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Many studies have been performed to quantify the formation and evolution of roughness on ice shapes created in Appendix C icing conditions, which exhibits supercooled liquid droplets ranging from 1-50 µm. For example Anderson and Shin (1997), Anderson et al. (1998), and Shin (1994) represent early studies of ice roughness during short-duration icing events measured in the Icing Research Tunnel at the NASA Glenn Research Center. In the historical literature, image analysis techniques were employed to characterize the roughness. Using multiple images of the roughness elements, these studies of roughness focused on extracting parametric representations of ice roughness elements. While the image analysis approach enabled many insights into icing physics, recent improvements in laser scanning approaches have revolutionized the process of ice accretion shape characterization.

Ice accretions that have formed inside gas turbine engines as a result of flight in clouds of high concentrations of ice crystals in the atmosphere have recently been identified as an aviation safety hazard. NASA's Aviation Safety Program (AvSP) has made plans to conduct research in this area to address the hazard. This paper gives an overview of NASA's engine ice-crystal icing research project plans. Included are the rationale, approach, and details of various aspects of NASA's research.

The International Space Station (ISS) is a complex spacecraft that will take several years to assemble in orbit. During many of the assembly and maintenance procedures, the space station’s large solar arrays must be locked, which can significantly reduce power generation. To date, power generation analyses have not included power generation from the backside of the solar cells in a desire to produce a conservative analysis. This paper describes the testing of ISS solar cell backside power generation, analytical modeling, and analysis results on an ISS assembly mission.

Increased nickel concentrations in the IATCS coolant prompted a study of the corrosion rates of nickel-brazed heat exchangers in the system. The testing has shown that corrosion is occurring in a silicon-rich intermetallic phase in the braze filler of coldplates and heat exchangers as the result of a decrease in the coolant pH brought about by cabin carbon dioxide permeation through polymeric flexhoses. Similar corrosion is occurring in the EMU de-ionized water loop. Certain heat exchangers and coldplates have more silicon-rich phase because of their manufacturing method, and those units produce more nickel corrosion product. Silver biocide additions did not induce pitting corrosion at silver precipitate sites.

This paper describes the DC bus regulation control algorithm for the NASA flywheel energy storage system during charge, charge reduction and discharge modes of operation. The algorithm was experimentally verified in [1] and this paper presents the necessary models for simulation. Detailed block diagrams of the controller algorithm are given. It is shown that the flywheel system and the controller can be modeled in three levels of detail depending on the type of analysis required. The three models are explained and then compared using simulation results.

The history and current status of spacecraft smoke detection is discussed including a review of the state of understanding of the effect of gravity on the resultant smoke particle size. The results from a spacecraft experiment (Comparative Soot Diagnostics (CSD)) which measured microgravity smoke particle sizes are presented. Five different materials were tested producing smokes with different properties including solid aerosol smokes and liquid droplets aerosol smokes. The particulate size distribution for the solid particulate smokes increased substantially in microgravity and the results suggested a corresponding increase for the smokes consisting of a liquid aerosol. A planned follow on experiment that will resolve the issues raised by CSD is presented. Early results from this effort have provided the first measurements of the ambient aerosol environment on the ISS (International Space Station) and suggest that the ISS has very low ambient particle levels.

Under a NASA sponsored technology development activity, Hamilton Sundstrand has designed, fabricated, tested and delivered a prototype solid amine-based carbon dioxide (CO2) and water (H2O) vapor removal system sized for Extravehicular Activity (EVA) operation. The prototype system employs two alternating and thermally-linked solid amine sorbent beds to continuously remove CO2 and H2O vapor from a closed environment. While one sorbent bed is exposed to the vent loop to remove CO2 and water vapor, the other bed is exposed to a regeneration circuit, defined as either vacuum or an inert sweep gas stream. A linear spool valve, coupled directly to the amine canister assembly, is utilized to simultaneously divert the vent loop flow and regeneration circuit flow between the two sorbent beds.

Under a cooperative agreement with NASA, Hamilton Sundstrand has successfully designed, fabricated, tested and delivered three, state-of-the-art, solid amine prototype systems capable of continuous CO2 and humidity removal from a closed, habitable atmosphere. Two prototype systems (CAMRAS #1 and #2) incorporated a linear spool valve design for process flow control through the sorbent beds, with the third system (CAMRAS #3) employing a rotary valve assembly that improves system fluid interfaces and regeneration capabilities. The operational performance of CAMRAS #1 and #2 has been validated in a relevant environment, through both simulated human metabolic loads in a closed chamber and through human subject testing in a closed environment.

