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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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151 records · Page 9
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NASA NTRS2008

Sixty-four-Channel Inline Cable Tester

Faults in wiring are a serious concern for the aerospace and aeronautics (commercial, military, and civil) industries. A number of accidents have occurred because faulty wiring created shorts or opens that resulted in the loss of control of the aircraft or because arcing led to fires and explosions. Some of these accidents have resulted in the massive loss of lives (such as in the TWA Flight 800 accident). Circuits on the Space Shuttle have also failed because of faulty insulation on wiring. STS-93 lost power when a primary power circuit in one engine failed and a second engine had a backup power circuit fault. Cables are usually tested on the ground after the crew reports a fault encountered during flight. Often such failures result from vibration and cannot be replicated while the aircraft is stationary. It is therefore important to monitor faults while the aircraft is in operation, when cables are more likely to fail. Work is in progress to develop a cable fault tester capable of monitoring up to 64 individual wires simultaneously. Faults can be monitored either inline or offline. In the inline mode of operation, the monitoring is performed without disturbing the normal operation of the wires under test. That is, the operations are performed unintrusively and are essentially undetectable for the test signal levels are below the noise floor. A cable can be monitored several times per second in the offline mode and once a second in the inline mode. The 64-channel inline cable tester not only detects the occurrence of a fault, but also determines the type of fault (short/open) and the location of the fault. This will enable the detection of intermittent faults that can be repaired before they become serious problems.

Source record↗
NASA NTRS2011

Metallographic Preparation of Space Shuttle Reaction Control System Thruster Electron Beam Welds for Electron Backscatter Diffraction

A Space Shuttle Reaction Control System (RCS) thruster failed during a firing test at the NASA White Sands Test Facility (WSTF), Las Cruces, New Mexico. The firing test was being conducted to investigate a previous electrical malfunction. A number of cracks were found associated with the fuel closure plate/injector assembly (Fig 1). The firing test failure generated a flight constraint to the launch of STS-133. A team comprised of several NASA centers and other research institutes was assembled to investigate and determine the root cause of the failure. The JSC Materials Evaluation Laboratory was asked to compare and characterize the outboard circumferential electron beam (EB) weld between the fuel closure plate (Titanium 6Al-4V) and the injector (Niobium C-103 alloy) of four different RCS thrusters, including the failed RCS thruster. Several metallographic challenges in grinding/polishing, and particularly in etching were encountered because of the differences in hardness, ductility, and chemical resistance between the two alloys and the bimetallic weld. Segments from each thruster were sectioned from the outboard weld. The segments were hot-compression mounted using a conductive, carbon-filled epoxy. A grinding/polishing procedure for titanium alloys was used [1]. This procedure worked well on the titanium; but a thin, disturbed layer was visible on the niobium surface by means of polarized light. Once polished, each sample was micrographed using bright field, differential interference contrast optical microscopy, and scanning electron microscopy (SEM) using a backscatter electron (BSE) detector. No typical weld anomalies were observed in any of the cross sections. However, areas of large atomic contrast were clearly visible in the weld nugget, particularly along fusion line interfaces between the titanium and the niobium. This prompted the need to better understand the chemistry and microstructure of the weld (Fig 2). Energy Dispersive X-Ray Spectroscopy (EDS) was used to confirm the chemical composition of the variations in contrast in these areas. Niobium alloys generally require exposure to more aggressive chemical reagents than titanium alloys for etching because of niobium s chemical resistance; therefore, the titanium portion of the sample was etched first. A five second immersion in Kroll s reagent revealed a general microstructure on the titanium portion of the sample; however, the titanium heat affected zone closest to the weld, was over-etched due to higher concentrations of refined grains and an increase in eta-phase. The Kroll s etchant also revealed some microstructure in the weld nugget itself; the niobium portion of the sample remained unetched.

