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At least 1,531 records · Page 85

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NASA NTRS1994

Advances in LO2 Propellant Conditioning

This paper describes the cryogenic testing and analysis that has recently been completed as part of a multi-year effort to develop a new, more robust and operable LO2 propellant conditioning system. Phase 1 of the program consisted of feasibility demonstrations ot four novel propellant conditioning concepts. A no-bleed, passive propellant conditioning option was shown for the first time to successfully provide desired propellant inlet conditions. The benefits of passive conditioning are reduced operations costs, decreased hardware costs, enhanced operability and increased reliability on future expendable launch vehicles In Phase 2 of the test program, effects of major design parameters were studied and design correlation for future vehicle design were developed. Simultaneously, analytical models were developed and validated. Over 100 tests were conducted with a full-scale feedline using LN2 as the test fluid. A circulation pump provided a range of pressure and flow conditions. The test results showed that the passive propellant conditioning system is insensitive to variations in many of the parameters. The test program provides the validation necessary to incorporate the passive conditioning system into the baseline of future vehicles. Modeling of these systems using computational fluid dynamics seems highly promising.

Mehta, Gopal↗

NASA NTRS2020

Maxar Electric Propulsion Development for Deep Space

Maxar has completed development activity for its deep-space electric propulsion system in support of NASA’s Psyche mission. Psyche’s electric propulsion system relies heavily on Maxar’s heritage EP-140 subsystem centered on the SPT-140. The evolution to the EP-140DS demonstrates an expanded thruster discharge capability from 0.9kW through 4.5kW and includes the addition of a constant keeper current function. This system, with development support from industry experts at JPL, will serve as the primary propulsion on board the Psyche spacecraft through the 3.6 year cruise to the Psyche asteroid. Subsystem level testing to validate the first EP-140DS design is complete. Major findings from system testing including xenon flow control loop timing, noise hardening, and constant keeper current and test activities are now being implemented in hardware and the status and plans are discussed.

Chaplin, Vernon↗

DOE OSTI2019

Time-Resolved PIV Measurements of Flow Field at the Exit of a Diffuser Mounted on a Rotating Detonation Combustor

Time-resolved particle image velocimetry at 30 kHz is applied to the exhaust flow field of a RDC system configured with a diffuser geometry and a converging-diverging (CD) nozzle. The RDC is operated with methane fuel and oxygen-enriched air, motivated by power generation applications. The measured flow-field is discretized into axial and non-axial velocity components. Results are shown to characterize the flow-field across a cord length at the exit plane. Instantaneous contours of the two velocity components, and 1D histograms of velocity components, and 2D histograms of total velocity and tilt angle are shown at multiple locations in the flow. Repeatability of the data is evaluated by comparing tests conducted at similar flow rates. Results are presented for a baseline test case, and test cases showing the effect of reactant higher flow rate and CD nozzle geometry. Finally, the results are compared to a previous RDC study without the diffuser to compare the effectiveness of the diffuser and CD nozzle designs in mitigating the flow periodicity at the combustor exit.

particle image velocimetry, pressure gain combusti↗

NASA NTRS1976

Langley facility for tests at Mach 7 of subscale, hydrogen-burning, airframe-integratable, scramjet models

Modifications to a 20-megawatt arc-heated facility for testing a hydrogen-burning, airframe-integratable, subscale, scramjet model are described. Arc-heated flow is mixed with unheated air to furnish a test flow duplicating Mach 7 flight. (Stagnation temperature is 2220 K.) Modifications to the commercially available heater to improve survivability and smoothness are described. Pitot profiles show uniform flow and a slightly thinner nozzle boundary layer than predicted. Comparison of the tunnel boundary layer, which will be ingested by the engine model, with the boundary layer that a flight engine might ingest from its vehicle forebody shows a difference in the density distribution through the boundary layer. Calculations of wall heating and transient wall temperatures of the engine model show that for a 30-sec burn, the heat sink model requires cooling at selected locations to avoid thermal-stress, cycle-life problems. Model performance predictions show that fuel equivalence ratio and nozzle exit area both have large effects on thrust. Average inlet entrance Mach number (as affected by boundary-layer ingestion) has little effect on thrust.

