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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Acoustic modes in fluid networks

Pressure and flow rate eigenvalue problems for one-dimensional flow of a fluid in a network of pipes are derived from the familiar transmission line equations. These equations are linearized by assuming small velocity and pressure oscillations about mean flow conditions. It is shown that the flow rate eigenvalues are the same as the pressure eigenvalues and the relationship between line pressure modes and flow rate modes is established. A volume at the end of each branch is employed which allows any combination of boundary conditions, from open to closed, to be used. The Jacobi iterative method is used to compute undamped natural frequencies and associated pressure/flow modes. Several numerical examples are presented which include acoustic modes for the Helium Supply System of the Space Shuttle Orbiter Main Propulsion System. It should be noted that the method presented herein can be applied to any one-dimensional acoustic system involving an arbitrary number of branches.

Michalopoulos, C. D.↗
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 NTRS2006

Numerical Modeling of Flow Control in a Boundary-Layer-Ingesting Offset Inlet Diffuser at Transonic Mach Numbers

This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe rake indicated that the circumferential distortion levels are very sensitive to the symmetry of the flow and that a miss alignment of the vanes in the experiment could have resulted in this difference. The numerical simulations of the BLI inlet with jets showed good agreement with the circumferential inlet distortion levels for a range of jet actuator mass flow ratios at a fixed inlet mass flow rate. The CFD simulations for the jet case also predicted an average total pressure recovery that was 0.01 lower than the experiment as was seen in the baseline. Comparison of the flow features the jet case revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Allan Brian G.↗
NASA NTRS1990

Computer code for predicting coolant flow and heat transfer in turbomachinery

A computer code was developed to analyze any turbomachinery coolant flow path geometry that consist of a single flow passage with a unique inlet and exit. Flow can be bled off for tip-cap impingement cooling, and a flow bypass can be specified in which coolant flow is taken off at one point in the flow channel and reintroduced at a point farther downstream in the same channel. The user may either choose the coolant flow rate or let the program determine the flow rate from specified inlet and exit conditions. The computer code integrates the 1-D momentum and energy equations along a defined flow path and calculates the coolant's flow rate, temperature, pressure, and velocity and the heat transfer coefficients along the passage. The equations account for area change, mass addition or subtraction, pumping, friction, and heat transfer.

Meitner, Peter L.↗
NASA NTRS2000

Hydrodynamic Suppression of Soot Formation in Laminar Coflowing Jet Diffusion Flames

Effects of flow (hydrodynamic) properties on limiting conditions for soot-free laminar non-premixed hydrocarbon/air flames (called laminar soot-point conditions) were studied, emphasizing non-buoyant laminar coflowing jet diffusion flames. Effects of air/fuel-stream velocity ratios were of particular interest; therefore, the experiments were carried out at reduced pressures to minimize effects of flow acceleration due to the intrusion of buoyancy. Test conditions included reactant temperatures of 300 K; ambient pressures of 3.7-49 8 kPa; methane-, acetylene-, ethylene-, propane-, and methane-fueled flames burning in coflowing air with fuel-port diameters of 1.7, 3.2, and 6.4 mm, fuel jet Reynolds numbers of 18-121; air coflow velocities of 0-6 m/s; and air/fuel-stream velocity ratios of 0.003-70. Measurements included laminar soot-point flame lengths, laminar soot-point fuel flow rates, and laminar liftoff conditions. The measurements show that laminar soot-point flame lengths and fuel flow rates can be increased, broadening the range of fuel flow rates where the flames remain soot free, by increasing air/fuel-stream velocity ratios. The mechanism of this effect involves the magnitude and direction of flow velocities relative to the flame sheet where increased air/fuel-stream velocity ratios cause progressive reduction of flame residence times in the fuel-rich soot-formation region. The range of soot-free conditions is limited by both liftoff, particularly at low pressures, and the intrusion of effects of buoyancy on effective air/fuel-stream velocity ratios, particularly at high pressures. Effective correlations of laminar soot- and smoke-point flame lengths were also found in terms of a corrected fuel flow rate parameter, based on simplified analysis of laminar jet diffusion flame structure. The results show that laminar smoke-point flame lengths in coflowing air environments are roughly twice as long as soot-free (blue) flames under comparable conditions due to the presence of luminous soot particles under fuel-lean conditions when smoke-point conditions are approached. This is very similar to earlier findings concerning differences between laminar smoke- and sootpoint flame lengths in still environments.

