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Matrix of pressure sensors determines shape and pressure distribution of object in contact with its surface. Output can be used to develop pressure map of objects' surface and displayed as array of alphanumeric symbols on video monitor.
The possibility of stratospheric HN3 condensing out of the gas-phase at low temperatures has become important in the chemical explanations of the rapid loss of Antarctic ozone in spring. Consequently, knowledge about the behavior of the vapor pressures of water and HNO3 over HNO3/water mixtures at stratospheric temperatures is needed to determine if HNO3 could condense, and by how much the HNO3 vapor pressure could be depressed. Laboratory investigations of vapor pressures above HNO3/water mixtures are described. Vapor pressures were initially measured over liquid and frozen bulk mixtures contained in a glass still which was attached to a stainless steel vacuum chamber. The total pressure in the chamber was monitored with a precision pressure sensor, and the vapor pressures of HNO3, water and impurities were analyzed with a mass spectrometer beam system. In a second set of experiments, vapor deposits were made to produce ice and the mono- and trihydrate of HNO3. Further data and conclusions are presented.
Cryogenic temperature sensing was studied using a pressurized fiber Bragg grating (PFBG). The PFBG was obtained by simply applying a small diametric load to a regular fiber Bragg grating (FBG), which was coated with polyimide of a thickness of 11 micrometers. The Bragg wavelength of the PFBG was measured at temperatures from 295 to 4.2 K. A pressure-induced transition occurred at 200 K during the cooling cycle. As a result the temperature sensitivity of the PFBG was found to be nonlinear but reach 24 pm/K below 200 K, more than three times the regular FBG. For the temperature change from 80 K to 10 K, the PFBG has a total Bragg wavelength shift of about 470 pm, 10 times more than the regular FBG. From room temperature to liquid helium temperature the PFBG gives a total wavelength shift of 3.78 nm, compared to the FBG of 1.51 nm. The effect of the coating thickness on the temperature sensitivity of the gratings is also discussed.
Cryogenic temperature sensing was studied using a pressurized fiber Bragg grating (PFBG). The PFBG was obtained by simply applying a small diametric load to a regular fiber Bragg grating (FBG), which was coated with polyimide of a thickness of 11 micrometers. The Bragg wavelength of the PFBG was measured at temperatures from 295 to 4.2 K. A pressure-induced transition occurred at 200 K during the cooling cycle. As a result the temperature sensitivity of the PFBG was found to be nonlinear but reach 24 pm/K below 200 K, more than three times the regular FBG. For the temperature change from 80 K to 10 K, the PFBG has a total Bragg wavelength shift of about 470 pm, 10 times more than the regular FBG. From room temperature to liquid helium temperature the PFBG gives a total wavelength shift of 3.78 nm, compared to the FBG of 1.51 nm. The effect of the coating thickness on the temperature sensitivity of the gratings is also discussed.
The International Space Station (ISS) United States Orbital Segment (USOS) Oxygen Generation System (OGS) has accumulated almost a year of operation at varied oxygen production rates within the US Laboratory Module (LAB) since it was first activated in July 2007. It was operated intermittently through 2009 and 2010, due to filter clogging and acid accumulation in the recirculation loop. Since the installation of a deionizing bed in the recirculation loop in May of 2011 the OGA has been operated continuously. Filters in the recirculation loop have clogged and have been replaced. Hydrogen sensors have drifted apart, and a power failure may have condensed water on a hydrogen sensor. A pump delta pressure sensor failed, and a replacement new spare pump failed to start. Finally, the voltage across the cell stack increased out of tolerance due to cation contamination, and the cell stack was replaced. This paper will discuss the operating experience and characteristics of the OGS, as well as operational issues and their resolution.
