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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Hydrogen-Oxygen PEM Regenerative Fuel Cell at NASA Glenn Research Center

The NASA Glenn Research Center has constructed a closed-cycle hydrogen-oxygen PEM regenerative fuel cell (RFC) to explore its potential use as an energy storage device for a high altitude solar electric aircraft. Built up over the last 2 years from specialized hardware and off the shelf components the Glenn RFC is a complete "brassboard" energy storage system which includes all the equipment required to (1) absorb electrical power from an outside source and store it as pressurized hydrogen and oxygen and (2) make electrical power from the stored gases, saving the product water for re-use during the next cycle. It consists of a dedicated hydrogen-oxygen fuel cell stack and an electrolyzer stack, the interconnecting plumbing and valves, cooling pumps, water transfer pumps, gas recirculation pumps, phase separators, storage tanks for oxygen (O2) and hydrogen (H2), heat exchangers, isolation valves, pressure regulators, nitrogen purge provisions, instrumentation, and other components. It specific developmental functions include: (1) Test fuel cells and fuel cell components under repeated closed-cycle operation (nothing escapes; everything is used over and over again). (2) Simulate diurnal charge-discharge cycles (3) Observe long-term system performance and identify degradation and loss mechanisms. (4) Develop safe and convenient operation and control strategies leading to the successful development of mission-capable, flight-weight RFC's.

Bents, David J.↗
NASA NTRS2007

The Dynamics of Shock Dispersion and Interactions in Supersonic Freestreams with Counterflowing Jets

An active flow control concept using counterflowing jets to significantly modify the external flowfields and strongly weaken or disperse the shock-waves of supersonic and hypersonic vehicles to reduce the aerothermal loads and wave drag was investigated. Experiments were conducted in a trisonic blow-down wind-tunnel, complemented by pre-test computational fluid dynamics (CFD) analysis of a 2.6% scale model of Apollo capsule, with and without counterflowing jets, in Mach 3.48 and 4.0 freestreams, to assess the potential aerothermal and aerodynamic benefits of this concept. The model was instrumented with heat flux gauges, thermocouples and pressure taps, and employed five counterflowing jet nozzles (three sonic and other two supersonic with design Mach numbers of 2.44 and 2.94) and nozzle exit diameters ranging from 0.25 to 0.5 inch. Schlieren data show that at low jet flow rates of 0.05 and 0.1lb(sub m)/sec, the interactions result in a long penetration mode (LPM) jet, while the short penetration mode (SPM) jet is observed at flow rates greater than 0.1 lb(sub m)/sec., consistent with the pre-test CFD predictions. For the LPM, the jet appears to be nearly fully-expanded, resulting in a very unsteady and oscillatory flow structure in which the bow shock becomes highly dispersed such that it is no longer discernable. Higher speed camera Schlieren data reveal the shock to be dispersed into striations of compression waves, which suddenly coalesce to a weaker bow shock with a larger standoff distance as the flow rate reached a critical value. The pronounced shock dispersion could significantly impact the aerodynamic performance (L/D) and heat flux reduction of spacecraft in atmospheric entry and re-entry, and could also attenuate the entropy layer in hypersonic blunt body flows. For heat transfer, the results show significant reduction in heat flux, even giving negative heat flux for some of the SPM interactions, indicating that the flow wetting the model is cooling, instead of heating the model, which could significantly impact the requirements and design of thermal protection system. These findings strongly suggest that the application of counterflowing jets as active flow control could have strong impact on supersonic and hypersonic vehicle design and performance.

