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At least 163 records · Page 9

Throttling capability of a 30 kW class ammonia arcjet

The throttling capabilities of a 30 kW class ammonia arcjet and its compatibility with a breadboard power conditioning unit (PCU) were tested in two series of tests. The first series was performed to determine the performance and operating characteristics of the arcjet and the PCU over a range of power levels and propellant flow rates. The power levels for the tests were nominally between 10 and 30 kW, with some operation below 10 kW at the lower flow rates. The ammonia flow rates varied between 0.16 and 0.35 g/s. The second series of tests was an extensive investigation of operation below 12 kW using three cathode spacings. The ammonia flow rates were between 0.115 and 0.335 g/s. Operation of the arcjet from 1.5 kW up to the 30 kW design point was demonstrated with the PCU.

Goodfellow, K. D.↗

Modeling of A Large-Scale Liquid Oxygen Propellant Transfer Test

Current planned NASA missions to the Moon and Mars involve the transfer of cryogenic propellants in orbit. Correctly predicting the flow rate during the propellant transfer operation could help with planning these operations. In lieu of expensive tests conducted on-orbit, accurate predictive computational models of these propellant transfer processes can be used to predict the propellant transfer flow rate. To gain confidence in these computational models, appropriate anchoring and validation to experimental data can be performed. Ground tests conducted at SpaceX involving the transfer of liquid oxygen were carried out to investigate the effect of the thermodynamic conditions including inlet pressure and temperature and downstream pressure on the flow rate. The results from the test were compared against flow rate predictions from a commercially available two-phase flow software. On average the model predicted the flow rates within 10% of the experimentally observed flow rates for the full range of thermodynamic inlet and outlet conditions.

Daniel Hauser↗

Hydrodynamics of Two-Phase Flows through Porous Media in Microgravity: Packed Bed Reactor Experiment onboard of the International Space Station

The objective of the Packed Bed Reactor Experiment was to develop a fundamental understanding of the hydrodynamics of two-phase flow with no phase transition through porous media in microgravity. This work describes two experimental campaigns flown on the International Space Station (ISS) that were designed to achieve the objective. This work presents the results on flow patterns observed, two-phase flow pressure drop in porous media, and the impact of capillary effects on pressure drop at various gas and liquid flow rates. This work is the first to present predictive correlations of the two-phase friction factor for the different regimes identified based on the superficial liquid and gas velocity in microgravity, the first to hypothesize the different regimes based on the change in slope of the pressure gradient versus flow rate plots, the first to address the effects of the capillarity on the pressure gradient and the first to measure column holdup and assess the effective porosity based on the packing and the trapped gas bubbles. Experimental data on pressure drop in gas (N2) – liquid (water) flow show the presence of four different flow regimes in microgravity. Dispersed bubble flow (DB) and pulse (P) flow regimes are detected at high liquid flow rates, whereas at low liquid flow rates “large bubble” or elongated bubble (LB) and “gas channeling” (GC) regimes are observed. For these different flow regimes, different two-phase flow friction factor and pressure gradient correlations are presented for the first time as a function of the gas and liquid modified Reynolds numbers and Suratman number. Within the viscous-capillary (V-C) regime, it is found that the capillary contribution is the dominant force that contributes to the pressure drop for the wetting case (glass). However, for the non-wetting packing (Teflon), the viscous contribution dominates. It was found that the gas hold-up and pressure drop are functions of bed history at low liquid and gas flow rates with the magnitude of the hysteresis decreasing with increasing flow rates. PBRE-2 results show that the capillary force is a strong function of the superficial liquid velocity but is a much weaker function of the superficial gas velocity and varies inversely with the particle diameter. Within the Viscous-Capillary (V-C) regime, over 90% of the pressure gradient is attributed to the capillary contribution in the gas continuous regime. However, in the large bubble regime, the viscous and capillary contributions were comparable. After the completion of PBRE-2 with glass spherical beads, another bed packed with alumina was installed. Pressure gradient data for the alumina packed bed (PBRE-Water Recovery) were transferred to the PI and were not analyzed by the authors of this work. The Packed Bed Reactor Experiment concluded its on-orbit operation after two successful campaigns of testing in 2017 and in 2021. The flight hardware was brought back from orbit and is being reconditioned for a series of future experiments referred to as PBRE-WRS in support of water recovery (WRS). The WRS series consists of testing different two-phase fluid system components and packed beds to assess their pressure gradient characteristics, which will be used for designing packed bed reactors for various applications relevant to life support. Although there is so much relevance of the experimental results obtained from the two PBRE campaign in microgravity to life support, these findings also apply to other applications that involve two phase flow in porous medium such as fuel cells, transport of nutrient to plants in space and other chemical and materials processing that involve two phase flows. These systems operate differently in microgravity because, due to the lack of buoyancy, the density difference between the phases becomes irrelevant and no longer leads to phase separation.

