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NASA NTRS2022

Design of Hypervelocity Flow Generator (HFG) and Its Diagnostics

Ground facilities for hypersonic research are a key element for successful development of aerodynamically proven hypervelocity vehicles. Design concepts and diagnostics of a hypervelocity flow generator (HFG) were made as a test platform for hypersonic/hypervelocity spacecraft models at the NASA Langley Research Center. The HFG is a hypersonic flow field generator using optically heated gas which is blown into an 80 cu m vacuum chamber. The vacuum chamber is kept at a stable vacuum pressure with a combination of three large vacuum pumps, while the HFG is in the test mode. The HFG provides a relatively small test section with approximately a 20 cm window. This facility was designed to generate 2.45 km/sec of flow speed, and potentially generate a continuous flow with the nozzle and vacuum system. The window in test section provides a direct view of the shock wave around a model in order to measure temperature, pressure, and density profile within the shock layer. One of the key test goals under this project is to understand why the emission spectra from the standing shockwave plasma predicted by Lora-Loran codes are significantly different from the measured emission spectra from the Flight Investigation Reentry Environment (FIRE II) Flight. The correct estimate of the thermal loading on the leading edge of hypersonic vehicles greatly affects the aerodynamic design, the material selection for the vehicle, and the cooling requirement and can be obtained by the precise modeling of emission spectra from shockwave plasma. However, the estimation of thermal loading is not an easy task due to complex non-equilibrium radiative process within high temperature shock layers that still falls into a category of cold plasma. Direct flight experiments are the most desirable, but not a cost-effective approach. Analysis by computational fluid dynamics (CFD) offers many test flexibilities. However, the CFD codes must be fully tested and validated with experimental data before the codes are effectively used for practical design. Large discrepancies between experiments and codes appear in hypersonic/hypervelocity flow regimes at high altitudes of 60 km ~ 90 km. This HFG facility offers some important parameters for CFD code validation, such as collision cross-sections, relaxation times, reaction rate coefficients and transportation coefficients. The HFG test facility is based on the ejection flow of high temperature gas heated over to 3500 K through a nozzle. The tungsten gas chamber of the HFG is heated up to a desired temperature by a 60 kW optical power beam source. This system consists of an optical power source, a thermal chamber, an expansion nozzle, a test section, and an 80 cu m vacuum tank. 60 kW optical input power is obtained from the 150 kW Vortek arc lamp system (by Vortek Industries, Vancouver, Canada). This optical beam is focused to heat the gas chamber within which a flow media is heated. The maximum achievable temperature of the flow medium reached approximately 3500 K or even higher but is limited by the melting point of the chamber material used. The exhaust velocity through the nozzle was determined by the stagnation temperature and the molecular weight of the working medium at the test section. To provide design parameters, a NASA Chemical Equilibrium with Application (CEA) computer program is used for the simulation of aerothermal data. This CEA program can calculate chemical equilibrium and properties of complex mixtures using shock tube parameters. For nitrogen gas at 2666 K stagnation temperature, the maximum achievable velocity at the test section is approximately 2.45 km/sec which is within the range of the thermal velocity of 8000 K shock layer. Based on the calculation through the CEA program, the design parameters of a HFG were determined and implemented for the test section that includes an expansion nozzle. The installed test section of the experimental facility can sustain a condition of a re-entry vehicle from the Space at an altitude of 60-90 km. After installation of the HFG, the system was fully tested and its operational parameters were measured. An 80 cu m vacuum chamber of HFG was set at 1 torr level to keep a stable downstream condition. The pumping time to reach the minimum vacuum pressure (~ 1 torr) at the test chamber from atmospheric pressure was approximately an hour using Kinney (MBV-14000/MB -1600/KT-300) and Stokes (Model 1772 and 412) pumps. Such a setup condition allowed a continued stable operation of the HFG experiment with flowrates through 1-, 3-, or 5-mm diameters of nozzle throat. The flow characteristics of the HFG for various operating conditions were performed using a focused Schlieren method. At 0.8 torr chamber pressure, a barrel shock was observed at the test section. With a sphere obstacle of 12 mm diameter in the flow, a bow shock (~ 2-mm thickness) was observed by a focused Schlieren visualization method.

Hypervelocity flow↗
NASA NTRS1985

Simulating Pressure Effects of High-Flow Volumes

Dynamic test stresses realized without high-volume pumps. Assembled in Sections in gas-flow passage, contoured mandrel restricts flow rate to valve convenient for testing and spatially varies pressure on passage walls to simulate operating-pressure profile. Realistic test pressures thereby achieved without extremely high flow volumes.