Daso, Endwell O.↗
NASA NTRS2014

Rotating Balances Used for Fluid Pump Testing

Marshall Space Flight Center has developed and demonstrated two direct read force and moment balances for sensing and resolving the hydrodynamic loads on rotating fluid machinery. These rotating balances consist of a series of stainless steel flexures instrumented with semiconductor type, unidirectional strain gauges arranged into six bridges, then sealed and waterproofed, for use fully submerged in degassed water at rotational speeds up to six thousand revolutions per minute. The balances are used to measure the forces and moments due to the onset and presence of cavitation or other hydrodynamic phenomena on subscale replicas of rocket engine turbomachinery, principally axial pumps (inducers) designed specifically to operate in a cavitating environment. The balances are inserted into the drive assembly with power to and signal from the sensors routed through the drive shaft and out through an air-cooled twenty-channel slip ring. High frequency data - balance forces and moments as well as extensive, flush-mounted pressures around the rotating component periphery - are acquired via a high-speed analog to digital data acquisition system while the test rig conditions are varied continuously. The data acquisition and correction process is described, including the in-situ verifications that are performed to quantify and correct for known system effects such as mechanical imbalance, "added mass," buoyancy, mechanical resonance, and electrical bias. Examples of four types of cavitation oscillations for two typical inducers are described in the laboratory (pressure) and rotating (force) frames: 1) attached, symmetric cavitation, 2) rotating cavitation, 3) attached, asymmetric cavitation, and 4) cavitation surge. Rotating and asymmetric cavitation generate a corresponding unbalanced radial force on the rotating assembly while cavitation surge generates an axial force. Attached, symmetric cavitation induces no measurable force. The frequency of the forces can be determined a priori from the pressure environment while the magnitude of the hydrodynamic force is proportional to the pressure unsteadiness.

Skelley, Stephen↗
DOE OSTI2025

Full Rod Length Gamma Spectroscopy Measurements of LWR Used Fuel Rods

Twenty-five high-burnup pressurized water reactor used fuel rods were scanned in the Irradiated Fuels Examination Laboratory at Oak Ridge National Laboratory with a high-resolution gamma spectrometer. A planar high-purity germanium detector was used across an energy range from 60 keV to 3 MeV. Gamma spectra of individual fuel rods were collected in 4.5 mm steps along the entire length of each rod. Additional spectra were collected on select rods in rotational increments of 30 degrees to provide data on the distribution of fission products along the rod perimeter.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS2014

Overview of the 6 Meter HIAD Inflatable Structure and Flexible TPS Static Load Test Series

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). At the beginning of 2014, a 6m HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify the designs structural performance. The 6m HIAD structure was constructed in a stacked toroid configuration using nine inflatable torus segments composed of fiber reinforced thin films, which were joined together using adhesives and high strength textile woven structural straps to help distribute the loads throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for displacement of the inflatable structure as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The centerbody of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to provide many data sets including: deformed shape, shoulder deflection, strap loads, cord loads, inflation pressures, and applied static load.In this overview, the 6m HIAD static load test series will be discussed in detail, including the 6m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally initial results and conclusions from the test series.

inflatable aerodynamic decelerator↗
NASA NTRS2014

Overview of the 6m HIAD Inflatable Structure and Flexible TPS Static Load Test Series

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). At the beginning of 2014, a 6m HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify the designs structural performace. The 6m HIAD structure was constructed in a stacked toroid configuration using nine inflatable torus segments composed of fiber reinforced thin films, which were joined together using adhesives and high strength textile woven structural straps to help distribute the loads throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for displacement of the inflatable structure as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The centerbody of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to provide many data sets including: deformed shape, shoulder deflection, strap loads, cord loads, inflation pressures, and applied static load.In this overview, the 6m HIAD static load test series will be discussed in detail, including the 6m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally initial results and conclusions from the test series.

TPS↗
DOE OSTI2024

Characterization of Intrinsic Fiber Optic Temperature Sensors by In-core and Furnace Testing

Increased research and development in nuclear technology has raised the demand for novel sensors and instrumentation to meet data objectives, and survive in different conditions and environments, beyond conventional light water reactor (LWR) environments. Expediting the deployment of advanced nuclear technologies by developing, demonstrating, and qualifying advanced reactor fuel forms necessitates a deeper understanding of how irradiation affects fuels and materials' performance. To achieve a more comprehensive understanding of fuels and materials performance, researchers require more specialized experiments and measurements. The demand for innovative sensors to support nuclear fuel development arises from the complexity of materials' behavior under irradiation and the challenges of deploying instrumentation in Material Test Reactors (MTRs) for irradiation tests. Additionally, measurements of material properties require integrated measurement systems to characterize thermal properties, mechanical properties, chemistry, and microstructure [1]. Among the potential measurement techniques, optical fiber-based sensors have been identified as potential sensors to measure different physical phenomena such as temperature, strain, pressure, and fluid level. Optical fiber sensors have the capability to provide multi-sensing and multiplexing instrumentation, allowing the measurement of different physical parameters within a single sensor configuration, and transmitting data collected at multiple locations through a single fiber. They offer immunity to electromagnetic interference, electrical passivity, compatibility with various sensing methodologies, and cost-effectiveness. Beyond their widespread use in telecommunications, silica fiber-based instruments are utilized in industrial applications, even at temperatures reaching 300?400°C, such as distributed temperature sensing in oil and gas recovery. The Department of Energy (DOE) is interested in using fiber optics to support fuel cycle development [2]. Fiber optics are an excellent candidate for harsh environment sensing, including sensing at very high temperatures (1900oC for sapphire optical fibers). Distributed fiber optic sensing has been deployed in other harsh environments like coal gasification plants [3]. Distributed strain sensing, which operates similarly to distributed temperature sensing, has been deployed to monitor underground mines and fibers have been imbedded in soil to monitor sink-hole development [4][5]. The application of fiber optic sensors to advanced reactor development is a promising area of research.