Packed Bed↗

Atmospheric Pressure Effects on Cryogenic Storage Tank Boil-Off

The Cryogenics Test Laboratory (CTL) at the Kennedy Space Center (KSC) routinely utilizes cryostat test hardware to evaluate comparative and absolute thermal conductivities of a wide array of insulation systems. The test method is based on measurement of the flow rate of gas evolved due to evaporative boil-off of a cryogenic liquid. The gas flow rate typically stabilizes after a period of a couple of hours to a couple of days, depending upon the test setup. The stable flow rate value is then used to calculate the thermal conductivity for the insulation system being tested. The latest set of identical cryostats, 1,000-L spherical tanks, exhibited different behavior. On a macro level, the flow rate did stabilize after a couple of days; however the stable flow rate was oscillatory with peak to peak amplitude of up to 25 percent of the nominal value. The period of the oscillation was consistently 12 hours. The source of the oscillation has been traced to variations in atmospheric pressure due to atmospheric tides similar to oceanic tides. This paper will present analysis of this phenomenon, including a calculation that explains why other cryostats are not affected by it.

Sass, J. P.↗

Thermodynamic Modeling of a Solid Oxide Fuel Cell to Couple with an Existing Gas Turbine Engine Model

The Controls and Dynamics Technology Branch at NASA Glenn Research Center are interested in combining a solid oxide fuel cell (SOFC) to operate in conjunction with a gas turbine engine. A detailed engine model currently exists in the Matlab/Simulink environment. The idea is to incorporate a SOFC model within the turbine engine simulation and observe the hybrid system's performance. The fuel cell will be heated to its appropriate operating condition by the engine s combustor. Once the fuel cell is operating at its steady-state temperature, the gas burner will back down slowly until the engine is fully operating on the hot gases exhausted from the SOFC. The SOFC code is based on a steady-state model developed by The U.S. Department of Energy (DOE). In its current form, the DOE SOFC model exists in Microsoft Excel and uses Visual Basics to create an I-V (current-voltage) profile. For the project's application, the main issue with this model is that the gas path flow and fuel flow temperatures are used as input parameters instead of outputs. The objective is to create a SOFC model based on the DOE model that inputs the fuel cells flow rates and outputs temperature of the flow streams; therefore, creating a temperature profile as a function of fuel flow rate. This will be done by applying the First Law of Thermodynamics for a flow system to the fuel cell. Validation of this model will be done in two procedures. First, for a given flow rate the exit stream temperature will be calculated and compared to DOE SOFC temperature as a point comparison. Next, an I-V curve and temperature curve will be generated where the I-V curve will be compared with the DOE SOFC I-V curve. Matching I-V curves will suggest validation of the temperature curve because voltage is a function of temperature. Once the temperature profile is created and validated, the model will then be placed into the turbine engine simulation for system analysis.