Kaufman, M.↗
NASA NTRS2016

Space Suit Portable Life Support System (PLSS) 2.0 Unmanned Vacuum Environment Testing

For the first time in more than 30 years, an advanced space suit Portable Life Support System (PLSS) design was operated inside a vacuum chamber representative of the flight operating environment. The test article, PLSS 2.0, was the second system-level integrated prototype of the advanced PLSS design, following the PLSS 1.0 Breadboard that was developed and tested throughout 2011. Whereas PLSS 1.0 included five technology development components with the balance the system simulated using commercial-off-the-shelf items, PLSS 2.0 featured first generation or later prototypes for all components less instrumentation, tubing and fittings. Developed throughout 2012, PLSS 2.0 was the first attempt to package the system into a flight-like representative volume. PLSS 2.0 testing included an extensive functional evaluation known as Pre-Installation Acceptance (PIA) testing, Human-in-the-Loop testing in which the PLSS 2.0 prototype was integrated via umbilicals to a manned prototype space suit for 19 two-hour simulated EVAs, and unmanned vacuum environment testing. Unmanned vacuum environment testing took place from 1/9/15-7/9/15 with PLSS 2.0 located inside a vacuum chamber. Test sequences included performance mapping of several components, carbon dioxide removal evaluations at simulated intravehicular activity (IVA) conditions, a regulator pressure schedule assessment, and culminated with 25 simulated extravehicular activities (EVAs). During the unmanned vacuum environment test series, PLSS 2.0 accumulated 378 hours of integrated testing including 291 hours of operation in a vacuum environment and 199 hours of simulated EVA time. The PLSS prototype performed nominally throughout the test series, with two notable exceptions including a pump failure and a Spacesuit Water Membrane Evaporator (SWME) leak, for which post-test failure investigations were performed. In addition to generating an extensive database of PLSS 2.0 performance data, achievements included requirements and operational concepts verification, as well as demonstration of vehicular interfaces, consumables sizing and recharge, and water quality control.

Watts, Carly↗
NASA NTRS1975

Far infrared maser communications technology

An optically pumped FIR laser was constructed and tested. Optimum operating conditions were determined with CH3OH as the lasing medium. The laser was found to operate equally well with flowing gas or in a sealed off configuration. The FIR cavity stability and pump laser stability were found to have significant problems. The absorption coefficient per unit pressure of 1-1 difluoroethylene at the P(22) and P(24) lines of the 10.4 micron CO2 band was measured. The FIR line pumped by P(22) occurs at approximately 890 microns, which may be in an atmospheric transmission window. It was found that significant Stark tuning of absorption lines of methanol and 1-1 difluoroethylene can be accomplished, even at the usual 100 to 300 mTorr operating pressures of FIR lasers. This means that the use of Stark tuning may enable more effective use of pump laser output.

Claspy, P. C.↗
NASA NTRS1974

Cavitation in liquid cryogens. 4: Combined correlations for venturi, hydrofoil, ogives, and pumps

The results of a series of experimental and analytical cavitation studies are presented. Cross-correlation is performed of the developed cavity data for a venturi, a hydrofoil and three scaled ogives. The new correlating parameter, MTWO, improves data correlation for these stationary bodies and for pumping equipment. Existing techniques for predicting the cavitating performance of pumping machinery were extended to include variations in flow coefficient, cavitation parameter, and equipment geometry. The new predictive formulations hold promise as a design tool and universal method for correlating pumping machinery performance. Application of these predictive formulas requires prescribed cavitation test data or an independent method of estimating the cavitation parameter for each pump. The latter would permit prediction of performance without testing; potential methods for evaluating the cavitation parameter prior to testing are suggested.

Hord, J.↗
NASA NTRS1975

Low-cost, portable fire hose tester

Availability of pumping unit permits scheduling and performing required periodic hose tests in proper manner while retaining full fire equipment readiness. Use of pumping unit preserves operating life and capability of pumper truck.

Jocke, R. F.↗
NASA NTRS1941

An Apparatus for Measuring Rates of Discharge of a Fuel-Injection System

A portable apparatus for rapidly determining rates of discharge of a fuel-injection system is described. Satisfactory operation of this apparatus with injection-pump speeds up to 2400 r.p.m was obtained. Rate-of-discharge tests were made with several cam-plunger-valve injection systems with long injection tubes. A check valve designed to reduce secondary discharges was tested. This check valve was operated with injection-pump speeds up to 2400 r.p.m without the occurrence of large secondary discharges.