This software is used in an automated cryogenic control system developed to monitor and control the operation of small-scale cryocoolers. The system was designed to automate the cryogenically cooled low-noise amplifier system described in "Automated Cryocooler Monitor and Control System" (NPO-47246), NASA Tech Briefs, Vol. 35, No. 5 (May 2011), page 7a. The software contains algorithms necessary to convert non-linear output voltages from the cryogenic diode-type thermometers and vacuum pressure and helium pressure sensors, to temperature and pressure units. The control function algorithms use the monitor data to control the cooler power, vacuum solenoid, vacuum pump, and electrical warm-up heaters. The control algorithms are based on a rule-based system that activates the required device based on the operating mode. The external interface is Web-based. It acts as a Web server, providing pages for monitor, control, and configuration. No client software from the external user is required.
A wind tunnel test was conducted to characterize the aeroacoustic environment of several configurations of the Space Launch System during ascent. The test was conducted in the 11-by-11 foot transonic and 9-by-7 foot supersonic test sections at NASA Ames research center. Throughout this experiment data was collected from several types of instrumentation including: dynamic and steady-state pressure sensors, unsteady and steady pressure sensitive paint, time-resolved shadowgraph and infrared imaging. The following details results and analysis from the time-resolved shadowgraph and infrared imaging data systems. The time-resolved shadowgraph provided a qualitative measurement of the near-field turbulent fluctuations. These results helped provide context to the relative magnitude and frequency content of the fluid-structure-interaction driving the surface pressure phenomena characterized by the discrete pressure transducers and unsteady pressure sensitive paint. The infrared imaging was used to verify boundary layer trip effectiveness and provide temperature correction for the unsteady pressure sensitive paint.
A system for performing distributed measurements of in-situ stress includes an expandable element with at least one fiber optic sensor. The expandable element can be positioned at various depths in a hole in a substrate. A pressurizing device expands (and contracts) the expandable element when the expandable element is inserted in the hole in the substrate to exert pressure on the hole wall. A pressure sensor provides a sensor output indicative of a pressure applied to the hole wall by the expandable element. The fiber optic sensor and an optical interrogator measure strain along a length of the sensor in a continuous, high spatial resolution manner Based on the measured strain and pressure sensor output, the system determines various properties of the substrate such as, minimum principal stress, maximum principal stress, and/or principal stress direction associated with one or more fractures in the substrate, as well as substrate modulus.
Optical fiber-based sensors are commonly used for industrial monitoring and control processes due to their small size, resistance to electromagnetic interference, and ability to perform a wide variety of high-precision measurements. However, implementing optical fiber sensors in harsh environments is challenging because they are small and fragile. Proper packaging of fiber-optic sensors could extend their use to harsh environments, including at high temperature and under high radiation. Furthermore, conventional fiber optic-based measurement systems often use computationally expensive signal processing algorithms which hinder their use in high-frequency dynamic applications. This work reports on the design of an optical fiber-pressure sensor system based on low-coherence interferometry that uses a metal-embedded optical fiber to provide a robust sensor package. Additionally, a novel phase demodulation scheme is proposed to extract length changes from the deformation of a thin diaphragm within a Fabry–Pérot cavity to measure pressure. The sensor has been tested to ±100 kPa, has a theoretical linear response over a 270 kPa dynamic range (24 dB maximum signal-to-noise ratio), and can resolve pressure transients up to 3910 kPa/s. Sampling at 100 kHz, the present sensor can resolve 2 kPa dynamic pressures at frequencies up to 2 kHz. Faster transients on the order of tens to hundreds of kHz can theoretically be resolved at the expense of decreasing the maximum resolvable amplitude. A methodology for designing a LCI pressure sensor for a given application is outlined based on the limitations imposed by the Nyquist criterion, the diaphragm's resonant frequency, the LCI optoelectronics, and the phase demodulation scheme. The sensor is the first to implement a low-coherence light source and a Fabry–Pérot interferometer designed to provide real-time high-fidelity pressure measurements using a metal-embedded optical fiber. The demonstrated sensor provides a platform for sensing in harsh conditions such as in nuclear and energy applications.