Brinson, Thomas E.↗

Ignition overpressure study from solid rocket motor firings

The objective was to investigate, through experimental means, the basic mechanisms influencing ignition overpressure and to determine ways to suppress ignition overpressure. Ignition overpressure was studied using solid rocket motors with geometry scaled at 1 percent of the Shuttle's Solid Rocket Boosters. Both water injection and aerosol foam were examined as a mean of reducing ignition overpressure. The results of the water injection tests indicate that a relatively small amount of water is sufficient to provide significant suppression. Of the flow rates tested, the lower water injection flow rates provided the best reduction of the ignition overpressure wave. Also, the test results show there is an optimum water flow rate range that provides the best suppression, and as this range is exceeded the effectiveness of water to reduce ignition overpressure is decreased. Aerosol foam provided very little reduction of ignition overpressure, but only small volumes of foam were used and further testing is necessary to determine its total effectiveness as a means of suppression.

Counter, Douglas D.↗

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

This paper will investigate the validation of the 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 a baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet experiment 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 free-stream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the fanface diameter. The numerical simulations with and without tunnel walls are performed, quantifying tunnel wall effects on the BLI inlet flow. A comparison is made between the numerical simulations and the BLI inlet experiment for the baseline and VG vane cases at various inlet mass flow rates. A comparison is also made to a BLI inlet jet configuration for varying actuator mass flow rates at a fixed inlet mass flow rate. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCP avg, very well within the designed operating range of the BLI inlet. A comparison of the average total pressure recovery showed that the simulations were able to predict the trends but had a negative 0.01 offset when compared to the experimental levels. 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 levels 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 misalignment 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 offset that was 0.01 lower than the experiment as was seen in the baseline. Comparisons of the flow features for the jet cases revealed that the CFD predicted a much larger vortex at the engine fan-face when compare to the experiment.

Allan, Brian G.↗

Effects of Convective Solute and Impurity Transport in Protein Crystal Growth

High-resolution optical interferometry was used to investigate the effects of forced solution convection on the crystal growth kinetics of the model protein lysozyme. Most experiments were conducted with 99.99% pure protein solutions. To study impurity effects, approx. 1% of lysozyme dimer (covalently bound) was added in some cases. We show that the unsteady kinetics, corresponding to bunching of growth steps, can be characterized by the Fourier components of time traces of the growth rate. Specific Fourier spectra are uniquely determined by the solution conditions (composition, temperature, and flow rate) and the growth layer source activity. We found that the average step velocity and growth rate increase by approx. I0% with increasing flow rate, as a result of the enhanced solute supply to the interface. More importantly, faster convective transport results in lower fluctuation amplitudes. This observation supports our rationale for system-dependent effects of transport on the structural perfection of protein crystals. We also found that solution flow rates greater than 500 microns/s result in stronger fluctuations while the average growth rate is decreased. This can lead to growth cessation at low supersaturations. With the intentionally contaminated solutions, these undesirable phenomena occurred at about half the flow rates required in pure solutions. Thus, we conclude that they are due to enhanced convective supply of impurities that are incorporated into the crystal during growth. Furthermore, we found that the impurity effects are reduced at higher crystal growth rates. Since the exposure time of terraces is inversely proportional to the growth rate, this observation suggests that the increased kinetics instability results from impurity adsorption on the interface. Finally, we provide evidence relating earlier observations of "slow protein crystal growth kinetics" to step bunch formation in response to nonsteady step generation.

Vekilov, Peter G.↗

Preliminary characterization of a water vaporizer for resistojet applications

A series of tests was conducted to explore the characteristics of a water vaporizer intended for application to resistojet propulsion systems. The objectives of these tests were to (1) observe the effect of orientation with respect to gravity on vaporizer stability, (2) characterize vaporizer efficiency and outlet conditions over a range of flow rates, and (3) measure the thrust performance of a vaporizer/resistojet thruster assembly. A laboratory model of a forced-flow, once-through water vaporizer employing a porous heat exchange medium was built and characterized over a range of flow rates and power levels of interest for application to water resistojets. In a test during which the vaporizer was rotated about a horizontal axis normal to its own axis, the outlet temperature and mass flow rate through the vaporizer remained steady. Throttlability to 30 percent of the maximum flow rate tested was demonstrated. The measured thermal efficiency of the vaporizer was near 0.9 for all tests. The water vaporizer was integrated with an engineering model multipropellant resistojet. Performance of the vaporizer/thruster assembly was measured over a narrow range of operating conditions. The maximum specific impulse measured was 234 s at a mass flow rate and specific power level (vaporizer and thruster combined) of 154 x 10(exp-6)kg/s and 6.8 MJ/kg, respectively.

