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NASA NTRS2015

Color Changing Hydrogen Sensors

During the Space Shuttle Program, one of the most hazardous operation that occurred was the loading of liquid hydrogen (LH2) during fueling operations of the spacecraft. Due to hydrogen's low explosive limit, any amount leaked could lead to catastrophic event. Hydrogen's chemical properties make it ideal as a rocket fuel; however, the fuel is deemed unsafe for most commercial use because of the inability to easily detect the gas leaking. The increased use of hydrogen over traditional fossil fuels would reduce greenhouse gases and America's dependency on foreign oil. Therefore a technology that would improve safety at NASA and in the commercial sector while creating a new economic sector would have a huge impact to NASA's mission. The Chemochromic Detector for sensing hydrogen gas leakage is a color-changing detector that is useful in any application where it is important to know not only the presence but also the location of the hydrogen gas leak. This technology utilizes a chemochromicpigment and polymer matrix that can be molded or spun into rigid or pliable shapes useable in variable temperature environments including atmospheres of inert gas, hydrogen gas, or mixtures of gases. A change in color of the detector material indicates where gaseous hydrogen leaks are occurring. The irreversible sensor has a dramatic color change from beige to dark grey and remains dark grey after exposure. A reversible pigment changes from white to blue in the presence of hydrogen and reverts back to white in the presence of oxygen. Both versions of the sensor's pigments were comprised of a mixture of a metal oxide substrate and a hydro-chromic compound (i.e., the compound that changed color in the presence of hydrogen) and immediately notified the operator of the presence of low levels of hydrogen. The detector can be used in a variety of formats including paint, tape, caulking, injection molded parts, textiles and fabrics, composites, and films. This technology brings numerous benefits over the traditional hydrogen sensors: The technology has excellent temperature stability (4K to 373 K), it can be used in cryogenic fluid applications, it is easy to apply and remove; it requires no power to operate; it has a quick response time; the leak points can be detected visually or electronically; it is nonhazardous, thus environmentally friendly; it can be reversible or irreversible; it does not require on-site monitoring; has a long shelf life; the detector is very durable; and the technology is inexpensive to manufacture.

Chemochromic sensor↗

NASA NTRS2015

Observation Platform for Dynamic Biomedical and Biotechnology Experiments Using the International Space Station (ISS) Light Microscopy Module (LMM)

Techshot, Inc., has developed an observation platform for the LMM on the ISS that will enable biomedical and biotechnology experiments. The LMM Dynamic Stage consists of an electronics module and the first two of a planned suite of experiment modules. Specimens and reagent solutions can be injected into a small, hollow microscope slide-the heart of the innovation-via a combination of small reservoirs, pumps, and valves. A life science experiment module allows investigators to load up to two different fluids for on-orbit, real-time image cytometry. Fluids can be changed to initiate a process, fix biological samples, or retrieve suspended cells. A colloid science experiment module conducts microparticle and nanoparticle tests for investigation of colloid self-assembly phenomena. This module includes a hollow glass slide and heating elements for the creation of a thermal gradient from one end of the slide to the other. The electronics module supports both experiment modules and contains a unique illuminator/condenser for bright and dark field and phase contrast illumination, power supplies for two piezoelectric pumps, and controller boards for pumps and valves. This observation platform safely contains internal fluids and will greatly accelerate the research and development (R&D) cycle of numerous experiments, products, and services aboard the ISS.

Kurk, Michael A. (Andy)↗

NASA NTRS2015

Observation Platform for Dynamic Biomedical and Biotechnology Experiments Using the International Space Station (ISS) Light Microscopy Module (LMM)

Kurk, Michael A. (Andy)↗

NASA NTRS2018

Hydrovolcanic Astromaterials in the Lab

Zag and Monahans (1998) are H chondrite regolith breccias that contain 4.5 GY old halite crystals which in turn contain abundant inclusions of aqueous fluids, solids and organics. We have previously proposed that these halites originated on a hydrovolcanically-active C class asteroid, probably Ceres, or a trans-neptunian object (TNO - or P- or D-class asteroid) injected into the inner solar system during giant planet migration. We have begun a detailed analysis of organics and other solids trapped within the halite, which we hypothesize sample the mantle of the halite parent object, and are examining a halite-bearing C1 chondrite clast also found in Zag, which is similar to the solids in the halite. These investigations will reveal the water-rock interactions on the hydrovolcanically-active parent world.

