The base pressure ratios are shown to be a function of Reynolds number based on momentum thickness. Profiles of the surface pressures ahead of and behind the step and the local boundary layer conditions are also presented.

Powers, S. G.↗
NASA NTRS1976

Experimental investigation to validate use of cryogenic temperatures to achieve high Reynolds numbers in boattail pressure testing

An investigation has been conducted in the Langley 1/3 meter transonic cryogenic tunnel to validate the use of cryogenic temperatures to achieve high Reynolds numbers in nozzle boattail pressure testing. Tests were conducted at 0 degree angle of attack and at Mach numbers of 0.60, 0.85, and 0.90 on two wing-body configurations with differing boattail geometries. Test data were obtained by using two different techniques, the cryogenic method and the conventional method, to obtain the same Reynolds number. Later, the test data obtained from the two techniques on boattail pressure coefficient distributions and pressure drag coefficients were compared; results from the comparisons show excellent repeatability for all test conditions and indicate no measurable errors when using cryogenic temperatures to achieve high Reynolds numbers for nozzle boattail pressure testing.

Reubush, D. E.↗
NASA NTRS1991

Turbulent boundary layer separation over a rearward facing ramp and its control through mechanical excitation

A vane oscillating about a fixed point at the inlet to a two-dimensional 20 degree rearward facing ramp has proven effective in delaying the separation of a turbulent boundary layer. Measurements of the ramp surface static pressure coefficient obtained under the condition of vane oscillation and constant inlet velocity revealed that two different effects occurred with surface distance along the ramp. In the vicinity of the oscillating vane, the pressure coefficients varied as a negative function of the vane's trailing edge rms velocity; the independent variable on which the rms velocity depends are the vane's oscillation frequency and its displacement amplitude. From a point downstream of the vane to the exit of the ramp; however, the pressure coefficient varied as a more complex function of the two independent variables. That is, it was found to vary as a function of the vane's oscillation frequency throughout the entire range of frequencies covered during the test, but over only a limited range of the trailing edge displacement amplitudes covered. More specifically, the value of the pressure coefficient was independent of increases in the vane's displacement amplitude above approximately 35 inner wall units of the boundary layer. Below this specific amplitude it varied as a function of the vane's trailing edge rms velocity. This height is close to the upper limit of the buffer layer. A parametric study was made to determine the variation of the maximum static pressure recovery as a function of the vane's oscillation frequency, for several ramp inlet velocities and a constant displacement amplitude of the vane's trailing edge. The results indicate that the phenomenon producing the optimum delay of separation may be Strouhal number dependent. Corona anemometer measurements obtained in the inner wall regions of the boundary layer for the excited case reveal a large range of unsteadiness in the local velocities. These measurements imply the existence of inflections in the profiles, which provide a mechanism for resulting inviscid flow instabilities to produce turbulence in the near wall region, thereby delaying separation of the boundary layer.

Mckinzie, Daniel J., Jr.↗