Dai, Z.↗
NASA NTRS1987

Amorphous silicon deposition diagnostics using coherent anti-Stokes Raman spectroscopy

This paper reports on an in situ silicon deposition process diagnostics, using CARS performed at state-of-the-art a-Si:H film deposition conditions in a reactor designed for a-Si:H solar cell fabrication. The diagnostics procedure measures the silane depletion in an RF plasma as a function of the silane flow rate (where the relationship is linear between 50-percent depletion, at a flow rate of 5.6 sccm, and about 8-percent depletion, at 80 sccm) and the RF power (where the silane depletion is linearly dependent on the RF power in the region of 4 to 12 W). The linear RF power dependence of the silane depletion is considered to be consistent with the mechanism of silane decomposition primarily by electron impact dissociations, while the flow rate dependence is interpreted in terms of the residence time of the SiH4 molecules in the glow discharge region, where an increase of the residence time at a low flow rate results in a high depletion ratio.

Shing, Y. H.↗
NASA NTRS2004

Spray Characteristics of a Hybrid Twin-Fluid Pressure-Swirl Atomizer

The spray performance of a fuel injection system applicable for use in main combustion chamber of an oxidizer-rich staged combustion (ORSC) cycles is presented. The experimental data reported here include mean drop size and drop size distribution, spray cone half-angle, and momentum rate (directly related to spray penetration). The maximum entropy formalism, MEF, method to predict drop size distribution is applied and compared to the experimental data. Geometric variables considered include the radius of the injector inlet orifice plate through which oxidizer flows (&) and the exposed length from the fuel inlet to the injector exit plane (L2). Operating conditions that were varied include the liquid mass flow rate and air mass flow rate. For orifices B and C there is a significant dependence of D3Z on both the air and liquid mass flow rates, as well as on L2. For the A orifice, the momentum rate of the air flow appears to exceed a threshold value above which a constant D32 is obtained. Using the MEF method, a semi-analytical process was developed to model the spray distribution using two input parameters (q = 0.4 and Dso). The momentum rate of the spray is directly related to the air and liquid mass flow rates. The cone half angle of the spray ranges from 25 to 17 degrees. The data resulting from this project will eventually be used to develop advanced rocket systems.

Durham, M. J.↗
NASA NTRS1993

Influence of heat transfer rates on pressurization of liquid/slush hydrogen propellant tanks

A multi-dimensional computational model of the pressurization process in liquid/slush hydrogen tank is developed and used to study the influence of heat flux rates at the ullage boundaries on the process. The new model computes these rates and performs an energy balance for the tank wall whereas previous multi-dimensional models required a priori specification of the boundary heat flux rates. Analyses of both liquid hydrogen and slush hydrogen pressurization were performed to expose differences between the two processes. Graphical displays are presented to establish the dependence of pressurization time, pressurant mass required, and other parameters of interest on ullage boundary heat flux rates and pressurant mass flow rate. Detailed velocity fields and temperature distributions are presented for selected cases to further illuminate the details of the pressurization process. It is demonstrated that ullage boundary heat flux rates do significantly effect the pressurization process and that minimizing heat loss from the ullage and maximizing pressurant flow rate minimizes the mass of pressurant gas required to pressurize the tank. It is further demonstrated that proper dimensionless scaling of pressure and time permit all the pressure histories examined during this study to be displayed as a single curve.