Zolensky, M. E.↗

NASA NTRS1987

Liquid Engine Jet Atomization Workshop report

The workshop was oriented to disclose information and unsettled problems to understand the fundamental physical mechanism of the droplet formation process. Based on presentation and discussion of results, recommendations were made which should lead to associated future activities. To accomplish this task, existing observations and experiments, contributing to the basic knowledge, providing data for analytical concept verification, and forming a basis for empirical correlations were solicited. Advanced analytical modeling methods or results from specific studies were requested as well as the experience and advice from injector designers. All effort is directed to advance current analytical techniques, simulating the flow behavior downstream of the injection elements in a liquid rocket combustion chamber. Such a tool can be used to optimize injector designs with respect to short length weight savings, wall material protection, or large heat energy transport to a regenerative cooling fluid, while simultaneously achieving the maximum specific impulse in performance. The liquid atomization process also forms a sound basis for combustion instability analysis.

Gross, Klaus W.↗

NASA NTRS1992

Analysis of a hydrocarbon scramjet with augmented preburning

This paper presents the results of a feasibility study of a hydrocarbon scramjet design utilizing an augmented preburner upstream of the main fuel injector locations. The combustor design evaluated here is for a small hypersonic research vehicle. It consists of a preburner into which a small amount of fuel is burned with on-board liquid oxygen and injected into the main airflow, upstream of the main fuel injector locations, thus ensuring that combustion is present and uninterrupted. Two degrees of analysis are presented including a one-dimensional cycle analysis and a complete computational fluid dynamic analysis with finite-rate chemistry and a two-equation turbulence model. Comparison of these analyses show good agreement when the CFD-predicted fuel consumption schedule is used in the cycle analysis.

Molvik, Gregory A.↗

NASA NTRS2016

Cold Flow Propulsion Test Complex Pulse Testing

When the propellants in a liquid rocket engine burn, the rocket not only launches and moves in space, it causes forces that interact with the vehicle itself. When these interactions occur under specific conditions, the vehicle's structures and components can become unstable. One instability of primary concern is termed pogo (named after the movement of a pogo stick), in which the oscillations (cycling movements) cause large loads, or pressure, against the vehicle, tanks, feedlines, and engine. Marshall Space Flight Center (MSFC) has developed a unique test technology to understand and quantify the complex fluid movements and forces in a liquid rocket engine that contribute strongly to both engine and integrated vehicle performance and stability. This new test technology was established in the MSFC Cold Flow Propulsion Test Complex to allow injection and measurement of scaled propellant flows and measurement of the resulting forces at multiple locations throughout the engine.

McDougal, Kris↗

NASA NTRS1976

Testing of molded high temperature plastic actuator road seals for use in advanced aircraft hydraulic systems

Molded high temperature plastic first and second stage rod seal elements were evaluated in seal assemblies to determine performance characteristics. These characteristics were compared with the performance of machined seal elements. The 6.35 cm second stage Chevron seal assembly was tested using molded Chevrons fabricated from five molding materials. Impulse screening tests conducted over a range of 311 K to 478 K revealed thermal setting deficiencies in the aromatic polyimide molding materials. Seal elements fabricated from aromatic copolyester materials structurally failed during impulse cycle calibration. Endurance testing of 3.85 million cycles at 450 K using MIL-H-83283 fluid showed poorer seal performance with the unfilled aromatic polyimide material than had been attained with seals machined from Vespel SP-21 material. The 6.35 cm first stage step-cut compression loaded seal ring fabricated from copolyester injection molding material failed structurally during impulse cycle calibration. Molding of complex shape rod seals was shown to be a potentially controllable technique, but additional molding material property testing is recommended.

