RP-1 Effects on Bearing Performance Investigated
Space Shuttle Main Engine Health-Monitoring System Algorithms Validated
Performance and Heat Load Prediction Improved for Multistage Turbines
Self-Diagnosing Accelerometer Fabricated
Prototype Real-Time Sensor Data Validation System Demonstrated
Metallized Propellant Combustion Tested
Rotor Coatings Tested for Cryogenic Brush Seal Applications
Electric Thruster Plume Impacts on Communications Signals Quantified
Fiber-Optic-Based Methods Improve Gas Analysis and Concentration Monitoring
High-Pressure, Compact Rockets Satisfy Small Satellite Requirements
Engineering Model Ion Thrusters and Power Processors Developed
RL-10 Turbopump Flight Cooldown Characterized
Small-Scale Hydrogen Test System Enables "Smaller, Faster, Cheaper"
Cryogenic Compression Mass Gage Passes Liquid Hydrogen Test
Cryogenic Two-Phase Nitrogen Flow Studied
RP-1 has been and continues to be used by a number of rocket engine propulsion systems. Among them are the F-1 engine used in the Apollo program, which is no longer in production, as well as the RS-27A engine used in the Delta II program and the MA-5A engine used in the Atlas program, both of which are still in production. As various concepts are considered for future launch systems, RP-1 continues to be a viable fuel option. Because more emphasis is being placed on reliability, durability, maintainability, and cost effectiveness in the new design concepts, more information is needed about the effects the various fuels have on rocket engine systems.
Liquid-fueled rocket engines typically utilize turbopumps to transfer the fuel and oxidizer from the storage tanks to the combustion chamber. To reduce the weight and complexity of these turbopumps, their bearings are often lubricated and cooled by the liquid being pumped rather than having a separate lubrication system. Although a few limited investigations of the lubricating capabilities of RP-1 had been performed (refs. 1 to 3), this test program was carried out to more thoroughly assess the impact of RP-1 on rocket engine turbopump bearing performance and reliability.
Early tests (ref. 4) indicated that bearing operating temperature was highly dependent on both shaft speed and RP-1 flow rate but only slightly dependent on applied load over the ranges of interest. Moreover, there was a lower threshold limit of RP-1 flow below which the bearing temperature increased rapidly, indicating that bearing failure was imminent. Conversely, high RP-1 flow rates greatly increased the amount of shaft power lost to the bearing due to churning of the liquid RP-1.
Further tests (ref. 5) subjected the bearings to combinations of speeds and loads resulting in Hertzian contact stresses as high as 2802 N/mm[2] (406,400 psi)--much higher than the generally accepted Hertzian contact stress of 2068 N/mm[2] (300,000 psi). The bearings ran well under these extreme conditions as long as the RP-1 flow remained adequate. If the temperature of the bearing outer ring rose above 150 deg.C (300 deg.F), the RP-1 would begin to boil, causing a loss of lubricant between the balls and races of the bearing and leading to an unstable situation where temperature runaway is likely.
The final tests of the program (ref. 6) investigated an alternative method of introducing the RP-1 to the bearing as well as several different cage materials to determine whether the lower limit of RP-1 flow for safe bearing operation could be reduced. Conventionally, lubricant is introduced to bearings in turbopumps through a stationary jet ring located adjacent to the bearing. This method was used as the baseline for these tests. The alternative method was to direct the RP-1 through a channel in the center of the shaft and then radially outward and through small holes machined in the bearing inner ring. This alternative method reduced bearing operating temperatures as much as 30 deg C (54 deg F). Three alternative cage materials (660 bronze, Vespel, and Bearium (bronze) were tested in addition to the baseline silver-coated 4340 steel cage. The bronze cages reduced bearing operating temperature as much as 10 deg C (18 deg F) more than the other two cage materials.
References
The recursive structural identification (RESID) algorithm is a nonlinear regression algorithm similar to conventional regression analysis. It results in a single equation that relates the input parameters selected by the RESID algorithm to the predicted output parameter. Once the equation is established, new data can be applied to it to estimate the output parameter. The error, which is the difference between the estimated and measured parameter values, is then thresholded to make the fault detection.
Whereas the RESID algorithm uses instantaneous relationships between sensor measurements to detect an off-nominal engine condition, the mixed autoregressive moving average (ARMA) approach focuses on the behavior of individual sensor values over blocks of time. Signal structural changes, rather than cross-sensor relationships, are the indicators of an off-nominal engine condition. The basic approach uses a set of training data to develop an equation for predicting the present value of a sensor by using previous sensor values and white noise inputs.