To test liquid rocket engines (LREs) for combustion instabilities, devices such as bombs are often used to create pressure wave disturbances. Bombs, while effective, are inherently dangerous, expensive, and difficult to procure. Over the years, pulse guns have been used as a safer and more cost-effective way to generate controlled pressure disturbances in engine chambers. In anticipation of the need for stability testing of prototype LREs at NASA Marshall Space Flight Center (MSFC), a set of pulse guns have been designed, fabricated, tested, and characterized. The pulse gun program is funded by the RS-25 Engine Program managed out of MSFC and funded by NASA’s Space Launch System (SLS) through the MSFC Liquid Engines Office (LEO). A pulse gun is a simple device – like a traditional gun, it has a breech and barrel section. However, unlike a traditional gun, there is no bullet, as the purpose of the pulse gun is strictly to fire a high pressure pulse. Instead of a firing pin and primer that would normally be used in a traditional gun, an initiator is used to activate the gun powder. The initiator is a highly reliable pyrotechnic initiator. For this study, clone versions of the NASA Standard Initiator (NSI) were used. The initiator is used to ignite a pre-measured amount of gun powder loaded into the breech. When the pressure of the burnt gun powder exceeds the set point of a downstream burst disk, the disk ruptures, allowing the high pressure pulse to travel downstream through the barrel section. A ballistic pressure transducer located in the breech section is used to measure the short duration, high pressure pulse. Some configurations of the pulse gun have barrel sections that include one or more additional ballistic pressure transducers. These additional pressure measurements help track the degradation and damping of the pulse as it travels out of the barrel section. The measurements may also be used to determine the velocity of pulse. The objective of this paper is to present the different variants of this newly-developed pulse gun and characterize performance over a range of parameters. The parameters include breech diameter, barrel length, amount of gun powder used, the way the gun powder is packed, and the pressure setting of the burst disk. There are a total of six pulse guns: three with the 0.25 inch breech and three different length barrels, and three with the 0.40 inch breech and three different length barrels. For both breech sizes, the associated barrels are referred to as standard, one-port, and two-port barrels. The standard barrel has no instrumentation and is likely the barrel that will be used for engine stability testing. The one- and two-port barrels were designed specifically for pulse gun component testing to allow measurements of the magnitude and timing of the high pressure pulse as it makes its way through the pulse gun. The burst disks tested were commercially-available burst disks designed to rupture at 8,000, 16,000, and 24,000 psid. Testing was accomplished by firing the pulse gun into a test chamber pressurized with nitrogen at about 2300 psig. A total of four Model 113B23 High frequency ICP® pressure sensors (10k psi) were mounted in the test chamber, in the same plane as the pulse gun. Two of the sensors had “trimmed” adapters, and two did not. Given the symmetrical configuration of the sensors within the test chamber, different amplitudes of pressure measurements are attributed to the use of trimmed versus untrimmed adapters. The untrimmed adapters, with their narrower passages, tended to amplify the pressure amplitudes by as much as 50%. In total, 41 pulse gun tests have been conducted.Data are still being analyzed, but some trends are apparent. For example, measurements taken within the pulse gun are shown in Figures 3 and 4 for Hot-fires (HFs) # 11 and 39, respectively. Both tests were identical in that the 0.40 inch ID breech, two-port barrel, 8,000 psid burst disk, and same amount of gunpowder (6.639 grains for HF# 11, and 6.576 grains for HF# 39) were used, and the back pressure in the nitrogen test chamber was ~2300 psig. Despite keeping all these variables constant, the results from these two tests look quite different. For both tests, the first pressure peak shown in the P2111 trace within the breech is the firing of the initiator. The second peak and any subsequent peaks are from the combustion of the gunpowder. It should be noted that because the sensors are dynamic, the ~2300 psig baseline static pressure is not shown in the figures. For HF# 11, the burst disk clearly ruptured in a little over 0.1 ms from the time the initiator was fired. This is evident from the abrupt rise in pressure for P2112 followed by another abrupt rise in pressure for P2113. The sensors track the movement of the high pressure pulse through the pulse gun. For HF# 39, however, there was nearly 1 ms delay between the initiator firing and the burst disk rupturing. The only difference between these two