Waterman, A. W.↗

NASA NTRS2005

Coherent Anti-Stokes Raman Spectroscopy (CARS) Measurements in Supersonic Combustors at NASA Langley Research Center

This paper describes the recent use of coherent anti-Stokes Raman spectroscopy (CARS) to study supersonic combustion at NASA Langley Research Center. CARS is a nonlinear optical measurement technique used to measure temperature and species mole fractions remotely in harsh environments. A CARS system has been applied to two different combustor geometries at NASA Langley. Both experiments used the same vitiated wind-tunnel facility to create an air flow that simulates flight at Mach numbers of 6 and 7 for the combustor inlet and both experiments used hydrogen fuel. In the first experiment, the hydrogen was injected supersonically at a 30-degree angle with respect to the incoming flow. In the second experiment, the hydrogen was injected sonically at normal incidence. While these injection schemes produced significantly different flow features, the CARS method provided mean temperature, N2, O2 and H2 maps at multiple downstream locations for both. The primary aim of these measurements was to provide detailed flowfield information for computational fluid dynamics (CFD) code validation.

Danehy, Paul M.↗

NASA NTRS2000

Thermocapillary Flow and Aggregation of Bubbles on a Solid Wall

During the electrolytic evolution of oxygen bubbles forming on a vertically oriented transparent tin oxide electrode, bubbles were found to be mutually attractive. The mechanism of the aggregation had never been explained satisfactorily until Guelcher et al. attributed it to thermocapillary flow. The gradient of surface tension of the liquid at the bubble's surface, which was established because of reaction heat and ohmic heat loss at the electrode wall, drives flow of the liquid adjacent to each bubble; the bubble "pumps" fluid along its surface away from the wall. Fluid flows toward the bubble to conserve mass and entrains nearby bubbles in the flow pattern. The same logic would apply when two bubbles of equal size are adjacent to each other on a warm wall. Each bubble drives thermocapillary flow and hence entrains the other in its flow pattern, which drives the aggregation. Our objective here is to perform experiments where the temperature gradient at the wall is well known and controlled. The theory can be quantitatively tested by studying aggregation of bubble pairs of equal size, and by varying system parameters such as temperature gradient, bubble size and fluid viscosity. The results are then compared with the theory in a quantitatively rigorous manner. We demonstrate that the theory without adjustable parameters is capable of quantitatively modeling the rate of aggregation of two bubbles. The equations governing the thermocapillary flow around a single stationary bubble on a heated or cooled wall in a semi-infinite domain were solved. Both Reynolds number and Marangoni number were much less than unity. The critical result is that liquid in the vicinity of a warm wall flows toward a stationary collector bubble. Consequently the thermocapillary flow around the stationary bubble entrains another bubble toward itself. The bubbles undergo hindered translation parallel to the wall with velocity U while the fluid flow field is described with u. Two velocities were equated by using a wall hindrance parameter q: U = qu [1] which shows the velocity of bubble is proportional to the entraining velocity. The hindrance parameter q can experimentally be measured independently. q can also be calculated by solving the equations of motion for a bubble translating parallel to a solid wall. The experimental cell is cylindrical with an ID of 10 cm and consists of a 1 cm deep main cell filled with silicone oil and flanked by two thermal reservoirs. The upper thermal reservoir was heated and the lower thermal reservoir was cooled so that the bubbles aggregate. Two types of silicone oil (eta = 0.02 and 0.50 Pa s) were used. Two equal sized air bubbles were injected into the cell with a syringe. The center-to-center distance of bubbles was observed through a microscope. Bubble radius ranged from 0.40 mm to 0.65 mm and the temperature gradients along with the cell ranged from 1400 to 5000 K/m. The bubbles aggregated when heat flows from the wall to the fluid. The velocities of bubbles were in the range of 1 - 10 microns/s. The separation r decreased more quickly when the temperature gradient was higher, bubble size was larger, and the oil viscosity was lower. r decreased more rapidly as the bubbles approached each other. Dimensionless time was arbitrarily set to be zero when the dimensionless center-to-center distance between the bubbles was 4. All the bubble trajectories fall onto one line, especially in the range of dimensionless distance from 4 to 3. This means the relative movement of the bubble pair is proportional to the temperature gradient and bubble size and it is inversely proportional to the viscosity of the oil. This result strongly suggests that the thermocapillary flow-based aggregation mechanism is correct. A value of q can be estimated by fitting the scaled data to Eq. [1]. A best fit value of q was obtained as q = 0.26 with a standard deviation of 0.03. Independent experimental results for q for a 0.5 mm radius bubble, give values of q in the range 0.11 to 0.23. The value of q obtained from solving the equations of motion reveals q has values in the range 0.23 - 0.30. Since the full scale of possible values of q is zero to one, the maximum deviation of independently determined values of q from the best fit value was 15% of this full scale. Thus reasonable quantitative agreement between theory and experiment has been obtained.

