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At least 1,189 records · Page 66
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Effects of initial delamination on CIc and GIth values from glass/epoxy double cantilever beam tests

The effect of insert thickness and method of precracking on mode I interlaminar fracture toughness, GIc, and delamination fatigue threshold, GIth, values were determined for a glass/epoxy double cantilever beam specimen. The results of the static tests showed that precracking in tension would cause fiber bridging and thus may yield unconservative values of GIc and GIth. Precracking in shear yields suitable values of GIc but overly conservative values of GIth. For the glass/epoxy composite used, an insert thickness of 0.5 mil was most suitable for determining GIc and GIth values, although an insert thickness up to 3 mil was acceptable. Inserts thicker than 3 mil were not acceptable for determining GIc and GIth values.

Martin, Roderick H.↗
NASA NTRS1994

Robot friendly probe and socket assembly

A probe and socket assembly for serving as a mechanical interface between structures is presented. The assembly comprises a socket having a housing adapted for connection to a first supporting structure and a probe which is readily connectable to a second structure and is designed to be easily grappled and manipulated by a robotic device for insertion and coupling with the socket. Cooperable automatic locking means are provided on the probe shaft and socket housing for automatically locking the probe in the socket when the probe is inserted a predetermined distance. A second cooperable locking means on the probe shaft and housing are adapted for actuation after the probe has been inserted the predetermined distance. Actuation means mounted on the probe and responsive to the grip of the probe handle by a gripping device, such as a robot for conditioning the probe for insertion and are also responsive to release of the grip of the probe handle to actuate the second locking means to provide a "hard" lock of the probe in the socket.

Karen L Nyberg↗
NASA NTRS1993

Optimum transfer to a large-amplitude halo orbit for the Solar and Heliospheric Observatory (SOHO) spacecraft

The Solar and Heliospheric Observatory (SOHO), built by the European Space Agency to study the Sun as part of the International Solar-Terrestrial Physics (ISTP) Program, will be launched in July 1995 into a transfer trajectory that terminates in a large-amplitude halo orbit. The spacecraft will enter the halo orbit by performing one insertion maneuver at a specified point on the halo orbit. The position on the halo orbit that requires the least fuel for the insertion maneuver is identified using the planar, circular restricted three-body problem as a model. Fuel costs for halo orbit insertion at other points in the orbit are also identified. Practical trajectories incorporating all significant accelerations are discussed. The use of a lunar swingby to avoid any insertion maneuver is mentioned.

Stalos, S.↗
NASA NTRS1998

Lunar prospector mission design and trajectory support

The Lunar Prospector mission is the first dedicated NASA lunar mapping mission since the Apollo Orbiter program which was flown over 25 years ago. Competitively selected under the NASA Discovery Program, Lunar Prospector was launched on January 7, 1998 on the new Lockheed Martin Athena 2 launch vehicle. The mission design of Lunar Prospector is characterized by a direct minimum energy transfer trajectory to the moon with three scheduled orbit correction maneuvers to remove launch and cislunar injection errors prior to lunar insertion. At lunar encounter, a series of three lunar orbit insertion maneuvers and a small circularization burn were executed to achieve a 100 km altitude polar mapping orbit. This paper will present the design of the Lunar Prospector transfer, lunar insertion and mapping orbits, including maneuver and orbit determination strategies in the context of mission goals and constraints. Contingency plans for handling transfer orbit injection and lunar orbit insertion anomalies are also summarized. Actual flight operations results are discussed and compared to pre-launch support analysis.

Lozier, David↗
NASA NTRS1997

Using Superconducting Thin Films in Microwave Lines

High temperature superconductors(HTS) and microwaves devices form the ideal partnership. The application of superconductors in microwave devices, components and systems allows the reduction in size, power consumption and insertion loss. The surface resistance of high-Tc superconductors has been found to be two orders of magnitude lower than normal conducting copper materials. The reduction in size and power requirements, which together both lead to a reduction in system mass, coupled with reasonably accessible operating temperatures, suggest that HTS microwave components should find ready application in satellite communications systems. At present, multi- channeling communication networks demand filters with narrow bandwidth in order to allow the available RF frequency spectrum to be partitioned into small frequency bands, -and possible variation of dielectric constant from substrate to substrate is undesirable. Microwave multiplexers demand the fabrication of two identical filters in each channel. Thus, the filter with tuning function is preferable. Tunable filters are the critical component for phased array antennas in order to electronically steer the radiated beam. To fabricate a tunable filter that uses an electric field for operation, one would like a material that provides a large change on dielectric constant for a given electric field, yet has a relatively low tangent in order to minimize the insertion loss of the device. Ferroelectrics have been the materials of choice. Their large dielectric constant sufficiently increases the coupling between microwave resonators and its dependence on electric field provides timability. Development of technology promises to diminish tangent loss. The use of thin ferroelectric films sufficiently decreases insertion losses keeping considerable potential for applications. NASA Lewis Research Center is the one of the leading centers in investigation of superconductors/ferroelectric tunable components for microwave devices. A large number of possible microwave devices were fabricated and tested on the basis of thin film multilayer superconductor-ferroelectric structures. In major cases the systems with edge-coupling scheme were investigated. Dr. Genkin has recently focused on the new potentialities which implements the using of thin ferroelectric films in filters fabricated with end-coupled microstrip lines. Numerical modeling shows that these systems have large potential for application in tunable narrow- and wide-bandpass filters in the frequency range 10-20 GHz. The phase shifter with end-coupled resonant sections was fabricated and tested. Experimental results show large tunability, particular in low voltages. The possible optimization of this structure promises to improve the obtained result and to reach the low level of insertion losses.