46 INSTRUMENTATION RELATED TO NUCLEAR SCIENCE AND ↗
NASA NTRS2011

Solar Dynamics Observatory Guidance, Navigation, and Control System Overview

The Solar Dynamics Observatory (SDO) was designed and built at the Goddard Space Flight Center, launched from Cape Canaveral on February 11, 2010, and reached its final geosynchronous science orbit on March 16, 2010. The purpose of SDO is to observe the Sun and continuously relay data to a dedicated ground station. SDO remains Sun-pointing throughout most of its mission for the instruments to take measurements of the Sun. The SDO attitude control system (ACS) is a single-fault tolerant design. Its fully redundant attitude sensor complement includes sixteen coarse Sun sensors (CSSs), a digital Sun sensor (DSS), three two-axis inertial reference units (IRUs), and two star trackers (STs). The ACS also makes use of the four guide telescopes included as a part of one of the science instruments. Attitude actuation is performed using four reaction wheels assemblies (RWAs) and eight thrusters, with a single main engine used to provide velocity-change thrust for orbit raising. The attitude control software has five nominal control modes, three wheel-based modes and two thruster-based modes. A wheel-based Safehold running in the attitude control electronics box improves the robustness of the system as a whole. All six modes are designed on the same basic proportional-integral-derivative attitude error structure, with more robust modes setting their integral gains to zero. This paper details the final overall design of the SDO guidance, navigation, and control (GN&C) system and how it was used in practice during SDO launch, commissioning, and nominal operations. This overview will include the ACS control modes, attitude determination and sensor calibration, the high gain antenna (HGA) calibration, and jitter mitigation operation. The Solar Dynamics Observatory mission is part of the NASA Living With a Star program, which seeks to understand the changing Sun and its effects on the Solar System, life, and society. To this end, the SDO spacecraft carries three Sun-observing instruments: Helioseismic and Magnetic Imager (HMI), led by Stanford University; Atmospheric Imaging Assembly (AIA), led by Lockheed Martin Space and Astrophysics Laboratory; and Extreme Ultraviolet Variability Experiment (EVE), led by the University of Colorado. The basic mission is to observe the Sun for a very high percentage of the 5-year mission (10-year goal) with long stretches of uninterrupted observations and with constant, high-data-rate transmission to a dedicated ground station to be located in White Sands, New Mexico. These goals guided the design of the spacecraft bus that will carry and service the three-instrument payload. Overarching design goals for the bus are geosynchronous orbit, near-constant Sun observations with the ability to fly through eclipses, and constant HGA contact with the dedicated ground station. A three-axis stabilized ACS is needed both to point at the Sun accurately and to keep the roll about the Sun vector correctly positioned with respect to the solar north pole. This roll control is especially important for the magnetic field imaging of HM I. The mission requirements have several general impacts on the ACS design. Both the AIA and HMI instruments are very sensitive to the blurring caused by jitter. Each has an image stabilization system (ISS) with some ability to filter out high frequency motion, but below the bandwidth of the ISS the control system must compensate for disturbances within the ACS bandwidth or avoid exciting jitter at higher frequencies. Within the ACS bandwidth, the control requirement imposed by AIA is to place the center of the solar disk no more than 2 arc sec, 3 , from a body-defined target based on one of the GTs that accompany the instrument. This body-defined target, called the science reference boresight (SRB), was determined from the postlaunch orientation of the GTs by averaging the bounding telescope boresights for pitch to get a pitch SRB coordinate, and by averaging the bounding boresights for yaw toet the yaw SRB coordinate. The location of this SRB in the 0.5-deg field-of-view for each GT then becomes the central target for each telescope; one GT is selected for use as the ACS controlling guide telescope (CGT) at any given time. Fine Sun-pointing is effected based on this SRB for all three instruments when the Sun is within the linear range of the CGT. In addition to limiting jitter, HMI science requires averaging several observations, making the instrument sensitive to low frequency motion that induces differential motion between each observation. This requires the spacecraft attitude to be stable about the roll axis to approximately 10 arcsec over a ten-minute period. Instrument calibrations require that the spacecraft point the SRB up to 2.5 degrees in pitch and yaw away from the center of the Sun, placing the Sun outside the field-of-view of the guide telescopes. In such instances, when the GTs cannot provide the definitive target for the ACS, on-board attitude determination combined with ephemeris prediction of the Sun direction must provide the definitive target. EVE is capable of observing the Sun with less dependence on attitude control. However, the ground data processing needs for calibrations result in the most strict attitude knowledge requirements for the mission: [35,70,70] arcsec, 3 , of knowledge with respect to the center of the solar disk. In addition to driving the ACS sensor selection, the knowledge requirements, which have their effect primarily during Inertial mode calibrations, drive the accuracy requirements for the solar ephemeris. The need to achieve and maintain geosynchronous orbit (GEO) drove the need for high-efficiency propulsive systems and appropriate attitude control. The main engine provided high specific impulse for the maneuvers to attain GEO, while the smaller ACS thrusters managed the disturbance torques of the larger engine and provided the capability for much smaller adjustment burns on orbit. SDO s large solar profile means that solar radiation pressure is a large torque disturbance, and the momentum buildup from this disturbance and the GEO altitude drives the ACS to use thrusters to manage vehicle momentum. The demanding data capture budget for the mission, however, requires SDO to avoid frequent thruster maneuvers, while concerns about on-orbit jitter restrict the maximum desired wheel speeds desired from the RWAs. The plan for on-orbit wheel speed and momentum management will be discussed as well as what is now being done in operation after the jitter environment was characterized. The SDO ACS hardware complement is single-fault tolerant. Two main processors carry virtually identical copies of the command and data handling and ACS software, and two identical attitude control electronics (ACE) boxes carry Coldfire processors with contingency ACS software and other hardware interface cards; the ACE structure allows reaction wheels to be commanded by the Sun-pointing Safehold independent of the Mil Std 1553 data bus. The sixteen Adcole CSSs are grouped into primary and backup sets of eight sensors, each set providing the ability to calculate a sun vector. Each set of eight eyes provides full 4 -steradian coverage. The Adcole DSS comprises an optics head and a separate electronics box providing a 1553 data interface. The electronics box is mounted inside the Faraday cage created by the spacecraft bus module. The DSS head with its 32- deg square FOV is mounted on the instrument module with its boresight along the spacecraft X axis, nearly aligned with the Sun during observations. Adcole has designed the DSS calibration parameters so that the accuracy is 0.24 arcminutes within 10 deg of the boresight, and diminishes to 3 arcminutes as the Sun moves towards the edges of its FOV . This DSS calibration scheme provides higher accuracy attitude determination over the range of the instrument calibration maneuvers.