Kasumi, Hiroki↗

NASA NTRS2014

Liquid Oxygen Thermodynamic Vent System Testing with Helium Pressurization

This report presents the results of several thermodynamic vent system (TVS) tests with liquid oxygen plus a test with liquid nitrogen. In all tests, the liquid was heated above its normal boiling point to 111 K for oxygen and 100 K for nitrogen. The elevated temperature was representative of tank conditions for a candidate lunar lander ascent stage. An initial test series was conducted with saturated oxygen liquid and vapor at 0.6 MPa. The initial series was followed by tests where the test tank was pressurized with gaseous helium to 1.4 to 1.6 MPa. For these tests, the helium mole fraction in the ullage was quite high, about 0.57 to 0.62. TVS behavior is different when helium is present than when helium is absent. The tank pressure becomes the sum of the vapor pressure and the partial pressure of helium. Therefore, tank pressure depends not only on temperature, as is the case for a pure liquid-vapor system, but also on helium density (i.e., the mass of helium divided by the ullage volume). Thus, properly controlling TVS operation is more challenging with helium pressurization than without helium pressurization. When helium was present, the liquid temperature would rise with each successive TVS cycle if tank pressure was kept within a constant control band. Alternatively, if the liquid temperature was maintained within a constant TVS control band, the tank pressure would drop with each TVS cycle. The final test series, which was conducted with liquid nitrogen pressurized with helium, demonstrated simultaneous pressure and temperature control during TVS operation. The simultaneous control was achieved by systematic injection of additional helium during each TVS cycle. Adding helium maintained the helium partial pressure as the liquid volume decreased because of TVS operation. The TVS demonstrations with liquid oxygen pressurized with helium were conducted with three different fluid-mixer configurations-a submerged axial jet mixer, a pair of spray hoops in the tank ullage, and combined use of the axial jet and spray hoops. A submerged liquid pump and compact heat exchanger located inside the test tank were used with all the mixer configurations. The initial series without helium and the final series with liquid nitrogen both used the axial jet mixer. The axial jet configuration successfully demonstrated the ability to control tank pressure; but in the normal-gravity environment, the temperature in the upper tank region (ullage and unwetted wall) was not controlled. The spray hoops and axial jet combination also successfully demonstrated pressure control as well as temperature control of the entire tank and contents. The spray-hoops-only configuration was not expected to be a reliable means of tank mixing because there was no direct means to produce liquid circulation. However, surprisingly good results also were obtained with the sprayhoops- only configuration (i.e., performance metrics such as cycle-averaged vent flowrate were similar to those obtained with the other configurations). A simple thermodynamic model was developed that correctly predicted the TVS behavior (temperature rise or pressure drop per TVS cycle) when helium was present in the ullage. The model predictions were correlated over a range of input parameters. The correlations show that temperature rise or pressure drop per cycle was proportional to both helium mole fraction and tank heat input. The response also depended on the tank fill fraction: the temperature rise or pressure drop (per TVS cycle) increased as the ullage volume decreased.

Thermodynamic vent system↗

NASA NTRS2014

Multi-Point Combustion System: Final Report

A low-NOx emission combustor concept has been developed for NASA's Environmentally Responsible Aircraft (ERA) program to meet N+2 emissions goals for a 70,000 lb thrust engine application. These goals include 75 percent reduction of LTO NOx from CAEP6 standards without increasing CO, UHC, or smoke from that of current state of the art. An additional key factor in this work is to improve lean combustion stability over that of previous work performed on similar technology in the early 2000s. The purpose of this paper is to present the final report for the NASA contract. This work included the design, analysis, and test of a multi-point combustion system. All design work was based on the results of Computational Fluid Dynamics modeling with the end results tested on a medium pressure combustion rig at the UC and a medium pressure combustion rig at GRC. The theories behind the designs, results of analysis, and experimental test data will be discussed in this report. The combustion system consists of five radially staged rows of injectors, where ten small scale injectors are used in place of a single traditional nozzle. Major accomplishments of the current work include the design of a Multipoint Lean Direct Injection (MLDI) array and associated air blast and pilot fuel injectors, which is expected to meet or exceed the goal of a 75 percent reduction in LTO NOx from CAEP6 standards. This design incorporates a reduced number of injectors over previous multipoint designs, simplified and lightweight components, and a very compact combustor section. Additional outcomes of the program are validation that the design of these combustion systems can be aided by the use of Computational Fluid Dynamics to predict and reduce emissions. Furthermore, the staging of fuel through the individually controlled radially staged injector rows successfully demonstrated improved low power operability as well as improvements in emissions over previous multipoint designs. Additional comparison between Jet- A fuel and a hydrotreated biofuel is made to determine viability of the technology for use with alternative fuels. Finally, the operability of the array and associated nozzles proved to be very stable without requiring additional active or passive control systems. A number of publications have been publish