Dutee, Francis J↗
NASA NTRS1993

The Influence of Swirl Brakes and a Tip Discharge Orifice on the Rotordynamic Forces Generated by Discharge-to-Suction Leakage Flows in Shrouded Centrifugal Pumps

Recent experiments conducted in the Rotor Force Test Facility at the California Institute of Technology have examined the effects of a tip leakage restriction and swirl brakes on the rotordynamic forces due to leakage flows on an impeller undergoing a prescribed circular whirl. The experiments simulate the leakage flow conditions and geometry of the Alternate Turbopump Design (ATD) of the Space Shuttle High Pressure Oxygen Turbopump and are critical to evaluating the pump's rotordynamic instability problems. Previous experimental and analytical results have shown that discharge-to-suction leakage flows in the annulus of a shrouded centrifugal pump contribute substantially to the fluid induced rotordynamic forces. Also, previous experiments have shown that leakage inlet (pump discharge) swirl can increase the cross-coupled stiffness coefficient and hence increase the range of positive whirl for which the tangential force is destabilizing. In recent experimental work, the present authors demonstrated that when the swirl velocity within the leakage path is reduced by the introduction of ribs or swirl brakes, then a substantial decrease in both the destabilizing normal and tangential forces could be achieved. Motivation for the present research is that previous experiments have shown that restrictions such as wear rings or orifices at pump inlets affect the leakage forces. Recent pump designs such as the Space Shuttle Alternate Turbopump Design (ATD) utilize tip orifices at discharge for the purpose of establishing axial thrust balance. The ATD has experienced rotordynamic instability problems and one may surmise that these tip discharge orifices may also have an important effect on the normal and tangential forces in the plane of impeller rotation. The present study determines if such tip leakage restrictions contribute to undesirable rotordynamic forces. Additional motivation for the present study is that the widening of the leakage path annular clearance and the installation of swirl brakes in the ATD has been proposed to solve its instability problems. The present study assesses the effect of such a design modification on the rotordynamic forces. The experimental apparatus consists of a solid or dummy impeller, a housing instrumented for pressure measurements, a rotating dynamometer and an eccentric whirl mechanism. The solid impeller is used so that leakage flow contributions to the forces are measured, but the main throughflow contributions are not experienced. The inner surface of the housing has been modified to accommodate meridional ribs or swirl brakes within the leakage annulus. In addition, the housing has been modified to accommodate a discharge orifice that qualitatively simulates one side of the balance piston orifice of the Space Shuttle ATD. Results indicate the detrimental effects of a discharge orifice and the beneficial effects of brakes. Plots of the tangential and normal forces versus whirl ratio show a substantial increase in these forces along with destabilizing resonances at some positive whirl ratios when a discharge orifice is added. When brakes are added, some of the detrimental effects of the orifice are reduced. For the tangential force, a plot versus whirl ratio shows a significant reduction and a destabilizing resonance appears to be eliminated. For the normal force, although the overall force is not reduced, again a destabilizing resonance appears to be eliminated.

Sivo, Joseph M.↗
NASA NTRS1998

Modeling of a Two-Phase Jet Pump with Phase Change, Shocks and Temperature-Dependent Properties

One of the primary motivations behind this work is the attempt to understand the physics of a two-phase jet pump which constitutes part of a flow boiling test facility at NASA-Marshall. The flow boiling apparatus is intended to provide data necessary to design highly efficient two-phase thermal control systems for aerospace applications. The facility will also be capable of testing alternative refrigerants and evaluate their performance using various heat exchangers with enhanced surfaces. The test facility is also intended for use in evaluating single-phase performance of systems currently using CFC refrigerants. Literature dealing with jet pumps is abundant and covers a very wide array of application areas. Example application areas include vacuum pumps which are used in the food industry, power station work, and the chemical industry; ejector systems which have applications in the aircraft industry as cabin ventilators and for purposes of jet thrust augmentation; jet pumps which are used in the oil industry for oil well pumping; and steam-jet ejector refrigeration, to just name a few. Examples of work relevant to this investigation includes those of Fairuzov and Bredikhin (1995). While past researchers have been able to model the two-phase flow jet pump using the one-dimensional assumption with no shock waves and no phase change, there is no research known to the author apart from that of Anand (1992) who was able to account for condensation shocks. Thus, one of the objectives of this work is to model the dynamics of fluid interaction between a two-phase primary fluid and a subcooled liquid secondary fluid which is being injected employing atomizing spray injectors. The model developed accounts for phase transformations due to expansion, compression, and mixing. It also accounts for shock waves developing in the different parts of the jet pump as well as temperature and pressure dependencies of the fluid properties for both the primary two-phase mixture and the secondary subcooled liquid. The research effort on which this document partly reports described a relatively simple model capable of describing the performance of a two-phase flow jet pump. The model is based on the isentropic homogeneous expansion/compression hypothesis and is capable of fully incorporating the effects of shocks in both the mixing chamber and the throat/diffuser parts of the pump. The physical system chosen is identical to that experimentally tested by Fairuzov and Bredikhin (1995) and should therefore be relatively easy to validate.