Sasmal, G. P.↗
NASA NTRS1994

Bubble Formation at a Submerged Orifice in Reduced Gravity

The dynamic regime of gas injection through a circular plate orifice into an ideally wetting liquid is considered, when successively detached bubbles may be regarded as separate identities. In normal gravity and at relatively low gas flow rates, a growing bubble is modeled as a spherical segment touching the orifice perimeter during the whole time of its evolution. If the flow rate exceeds a certain threshold value, another stage of the detachment process takes place in which an almost spherical gas envelope is connected with the orifice by a nearly cylindrical stem that lengthens as the bubble rises above the plate. The bubble shape resembles then that of a mushroom and the upper envelope continues to grow until the gas supply through the stem is completely cut off. Such a stage is always present under conditions of sufficiently low gravity, irrespective of the flow rate. Two major reasons make for bubble detachment: the buoyancy force and the force due to the momentum inflow into the bubble with the injected gas. The former force dominates the process at normal gravity whereas the second one plays a key role under negligible gravity conditions. It is precisely this fundamental factor that conditions the drastic influence on bubble growth and detachment that changes in gravity are able to cause. The frequency of bubble formation is proportional to and the volume of detached bubbles is independent of the gas flow rate in sufficiently low gravity, while at normal and moderately reduced gravity conditions the first variable slightly decreases and the second one almost linearly increases as the flow rate grows. Effects of other parameters, such as the orifice radius, gas and liquid densities, and surface tension are discussed.

Buyevich, Yu A.↗
NASA NTRS1995

Effects of Gravity on Bubble Formation at a Plate Orifice

A model of the dynamic regime of gas injection through a submerged plate orifice into an ideally wetting liquid is developed in the circumstance when successively detached bubbles may be regarded as independent objects. Two major factors favor bubble detachment: the buoyancy force and a force due to the momentum inflow into the bubble with injected gas. In normal and moderately reduced gravity, the first factor dominates. At relatively low flow rates, a growing bubble is modeled as a spherical segment touching the orifice perimeter during the whole period of its evolution till detachment. If the flow rate exceeds a critical value, another stage of bubble evolution occurs in which an almost spherical gas envelope is connected with the orifice by a nearly cylindrical gaseous stem that lengthens as the bubble rises above the plate. The bubble continues to grow until the gas supply through the stem is completely cut off, after which back flow of gas into the stem from the bubble becomes possible. In microgravity, the second factor prevails, and the latter stage is always present irrespective of the flow rate. However, the gas envelope rises and the stem lengthens very slowly. This difference in the underlying physical mechanisms provides for key properties of bubble growth and detachment being drastically different in appreciable and sufficiently reduced gravity. The frequency of bubble formation slightly decreases with and the detachment volume is almost proportional to the gas flow rate in the first case, in accordance with familiar relations. In the second case, the first variable is proportional to the flow rate whereas the second one is independent of it. Effects of other parameters, such as the orifice radius, gas and liquid densities, and surface tension are discussed.

Webbon, Bruce W.↗
NASA NTRS2001

Investigation of the Flow-Induced Vibration in the E2 Test Facility

An investigation of flow induced vibration due to coupling between the fluid flow and the propellants lines (LOX and RP-1) was performed. Various flow rate conditions were studied to check whether flow induced vibration was possible due to vortex shedding in both valves and pipe lines. Resonance test was conducted for all segments of the LOX-feedline for the preburner under test. In addition, critical values of frequency and velocity are calculated using a mass damping model. A simple chart characterizing the relation between frequency and velocity is developed for each component; i.e. propellant lines, valves and flow meters. It was found that flow induced vibration occurs for various segments with flow rates of 113 lb/s, 275 lb/s and 40 lb/s. Even more interesting using critical conditions for buckling, it was found that the valve or pipe may collapse for a flow rate of 275 lb/s and valve height of 10% of pipe diameter. Furthermore, two models for the acoustic pressure acting on the segments particularly for the valve are proposed.

Castillo, Luciano↗
NASA NTRS2001

Investigation of the Flow-Induced Vibration in the E2 Test Facility

An investigation of flow induced vibration due to coupling between the fluid flow and the propellants lines (LOX and RP-1) was performed. Various flow rate conditions were studied to check whether flow induced vibration was possible due to vortex shedding in both valves and pipe lines. Resonances test was conducted for all segments of the LOX-feedline for the preburner under test. In addition, critical values of frequency and velocity are calculated using a mass damping model. A simple chart characterizing the relation between frequency and velocity is developed for each component; i.e. propellant lines, valves and flow meters. It was found that flow induced vibration occurs for various segments with flow rates of 113 1b/s, 275 lb/s and 40 lb/s. Even more interesting using critical conditions for buckling, it was found that the valve or pipe may collapse for a flow rate of 275 lb/s and valve height of 10% of pipe diameter. Furthermore, two models for the acoustic pressure acting on the segments particularly for the valve are proposed.