The CLUSTER algorithm is a sensor fusion technique based on a classical pattern-recognition approach to data classification. In the classical clustering approach a training set consisting of data from n different sensors is selected and processed by a cluster-defining algorithm. This training algorithm uses the interrelationships between the sensors to group them into n -dimensional clusters that indicate the operating states of the system described by the training data. Once the clusters have been defined in the training stage, new data are processed and evaluated either as belonging to one of the predefined operating states (clusters) or as being outside the experience base constructed during training.
The three algorithms were validated over a wide range of SSME configurations and over 150 SSME test firings. During this validation the algorithms demonstrated both robust fault detection and false-alarm rejection; 24 of 30 off-nominal conditions were detected and 21 of these 24 were detected significantly earlier than by the SSME redline system. Also, the algorithms demonstrated operational feasibility during real-time SSME hot-fire tests on the technology test-bed engine.
The algorithms are undergoing continual in-house improvement to expand the fault-detection capability and to add fault-diagnosis capability. When completed, these algorithms can be used to improve SSME safety both on the test stand and in flight; they can also be used as baseline algorithms for future NASA missions.
Bibliography
NASA Lewis and Pratt & Whitney (under Contract NAS3-25804) developed "closure models" to predict time-averaged effects of unsteadiness in multistage turbines. The predictive capabilities of these closure models were verified through design and by testing hardware in a large-scale rotating rig at United Technologies Research Center (UTRC). Generalized formulations of these closure models will enhance the state of the art of turbine design procedures, allowing designers to cost-effectively optimize the performance, life, and structural integrity of turbines used in air-breathing and rocket propulsion systems.
Closure models were formulated by using both new and existing unique experimental and numerical data. These closure models provided numerical values needed for "average passage" solvers developed by scientists at NASA to predict effects of periodic unsteadiness on time-averaged flows in multistage machines. Computational fluid dynamics codes with these closure models were used to redesign airfoil flow for the UTRC large-scale rotating rig so as to reduce heat loads and improve the performance of the second stator. An experimental program was conducted to define the distribution of the second stator's heat load and aerodynamic performance and to allow verification of the predictive capabilities of the closure models. The closure models were then assessed for their predictive capabilities and contribution to the enhancement of the current design system.
With the aid of this contract, NASA Lewis has developed a new understanding of turbine flow physics in low-aspect-ratio multistage turbines and is sharing the results of this contract with the U.S. turbomachinery community.
Bibliography
Lewis contact: Dr. John J. Adamczyk, (216) 433-5829
Headquarters program office: OSAT
Sensor failure can be detected by providing a known input and observing the response. For piezoelectric devices an input could be generated by supplying an electrical excitation signal to the piezoelement itself. When triggered, the excitation signal would produce a measurable vibrational response in the sensor, with its frequency content dependent on the health of the sensor element, the electronics, the mounting impedance, and the dynamic activity from other sources. Comparing the response with a known "good" response will provide information on sensor health.
NASA Lewis is supporting research to develop self-diagnostic and self-calibration technology. A self-diagnosing piezoelectric accelerometer has been fabricated and is scheduled for testing on the technology test-bed engine at NASA Marshall Space Flight Center. This technology, whether incorporated into a rocket engine health management and control system or other applications requiring highly reliable measurements, could improve system safety and cost effectiveness.
Lewis contact: George C. Madzsar, (216) 977-7434
Headquarters program office: OSAT
We combined analytical redundancy with Bayesian belief networks to provide a solution that has well-defined, real-time characteristics and well-defined error rates and can be scaled to validate any number of engine sensors. Analytical redundancy is a technique in which a sensor's value is predicted by using values from other, usually nonredundant, sensors and known or empirically derived mathematical relations. For the example engine plant diagram, fuel flow can be related to either the low-pressure-pump speed or the high-pressure-pump speed by a pump affinity equation (assuming constant fuel density). As shown, a set of sensors and a set of relationships among them form a network of cross-checks that can be used to periodically validate all sensors in the network. In addition to simple binary empirical correlations and engine characteristic equations, we are developing neural network models to model parameters during the highly nonlinear start transient and those parameters difficult to approximate during main-stage engine operation.
Bayesian belief networks are a mathematically sound approach to information fusion--the combination of evidence from several sources into a single consistent solution regarding the status of all sensors in a network. In information fusion, uncertainties in the sources of evidence (i.e., inaccuracies in the sensors or uncertainties in the models themselves) are explicitly modeled and accounted for.