Goeke, Jerry↗

NASA NTRS2000

Mixing and NOx Emission Calculations of Confined Reacting Jet Flows in Cylindrical and Annular Ducts

Rapid mixing of cold lateral jets with hot cross-stream flows in confined configurations is of practical interest in gas turbine combustors as it strongly affects combustor exit temperature quality, and gaseous emissions in for example rich-lean combustion. It is therefore important to further improve our fundamental understanding of the important processes of dilution jet mixing especially when the injected jet mass flow rate exceeds that of the cross-stream. The results reported in this report describe some of the main flow characteristics which develop in the mixing process in a cylindrical duct. A three-dimensional computational fluid dynamics (CFD) code has been used to predict the mixing flow field characteristics and NOx emission in a quench section of a rich-burn/quick-mix/lean-burn (RQL) combustor. Sixty configurations have been analyzed in both circular and annular geometries in a fully reacting environment simulating the operating condition of an actual RQL gas turbine combustion liner. The evaluation matrix was constructed by varying the number of orifices per row and orifice shape. Other parameters such as J (momentum-flux ratio), MR (mass flowrate ratio), DR (density ratio), and mixer sector orifice ACd (effective orifice area) were maintained constant throughout the entire study. The results indicate that the mixing flow field can be correlated with the NOx production if they are referenced with the stoichiometric equivalence ratio value and not the equilibrium value. The mixing flowfields in both circular and annular mixers are different. The penetration of equal jets in both annular and circular geometries is vastly different which significantly affects the performance of the mixing section. In the computational results with the circular mixer, most of the NOx formation occurred behind the orifice starting at the orifice wake region. General trends have been observed in the NOx production as the number of orifices is changed and this appears to be common for all hole configurations and mixer types (circular or annular). The performance of any orifice shape (in producing minimum NOx) appears to be acceptable if the number of orifices can be freely varied in order to attain the optimum jet penetration.

Oechsle, Victor L.↗

NASA NTRS1985

Characterization of flow in a scroll duct

A quantitative, flow visualization study was made of a partially elliptic cross section, inward curving duct (scroll duct), with an axial outflow through a vaneless annular cutlet. The working fluid was water, with a Re(d) of 40,000 at the inlet to the scroll duct, this Reynolds number being representative of the conditions in an actual gas turbine scroll. Both still and high speed moving pictures of fluorescein dye injected into the flow and illuminated by an argon ion laser were used to document the flow. Strong secondary flow, similar to the secondary flow in a pipe bend, was found in the bottom half of the scroll within the first 180 degs of turning. The pressure field set up by the turning duct was strong enough to affect the inlet flow condition. At 90 degs downstream, the large scale secondary flow was found to be oscillatory in nature. The exit flow was nonuniform in the annular exit. By 270 degs downstream, the flow appeared unorganized with no distinctive secondary flow pattern. Large scale structures from the upstream core region appeared by 90 degs and continued through the duct to reenter at the inlet section.

Begg, E. K.↗

NASA NTRS2005

Active control of Boundary Layer Separation & Flow Distortion in Adverse Pressure Gradient Flows via Supersonic Microjets