Castillo, Luciano↗
NASA NTRS1999

Permeation Resistance of Chlorinated Polyethylene Against Hydrazine Fuels

The permeation resistance of chlorinated polyethylene (CPE) used in chemical protective clothing against the aerospace fuels hydrazine, monomethylhydrazine (MMH), and uns-dimethylhydrazine (UDMH) was determined by measuring breakthrough times and time-averaged vapor transmission rates using an ASTM F 739 permeation cell. Two exposure scenarios were simulated: a 2 hour (h) fuel vapor exposure, and a liquid fuel "splash" followed by a 2 h vapor exposure. To simulate internal suit pressure during operation, a positive differential pressure of 0.3 in. water (75 Pa) on the collection side of the permeation apparatus was used. Breakthrough was observed after exposure to liquid MMH, and to vapor and liquid UDMH. No breakthrough was observed after exposure to vapor and liquid hydrazine, or vapor MMH. A model was then used to calculate propellant concentrations inside a totally encapsulating chemical protective suit based on the ASTM permeation data obtained in the present study. Concentrations were calculated under conditions of fixed vapor transmission rate, variable breathing air flow rate, and variable splash exposure area. Calculations showed that the maximum allowable permeation rates of hydrazine fuels through CPE were of the order of 0.05 to 0.08 ng sq cm/min for encapsulating suits with low breathing air flow rates (of the order of 5 scfm or 140 L/min). Above these permeation rates, the 10 parts per billion (ppb) threshold limit value time - weighted average could be exceeded for chemical protective suits having a CPE torso. To evaluate suit performance at ppb level concentrations, use of a sensitive analytical method such as cation exchange high performance liquid chromatography with amperometric detection was found to be essential.

Waller, J. M.↗
NASA NTRS1999

Sensitive Procedures for Determining the Permeation Resistance of Chlorinated Polyethylene Against Liquid Propellants

The permeation resistance of chlorinated polyethylene (CPE) used in totally encapsulating chemical protective suits against the aerospace fuels hydrazine, monomethylhydrazine, and uns-dimethylhydrazine was determined by measuring the breakthrough time (BT) and time-averaged vapor transmission rate (VTR) using procedures consistent with ASTM F 739 and ASTM F 1383. Two exposure scenarios were simulated: a 2 hour (h) fuel vapor exposure, and a liquid fuel "splash" followed by a 2 h vapor exposure. To simulate internal suit pressure during operation, a positive differential pressure of 0.3 in. water (75 Pa) on the collection side of the permeation apparatus was used. Using the available data, a model was developed to estimate propellant concentrations inside an air-line fed, totally encapsulating chemical protective suit. Concentrations were calculated under simulated conditions of fixed vapor transmission rate, variable breathing air flow rate, and variable splash exposure area. Calculations showed that the maximum allowable permeation rates of hydrazine fuels through CPE were of the order of 0.05 to 0.08 ng/sq cm min for encapsulating suits with low breathing air flow rates (of the order of 5 scfm or 140 L min-1). Above these permeation rates, the 10 parts-per-billion (ppb) threshold limit value time-weighted average could be exceeded. To evaluate suit performance at 10 ppb threshold-limiting value/time-weighted average level concentrations, use of a sensitive analytical method such as cation exchange high performance liquid chromatography with amperometric detection was found to be essential. The analytical detection limit determines the lowest measurable VTR, which in turn governed the lowest per meant concentration that could be calculated inside the totally encapsulating chemical protective suit.

Waller, Jess M.↗
NASA NTRS1986

Bearing tester data compilation analysis, and reporting and bearing math modeling

Integration of heat transfer coefficients, modified to account for local vapor quality, into the 45 mm bearing model has been completed. The model has been evaluated with two flow rates and subcooled and saturated coolant. The evaluation showed that by increasing the flow from 3.6 to 7.0 lbs/sec the average ball temperature was decreased by 102 F, using a coolant temperature of -230 F. The average ball temperature was decreased by 63 F by decreasing the inlet coolant temperature from saturated to -230 F at a flow rate of 7.0 lbs/sec. Since other factors such as friction, cage heating, etc., affect bearing temperatures, the above bearing temperature effects should be considered as trends and not absolute values. The two phase heat transfer modification has been installed in the 57 mm bearing model and the effects on bearing temperatures have been evaluated. The average ball temperature was decreased by 60 F by increasing the flow rate from 4.6 to 9.0 lbs/sec for the subcooled case. By decreasing the inlet coolant temperature from saturation to -24 F, the average ball temperature was decreased 57 F for a flow rate of 9.0 lbs/sec. The technique of relating the two phase heat transfer coefficient to local vapor quality will be applied to the tester model and compared with test data.