This approach has been used to validate a six-sensor network on Aerojet's advanced rocket engine controller. Data were received in real time from the technology test-bed engine at NASA Marshall Space Flight Center. Current efforts are focused on extending the demonstration system to provide real-time validation capability for additional critical sensors on the space shuttle main engines.
Bibliography
Lewis contact: Claudia M. Meyer, (216) 977-7511
Headquarters program office: OSAT
Metallized gelled propellants have been studied analytically and experimentally for over 60 years. The historical work has focused on the benefits of high specific impulse, high density, and safety (refs. 1 and 2). Several mission studies have indicated that O2/RP-1/Al can have significant benefits by increasing propellant density (refs. 2 to 5). Testing was therefore conducted with O2/RP-1/Al propellants using gelled RP-1 with various loadings of aluminum particles and with liquid propellants as a baseline.
Rocket performance and heat transfer measurements were desired in this test program. During the combustion of metal particles the different densities of the gas, liquid, and metal flows create a mismatch in their speeds. The heat transfer measurements were made to estimate the differences in ignition time for the parts of the combustion flow. The combustion temperatures and heat flow profiles of the different flows will be compared.
Both heat-sink and calorimeter experiments were conducted. The heat-sink combustion chamber has a 2.6-in. inside diameter and is 6 in. long, and the nozzle has a 0.6-in.-diameter throat. A 22-channel cooling passage calorimeter chamber was used. The associated calorimeter nozzle has nine cooling channels. Numerous thermocouples are located in the cooling passages of the calorimeter combustion chamber. The injectors use an oxidizer manifold within the injector body and have a fuel dome set atop it. The injector elements are an O-F-O design and both four- and eight-element patterns were tested. A wide range of oxidizer-to-fuel ratios (O/F) were investigated. The injectors were designed for O/F ranges of 1.2 to 4.2 for O2/RP-1 and 1.4 to 3.7 for O2/RP-1/Al.
All these preliminary results are for the heat-sink engine. With O2/RP-1 the maximum combustion efficiency occurred near O/F = 3.0. With the gelled RP-1 (0 wt% RP-1/Al) the O/F for the maximum combustion efficiency was nearly the same. Both gelled and ungelled RP-1 demonstrated high combustion efficiencies (to 98%). Using O2/RP-1/Al (5 wt% RP-1/Al) delivered 90 to 95% combustion efficiency, and the efficiency curve had no obvious strong peak. At a 55 wt% RP-1/Al loading the combustion efficiency was similar to the 5 wt% RP-1, and the peak occurred near 1.5 to 2.0 O/F.
During the testing with gelled RP-1 and the 5 wt% RP-1/Al, some residual propellant was found in the rocket chamber, coating the injector face and chamber walls. When this thin layer was wiped off, the metal surfaces exhibited minimal erosion--a marked contrast to the blackening of the O2/RP-1 injector faces and injector-face erosion that occurred with the 55 wt% RP-1/Al. An improved cooling technique might be derived from this effect. In the photograph the contrast of the shiny, almost unblemished 5 wt% RP-1/Al injector surface and the partially eroded 55 wt% RP-1/Al injector surface is clearly evident.
High-efficiency combustion of metallized gelled propellants was realized with even simple four- and eight-element triplet injectors. Combustion efficiencies above 90% were achieved, with the highest efficiency (95%) for the 55 wt% RP-1/Al. Although engine runs with that high metal loading experienced some agglomeration and erosion difficulties, the 0 and 5 wt% tests ran well, with a high combustion efficiency, and demonstrated a self-protective layer of gelled propellants and combustion products.
References
Lewis contact: Bryan A. Palaszewski, (216) 977-7493
Headquarters program office: OSAT
A brush seal is simply a ring of wire bristles sandwiched between a front and back washer. The bristles usually have a 5- to 10-mil interference fit with the shaft and are installed at a 30deg. to 60deg. angle to the radius so that the bristles can bend as cantilever beams during shaft perturbations. The back washer is on the low-pressure side of the seal and has an inside diameter just slightly larger than the bristle bore diameter. It supports the bristles, preventing them from blowing downstream, and acts as a fixed-clearance seal should the bristles fail.