Special coatings are being developed and tested to contend with the effects of dust on the lunar surface. These coatings will have wide applicability ranging from prevention of dust buildup on solar arrays and radiator surfaces to protection of EVA space suit fabrics and visors. They will be required to be durable and functional based on application. We have started preparing abrasion-resistant transparent conductive coatings ∼40 nm thick were formed by co-deposition of titanium dioxide (TiO2) and titanium (Ti) on room-temperature glass and polycarbonate substrates using two RF magnetron sputtering sources. By adjusting Ti content, we obtained sheet resistivities in the range 104-1010 ohms/square. We have also started conducting a series of environmental tests that simulate the exposure of coated samples to dust under relevant conditions, beginning with abrasion tests using regolith simulant materials.

A High Heat Flux Demonstration Program has been initiated to investigate and demonstrate the performance of a number of candidate cooling technologies to address the need of dissipating the large thermal loads and high heat fluxes associated with Directed Energy Weapons (DEW) systems. The technologies selected for these investigations utilize both single-phase and two-phase cooling concepts. The single-phase devices investigated are based upon the concept of jet impingement with and without extended surface areas. The two-phase devices investigated extend the jet impingement concepts into the liquid-vapor phase change regime, as well as a device based upon vapor injection spray cooling technology. In addition, all devices must demonstrate scalability. For each device a unit cooling cell has been defined and greater surface area capability is to be achieved with the addition of adjacent cells without significantly affecting the performance of neighboring cells.

The influence of freestream static temperature on roughness temporal evolution and spatial variation was investigated in the Icing Research Tunnel (IRT) at NASA Glenn Research Center. A 53.34 cm (21-in.) NACA 0012 airfoil model and a 152.4 cm (60-in.) HAARP-II business jet airfoil model were exposed to Appendix C clouds for fixed exposure times and thus fixed ice accumulation parameter. For the base conditions, the static temperature was varied to produce different stagnation point freezing fractions. The resulting ice shapes were then scanned using a ROMER Absolute Arm system and analyzed using the self-organizing map approach of McClain and Kreeger. The ice accretion prediction program LEWICE was further used to aid in interrogations of the ice accretion point clouds by using the predicted surface variations of local collection efficiency.

Innovative multi-environment multimode thermal management architecture has been described that is capable of meeting widely varying thermal control requirements of various exploration mission scenarios currently under consideration. The proposed system is capable of operating in a single-phase or two-phase mode rejecting heat to the colder environment, operating in a two-phase mode with heat pump for rejecting heat to a warm environment, as well as using evaporative phase-change cooling for the mission phases where the radiator is incapable of rejecting the required heat. A single fluid loop can be used internal and external to the spacecraft for the acquisition, transport and rejection of heat by the selection of a working fluid that meets NASA safety requirements. Such a system may not be optimal for each individual mode of operation but its ability to function in multiple modes may permit global optimization of the thermal control system.

In this study, lightning effects on hydraulic transport elements in composite aircraft have been considered for the first time. Based on recent test results and analysis, several forms of possible structural damage and system component failures are presented. A unique approach in analysis has been taken to account that hydraulic transport elements, as a part of several aircraft systems, have a common interface with electrical wiring, and become complex electric networks. When an aircraft is exposed to a direct lightning strike, a metal skin on the wings and fuselage will conduct lightning currents in a way that only a small amount of induced electromagnetic energy will be present on hydraulic transport elements. So, in the past, hydraulic tubes, actuators, manifolds, and all other hydro-mechanical devices, as parts of various aircraft systems, have never been considered as lightning sensitive components.

The International Space Station Carbon Dioxide Removal Assembly (CDRA) uses regenerable adsorption technology to remove carbon dioxide (CO2) from cabin air. CO2 product water vapor measurements from a CDRA test bed unit at the NASA Marshall Space Flight Center were made using a tunable infrared diode laser differential absorption spectrometer (TILDAS) provided by NASA Glenn Research Center. The TILDAS instrument exceeded all the test specifications, including sensitivity, dynamic range, time response, and unattended operation. During the CO2 desorption phase, water vapor concentrations as low as 5 ppmv were observed near the peak of CO2 evolution, rising to levels of ∼40 ppmv at the end of a cycle. Periods of high water concentration (>100 ppmv) were detected and shown to be caused by an experimental artifact.