Cody, J. C.↗
NASA NTRS2002

Rudder/Fin Seals Investigated for the X-38 Re-Entry Vehicle

NASA is developing the X-38 vehicle that will demonstrate the technologies required for a potential crew return vehicle for the International Space Station. This vehicle would serve both as an ambulance for medical emergencies and as an evacuation vehicle for the space station. Control surfaces on the X-38 (body flaps and rudder/fin assemblies) require high temperature seals to limit hot gas ingestion and the transfer of heat to underlying low temperature structures. Working with the NASA Johnson Space Center, the Seals Team at the NASA Glenn Research Center completed a series of tests to further characterize baseline seal designs for the rudder/fin interfaces of the X-38. The structures of the rudder/fin assembly and its associated seals are shown in the the preceding illustration. Tests performed at Glenn indicated that exposure of the seals in a compressed state at simulated seal re-entry temperatures resulted in a large permanent set and loss of seal resiliency. This could be of concern because the seals are required to maintain contact with the sealing surfaces while the vehicle goes through the maximum re-entry heating cycle to prevent hot gases from leaking past the seals and damaging interior low-temperature structures. To simulate conditions in which the seals may become unloaded during use, such as when they take on a large permanent set, Glenn researchers performed room temperature flow and compression tests to determine seal flow rates, resiliency, and unit loads under minimal loads. Flow rates through an unloaded (i.e., 0-percent compression) double seal arrangement were twice those of a double seal compressed to the 20-percent design compression level. These flow rates are being used in thermal analyses to predict the effect of flow through the seals on over-all seal temperatures. Compression test results showed that seal unit loads and contact pressures were below the limits that Johnson had set as goals for the seals. In the rudder/fin seal location, the seals are in contact with shuttle thermal tiles and are moved across the tiles as the rudder is rotated during re-entry. Low seal unit loads and contact pressures are required to limit the loads on these tiles and minimize any damage that the seals could cause. A series of tests were performed on these seals in NASA Ames Research Center's arc jet facility. The arc jet facility approximates relevant thermal environments that a seal or other structure would be subjected to during extreme heating conditions such as those experienced during space vehicle re-entry. Eleven tests were completed, including one test in which no seal was installed in the gap to examine the flow of heat down into the gap. The seal was compressed between stationary insulation tiles and a movable elevon that was rotated during the test to deflect the arc jet exhaust into the seal gap. Peak seal temperatures as high as 2000 F were reached during the 5-min tests. Results of these tests indicate satisfactory performance of the seal for single-use (e.g., X-38) applications. The results of these tests were shared with the NASA Johnson Space Center and are being used to validate aerothermostructural analysis codes that predict seal temperatures under these conditions. The tests performed at Glenn have provided valuable information to Johnson about the performance of the seals that they are considering using in the rudder/fin location of the X-38 vehicle. Glenn and Johnson are currently defining what additional work needs to be done to develop the final rudder/fin seal design for the X-38 vehicle.

Dunlap, Patrick H., Jr.↗
NASA NTRS1993

Two stage turbine for rockets

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The advanced turbine airfoils were designed and tested by Pratt & Whitney. The technology represented by this turbine is being developed for a turbopump to be used in an advanced upper stage rocket engine. The advanced engine will use a hydrogen expander cycle and achieve high performance through efficient combustion of hydrogen/oxygen propellants, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low mass flow rates and high operating pressures result in very small airfoil heights and diameters. The high efficiency and small size requirements present a challenging turbine design problem. The shrouded axial turbine blades are 50 percent reaction with a maximum thickness to chord ratio near 1. At 6 deg from the tangential direction, the nozzle and blade exit flow angles are well below the traditional design minimum limits. The blade turning angle of 160 deg also exceeds the maximum limits used in traditional turbine designs.

Veres, Joseph P.↗