Boatright, W. B.↗

DOE OSTI2019

Time-Resolved PIV Measurements of Flow Field at Diffuser Exit Downstream of a Rotating Detonation Combustor

03 NATURAL GAS↗

NASA NTRS1972

Experimental and Analytical Investigation of the Coolant Flow Characteristics in Cooled Turbine Airfoils

The flow characteristics of turbine airfoil cooling system components were experimentally investigated. Flow models representative of leading edge impingement, impingement with crossflow (midchord cooling), pin fins, feeder supply tube, and a composite model of a complete airfoil flow system were tested. Test conditions were set by varying pressure level to cover the Mach number and Reynolds number range of interest in advanced turbine applications. Selected geometrical variations were studied on each component model to determine these effects. Results of these tests were correlated and compared with data available in the literature. Orifice flow was correlated in terms of discharge coefficients. For the leading edge model this was found to be a weak function of hole Mach number and orifice-to-impinged wall spacing. In the impingement with crossflow tests, the discharge coefficient was found to be constant and thus independent of orifice Mach number, Reynolds number, crossflow rate, and impingement geometry. Crossflow channel pressure drop showed reasonable agreement with a simple one-dimensional momentum balance. Feeder tube orifice discharge coefficients correlated as a function of orifice Mach number and the ratio of the orifice-to-approach velocity heads. Pin fin data was correlated in terms of equivalent friction factor, which was found to be a function of Reynolds number and pin spacing but independent of pin height in the range tested.

Damerow, W. P.↗

NASA NTRS1999

Rotor Re-Design for the SSME Fuel Flowmeter

The present report describes the process of redesigning a new rotor for the SSME Fuel Flowmeter. The new design addresses the specific requirement of a lower rotor speed which would allow the SSME operation at 1 15% rated power level without reaching a blade excitation by the wakes behind the hexagonal flow straightener upstream at frequencies close to the blade natural frequency. A series of calculations combining fleet flowmeters test data, airfoil fluid dynamics and CFD simulations of flow patterns behind the flowmeter's hexagonal straightener has led to a blade twist design alpha = alpha (radius) targeting a kf constant of 0.8256. The kf constant relates the fuel volume flow to the flowmeter rotor speed, for this particular value 17685 GPM at 3650 RPM. Based on this angle distribution, two actual blade designs were developed. A first design using the same blade airfoil as the original design targeted the new kf value only. A second design using a variable blade chord length and airfoil relative thickness targeted simultaneously the new kf value and an optimum blade design destined to provide smooth and stable operation and a significant increase in the blade natural frequency associated with the first bending mode, such that a comfortable margin could be obtained at 115% RPL. The second design is a result of a concurrent engineering process, during which several iterations were made in order to achieve a targeted blade natural frequency associated with the first bending mode of 1300 Hz. Water flow tests preliminary results indicate a kf value of 0.8179 for the f-irst design, which is within 1% of the target value. The second design rotor shows a natural frequency associated with the first bending mode of 1308 Hz, and a water-flow calibration constant of kf 0.8169.

Marcu, Bogdan↗

NASA NTRS1991

Flow Visualization Study of a 1/48-Scale AFTI/F111 Model to Investigate Horizontal Tail Flow Disturbances

During flight testing of the AFTI/F111 aircraft, horizontal tail buffet was observed. Flutter analysis ruled out any aeroelastic instability, so a water-tunnel flow visualization study was conducted to investigate possible flow disturbances on the horizontal tail which might cause buffet. For this study, a 1/48-scale model was used. Four different wing cambers and one horizontal tail setting were tested between 0 and 20 deg angle of attack. These wing cambers corresponded to the following leading training edge deflections: 0/2, 10/10, 10/2, and 0/10. Flow visualization results in the form of still photographs are presented for each of the four wing cambers between 8 and 12 deg angle of attack. In general, the horizontal tail experiences flow disturbances which become more pronounced with angle of attack or wing trailing-edge deflection.

Bjarke, Lisa J.↗

DOE OSTI2016

WHEN SHOCK WAVES COLLIDE

Supersonic outflows from objects as varied as stellar jets, massive stars, and novae often exhibit multiple shock waves that overlap one another. When the intersection angle between two shock waves exceeds a critical value, the system reconfigures its geometry to create a normal shock known as a Mach stem where the shocks meet. Mach stems are important for interpreting emission-line images of shocked gas because a normal shock produces higher postshock temperatures, and therefore a higher-excitation spectrum than does an oblique shock. In this paper, we summarize the results of a series of numerical simulations and laboratory experiments designed to quantify how Mach stems behave in supersonic plasmas that are the norm in astrophysical flows. The experiments test analytical predictions for critical angles where Mach stems should form, and quantify how Mach stems grow and decay as intersection angles between the incident shock and a surface change. While small Mach stems are destroyed by surface irregularities and subcritical angles, larger ones persist in these situations and can regrow if the intersection angle changes to become more favorable. The experimental and numerical results show that although Mach stems occur only over a limited range of intersection angles and size scales, within these ranges they are relatively robust, and hence are a viable explanation for variable bright knots observed in Hubble Space Telescope images at the intersections of some bow shocks in stellar jets.