Morgenstern, Wendy M.↗
NASA NTRS1950

Calibration of Instruments for Measuring Wind Velocity and Direction

Signal Corps wind equipment AN/GMQ-1 consisting of a 3-cup anemometer and wind vane was calibrated for wind velocities from 1 to 200 miles per hour. Cup-shaft failure prevented calibration at higher wind velocities. The action of the wind vane was checked and found to have very poor directional accuracy below a velocity of 8 miles per hour. After shaft failure was reported to the Signal Corps, the cup rotors were redesigned by strengthening the shafts for better operation at high velocities. The anemometer with the redesigned cup rotors was recalibrated, but cup-shaft failure occurred again at a wind velocity of approximately 220 miles per hour. In the course of this calibration two standard generators were checked for signal output variation, and a wind-speed meter was calibrated for use with each of the redesigned cup rotors. The variation of pressure coefficient with air-flow direction at four orifices on a disk-shaped pitot head was obtained for wind velocities of 37.79 53.6, and 98.9 miles per hour. A pitot-static tube mounted in the nose of a vane was calibrated up to a dynamic pressure of 155 pounds per square foot, or approximately 256 miles per hour,

Vogler, Raymond D.↗
NASA NTRS2019

Application and Development of Atomic Layer Deposition Techniques to Improve Thermo-Optical Coatings for Spacecraft Thermal Control and Advanced Optical Instruments

A key technology development driver in environmental control systems and next generation optics are discussed utilizing thin film development borrowed from the semiconductor industry. The optical and physical properties of spacecraft radiator coatings are dictated by orbital environmental conditions. For example, coatings must adequately dissipate charge buildup when orbital conditions, such as polar, geostationary or gravity neutral, result in surface charging. Current dissipation techniques include depositing a layer of ITO (indium tin oxide) on the radiator surface in a high temperature process. The application of these enhanced coatings must be such that the properties in question are tailored to mission-specific requirements. The multi-billion-dollar semiconductor industry has adopted Atomic Layer Deposition (ALD) for self-assembly and atomic-scale placement. ALD is a cost-effective nanoadditive-manufacturing technique that allows for the conformal coating of substrates with atomic control in a benign temperature and pressure environment. By using ALD, modification of these coatings can be accomplished during coating application preprocessing. The preprocessing is rendered directly on the coating dry pigment before binding. Through the introduction of paired precursor gases, thin films can be deposited on a myriad of substrates ranging from glass, polymers, aerogels, metals, powders, and other high aspect-ratio micro- and nano- structures. By providing atomic-level control, where single layers of atoms can be deposited, the fabrication of metal transparent films, precise nano-laminates, and coatings of nano-channels and pores is achievable. A method has been demonstrated for the ALD of In2O3 and films on a variety of substrates from Si(100) wafers, glass slides, and on Z93P pigments resulting in a direct spaceflight application. Results will be presented that verify the chemical composition of ALD pigments and charge dissipation properties when the pigment goes through its binding and coating process and we present early results of ALD for carbon nanotube formation and encapsulation.

Dwivedi, Vivek H.↗
NASA NTRS1998

Structure and Dynamics of Coronal Plasma

Brief summaries of the four published papers produced within the present performance period of NASA Grant NAGW-4081 are presented. The full text of the papers are appended to the report. The first paper titled "Coronal Structures Observed in X-rays and H-alpa Structures" was published in the Kofu Symposium proceedings. The study analyzes cool and hot behavior of two x-ray events, a small flare and a surge. It was found that a large H-alpha surge appears in x-rays as a very weak event, while a weak H-alpha feature corresponds to the brightest x-ray emission on the disk at the time of the observation. Calculations of the heating necessary to produce these signatures, and implications for the driving and heating mechanisms of flares vs. surges are presented. The second paper "Differential Magnetic Field Shear in an Active Region" has been published in The Astrophysical Journal. The study compared the three dimensional extrapolation of magnetic fields with the observed coronal structure in an active region. Based on the fit between observed coronal structure throughout the volume of the region and the calculated magnetic field configurations, the authors propose a differential magnetic field shear model for this active region. The decreasing field shear in the outer portions of the AR may indicate a continual relaxation of the magnetic field with time, corresponding to a net transport of helicity outward. The third paper "Difficulties in Observing Coronal Structure" has been published in the journal Solar Physics. This paper discusses the evidence that the temperature and density structure of the corona are far more complicated than had previously been thought. The discussion is based on five studies carried out by the group on coronal plasma properties, showing that any one x-ray instrument does see all of the plasma present in the corona, that hot and cool material may appear to be co-spatial at a given location in the corona, and that simple magnetic field extrapolations provide only a poor fit to the observed structure. The fourth paper "Analysis and Comparison of Loop Structures Imaged with NIXT and Yohkoh/SXT" has been published in Astronomy and Astrophysics. This paper analyzes and compares a variety of coronal loops, deriving loop pressure and emission measure from loop models. They are able to determine the volume filling factor in the corona, which is found to be in the range 0.001 to 0.01 for compact loops, and of order 1 for large structures. The small values suggest highly filamented structures, especially at lower temperatures.