Genkin, Varery↗
NASA NTRS2000

Mission Design Overview for Mars 2003/2005 Sample Return Mission

In May 2003, a new and exciting chapter in Mars exploration will begin with the launch of the first of three spacecraft that will collectively contribute toward the goal of delivering samples from the Red Planet to Earth. This mission is called Mars Sample Return (MSR) and will utilize both the 2003 and 2005 launch opportunities with an expected sample return in October 2008. NASA and CNES are major partners in this mission. The baseline mission mode selected for MSR is Mars orbit rendezvous (MOR), analogous in concept to the lunar orbit rendezvous (LOR) mode used for Apollo in the 1960s. Specifically, MSR will employ two NASA-provided landers of nearly identical design and one CNES-provided orbiter carrying a NASA payload of rendezvous sensors, orbital capture mechanisms, and an Earth entry vehicle (EEV). The high-level concept is that the landers will launch surface samples into Mars orbit, and the orbiter will retrieve the samples in orbit and then carry them back to Earth. The first element to depart for Mars will be one of the two landers. Currently, it is proposed that an intermediate class launch vehicle, such as the Boeing Delta 3 or Lockheed Martin Atlas 3A, will launch this 1800-kg lander from Cape Canaveral during the May 2003 opportunity. The lander will utilize a Type-1 transfer trajectory with an arrival at Mars in mid-December 2003. Landing will be aided by precision approach navigation and a guided hypersonic entry to achieve a touchdown accuracy of 10 km or better. Although the exact landing site has not yet been determined, it is estimated that lander resource constraints will limit the site to between 15 degrees north and south latitudes. Following touchdown, the lander will deploy a six-wheeled, 60-kg rover carrying an extensive suite of instruments designed to aid in the analysis of the local terrain and collection of core samples from selected rocks. The surface mission is currently designed around a concept called the surface traverse. Each traverse will involve the rover exploring a selected area of terrain up to 100 meters from the lander, the collection of rock core samples, and the delivery of the samples from the traverse back to a sample canister on the lander. Planning estimates indicate that up to three traverses may be possible during the expected 90-sol lifetime of the lander. The canister that will receive the samples from the rover will be attached to the top stage of a small solid-fueled rocket mounted to the deck of the lander. This rocket is called the Mars Ascent Vehicle (MAV) and consists of three stages weighing a total of about 140 kg. After the conclusion of the surface mission, the MAV will lift-off and insert the sample canister into a near-circular orbit with an altitude of about 600 km and inclination of 45 degrees. The sample canister will wait in this orbit until it is retrieved by the orbiter sometime in early 2007. In August 2005, the second lander and a CNES-provided orbiter weighing 2700 kg will depart for Mars. Currently, it is proposed that a single Ariane 5 provided by CNES will launch both of these two elements onto a Type-2 transfer trajectory. Although the orbiter and lander will be launched together, they will separate shortly after injection and will fly to Mars as two independent spacecraft. However, both spacecraft will perform a maneuver between 10 and 15 days after launch so that their arrival times at Mars differ by between 12 and 24 hours. This scheme will reduce the operational complexity at the encounter date. A set of four 60-kg surface probes will ride piggyback on the orbiter to Mars. These CNES-provided probes are called Netlanders and will serve as surface stations for scientific investigations independent of the Mars Sample Return goals. Starting approximately one month prior to arrival at Mars, the orbiter will begin to release the Netlanders one at a time. Each release cycle will take several days, and will include time for precision navigation to execute one or two maneuvers that will target the Netlanders to their proper landing site. All four deployment cycles will be completed prior to 10 days before arrival. Both the orbiter and lander will arrive in late-July 2006. Upon arrival, the lander will perform a precision landing and surface mission similar in concept to the one that was executed during the 2003 opportunity. Although the landing site for the 2005 opportunity has not been selected, it is expected to be different from the 2003 site to enhance the diversity of the collected samples. The orbiter's arrival at Mars will be highlighted by the first use of aerocapture to insert a spacecraft into a capture orbit around another planet. The choice of aerocapture, as opposed to a propulsive orbit insertion, was considered mission enabling due to a reduction of over 2000 m/s in mission AV. Aerocapture will be targeted to produce a 250 km x 1400 km capture orbit with an inclination of 45 degrees. Current analysis indicates that achieving this goal will require approximately six minutes of flight deep in the atmosphere with a targeted periapsis of approach of about 43 km. After factoring into account the penalty for carrying a heat shield to survive aerocapture, the net savings compared to a propulsive orbital insertion amounts to several hundred kilograms.