Morren, W. Earl↗

A soft-start circuit for arcjet ignition

The reduced propellant flow rates associated with high performance arcjets have placed new emphasis on electrode erosion, especially at startup. A soft-start current profile was defined which limited current overshoot during the initial 30 to 50 ms of operation, and maintained significantly lower than the nominal arc current for the first eight seconds of operation. A 2-5 kW arcjet PPU was modified to provide this current profile, and a 500 cycle test using simulated fully decomposed hydrazine was conducted to determine the electrode erosion during startup. Electrode geometry and mass flow rates were selected based on requirements for a 600 sec specific impulse (mission average) arcjet system. The flow rate was varied throughout the test to simulate the blow down of a flight propellant system. Electrode damage was negligible at flow rates above 33 mg/s, and minor chamfering of the constrictor occurred at flow rates of 33 to 30 mg/s, corresponding to flow rates expected in the last 40 percent of the mission. The soft-start current profile significantly reduced electrode damage when compared to state-of-the-art starting techniques.

Hamley, John A.↗

A Soft-Start Circuit for Arcjet Ignition

The reduced propellant flow rates associated with high performance arcjets have placed new emphasis on electrode erosion, especially at startup. A soft-start current profile was defined which limited current overshoot during the initial 30 to 50 ms of operation, and maintained significantly lower than the nominal arc current for the first eight seconds of operation. A 2-5 kW arcjet PPU was modified to provide this current profile, and a 500 cycle test using simulated fully decomposed hydrazine was conducted to determine the electrode erosion during startup. Electrode geometry and mass flow rates were selected based on requirements for a 600 second specific impulse mission average arcjet system. The flow rate was varied throughout the test to simulate the blow down of a flight propellant system. Electrode damage was negligible at flow rates above 33 mg/s, and minor chamfering of the constrictor occurred at flow rates of 33 to 30 mg/s, corresponding to flow rates expected in the last 40 percent of the mission. Constrictor diameter remained unchanged and the thruster remained operable at the completion of the test. The soft-start current profile significantly reduced electrode damage when compared to state of the art starting techniques.

Hamley, John A.↗

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↗

Benchmark Performance Metrics of a Vane Propellant Management Device for a 0.15 m3 Liquid Hydrogen Tank

The use of cryogenic propellants has and will continue to play an integral role in manned-space exploration due the high specific impulses offered and its ubiquity through in-situ resource utilization. But guaranteeing vapor-free transfer of such low-surface tension liquids is difficult for traditional capillary-action propellant management devices (PMDs). Screen-channel liquid acquisition devices and compliant origami bladders are potential solutions, but to quantify the benefits these technologies offer, this paper present a cast study analyzing the performance metrics of an orthodox vane PMD as a benchmark for comparison. A 0.15 m3 liquid hydrogen tank at 20.3 K and 103 kPa was selected for study. Then assuming no body forces and no heat transfer (for simplicity), a steady-state 1-D differential equation was numerically solved in tandem with four possible wetted area configurations to yield an expulsion efficiency for a theoretically maximum inputted expulsion flow rate. Maximization of the flow rate was constrained by the onset of choked flow. Additional inputs, including vane height and vane number, were parametrically varied between 0.1 – 10 cm and 4 – 28 vanes, respectively.