Golub, Leon↗
NASA NTRS2002

An Improved Retrieval of Tropospheric Nitrogen Dioxide from GOME

We present a retrieval of tropospheric nitrogen dioxide (NO2) columns from the Global Ozone Monitoring Experiment (GOME) satellite instrument that improves in several ways over previous retrievals, especially in the accounting of Rayleigh and cloud scattering. Slant columns, which are directly fitted without low-pass filtering or spectral smoothing, are corrected for an artificial offset likely induced by spectral structure on the diffuser plate of the GOME instrument. The stratospheric column is determined from NO2 columns over the remote Pacific Ocean to minimize contamination from tropospheric NO2. The air mass factor (AMF) used to convert slant columns to vertical columns is calculated from the integral of the relative vertical NO2 distribution from a global 3-D model of tropospheric chemistry driven by assimilated meteorological data (Global Earth Observing System (GEOS)-CHEM), weighted by altitude dependent scattering weights computed with a radiative transfer model (Linearized Discrete Ordinate Radiative Transfer), using local surface albedos determined from GOME observations at NO2 wavelengths. The AMF calculation accounts for cloud scattering using cloud fraction, cloud top pressure, and cloud optical thickness from a cloud retrieval algorithm (GOME Cloud Retrieval Algorithm). Over continental regions with high surface emissions, clouds decrease the AMT by 20- 30% relative to clear sky. GOME is almost twice as sensitive to tropospheric NO2 columns over ocean than over land. Comparison of the retrieved tropospheric NO2 columns for July 1996 with GEOS-CHEM values tests both the retrieval and the nitrogen oxide radical (NOx) emissions inventories used in GEOS-CHEM. Retrieved tropospheric NO2 columns over the United States, where NOx emissions are particularly well known, are within 18% of GEOS-CHEM columns and are strongly spatially correlated (r = 0.78, n = 288, p less than 0.005). Retrieved columns show more NO2 than GEOS-CHEM columns over the Transvaal region of South Africa and industrial regions of the northeast United States and Europe. They are lower over Houston, India, eastern Asia, and the biomass burning region of central Africa, possibly because of biases from absorbing aerosols.

Martin, Randall V.↗
NASA NTRSDate not supplied

Design and Technology Maturation of the Stratospheric Projectile Experiment of Entry Dynamics

The supersonic and transonic dynamic stability of blunt-body reentry vehicles currently poses large risks in all of NASA’s ongoing entry missions (MSR SRL, MSR EES, and Dragonfly). These projects have allocated millions of dollars to testing and modeling efforts to buy down risk by using the current state-of-the-art (SoA) facilities at NASA’s disposal. While these facilities have heritage in supplying dynamics data to reentry missions, their availability is severely limited – particularly with the high number of concur-rent projects requesting simultaneous testing– and are costly when considering the science density per dollar. None of the current SoA facility methodologies allow the test model to have the dynamics fully develop through a flight relevant free-stream profile and as such require extrapolations with resultant high uncertainties in order to relate the test dynamics to flight expectations. SPEED is a NASA Ames Center Innovation Fund (CIF) project that is developing a highly tailorable and cost-effective test methodology to better assess the dynamic stability of blunt-body reentry vehicles via a stratospheric balloon flight. This is accomplished by dropping a suite of instrumented capsules from a stratospheric balloon to gain a statistically relevant dataset of scaled reentry vehicles in mission relevant free-flight conditions. This presentation will walk through how the test methodology is being implemented specifically for the Mars Sample Return (MSR) Earth Entry System (EES) geometry in an