In a cryogenic turbopump brush seals may be used to seal either liquid hydrogen or liquid oxygen at locations near the pump or the bearings, or they may be used to seal hot gaseous hydrogen, warm gaseous oxygen, or helium at locations near the turbine or purge seals. In this environment large temperature gradients, oxygen compatibility, and hydrogen embrittlement are concerns. Also, the shaft speeds attained in cryogenic turbopumps for rocket engine systems are high, up to 200,000 rpm for a liquid hydrogen turbopump. Because brush seals are contact seals, their wear rates under these conditions are important.
To ensure that long life requirements will be met, several rotor coatings were tested against a brush seal in liquid hydrogen. The coatings tested were zirconium oxide, chromium carbide, and a Teflon-impregnated chromium. A bare Inconel 718 rotor was tested as the baseline material. The 2-in.-diameter bore brush seals had an initial interference with the rotor of 5 mils and were made with 2.8-mil-diameter, Haynes-25 bristles. Testing was done at 35,000 and 65,000 rpm to obtain surface velocities representative of those found in turbopumps for launch vehicles and orbital transfer vehicles, respectively. Pressure drops across the seal were as high as 175 psid. Leakage data were also taken. Although the data are still being analyzed, initial indications are that the bristle material transferred onto the bare Inconel 718 rotor, which did not happen in previous liquid nitrogen tests; that the zirconium oxide coating had the deepest wear groove; and that the chromium carbide and the Teflon-impregnated chromium wore somewhat better.
Further analysis is being done to quantify the wear rate and determine the wear mechanisms and the effects of coating techniques and coating density. The leakage data will be used to update and calibrate an analytical computer code developed at NASA Lewis that models brush seal performance.
Lewis contacts: Margaret P. Proctor, (216) 977-7526;
James F. Walker, (216) 977-7465
Headquarters program office: OSAT
A specific area of increasing interest to users is the impact of electric thruster plumes on spacecraft communications. Until recently, the state of the art in this area was represented by a model developed at NASA Lewis. Intrusive diagnostics to quantify plume electrical characteristics were developed in-house (refs. 1 and 2), and an analytical model was developed under grant at the University of Texas (ref. 3). The UT code used experimental data to generate a plume model and then an analytic approach to determine both beam attenuation and squint for specific cases. Following the first demonstration of this code, General Electric's Astro-Space Division (now Martin Marietta) funded a study to evaluate arcjet plume impacts on communications for the specific case of the 7000 Series satellite. This approach was sufficient for low-power arcjets, which have very tenuous and well-behaved plumes. The simple arcjet model will not, however, be adequate for plumes associated with higher performance electric thrusters now under development for use by industry and in NASA programs.
To address this shortfall, NASA Lewis has developed a large communications impacts test bed (see photograph). In the arrangement shown, transmitting and receiving horns are positioned across the tank, and the thruster is on a movable stand and oriented so that the exhaust plume is in the path of the transmitted signal. Nanosecond gating electronics are used to mitigate effects from signals reflected from the tank walls. Attenuation and phase shift can be monitored with respect to thruster position and operating condition at frequencies to 18 GHz. Other experiments, using a movable receiving antenna, will allow the determination of beam squint. To date, the new test-bed capability has been successfully used to examine the impacts of a NASA 30-cm-diameter xenon ion engine plume on communications signals. A cooperative program to use this new capability to characterize a commercial electric thruster system is being planned.
References
NASA Lewis has evaluated a Raman scattering system that uses fiber optics for species concentration determination. An optical probe head was designed to facilitate remote mounting. A single-mode fiber-optic delivers a 0.5-W green laser beam from an argon-ion laser to the probe, which focuses the beam into a gas sample. A photodiode inside the probe wall measures the laser beam intensity for reference purposes. The same probe collects the light scattered back from the molecules and couples this observed light into a second fiber-optic, which guides the light to the detection equipment, where the light intensity is measured. Calibration charts allow a direct readout of the species concentration.
Initial measurements were performed on a gas mixture in an enclosed vessel at ambient temperature and containing 20% oxygen, 78% nitrogen, and 2% hydrogen. The pressure in the vessel was increased from 0 to 0.1 MPa while the scattered intensities of nitrogen, hydrogen, and oxygen molecules were monitored. A tradeoff existed between the speed and accuracy that could be obtained. In this example, a 30-sec time constant had been set, which caused a significant uncertainty. The lower limit of hydrogen detectability using the nonoptimal apparatus was 1 kPa. An improved, simpler design and upgraded components will greatly enhance the capability of this system. This approach to gas analysis opens exciting new possibilities in many areas of leak detection and process control.