tests was the way in which the gunpowder was packed within the breech. For HF# 11, the gunpowder was poured directly into the breech, atop the initiator, and held in place with a vegetable fiber wad appropriately sized for the barrel diameter. Once in place, the wad firmly held the gunpowder in place. For HF# 39, the gunpowder was rolled within cigarette paper with the ends of the paper twisted. The rolled gunpowder was gently pushed down into the breech, toward the initiator, and no wad was used. This method of loading the gunpowder consistently produced longer delays in the rupture of the burst disk, presumably because there was a steady, fuller burning of the gunpowder. The burst disk was rated for 8,000 psid, which means that the burst disk should not have opened until the pressure in the breech reached ~10,300 since there was ~2300 psig back pressure in the test chamber. In the case of HF# 11, the breech pressure only reached roughly 4200 psi before there was leakage either through or past the burst disk, but for HF#39, the breech pressure reached 10,600 psi prior to the burst disk opening. The temperature of the burst disk is certainly a factor in determining when the burst disk will rupture. A sharp rise in the burst disk temperature could weaken it and cause it to stray from the designed set point burst pressure. The burst disks are manufactured of Inconel to try to reduce their sensitivity to temperature, but how the burst disks perform above 900°F is not documented. All burst disks were engineered and tested by the manufacturer for ambient temperature conditions. One theory for why rolling the gunpowder in cigarette paper may produce better results is related to a more controlled temperature environment. The cigarette paper may act as a sheath that protects the burst disk from excessive temperatures during the rapid buildup of pressure in the breech. The ultimate benefit of not breeching the burst disk prematurely is shown in Figures 5 and 6, which graphs the resultant pressure pulse produced in the test chamber. For HF #11, the peak pressure on the P2123 trimmed adapter is about 270 psi, while for HF# 39 it is 343 psi, which is about 1.27 times greater in magnitude. In either case, the pressure pulse is ~20 μs in width once it reaches the test chamber. Another pressure rise about 100 μs later can be seen on the sensors located across from the pulse gun port, P2124 and P2121. These pressures are lower due to attenuation of the pressure wave as it moves across the test chamber. After the wave encounters the wall, it reflects back and forth within the chamber until it completely dissipates.
A test campaign was conducted in the open-jet test section of the NASA Langley 14- by 22-Foot Subsonic Tunnel on a 10%-scale, semispan version of a High-Lift Common Research Model (CRM-HL) incorporating a leading-edge slat, trailing-edge flap and removable high-fidelity main landing gear. Aerodynamic and far field acoustic measurements were obtained on baseline and acoustically treated model configurations using an ensemble of time-averaged and unsteady surface pressure sensors paired with a traversing 97-microphone phased array viewing the pressure side of the airframe. A modified version of the DAMAS deconvolution method incorporating corrections for shear layer acoustic wave decorrelation was employed to determine the locations and strengths of relevant noise sources along the span of the slat and in the vicinity of the flap edges and landing gear. The main goal of the test campaign was to use the surface pressure and array data to evaluate the effectiveness of various slat noise reduction concepts. It was found that slat-gap fillers produced significant broadband noise reduction with minimal aerodynamic performance penalties for all Mach numbers and airfoil angles of attack that were investigated. The test campaign clearly demonstrated the noise reduction benefits that can be obtained by applying appropriate treatments to leading-edge slats on commercial transport-class aircraft.
A multi-scale representation of the combustor dynamics was used to create a self-tuning, scalable controller to suppress multiple instability modes in a liquid-fueled aero engine-derived combustor operating at engine-like conditions. Its self-tuning features designed to handle the uncertainties in the combustor dynamics and time-delays are essential for control performance and robustness. The controller was implemented to modulate a high-frequency fuel valve with feedback from dynamic pressure sensors. This scalable algorithm suppressed pressure oscillations of different instability modes by as much as 90 percent without the peak-splitting effect. The self-tuning logic guided the adjustment of controller parameters and converged quickly toward phase-lock for optimal suppression of the instabilities. The forced-response characteristics of the control model compare well with those of the test rig on both the frequency-domain and the time-domain.