awarded Flight Opportunities Program (FOP) test flight in early CY24. SPEED Application to MSR: SPEED consists of three main mechanical systems: the Drop Platform, the Projectile, and the test Capsule. SPEED is being developed as a set of guidelines and recommendations for how to test with the proposed Concept of Operations (Conops) since the specific design parameters will vary depending on the specific project’s reference trajectory and entry vehicle design. As such, this presentation will walk through the development time-line as shown in Fig. 2. This is meant to serve as a blueprint for further missions as desired. Mechanical and Avionics Design. The SPEED test platform designed for the MSR-EES capsule geometry with nominal entry parameters has the ability to carry 10 Capsules to altitude instrumented with: 1. 3-Axis Accelerometer 2. IMU 3. Gyroscope 4. Magnetometer 5. Pressure Transducer cruciform 6. Uplook and Horizon Cameras To package the avionics/instrumentation suite, the capsule is approximately 1’ in diameter with the Outer Mold Line (OML) centroid-scaled from the full EES design. The internal volume is gutted and custom-shaped to fit the desired instrumentation suite, as well as to allow for the positioning of ballast mass such that the Center of Gravity is analogous to the flight vehicle. All structural components in the Capsule and Projectile are 3D printed, which significantly reduces the cost of each flight unit to around $1500 including all instrumentation, avionics, and structural components. Flight Conops. The test Capsule is accelerated to the desired altitude and Mach number while stowed in the Projectile, a missile-like vehicle consisting of steel ballast in the nose, a low-drag OML, and an Ejection Mechanism to reliably release the Capsule into the free-flow supersonic conditions. For the MSR-EES design, the capsule employs ~3kg of ballast mass at the nose to accelerate the 1.25kg test Capsule to ~Mach 1.7 at 23km altitude. This requires an initial release altitude of 40km, the quoted limit of a 80kg payload by the FOP-contracted balloon provider. Once the Ejection Mechanism avionics detect the proper conditions, the spring-loaded Ejection Mechanism will release and – guided by the sabot – expose the test Capsule to the desired test conditions for ~5 seconds of free-flight in the supersonic/transonic regimes. Dynamics in the subsonic regime will also be captured with the instrumentation suite with post-flight recovery operations aimed at recovering the high-G-load capable SD cards after the planned hard impact landings. Testing and Development: In the few months the SPEED project has worked the development of MSR-EES flight test, the team has performed lab and drone based testing which this presentation will overview. After the first design phase, the team fabricated Engineering Demonstration Units (EDUs) of all subsystems to perform validation testing shown in Fig. 5. After validation was completed on the subsystem level, a drone-drop test was performed at the recreational flight ceiling of 400ft altitude to assess the SPEED systems in a flight environment. Parameters such as in-flight stability, hard impact landing performance, and avionics performance were quantified and qualified. The FY23 CIF will culminate in a helicopter drop test aboard an Air National Guard Blackhawk. This will prepare the team for the CY24 FOP stratospheric balloon flight that should provide the final verification to begin offering the test platform for mission support. Focus of Presentation: This presentation will outline the technology maturation path of the SPEED implementation to the MSR-EES capsule baseline as well as the details regarding the mechanical system, avionics and instrumentation, and flight operations. Note that a complementary presentation is being submitted for a methodology overview of the SPEED test platform, introducing the testing technique and benefits as well as the full application space of the technology.