79 ASTRONOMY AND ASTROPHYSICS↗

DOE OSTI2018

The dissolution behavior of borosilicate glasses in far-from equilibrium conditions

In far-from-equilibrium conditions, the dissolution of borosilicate glasses used to immobilize nuclear waste is known to be a function of both temperature and pH. The aim of this paper is to study effects of these variables on three model waste glasses (SON68, ISG, AFCI). To do this, experiments were conducted at temperatures of 23, 40, 70, and 90 °C and pH(RT) values of 9, 10, 11, and 12 with the single-pass flow-through (SPFT) test method. The results from these tests were then used to parameterize a kinetic rate model based on transition state theory. Both the absolute dissolution rates and the rate model parameters are compared with previous results. Discrepancies in the absolute dissolution rates as compared to those obtained using other test methods are discussed. Rate model parameters for the three glasses studied here are nearly equivalent within error and in relative agreement with previous studies. The results were analyzed with a linear multivariate regression (LMR) and a nonlinear multivariate regression performed with the use of the Glass Corrosion Modeling Tool (GCMT), which is capable of providing a robust uncertainty analysis. This robust analysis highlights the high degree of correlation of various parameters in the kinetic rate model. As more data are obtained on borosilicate glasses with varying compositions, the effect of glass composition on the rate parameter values could possibly be obtained. This would allow for the possibility of predicting the forward dissolution rate of glass based solely on composition

Neeway, James J.↗

NASA NTRSDate not supplied

Reference-Free, Projection Background-Oriented Schlieren

A projection background-oriented schlieren (P-BOS) system is developed and demonstrated. Instead of a background that has a speckle pattern printed, painted, or otherwise deposited onto its surface, and can thus not be altered, the pattern here is projected onto the background. This allows for changes to the speckle pattern without replacement of the background material, which can be time-consuming and expensive. Reference images are acquired simultaneously with flow images. A pre-test transformation between the reference and flow images allows for on-the-fly changes to the speckle pattern during a test without requiring a stoppage of the flow, allowing for optimization of the BOS signal. Using a programmable LCD screen as the speckled optic allows for remote control of the speckle pattern. Because the background is not speckled, the system can be easily transformed to acquired shadowgraph images by removing the speckled optic and closing the aperture of the light source, which is useful for achieving measurements with higher spatial resolution. The system can be nearly as compact as a conventional BOS system, and can be assembled with polarized optics to reduce or eliminate window reflections and glare, making this system particularly well-suited for wind tunnel testing.

Joshua M Weisberger↗

NASA NTRS1969

Pressure measurements and gas-flow analysis in chambers A and B during thermal vacuum tests of spacecraft 2TV-1 and LTA-8.

Pressure measurements and gas flow analysis during thermal vacuum tests of manned spacecraft indicating adequate space vacuum simulation

Ehlers, H. K. F.↗

NASA NTRS1990

Airflow Model Testing to Determine the Distribution of Hot Gas Flow and O/F Ratio Across the Space Shuttle Main Engine Main Injector Assembly

Engine 0209, the certification engine for the new Phase 2+ Hot Gas Manifold (HGM), showed severe deterioration of the Main Combustion Chamber (MCC) liner during hot fire tests. One theory on the cause of the damage held that uneven local distribution of the fuel rich hot gas flow through the main injector assembly was producing regions of high oxidizer/fuel (O/F) ratio near the wall of the MCC liner. Airflow testing was proposed to measure the local hot gas flow rates through individual injector elements. The airflow tests were conducted using full scale, geometrically correct models of both the current Phase 2 and the new Phase 2+ HGMs. Different main injector flow shield configurations were tested for each HGM to ascertain their effect on the pressure levels and distribution of hot gas flow. Instrumentation located on the primary faceplate of the main injector measured hot gas flow through selected injector elements. These data were combined with information from the current space shuttle main engine (SSME) power balances to produce maps of pressure, hot gas flow rate, and O/F ratio near the main injector primary plate. The O/F distributions were compared for the different injector and HGM configurations.