Martinez, James↗
NASA NTRS2009

Design Evolution and Verification of the A-3 Chemical Steam Generator

Following is an overview of the Chemical Steam Generator system selected to provide vacuum conditions for a new altitude test facility, the A-3 Test Stand at Stennis Space Center (SSC) in Bay St. Louis, MS. A-3 will serve as NASA s primary facility for altitude testing of the J-2X rocket engine, to be used as the primary propulsion device for the upper stages of the Ares launch vehicles. The Chemical Steam Generators (CSGs) will produce vacuum conditions in the test cell through the production and subsequent supersonic ejection of steam into a diffuser downstream of the J-2X engine nozzle exit. The Chemical Steam Generators chosen have a rich heritage of operation at rocket engine altitude test facilities since the days of the Apollo program and are still in use at NASA White Sands Test Facility (WSTF) in New Mexico. The generators at WSTF have been modified to a degree, but are still very close to the heritage design. The intent for the A-3 implementation is to maintain this heritage design as much as possible, making minimal updates only where necessary to substitute for obsolete parts and to increase reliability. Reliability improvements are especially desired because the proposed system will require 27 generators, which is nine times the largest system installed in the 1960s. Improvements were suggested by the original design firm, Reaction Motors, by NASA SSC and NASA WSTF engineers, and by the A-3 test stand design contractor, Jacobs Technology, Inc. (JTI). This paper describes the range of improvements made to the design to date, starting with the heritage generator and the minor modifications made over time at WSTF, to the modernized configuration which will be used at A-3. The paper will discuss NASA s investment in modifications to SSC s E-2 test facility fire a full-scale Chemical Steam Generator in advance of the larger steam system installation at A-3. Risk mitigation testing will be performed in early 2009 at this test facility to verify that the CSGs operate as expected. The generator which will undergo this testing is of the most recent A-3 configuration, and will be instrumented far in excess of what is normally required for operation. The extra data will allow for easier troubleshooting and more complete knowledge of expected generator performance. In addition, the early testing will give SSC personnel experience in operating the CSG systems, which will expedite the process of installation and activation at A-3. Each Chemical Steam Generator is supported by a complement of valves, instruments, and flow control devices, with the entire assembly called a "module." The generators will be installed in groups of three, historically called "units". A module is so called because of its modular ability to be replaced or serviced without disturbing the other two modules installed on the same unit. A module is pictured in Figure 1, shown with its generator secured by white bands in its shipping (vs. installed) configuration. The heritage system at WSTF is composed of a single unit (three generator modules), pictured in Figure 2 as it was installed in 1965. In contrast, A-3 will have nine units operating in parallel to achieve vacuum conditions appropriate for testing the J-2X engine. Each of the combustors operates in two modes and achieves the so-called "full-steam" mode after all three of its stages ignite. Ignition of the first stage is achieved by exciting a spark plug; the second stage and main stage are lit by the flame front of the previous stage. The main stage burns approximately 97% of the total propellant flow and uses the heat energy to vaporize water into superheated steam. While the main stage remains unlit, the combustor is in so-called "idle" mode. In the WSTF system, this idle mode is not optimized for water usage, and does not need to be, as the water is pumped from a large reservoir. The water supply at A-3 will be contained in tanks with finite volume, so water optimization is preferred for the modnized configuration. Multiple solutions for this issue have been proposed, with the leading concept being a change to the operational definition of "idle mode," with the generator running in a lower heat flux condition.

Kirchner, Casey K.↗
NASA NTRSDate not supplied

Development and Testing of Pulse Guns for Combustion Instability Testing

Combustion stability requirements of a liquid propellant rocket engine are most commonly verified by evaluating the oscillatory pressure or vibration characteristics following an artificial disturbance of the combustion in the combustion chamber. Two types of devices are most commonly used for dynamic stability rating – nondirectional bombs and pulse guns. Bombs have historically been used most often, especially for the large engines developed around the Apollo era. However, bombs in recent years have become much more expensive and difficult to procure, and their transportation and handling requirements have become more demanding. Pulse guns, with inherently less expensive materials, and far less restrictive handling and shipping requirements, have become an attractive alternative. The NASA Marshall Space Flight Center has developed a pulse gun for use during future development programs at the center, and as a design reference for interested commercial users. The objective of this paper is to present the different variants of the pulse gun and characterize performance over a range of parameters. The parameters include the breech diameter and corresponding volume, barrel length, amount of gun powder used, the method of containing the gunpowder within the breech, and the pressure rating of the burst disk. In total, 44 tests were conducted with the pulse gun firing into a test chamber pressurized to 2300 psig with gaseous nitrogen. The best experimental results – including the highest and most consistent peak pressure amplitudes recorded by dynamic pressure sensors in the test chamber – were from a configuration consisting of a 0.40” inner-diameter breech, 15-16 grains of gunpowder wrapped in cigarette paper, and a 24,000 psid burst disk. With this configuration, the zero-to-peak overpressures recorded in the test chamber ranged from about 37% to 58% of the 2300 psig mean pressure, which is adequate for typical combustion stability rating. To complement the data analysis, the acoustics of the pressure waves as they propagated through the pulse gun and into the test chamber were modeled. A more comprehensive final report of this pulse gun development program is planned to be available in the near future.