pitch damping coefficient↗
NASA NTRS1980

Pioneer Venus Sounder Probe Neutral Gas Mass Spectrometer

A neutral gas mass spectrometer was flown to Venus as part of the Pioneer Venus Multiprobe to measure the composition of its lower atmosphere. The instrument, mounted in the Sounder Probe, was activated after the probe entered the top of the atmosphere, and it obtained data during the descent from 62 km to the surface. Atmospheric gases were sampled through a pair of microleaks, the effluent from which was pumped by a combination of ion and getter pumping. A pneumatically operated valve, controlled by the ambient atmospheric pressure, maintained the ion source pressure at a nearly constant value during descent while the atmospheric pressure varied by three orders of magnitude. A single focusing magnetic sector field mass spectrometer with mass resolution sufficient to reasonably separate argon from C3H4 at 40 amu provided the mass analysis and relative abundance measurements. A microprocessor controlled the operation of the mass spectrometer through a highly efficient peak-tip stepping routine and data compression algorithm that effected a scan of the mass spectrum from 1 to 208 amu in 64 sec while requiring an information rate of only 40 bits/sec to return the data to earth. A subscale height altitude resolution was thus obtained. Weight, size, and power requirements were minimized to be consistent with interplanetary flight contraints.

Hoffman, J. H.↗
NASA NTRS1993

A method for the modelling of porous and solid wind tunnel walls in computational fluid dynamics codes

Porous wall wind tunnels have been used for several decades and have proven effective in reducing wall interference effects in both low speed and transonic testing. They allow for testing through Mach 1, reduce blockage effects and reduce shock wave reflections in the test section. Their usefulness in developing computational fluid dynamics (CFD) codes has been limited, however, by the difficulties associated with modelling the effect of a porous wall in CFD codes. Previous approaches to modelling porous wall effects have depended either upon a simplified linear boundary condition, which has proven inadequate, or upon detailed measurements of the normal velocity near the wall, which require extensive wind tunnel time. The current work was initiated in an effort to find a simple, accurate method of modelling a porous wall boundary condition in CFD codes. The development of such a method would allow data from porous wall wind tunnels to be used more readily in validating CFD codes. This would be beneficial when transonic validations are desired, or when large models are used to achieve high Reynolds numbers in testing. A computational and experimental study was undertaken to investigate a new method of modelling solid and porous wall boundary conditions in CFD codes. The method utilized experimental measurements at the walls to develop a flow field solution based on the method of singularities. This flow field solution was then imposed as a pressure boundary condition in a CFD simulation of the internal flow field. The effectiveness of this method in describing the effect of porosity changes on the wall was investigated. Also, the effectiveness of this method when only sparse experimental measurements were available has been investigated. The current work demonstrated this approach for low speed flows and compared the results with experimental data obtained from a heavily instrumented variable porosity test section. The approach developed was simple, computationally inexpensive, and did not require extensive or intrusive measurements of the boundary conditions during the wind tunnel test. It may be applied to both solid and porous wall wind tunnel tests.