Liquid Acquisition Device↗

Liquid Acquisition Device Design Sensitivity Study

In-space propulsion often necessitates the use of a capillary liquid acquisition device (LAD) to assure that gas-free liquid propellant is available to support engine restarts in microgravity. If a capillary screen-channel device is chosen, then the designer must determine the appropriate combination screen mesh and channel geometry. A screen mesh selection which results in the smallest LAD width when compared to any other screen candidate (for a constant length) is desirable; however, no best screen exists for all LAD design requirements. Flow rate, percent fill, and acceleration are the most influential drivers for determining screen widths. Increased flow rates and reduced percent fills increase the through-the-screen flow pressure losses, which drive the LAD to increased widths regardless of screen choice. Similarly, increased acceleration levels and corresponding liquid head pressures drive the screen mesh selection toward a higher bubble point (liquid retention capability). After ruling out some screens on the basis of acceleration requirements alone, candidates can be identified by examining screens with small flow-loss-to-bubble point ratios for a given condition (i.e., comparing screens at certain flow rates and fill levels). Within the same flow rate and fill level, the screen constants inertia resistance coefficient, void fraction, screen pore or opening diameter, and bubble point can become the driving forces in identifying the smaller flow-loss-to-bubble point ratios.

VanDyke, M. K.↗

Validation of High Aspect Ratio Cooling in a 89 kN (20,000 lb(sub f)) Thrust Combustion Chamber

In order to validate the benefits of high aspect ratio cooling channels in a large scale rocket combustion chamber, a high pressure, 89 kN (20,000 lbf) thrust, contoured combustion chamber was tested in the NASA Lewis Research Center Rocket Engine Test Facility. The combustion chamber was tested at chamber pressures from 5.5 to 11.0 MPa (800-1600 psia). The propellants were gaseous hydrogen and liquid oxygen at a nominal mixture ratio of six, and liquid hydrogen was used as the coolant. The combustion chamber was extensively instrumented with 30 backside skin thermocouples, 9 coolant channel rib thermocouples, and 10 coolant channel pressure taps. A total of 29 thermal cycles, each with one second of steady state combustion, were completed on the chamber. For 25 thermal cycles, the coolant mass flow rate was equal to the fuel mass flow rate. During the remaining four thermal cycles, the coolant mass flow rate was progressively reduced by 5, 6, 11, and 20 percent. Computer analysis agreed with coolant channel rib thermocouples within an average of 9 percent and with coolant channel pressure drops within an average of 20 percent. Hot-gas-side wall temperatures of the chamber showed up to 25 percent reduction, in the throat region, over that of a conventionally cooled combustion chamber. Reducing coolant mass flow yielded a reduction of up to 27 percent of the coolant pressure drop from that of a full flow case, while still maintaining up to a 13 percent reduction in a hot-gas-side wall temperature from that of a conventionally cooled combustion chamber.

Wadel, Mary F.↗

Electrophoresis experiment, experiment MA-014

A continuous, free flow electrophoresis study was conducted during the Apollo Soyuz Test Project mission to investigate and evaluate the increase in sample flow rate and sample resolution achievable in space. The electrophoresis equipment was designed for the separation of four mixtures of biological cells with variable sample flow rates, buffer flow rates, and electric field gradients. Separation quality was assessed by measuring the light from a quartz lamp through the electrophoresis channel and on to a photodiode system. The preliminary results indicate that all monitored systems operated correctly during the experiment. The optical system produced a light that was too bright to discern true cell distributions, but data were received that show a distribution of separated cells.

Hannig, K.↗

Electrophoresis experiment. Experiment MA-014

A continuous free-flow electrophoresis study was conducted during the Apollo-Soyuz Test Project Mission to investigate and evaluate the increase in sample flow rate and sample resolution achievable in space. The electrophoresis equipment was designed for the separation of four mixtures of biological cells with variable sample flow rates, buffer flow rates, and electric field gradients. Separation quality was assessed by measuring the light from a quartz lamp through the electrophoresis channel and onto a photodiode system. The data evaluation indicates that all monitored systems operated correctly during the experiment. The optical system produced a light that was too bright to discern true cell distributions, but final analysis of scientific data by computer processing shows the expected distribution of separated cells.

Hannig, K. H.↗