Sherif, S. A.↗
NASA NTRS1990

Mixer-ejector nozzle for jet noise suppression

An aero/acoustic model test of a mixer-ejector nozzle was conducted at the 9 x 15 foot low-speed acoustic wind tunnel at NASA Lewis Research Center. The objective of the test was to get a preliminary assessment of ejector pumping and noise reduction potential of this device for possible application in the exhaust system of an advanced supersonic civil transport. The results of the test showed that goal levels of pumping were achieved. Exit pressure/temperature traverse data showed that there was good mixing between the primary and secondary streams. Acoustics data were dominated by shock noise; jet mixing noise levels were low because of a facility limit on primary temperature. The mixer-ejector did significantly reduce shock noise relative to the baseline conic nozzle. Because the relative magnitudes of jet mixing noise and shock noise were not in the correct proportion to properly model engine noise, an Effective Perceived Noise Level (EPNL) assessment was precluded.

Lord, W. K.↗
NASA NTRS2001

Time-Dependent Simulation of Incompressible Flow in a Turbopump Using Overset Grid Approach

This paper reports the progress being made towards complete unsteady turbopump simulation capability by using overset grid systems. A computational model of a turbo-pump impeller is used as a test case for the performance evaluation of the MPI, hybrid MPI/Open-MP, and MLP versions of the INS3D code. Relative motion of the grid system for rotor-stator interaction was obtained by employing overset grid techniques. Unsteady computations for a turbo-pump, which contains 114 zones with 34.3 Million grid points, are performed on Origin 2000 systems at NASA Ames Research Center. The approach taken for these simulations, and the performance of the parallel versions of the code are presented.

Kiris, Cetin↗
NASA NTRSDate not supplied

Uncertainty Quantification of Inducer Natural Frequency using Conditional Assessment of Modeling and Modal Testing of Simpler Systems

The low pressure fuel pump inducer of the new Space Launch System RS25 core stage engine operates in a highly complex environment that substantially affects its modal characteristics. Some of the more important effects are fluid-added mass resulting from operation within a light liquid (Hydrogen), and the magnification of this effect due to tight tip clearance. Since higher order cavitation has been identified as a significant harmonic driver, knowledge of the natural frequency of potentially excitable modes is critical for safe operation, but this frequency cannot be measured during the severe operational environment. A comprehensive testing and analysis program has therefore been performed over the last four years to identify the nominal value and uncertainty of the frequency by modeling and testing four simpler structures which share some of the characteristics of the operational inducer. This testing was used to assess and adjust modeling techniques and excellent correlation was achieved. Identification of the uncertainty in the inducer frequency itself was still problematic, however. This difficulty led to an investigation of Bayesian uncertainty quantification techniques, and to the application of the relatively simple technique of Multi-Variate Normal conditional distributions to calculate the inducer natural frequency uncertainty. Assumptions on prior distributions of uncertainty of the fluid-added mass and tip clearance effect are initially applied to models of each of the simple structures and the inducer itself, and these uncertainties are propagated to generate natural frequencies using design of experiments. Simple response surfaces are then created from this data in order to calculate a Covariance Matrix relating all of these natural frequencies. Finally, the results from modal test of the simple structures are considered to be observations and used to calculate the conditional variance of the desired inducer frequencies. As this method is less rigorous than more complicated Bayesian methods reported in the literature, a conservative factor is applied to the result, but the resulting uncertainty is still significantly less than originally estimated and will greatly assist certification of the inducer for use in the engine.