Beutner, Thomas John↗
NASA NTRS1995

Design and development of a high-speed bearing test rig

The development of a high-speed test rig, to be used for compiling an experimental data base of bearing signatures for bearings with known faults, is described. This bearing test rig can be adapted to test oil-film bearings as well as rolling element bearings. This is achieved by mounting the test bearing in one of two special test housings, either of which can be mounted onto a common test shaft which can be driven up to 30,000 rpm. The test bearing housing for rolling element bearings can accommodate proximity displacement transducers, accelerometers, thermocouples, and acoustic emission sensors. The test bearing housing for the fluid-film bearings can accommodate the same instrumentation as well as Bourdon tube-type transducers for measuring oil film pressures around the bearing circumference.

Cockburn, J. A.↗
NASA NTRS2004

Exhaust Simulation Testing of a Hypersonic Airbreathing Model at Transonic Speeds

An experimental study was performed to examine jet-effects for an airframe-integrated, scramjet-rocket combined-cycle vehicle configuration at transonic test conditions. This investigation was performed by testing an existing exhaust simulation wind tunnel model, known as Model 5B, in the NASA Langley 16-Ft. Transonic Tunnel. Tests were conducted at freestream Mach numbers from 0.7 to 1.2, at angles of attack from 2 to +14 degrees, and at up to seven nozzle static pressure ratio values for a set of horizontal-tail and body-flap deflections. The model aftbody, horizontal tails, and body flaps were extensively pressure instrumented to provide an understanding of jet-effects and control-surface/plume interactions, as well as for the development of analytical methodologies and calibration of computational fluid dynamic codes to predict this type of flow phenomenon. At all transonic test conditions examined, the exhaust flow at the exit of the internal nozzle was over-expanded, generating an exhaust plume that turned toward the aftbody. Pressure contour plots for the aftbody of Model 5B are presented for freestream transonic Mach numbers of 0.70, 0.95, and 1.20. These pressure data, along with shadowgraph images, indicated the impingement of an internal plume shock and at least one reflected shock onto the aftbody for all transonic conditions tested. These results also provided evidence of the highly three-dimensional nature of the aftbody exhaust flowfield. Parametric testing showed that angle-of-attack, static nozzle pressure ratio, and freestream Mach number all affected the exhaust-plume size, exhaust-flowfield shock structure, and the aftbody-pressure distribution, with Mach number having the largest effect. Integration of the aftbody pressure data showed large variations in the pitching moment throughout the transonic regime.

Huebner, Lawrence D.↗
NASA NTRS1999

CF2 Detection in Radio-Frequency Ar/CHF3 Plasmas by Fourier Transform Infrared Spectroscopy

CFx radicals, in particular CF2, are instrumental in anisotropic etching of SiO2. In order to optimize the CFx radical population in a given process environment, it is imperative that we understand their production mechanism. Towards this goal, we have conducted a series of quantitative measurements of CF2 radicals in low pressure RF plasmas similar to those used in SiO2 etching. In this study, we present preliminary results for Ar/CHF3 plasmas operating at pressures ranging from 10-50 mTorr and powers ranging from 100-500 W in the GEC reference cell, modified for inductive (transformer) coupling. Fourier transform infrared (FTIR) spectroscop) is used to observe the absorption features of the CF2 radical in the 1114 cm-1 and 1096 cm-1 spectral regions. The FTIR spectrometer is equipped with a high-sensitivity mercury cadmium telluride (MCT) detector and has afixed resolution of 0.125 cm- 1. The CF2 concentrations are measured for a range of operating pressures and discharge power levels, and are compared to measurements of the relative CF2 concentrations made by mass spectrometry using the method of appearance potential for radical selectivity.

Kim, J. S.↗