ISHM capability includes: detection of anomalies, diagnosis of causes of anomalies, prediction of future anomalies, and user interfaces that enable integrated awareness (past, present, and future) by users. This is achieved by focused management of data, information and knowledge (DIaK) that will likely be distributed across networks. Management of DIaK implies storage, sharing (timely availability), maintaining, evolving, and processing. Processing of DIaK encapsulates strategies, methodologies, algorithms, etc. focused on achieving high ISHM Functional Capability Level (FCL). High FCL means a high degree of success in detecting anomalies, diagnosing causes, predicting future anomalies, and enabling health integrated awareness by the user. A model that enables ISHM capability, and hence, DIaK management, is denominated the ISHM Model of the System (IMS). We describe aspects of the IMS that focus on processing of DIaK. Strategies, methodologies, and algorithms require proper context. We describe an approach to define and use contexts, implementation in an object-oriented software environment (G2), and validation using actual test data from a methane thruster test program at NASA SSC. Context is linked to existence of relationships among elements of a system. For example, the context to use a strategy to detect leak is to identify closed subsystems (e.g. bounded by closed valves and by tanks) that include pressure sensors, and check if the pressure is changing. We call these subsystems Pressurizable Subsystems. If pressure changes are detected, then all members of the closed subsystem become suspect of leakage. In this case, the context is defined by identifying a subsystem that is suitable for applying a strategy. Contexts are defined in many ways. Often, a context is defined by relationships of function (e.g. liquid flow, maintaining pressure, etc.), form (e.g. part of the same component, connected to other components, etc.), or space (e.g. physically close, touching the same common element, etc.). The context might be defined dynamically (if conditions for the context appear and disappear dynamically) or statically. Although this approach is akin to case-based reasoning, we are implementing it using a software environment that embodies tools to define and manage relationships (of any nature) among objects in a very intuitive manner. Context for higher level inferences (that use detected anomalies or events), primarily for diagnosis and prognosis, are related to causal relationships. This is useful to develop root-cause analysis trees showing an event linked to its possible causes and effects. The innovation pertaining to RCA trees encompasses use of previously defined subsystems as well as individual elements in the tree. This approach allows more powerful implementations of RCA capability in object-oriented environments. For example, if a pressurizable subsystem is leaking, its root-cause representation within an RCA tree will show that the cause is that all elements of that subsystem are suspect of leak. Such a tree would apply to all instances of leak-events detected and all elements in all pressurizable subsystems in the system. Example subsystems in our environment to build IMS include: Pressurizable Subsystem, Fluid-Fill Subsystem, Flow-Thru-Valve Subsystem, and Fluid Supply Subsystem. The software environment for IMS is designed to potentially allow definition of any relationship suitable to create a context to achieve ISHM capability.
A series of wind tunnel tests were conducted to characterize the force-and-moment, and aeroacoustic environment of several configurations of the Space Launch System during ascent. The tests were conducted in the 11-by-11 foot transonic and 9-by-7 foot supersonic test sections at NASA Ames research center. Throughout these experiments data was collected from several types of instrumentation including: multicomponent force-and-moment strain gage balances, dynamic and steady-state pressure sensors, unsteady and steady pressure-sensitive paint, time-resolved shadowgraph and infrared imaging. The following details results and analysis from the time-resolved shadowgraph and infrared imaging data systems. The time-resolved shadowgraph and infrared imaging provided a qualitative measurement of the near-field turbulent fluctuations. These results helped provide context to the relative magnitude and frequency content of the fluid-structure-interaction driving the surface pressure phenomena characterized by the discrete pressure transducers and unsteady pressure sensitive paint.
Techniques for hydrogen sensing and mitigation are provided. As one example, a device includes a chamber and a membrane that is permeable to a first gas and is impermeable to a second gas. The membrane separates the chamber from a gas mixture that contains the first gas, such that the first gas in the gas mixture can only enter the chamber via the membrane. The device also includes a pressure sensor configured to measure a pressure within the chamber.