Inlets to aircraft propulsion systems must supply flow to the compressor with minimal pressure loss, flow distortion or unsteadiness. Flow separation in internal flows such as inlets and ducts in aircraft propulsion systems and external flows such as over aircraft wings, is undesirable as it reduces the overall system performance. The aim of this research has been to understand the nature of separation and more importantly, to explore techniques to actively control this flow separation. In particular, the use of supersonic microjets as a means of controlling boundary layer separation was explored. The geometry used for the early part of this study was a simple diverging Stratford ramp, equipped with arrays of supersonic microjets. Initial results, based on the mean surface pressure distribution, surface flow visualization and Planar Laser Scattering (PLS) indicated a reverse flow region. We implemented supersonic microjets to control this separation and flow visualization results appeared to suggest that microjets have a favorable effect, at least to a certain extent. However, the details of the separated flow field were difficult to determine based on surface pressure distribution, surface flow patterns and PLS alone. It was also difficult to clearly determine the exact influence of the supersonic microjets on this flow. In the latter part of this study, the properties of this flow-field and the effect of supersonic microjets on its behavior were investigated in further detail using 2-component (planar) Particle Image Velocimetry (PIV). The results clearly show that the activation of microjets eliminated flow separation and resulted in a significant increase in the momentum of the fluid near the ramp surface. Also notable is the fact that the gain in momentum due to the elimination of flow separation is at least an order of magnitude larger (two orders of magnitude larger in most cases) than the momentum injected by the microjets and is accomplished with very little mass flow through the microjets.

Alvi, Farrukh S.↗

NASA NTRS2006

Liquid Hydrogen Sensor Considerations for Space Exploration

The on-orbit management of liquid hydrogen planned for the return to the moon will introduce new considerations not encountered in previous missions. This paper identifies critical liquid hydrogen sensing needs from the perspective of reliable on-orbit cryogenic fluid management, and contrasts the fundamental differences in fluid and thermodynamic behavior for ground-based versus on-orbit conditions. Opportunities for advanced sensor development and implementation are explored in the context of critical Exploration Architecture operations such as on-orbit storage, docking, and trans-lunar injection burn. Key sensing needs relative to these operations are also examined, including: liquid/vapor detection, thermodynamic condition monitoring, mass gauging, and leak detection. Finally, operational aspects of an integrated system health management approach are discussed to highlight the potential impact on mission success.

Moran, Matthew E.↗

NASA NTRS2008

Cryogenic Pressure Control Modeling for Ellipsoidal Space Tanks in Reduced Gravity

A computational fluid dynamics (CFD) model is developed to simulate pressure control of an ellipsoidal-shaped liquid hydrogen tank under external heating in low gravity. Pressure control is provided by an axial jet thermodynamic vent system (TVS) centered within the vessel that injects cooler liquid into the tank, mixing the contents and reducing tank pressure. The two-phase cryogenic tank model considers liquid hydrogen in its own vapor with liquid density varying with temperature only and a fully compressible ullage. The axisymmetric model is developed using a custom version of the commercially available FLOW-3D software and simulates low gravity extrapolations of engineering checkout tests performed at Marshall Space Flight Center in 1999 in support of the Solar Thermal Upper Stage Technology Demonstrator (STUSTD) program. Model results illustrate that stable low gravity liquid-gas interfaces are maintained during all phases of the pressure control cycle. Steady and relatively smooth ullage pressurization rates are predicted. This work advances current low gravity CFD modeling capabilities for cryogenic pressure control and aids the development of a low cost CFD-based design process for space hardware.

Hedayat, Ali↗

NASA NTRS2007

Cryogenic Pressure Control Modeling for Ellipsoidal Space Tanks

A computational fluid dynamics (CFD) model is developed to simulate pressure control of an ellipsoidal-shaped liquid hydrogen tank under external heating in normal gravity. Pressure control is provided by an axial jet thermodynamic vent system (TVS) centered within the vessel that injects cooler liquid into the tank, mixing the contents and reducing tank pressure. The two-phase cryogenic tank model considers liquid hydrogen in its own vapor with liquid density varying with temperature only and a fully compressible ullage. The axisymmetric model is developed using a custom version of the commercially available FLOW-31) software. Quantitative model validation is ,provided by engineering checkout tests performed at Marshall Space Flight Center in 1999 in support of the Solar Thermal Upper Stage_ Technology Demonstrator (STUSTD) program. The engineering checkout tests provide cryogenic tank self-pressurization test data at various heat leaks and tank fill levels. The predicted self-pressurization rates, ullage and liquid temperatures at discrete locations within the STUSTD tank are in good agreement with test data. The work presented here advances current CFD modeling capabilities for cryogenic pressure control and helps develop a low cost CFD-based design process for space hardware.

Lopez, Alfredo↗