Bibliography
NASA Lewis has established a research effort with two contractors, GENCORP Aerojet Propulsion Division and TRW Space and Technology Division, to develop high-pressure, Earth-storable-liquid-fueled rocket technology for lightweight satellites. Basic data on combustion efficiency and heat transfer were obtained in rocket test-bed hardware in preparation for flight rocket development. TRW tested a 50-lbf-class heat sink engine. The estimated performance was 338 sec at 500-psia chamber pressure and 150:1 area ratio using nitrogen tetroxide (NTO)-hydrazine propellants. Aerojet tested a 100-lbf-class chamber with a projected performance of 334 sec at 250-psia chamber pressure and a 300:1 area ratio also using nitrogen tetroxide-hydrazine propellants. High-pressure engines developed in this program are projected to fit into the tight volume and length constraints of small satellites. TRW estimates that as much as 80 lbm can be saved on the NASA Goddard Space Flight Center/TRW Total Ozone Mapping Spectrometer-Earth Probe mission by using this rocket technology.
Lewis contact: Dr. Steven J. Schneider, (216) 977-7484
Headquarters program office: OSAT
Ground and spaceflight validations of the ion propulsion technology will be conducted under the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) Program. Thirty-centimeter-diameter xenon thrusters have been developed to a high level of maturity to satisfy near-term requirements for planetary and Earth-orbiting satellites. A functional engineering model thruster has undergone diagnostic vibration tests, performance tests using a thrust stand, and plume diagnostic tests. These data have been used to define the design of an engineering model thruster (EMT). The first EMT has been assembled and has completed a short-term design verification test. The EMT is now undergoing a 2000-hr wear test at 2.3 kW in a 0.3-mPa test chamber.
A breadboard power processor is being developed to operate from an 80- to 120-V power bus and provide 0.55 to 2.3 kW to the xenon thruster. A 50-kHz switching frequency was selected in a tradeoff aimed at reduced magnetics mass and high efficiency. The breadboard power processor design incorporates three power supplies for all thruster functions. Dual-use neutralizer and discharge power supplies each provide outputs for both a heater and an anode electrode. The beam/accelerator stage supplies voltages to the ion accelerator grids and also accommodates recovery from high-voltage faults or arcs. The two low-voltage power supplies employ a push-pull design, but the beam/accelerator stage utilizes a full-bridge topology. The breadboard power processor topology, a pathfinder, will be evaluated by NASA and an industrial contractor. A breadboard controller, using a microprocessor, completes the power processor assembly. It provides for startup, throttling, shutdown, and fault recovery and supplies subsystem data.
The low-voltage power supplies have been fabricated and are undergoing functional tests on resistive loads and ion thrusters. The high-voltage supply is being assembled. The breadboard power supplies will be separately integrated with an ion thruster before final tests with a full-up breadboard power processor. Projected specific mass of a flight-packaged power processor is about 5 kg/kW at an input power of 2.5 kW and an efficiency of 0.92.
The NSTAR ground-test program has developed, integrated, and tested three engineering model ion thrusters and two breadboard power processors. Thruster wear tests of 2000, 4000, and 12,000 hr will verify the thruster design, establish wear rates and symptoms of early and random failures, and demonstrate compatibility of the ion thruster and power processor with user requirements.
Bibliography
Lewis contact: James S. Sovey, (216) 977-7454
Headquarters program office: OSAT
To gain an understanding of the percolation process, NASA Lewis (with the support of the AREP team) has conducted a series of percolation cooldowns with an instrumented turbopump at the Lewis Cryogenic Components Laboratory. Tests included cooling the oxidizer side of an instrumented turbopump by percolating with liquid nitrogen; cooling the fuel side of an instrumented turbopump by percolating with liquid hydrogen; studying the effects of the initial turbopump temperature and the supply pressure; and accelerating the cooling process by allowing some of the flow to continue through the turbopump and out a small valve (trickle-flow chilldown).
The cryogen is supplied from a test tank 11 in. in diameter by 8 ft high to simulate the Centaur tanks. A mockup of the Centaur oxygen feedline exits approximately 3 ft up the side of the test tank. A section of quartz glass tube was included in the line mockup so that flow within it could be observed and video taped. Temperature sensors were installed at more than 70 locations, both internal and external to the turbopump. The turbopump was mounted in an enclosure purged with helium to prevent frost formation.
Tests with liquid nitrogen resulted in a cooling time of 3 to 6 min (based on time to reach steady-state temperature). Cooling time decreased with increasing supply pressure and with the use of trickle flow. Observation of the flow through the quartz showed a stratified flow throughout the test without any backflow of liquid. The test with hydrogen was inconclusive because the hydrogen depleted before steady state was achieved. Percolation looks feasible, at least for cooling the liquid oxygen side, but further testing is planned to refine the cooling procedure.