Lean combustion concepts for aircraft engine combustors are prone to combustion instabilities. Mitigation of instabilities is an enabling technology for these low-emissions combustors. NASA Glenn Research Center s prior activity has demonstrated active control to suppress a high-frequency combustion instability in a combustor rig designed to emulate an actual aircraft engine instability experience with a conventional, rich-front-end combustor. The current effort is developing further understanding of the problem specifically as applied to future lean-burning, very low-emissions combustors. A prototype advanced, low-emissions aircraft engine combustor with a combustion instability has been identified and previous work has characterized the dynamic behavior of that combustor prototype. The combustor exhibits thermoacoustic instabilities that are related to increasing fuel flow and that potentially prevent full-power operation. A simplified, non-linear oscillator model and a more physics-based sectored 1-D dynamic model have been developed to capture the combustor prototype s instability behavior. Utilizing these models, the NASA Adaptive Sliding Phasor Average Control (ASPAC) instability control method has been updated for the low-emissions combustor prototype. Active combustion instability suppression using the ASPAC control method has been demonstrated experimentally with this combustor prototype in a NASA combustion test cell operating at engine pressures, temperatures, and flows. A high-frequency fuel valve was utilized to perturb the combustor fuel flow. Successful instability suppression was shown using a dynamic pressure sensor in the combustor for controller feedback. Instability control was also shown with a pressure feedback sensor in the lower temperature region upstream of the combustor. It was also demonstrated that the controller can prevent the instability from occurring while combustor operation was transitioning from a stable, low-power condition to a normally unstable high-power condition, thus enabling the high-power condition.
As an alternative to conventional tubing instrumentation for measuring airflow, designers and technicians at the NASA Glenn Research Center have been fabricating packaging components and assembling a set of unique probes using commercial-off-the-shelf microelectromechanical systems (MEMS) integrated circuits (computer chips). Using MEMS as an alternative has some compelling advantages over standard measurement devices. Sensor technologies have matured through high-production usage in industries such as automotive and aircraft manufacturers. Currently, MEMS are the choice in applications such as tire pressure monitors, altimeters, pneumatic controls, cable leak detectors, and consumer appliances. Conventional instrumentation uses tubing buried in the model aerodynamic surfaces or wind tunnel walls. The measurements are made when pressure is introduced at the tube opening. The pressure then must travel the tubing for lengths ranging from 20 to hundreds of feet before reaching an electronic signal conditioner. This condition causes a considerable amount of damping and requires measurements to be made only after the test rig has reached steady-state operation. The electronic MEMS pressure sensor is able to take readings continuously under dynamic states in nearly real time. The use of stainless steel tubing for pressure measurements requires many tubes to be cleaned, cut to length, carefully installed, and delicately deburred and spliced for use. A cluster of a few hundred 1/16-in.- (0.0625-in.-) diameter tubes (not uncommon in research testing facilities) can be several inches in diameter and may weigh enough to require two men to handle. Replacing hard tubing with electronic chips can eliminate much of the bulk. Each sensor would fit on the tip of the 1/16-in. tubing with room to spare. The P592 piezoresistive silicon pressure sensor (Lucas NovaSensor, Fremont, CA) was chosen for this project because of its cost, availability, and tolerance to extreme ambient conditions. The chip is 1 mm square by 0.6 mm thick (0.039 by 0.023 in.) with 0.12-mm (0.005-in.) wire connection tabs. Three MEMS chips were used to build the first type of flow-angularity probe. This MEMS probe will be demonstrated as an alternative to a standard tube type "Cobra Probe" now used routinely in wind tunnel and aeronautical hardware applications. The response time and accuracy would allow the probe to be translated on an actuator across a flow field, yielding precision dynamic measurements not possible with conventional instrumentation. The low profile, the minimal power requirement, the rugged construction, and the moderate cost all contribute to making MEMS sensors the enticing choice instrument in future research measurement needs. The MEMS probe efforts are a continuation of work initiated by Brian Willis, without whose foresight and efforts this project would never have been realized. This task was funded through cooperation with the NASA Electronic Parts and Packaging (NEPP) program at the Jet Propulsion Laboratory.
Duty cycle of switched, constant current circuit controlled by pressure sensor provides oxygen at a constant pressure from a cell which electrolyzes water, eliminating possibility of cell damage by excessive oxygen demand.