Uncertainty Quantification↗
NASA NTRSDate not supplied

Uncertainty Quantification of Inducer Eigenvalues using Conditional Assessment of Models and Modal Test of Simpler Systems

The low pressure fuel pump inducer of the new Space Launch System RS25 core stage engine operates in a highly complex environment that substantially affects its modal characteristics. Some of the more important effects are fluid-added mass resulting from operation within a light liquid (Hydrogen), and the magnification of this effect due to tight tip clearance. Since higher order cavitation has been identified as a significant harmonic driver, knowledge of the natural frequency of potentially excitable modes is critical for safe operation, but this frequency cannot be measured during the severe operational environment. A comprehensive testing and analysis program has therefore been performed over the last four years to identify the nominal value and uncertainty of the frequency by modeling and testing four simpler structures which share some of the characteristics of the operational inducer. This testing was used to assess and adjust modeling techniques and excellent correlation was achieved. Identification of the uncertainty in the inducer frequency itself was still problematic, however. This difficulty led to an investigation of Bayesian uncertainty quantification techniques, and to the application of the relatively simple technique of Multi-Variate Normal conditional distributions to calculate the inducer natural frequency uncertainty. Assumptions on prior distributions of uncertainty of the fluid-added mass and tip clearance effect are initially applied to models of each of the simple structures and the inducer itself, and these uncertainties are propagated to generate natural frequencies using design of experiments. Simple response surfaces are then created from this data in order to calculate a Covariance Matrix relating all of these natural frequencies. Finally, the results from modal test of the simple structures are considered to be observations and used to calculate the conditional variance of the desired inducer frequencies. As this method is less rigorous than more complicated Bayesian methods reported in the literature, a conservative factor is applied to the result, but the resulting uncertainty is still significantly less than originally estimated and will greatly assist certification of the inducer for use in the engine.

Uncertainty Quantification↗
NASA NTRS1985

Levitating Furnance for Containerless Processing in Space

The purpose of this project is to design, construct and test a levitator which would function reliably in a radiant heated cold wall furnace in a reduced gravity environment. A prototype levitating/positioning device termed the Sonic Pump Levitator is designed, built and successfully tested in full gravity and in the reduced gravity of the parabolic flight regime of the KC-135. Positioning is achieved by timely and appropriate application of gas momentum from one or more of six sonic pumps. The sonic pumps, which are arranged orthogonally in opposed pairs about the levitation region, are activated by an electro-optical, computer controlled, feedback system. The Sonic Pump Levitator is a promising candidate for containerless processing in microgravity. Being a null point device, it imposes minimal perturbation on the levitated, or positioned, target. Functioning as it does by optical feedback controlled jets of air, it would be expected to be free from disturbance caused by high temperature. This expectation is now being examined experimentally.

Dunn, S. A.↗
NASA NTRS1981

Analytical and experimental study of flow phenomena in noncavitating rocket pump inducers

The flow processes in rocket pump inducers are summarized. The experimental investigations were carried out with air as the test medium. The major characteristics features of the rocket pump inducers are low flow coefficient (0.05 to 0.2) large stagger angle (70 deg to 85 deg) and high solidity blades of little or no camber. The investigations are concerned with the effect of viscosity not the effects of cavitation. Flow visualization, conventional and hot wire probe measurement inside and at the exit of the blade passage, were the analytical methods used. The experiment was carried out using four three and two bladed inducers with cambered blades. Both the passage and the exit flow were measured. The basic research and boundary layer investigation was carried out using a helical flat plate (of some dimensions as the inducer blades tested), and flat plate helical inducer (four bladed). Detailed mean and turbulence flow field inside the passage as well as the exit of the rotor were derived from these measurement. The boundary layer, endwall, and other passage data reveal extremely complex nature of the flow, with major effects of viscosity present across the entire passage. Several analyses were carried out to predict the flow field in inducers. These included an approximate analysis, the shear pumping analysis, and a numerical solution of exact viscous equations with approximate modeling for the viscous terms.