Mahorter, L.↗

NASA NTRS2009

Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

Melcher, John C., IV↗

NASA NTRS2011

Large-Scale Low-Boom Inlet Test Overview

This presentation provides a high level overview of the Large-Scale Low-Boom Inlet Test and was presented at the Fundamental Aeronautics 2011 Technical Conference. In October 2010 a low-boom supersonic inlet concept with flow control was tested in the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). The primary objectives of the test were to evaluate the inlet stability and operability of a large-scale low-boom supersonic inlet concept by acquiring performance and flowfield validation data, as well as evaluate simple, passive, bleedless inlet boundary layer control options. During this effort two models were tested: a dual stream inlet intended to model potential flight hardware and a single stream design to study a zero-degree external cowl angle and to permit surface flow visualization of the vortex generator flow control on the internal centerbody surface. The tests were conducted by a team of researchers from NASA GRC, Gulfstream Aerospace Corporation, University of Illinois at Urbana-Champaign, and the University of Virginia

Hirt, Stefanie↗

DOE OSTI2016

ACCELEROMETERS IN FLOW FIELDS: A STRUCTURAL ANALYSIS OF THE CHOPPED DUMMY INPILE TUBE

Four tests characterizing the structural response of the Chopped-Dummy In-Pile tube (CDIPT) experiment design were measured in the Hydro-Mechanical Fuel Test Facility (HMFTF). Four different test configurations were tried. These configurations tested the pressure drop and flow impact of various plate configurations and flow control orifices to be used later at different reactor power levels. Accelerometers were placed on the test vehicle and flow simulation housing. A total of five accelerometers were used with one on the top and bottom of the flow simulator and vehicle, and one on the outside of the flow simulator. Data were collected at a series of flow rates for 5 seconds each at an acquisition rate of 2 kHz for a Nyquist frequency of 1 kHz. The data were then analyzed using a Fast Fourier Transform (FFT) algorithm. The results show very coherent vibrations of the CDIPT experiment on the order of 50 Hz in frequency and 0.01 m/s2 in magnitude. The coherent vibrations, although small in magnitude pose a potential design problem if the frequencies coincide with the natural frequency of the fueled plates or test vehicle. The accelerometer data was integrated and combined to create a 3D trace of the experiment during the test. The merits of this data as well as further anomalies and artifacts are also discussed as well as their relation to the instrumentation and experiment design.

11 NUCLEAR FUEL CYCLE AND FUEL MATERIALS↗

NASA NTRS2004

Performance of Thermal Mass Flow Meters in a Variable Gravitational Environment

The performance of five thermal mass flow meters, MKS Instruments 179A and 258C, Unit Instruments UFM-8100, Sierra Instruments 830L, and Hastings Instruments HFM-200, were tested on the KC-135 Reduced Gravity Aircraft in orthogonal, coparallel, and counterparallel orientations relative to gravity. Data was taken throughout the parabolic trajectory where the g-level varied from 0.01 to 1.8 times normal gravity. Each meter was calibrated in normal gravity in the orthogonal position prior to flight followed by ground testing at seven different flow conditions to establish a baseline operation. During the tests, the actual flow rate was measured independently using choked-flow orifices. Gravitational acceleration and attitude had a unique effect on the performance of each meter. All meters operated within acceptable limits at all gravity levels in the calibrated orthogonal position. However, when operated in other orientations, the deviations from the reference flow became substantial for several of the flow meters. Data analysis indicated that the greatest source of error was the effect of orientation, followed by the gravity level. This work emphasized that when operating thermal flow meters in a variable gravity environment, it is critical to orient the meter in the same direction relative to gravity in which it was calibrated. Unfortunately, there was no test in normal gravity that could predict the performance of a meter in reduced gravity. When operating in reduced gravity, all meters indicated within 5 percent of the full scale reading at all flow conditions and orientations.

Brooker, John E.↗

NASA NTRS2004

Numerical Simulation of 3-D Supersonic Viscous Flow in an Experimental MHD Channel

The 3-D supersonic viscous flow in an experimental MHD channel has been numerically simulated. The experimental MHD channel is currently in operation at NASA Ames Research Center. The channel contains a nozzle section, a center section, and an accelerator section where magnetic and electric fields can be imposed on the flow. In recent tests, velocity increases of up to 40% have been achieved in the accelerator section. The flow in the channel is numerically computed using a new 3-D parabolized Navier-Stokes (PNS) algorithm that has been developed to efficiently compute MHD flows in the low magnetic Reynolds number regime. The MHD effects are modeled by introducing source terms into the PNS equations which can then be solved in a very e5uent manner. To account for upstream (elliptic) effects, the flowfield can be computed using multiple streamwise sweeps with an iterated PNS algorithm. The new algorithm has been used to compute two test cases that match the experimental conditions. In both cases, magnetic and electric fields are applied to the flow. The computed results are in good agreement with the available experimental data.

Kato, Hiromasa↗