Lewis contact: David J. Chato, (216) 977-7488
Headquarters program office: OSAT
The test system is located at NASA Lewis' Plum Brook Station K-site Facility in Sandusky, Ohio. The system's two parts (see photograph) are the 200-gallon, vacuum-jacketed test dewar (top) and the fluid transfer system (bottom), connected to the test dewar with 4-in. vacuum-jacketed piping. The test dewar is unique in that the test fluid can be viewed through several ports on its lid and sides. The fluid transfer system consists of the vacuum-jacketed transfer lines and an instrumentation box, an evacuated aluminum box with a 2-in.-diameter, 1-ft-long clear test section for flow visualization. Advanced instrumentation to measure flow rates or fluid density can be installed in place of, or in series with, the clear test section.
Other test system capabilities include a 778-ft[3]/min vacuum pump, a vent gas measurement system, a test dewar pressure control system, and a pressurant gas supply system. Fluids available at the facility include liquid hydrogen and nitrogen and small quantities of liquid helium and gaseous helium, hydrogen, and nitrogen. The data acquisition system can monitor 512 channels of instrumentation.
Instrumentation available in the test system includes a continuous liquid-level measurement using a capacitance liquid-level probe, a capacitance densimeter, a silicone diode and type E, T, and J thermocouple temperature sensors, various types of pressure transducer, orifice and thermoconductivity types of flowmeter for measuring gas flow, and several sizes of venturi for measuring liquid flow.
The test system has been operational since April 1994, when the slush hydrogen production optimization (SHPO) experiments began. (Slush hydrogen is composed of liquid and solid hydrogen and is 16% denser than liquid hydrogen.) A new slush hydrogen production technique reduced the previous slush production time by 50%, potentially lowering operating and capital costs of a full-scale slush hydrogen production plant. A new test program, sponsored by Martin Marietta of Denver, involving the chilldown of an RL-10 turbopump assembly is slated to begin in September 1994.
These two test programs have direct impacts on the NASA strategic plan. The SHPO program supports several studies showing that propellant densification (i.e., slush hydrogen) can reduce the weight of single-stage-to-orbit vehicles by 30 to 50%. NASA is seriously considering the use of propellant densification in the design of the next generation of reusable launch vehicles. Also, the industry-sponsored RL-10 turbopump chilldown tests are a perfect example of government/industry partnership.
Lewis contact: Nancy B. McNelis, (216) 977-7474
Headquarters program office: OSAT
The CMG concept is based on the principle that the pressure of a gas changes when its volume changes. In operation the tank volume is changed slightly by an oscillating bellows, and the corresponding change in tank pressure is measured.
Preliminary testing in a normal-gravity environment using a cryogenic simulant has concluded. Results from these tests have led to the development of a cryogenic compression mass gage. This gage was built by Southwest Research Institute under terms of a Space Act Agreement with NASA Lewis. It was tested in a liquid hydrogen dewar by Lewis in their Cryogenic Components Laboratory Cell 7. These tests verified the ability to operate the CMG in an actual cryogenic vessel while obtaining accurate results. Further enhancements to the CMG are currently being evaluated, as well as improvements to the algorithm used to calculate mass. Testing to date has met several critical milestones in proving out the design for use in future space systems.
Lewis contact: John M. Jurns, (216) 977-7416
Headquarters program office: OSAT
NIST's unique apparatus can obtain flow pattern and heat transfer data at mass fluxes from 0.05 to 5 kg/m[2]-sec and heat fluxes from 3 to 300 W/m[2]. The apparatus has a horizontal flow orientation and is used for ground-based experimentation. It has view ports for visual observation of the two-phase flow.
Successfully determining two-phase flow patterns is the first step in correctly modeling the heat transfer to the fluid. A joint research effort with NASA Lewis has obtained some initial experimental results. The work is ongoing. Visual observations and videotapes have been made of the two-phase flow patterns. Preliminary measurements of heat transfer phenomena have also been obtained. These measurements include convective heat transfer coefficients, wall superheat at inception of boiling, and transient wall temperature histories during initial chilldown. This work is precursory to follow-on experiments in low gravity and will guide the development of these experiments.
Lewis contact: Neil T. Van Dresar, (303) 497-7553
Headquarters program office: OSAT
Last updated 1995
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