Pulse Gun↗
NASA NTRS2004

Thin-Filament Pyrometry Developed for Measuring Temperatures in Flames

Many valuable advances in combustion science have come from observations of microgravity flames. This research is contributing to the improved efficiency and reduced emissions of practical combustors and is benefiting terrestrial and spacecraft fire safety. Unfortunately, difficulties associated with microgravity have prevented many types of measurements in microgravity flames. In particular, temperature measurements in flames are extremely important but have been limited in microgravity. A novel method of measuring temperatures in microgravity flames is being developed in-house at the National Center for Microgravity Research and the NASA Glenn Research Center and is described here. Called thin-filament pyrometry, it involves using a camera to determine the local gas temperature from the intensity of inserted fibers glowing in a flame. It is demonstrated here to provide accurate measurements of gas temperatures in a flame simultaneously at many locations. The experiment is shown. The flame is a laminar gas jet diffusion flame fueled by methane (CH4) flowing from a 14-mm round burner at a pressure of 1 atm. A coflowing stream of air is used to prevent flame flicker. Nine glowing fibers are visible. These fibers are made of silicon carbide (SiC) and have a diameter of 15 m (for comparison, the average human hair is 75 m in diameter). Because the fibers are so thin, they do little to disturb the flame and their temperature remains close to that of the local gas. The flame and glowing filaments were imaged with a digital black-and-white video camera. This camera has an imaging area of 1000 by 1000 pixels and a wide dynamic range of 12 bits. The resolution of the camera and optics was 0.1 mm. Optical filters were placed in front of the camera to limit incoming light to 750, 850, 950, and 1050 nm. Temperatures were measured in the same flame in the absence of fibers using 50-m Btype thermocouples. These thermocouples provide very accurate temperatures, but they generally are not useful in microgravity tests because they measure temperature at only one location at a time. Thermocouple measurements at a height of 11 mm above the burner were used to calibrate the thin-filament pyrometry system at all four wavelengths. This calibration was used to perform thin-filament pyrometry at other heights above the burner. One such profile is shown in this graph; this is for a height of 21 mm. The agreement between the pyrometry measurements and thermocouple results at this height is excellent in the range of 1000 to 2000 K, with an estimated uncertainty of 50 K and an estimated upper limit of 2500 K. Neither the thermocouple nor the thin-filament pyrometry temperatures have been corrected for radiation, but the correction is expected to be nearly the same for both methods. We anticipate that thin-filament pyrometry similar to that developed here will become an important diagnostic for studies of microgravity flames owing to its accuracy and its ability to simultaneously measure finely spaced temperatures.