Lakshminarayana, B.↗
NASA NTRS2005

Large Liquid Rocket Testing: Strategies and Challenges

Rocket propulsion development is enabled by rigorous ground testing in order to mitigate the propulsion systems risks that are inherent in space flight. This is true for virtually all propulsive devices of a space vehicle including liquid and solid rocket propulsion, chemical and non-chemical propulsion, boost stage and in-space propulsion and so forth. In particular, large liquid rocket propulsion development and testing over the past five decades of human and robotic space flight has involved a combination of component-level testing and engine-level testing to first demonstrate that the propulsion devices were designed to meet the specified requirements for the Earth to Orbit launchers that they powered. This was followed by a vigorous test campaign to demonstrate the designed propulsion articles over the required operational envelope, and over robust margins, such that a sufficiently reliable propulsion system is delivered prior to first flight. It is possible that hundreds of tests, and on the order of a hundred thousand test seconds, are needed to achieve a high-reliability, flight-ready, liquid rocket engine system. This paper overviews aspects of earlier and recent experience of liquid rocket propulsion testing at NASA Stennis Space Center, where full scale flight engines and flight stages, as well as a significant amount of development testing has taken place in the past decade. The liquid rocket testing experience discussed includes testing of engine components (gas generators, preburners, thrust chambers, pumps, powerheads), as well as engine systems and complete stages. The number of tests, accumulated test seconds, and years of test stand occupancy needed to meet varying test objectives, will be selectively discussed and compared for the wide variety of ground test work that has been conducted at Stennis for subscale and full scale liquid rocket devices. Since rocket propulsion is a crucial long-lead element of any space system acquisition or development, the appropriate plan and strategy must be put in place at the outset of the development effort. A deferment of this test planning, or inattention to strategy, will compromise the ability of the development program to achieve its systems reliability requirements and/or its development milestones. It is important for the government leadership and support team, as well as the vehicle and propulsion development team, to give early consideration to this aspect of space propulsion and space transportation work.

Rahman, Shamim A.↗
NASA NTRS2008

J-2X Fuel Turbopump Point of Departure: The Performance of the J-2s Fuel Turbopump Inducer

To aid the J-2X program design effort with detailed performance and environment information, the J-2S fuel turbopump (FTP) inducer has undergone a thorough test series in both water and hydrogen. The test series utilizes both inducer only and a complete pump configuration to assess the inducer interaction to the overall turbopump system. The test goals include verification of suction performance against heritage J-2S data, head production, effects of thermodynamic suppression head (TSH), and evaluation of the inducer dynamic pressure caused by cavitation instabilities. Test facilities at both Pratt & Whitney Rocketdyne (PWR) and NASA s Stennis Space Center (SSC) are employed for the testing. The inducer only water test effort conducted at PWR established performance curves for suction performance, head production, and efficiency over a wide operating range. Because the heritage J-2S suction performance data set is in hydrogen, it is desired to obtain current suction performance data in hydrogen as well, thus avoiding the reliance on a theoretical TSH correlation for direct comparison. This data then provides an empirically based TSH correlation allowing for the comparison of water test suction data to system suction requirements. The FTP testing performed at SSC provides these suction performance relationships as well as inlet duct dynamic pressures during liquid hydrogen operation. The test effort successfully confirms the heritage J-2S suction performance and establishes the amount of TSH between water and hydrogen operation at the design flow coefficient. Correlating data is also obtained for cavitating instability frequency content, illustrating the validity of using the wide flow range water test data to predict hydrogen performance.

Sargent, S. R.↗
NASA NTRS2014

Cascade Distiller System Performance Testing Interim Results

The Cascade Distillation System (CDS) is a rotary distillation system with potential for greater reliability and lower energy costs than existing distillation systems. Based upon the results of the 2009 distillation comparison test (DCT) and recommendations of the expert panel, the Advanced Exploration Systems (AES) Water Recovery Project (WRP) project advanced the technology by increasing reliability of the system through redesign of bearing assemblies and improved rotor dynamics. In addition, the project improved the CDS power efficiency by optimizing the thermoelectric heat pump (TeHP) and heat exchanger design. Testing at the NASA-JSC Advanced Exploration System Water Laboratory (AES Water Lab) using a prototype Cascade Distillation Subsystem (CDS) wastewater processor (Honeywell d International, Torrance, Calif.) with test support equipment and control system developed by Johnson Space Center was performed to evaluate performance of the system with the upgrades as compared to previous system performance. The system was challenged with Solution 1 from the NASA Exploration Life Support (ELS) distillation comparison testing performed in 2009. Solution 1 consisted of a mixed stream containing human-generated urine and humidity condensate. A secondary objective of this testing is to evaluate the performance of the CDS as compared to the state of the art Distillation Assembly (DA) used in the ISS Urine Processor Assembly (UPA). This was done by challenging the system with ISS analog waste streams. This paper details the results of the AES WRP CDS performance testing.

Callahan, Michael R.↗