To support NASA's goal to improve aviation safety, the Aircraft Icing Project of the Aviation Safety Program has developed a number of education and training aids for pilots and operators on the hazards of atmospheric icing. A review of aircraft incident and accident investigations has revealed that flight crews have not always understood the effects of ice contamination on their aircraft. To increase this awareness, NASA has partnered with regulatory agencies and pilot trade organizations to assure relevant and practical materials that are focused toward the intended pilot audience. A number of new instructional design approaches and media delivery methods have been introduced to increase the effectiveness of the training materials by enhancing the learning experience, expanding user interactivity and participation, and, hopefully, increasing learner retention rates.

Manufacturing readiness of composite rotors and certification of flywheels depend in part on the maturity of nondestructive evaluation (NDE) technology for process optimization and quality assurance, respectively. Capabilities and limitations of x-ray-computed tomography and radiography, as well as advanced ultrasonics were established on NDE ring and rotor standards with EDM notches and drilled holes. Also, intentionally seeded delamination, tow break, and insert of bagging material were introduced in hydroburst-rings to study the NDE detection capabilities of such anomalies and their effect on the damage tolerance and safe life margins of subscale rings and rotors. Examples of possible occurring flaws or anomalies in composite rings as detected by NDE and validated by destructive metallography are shown. The general NDE approach to ensure quality of composite rotors and to help in the certification of flywheels is briefly outlined.

Following the Columbia accident, the EMUs (Extravehicular Mobility Units) onboard the ISS (International Space Station) went unused for an extended period of time. Upon startup, the units experienced a failure in the coolant systems. The failure resulted in a loss of EVA (Extravehicular Activity) capability from the US segment of the ISS. A failure investigation determined that chemical and biological contaminants and byproducts from the ISS Airlock Heat Exchanger, and the EMU itself, fouled the magnetically coupled pump in the EMU Transport Loop Fan/Pump Separator leading to a lack of coolant flow. Remediation hardware (the Airlock Coolant Loop Remediation water processing kit) and a process to periodically clean the EMU coolant loops on orbit were devised and implemented. The intent of this paper is to report on the successful implementation of the resultant hardware and process, and to highlight the go-forward plan.

An EMU water processing kit (Airlock Coolant Loop Recovery – A/L CLR) was developed as a corrective action to Extravehicular Mobility Unit (EMU) coolant flow disruptions experienced on the International Space Station (ISS) in May of 2004 and thereafter. Conservative schedules for A/L CLR use and component life were initially developed and implemented based on prior analysis results and analytical modeling. The examination of post-flight samples and EMU hardware in November of 2006 indicated that the A/L CLR kits were functioning well and had excess capacity that would allow a relaxation of the initially conservative schedules of use and component life. A relaxed use schedule and list of component lives were implemented thereafter. Since the adoption of the relaxed A/L CLR schedules of use and component lives, several A/L CLR kit items, transport loop water samples and sensitive EMU transport loop components have been examined to gage the impact of the relaxed requirements.

The problem of determining equilibrium temperatures for re-radiating surfaces in space vacuum was analyzed and the resulting mathematical relationships were incorporated in a code to determine space sink temperatures in the solar system. A brief treatment of planetary atmospheres is also included. Temperature values obtained with the code are in good agreement with available spacecraft telemetry and meteorological measurements for Venus and Earth. The code has been used in the design of space power system radiators for future interplanetary missions.

Two separate experimental rigs used in tests on NASA and Zero-G Corporation aircrafts flying low-gravity trajectories, and in the NASA 2.2 Second Drop Tower have been developed to test the functioning of the Flexible Membrane Commode (FMC) concept under reduced gravity conditions. The first rig incorporates the flexible, optically opaque membrane bag and the second rig incorporates a transparent chamber with a funnel assembly for evacuation that approximates the size of the membrane bag. Different waste dispensers have been used including a caulking gun and flexible hose assembly, and an injection syringe. Waste separation mechanisms include a pair of wire cutters, an iris mechanism, as well as discrete slug injection. The experimental work is described in a companion paper. This paper focuses on the obtained results and analysis of the data.