Sunderland, Peter B.↗
NASA NTRS2005

STS-107 Microgravity Environment Summary Report

This summary report presents the results of the processed acceleration data measured aboard the Columbia orbiter during the STS-107 microgravity mission from January 16 to February 1, 2003. Two accelerometer systems were used to measure the acceleration levels due to vehicle and science operations activities that took place during the 16-day mission. Due to lack of precise timeline information regarding some payload's operations, not all of the activities were analyzed for this report. However, a general characterization of the microgravity environment of the Columbia Space Shuttle during the 16-day mission is presented followed by a more specific characterization of the environment for some designated payloads during their operations. Some specific quasi-steady and vibratory microgravity environment characterization analyses were performed for the following payloads: Structure of Flame Balls at Low Lewis-number-2, Laminar Soot Processes-2, Mechanics of Granular Materials-3 and Water Mist Fire-Suppression Experiment. The Physical Science Division of the National Aeronautics and Space Administration sponsors the Orbital Acceleration Research Experiment and the Space Acceleration Measurement System for Free Flyer to support microgravity science experiments, which require microgravity acceleration measurements. On January 16, 2003, both the Orbital Acceleration Research Experiment and the Space Acceleration Measurement System for Free Flyer accelerometer systems were launched on the Columbia Space Transportation System-107 from the Kennedy Space Center. The Orbital Acceleration Research Experiment supported science experiments requiring quasi-steady acceleration measurements, while the Space Acceleration Measurement System for Free Flyer unit supported experiments requiring vibratory acceleration measurement. The Columbia reduced gravity environment analysis presented in this report uses acceleration data collected by these two sets of accelerometer systems: The Orbital Acceleration Research Experiment is a low frequency sensor, which measures acceleration up to 1 Hz, but the 1 Hz acceleration data is trimmean filtered to yield much lower frequency acceleration data up to 0.01 Hz. This filtered data can be mapped to other locations for characterizing the quasi-steady environment for payloads and the vehicle. The Space Acceleration Measurement System for Free Flyer measures vibratory acceleration in the range of 0.01 to 200 Hz at multiple measurement locations. The vibratory acceleration data measured by this system is used to assess the local vibratory environment for payloads as well as to measure the disturbance causes by the vehicle systems, crew exercise devices and payloads operation disturbances. This summary report presents analysis of selected quasi-steady and vibratory activities measured by these two accelerometers during the Columbia 16-day microgravity mission from January 16 to February 1, 2003.

Jules, Kenol↗
NASA NTRS1962

NASA Scout ST-1 Flight-Test Results and Analyses, Launch Operations, and Test Vehicle Description

The first of a series of flight tests for the development of the four-stage, solid-propellant Scout vehicle was conducted at the NASA Wallops Station under the direction of the Langley Research Center. Vehicle designation for the test was NASA Scout ST-1. Performance characteristics of the vehicle and components were recorded during a high-altitude probe mission. Flight-simulation studies are presented and show that the accuracy of the guidance system during the flight was within control-system design specifications. The control system functioned normally during the flight with the exception of an overpowering of the reaction-control roll jets near burnout of the third-stage rocket motor. The resulting roll displacement of the vehicle is shown to have caused the monitor tracking radar which had been erroneously tracking a radar beacon in the vehicle on a side lobe to reorient to the major lobe of the receiving antenna. This tracking switch falsely indicated a violent turning maneuver on the monitor plot board and resulted in a hold-fire decision for the fourth-stage rocket motor. Although data for the final thrusting and coast phase of the flight were not obtained, the majority of the test objectives were achieved. In-flight thrust misalignment angles for the second- and third-stage rocket motors derived from control-system error data and for the first-stage motor determined from flight-simulation studies are presented. All rocket-motor thrust misalignment angles were well within the tolerances used for control-system design. Rocket-motor flight performance is presented, and velocity increments attained from the first three stages substantiated the predicted nominal performance. Operation of the rocket motors was satisfactory with the exception of high-level vibrations which were encountered during third-stage motor burning. Rolling moments which overpowered the reaction-control jets are also attributed to the burning characteristics of the third-stage motor. A discussion of the premature loss of the third-stage heat shield is given and shows that the heat-shield latching mechanism failed from pressure loads as the vehicle entered the transonic speed range. Although venting was provided to relieve the high negative pressures known to exist on the heat shield at these speeds, a field modification of the wiring tunnel had the same effect as opening the inside of the heat shield to ambient pressures. Consequently, the latching mechanism failed from pressure loads which were of about the same magnitude as the latching-mechanism yield loads. Skin temperatures were recorded at several locations on the vehicle and were generally in good agreement with theoretical values. Aerodynamic heating presented no problem during the flight since the maximum temperatures recorded during the flight were only about half the design values because of the high-launch-angle trajectory. Environmental vibrations recorded in the vicinity of the guidance package showed that no significant continuous amplitude levels above the general instrumentation noise level were present during first- and second-stage burning. Large vibration amplitudes were recorded during third-stage burning which coincided with the large roll disturbance experienced by the vehicle near burnout of the third-stage motor.

Scout Launch Vehicle↗