Lee, Wayne J.↗
NASA NTRS1993

Mars Observer Press Conference JPL

The Mars Observer mission spacecraft was primarily designed for exploring Mars and the Martian environment. The Mars Observer was launched on September 25, 1992. The spacecraft was lost in the vicinity of Mars on August 21, 1993 when the spacecraft began its maneuvering sequence for Martian orbital insertion. This videotape shows a press briefing, held after the spacecraft had not responded to attempts to communicate with it, to explain to the press the problems and the steps that were being taken to re-establish communication with the spacecraft. The communications had been shutdown prior to the orbital insertion burn to protect the instruments. At the time of the press conference, the communications system was still not operational, and attempts were being made to re-establish communication. Bob McMillan of the Public Affairs Office at JPL gives the initial announcement of the continuing communication problem with the spacecraft. Mr. McMillan introduces William Piotrowski, acting director of solar system exploration, who reiterates that there is indeed no communication with the Observer spacecraft. He is followed by Glenn Cunningham, the Project Manager of the Mars Observer who speaks about the attempts to re-establish contact. Mr. Cunningham is followed by Satenios Dallas, the Mission Manager for the Mars Observer Project, who speaks about the sequence of events leading up to the communication failure, and shows an animated video presenting the orbital insertion maneuvers. The briefing was then opened up for questions from the assembled press, both at JPL and at the other NASA Centers. The questions are about the possible reasons for the communication failure, and the attempts to restore communications with the spacecraft. Dr. Arden L. Albee, chief scientist for the Mars Observer Mission, joins the other panel members to answer questions. At the end of the press briefing the animation of the Mars orbital insertion is shown again.

Source record↗
NASA NTRS2000

Use Computer-Aided Tools to Parallelize Large CFD Applications

Porting applications to high performance parallel computers is always a challenging task. It is time consuming and costly. With rapid progressing in hardware architectures and increasing complexity of real applications in recent years, the problem becomes even more sever. Today, scalability and high performance are mostly involving handwritten parallel programs using message-passing libraries (e.g. MPI). However, this process is very difficult and often error-prone. The recent reemergence of shared memory parallel (SMP) architectures, such as the cache coherent Non-Uniform Memory Access (ccNUMA) architecture used in the SGI Origin 2000, show good prospects for scaling beyond hundreds of processors. Programming on an SMP is simplified by working in a globally accessible address space. The user can supply compiler directives, such as OpenMP, to parallelize the code. As an industry standard for portable implementation of parallel programs for SMPs, OpenMP is a set of compiler directives and callable runtime library routines that extend Fortran, C and C++ to express shared memory parallelism. It promises an incremental path for parallel conversion of existing software, as well as scalability and performance for a complete rewrite or an entirely new development. Perhaps the main disadvantage of programming with directives is that inserted directives may not necessarily enhance performance. In the worst cases, it can create erroneous results. While vendors have provided tools to perform error-checking and profiling, automation in directive insertion is very limited and often failed on large programs, primarily due to the lack of a thorough enough data dependence analysis. To overcome the deficiency, we have developed a toolkit, CAPO, to automatically insert OpenMP directives in Fortran programs and apply certain degrees of optimization. CAPO is aimed at taking advantage of detailed inter-procedural dependence analysis provided by CAPTools, developed by the University of Greenwich, to reduce potential errors made by users. Earlier tests on NAS Benchmarks and ARC3D have demonstrated good success of this tool. In this study, we have applied CAPO to parallelize three large applications in the area of computational fluid dynamics (CFD): OVERFLOW, TLNS3D and INS3D. These codes are widely used for solving Navier-Stokes equations with complicated boundary conditions and turbulence model in multiple zones. Each one comprises of from 50K to 1,00k lines of FORTRAN77. As an example, CAPO took 77 hours to complete the data dependence analysis of OVERFLOW on a workstation (SGI, 175MHz, R10K processor). A fair amount of effort was spent on correcting false dependencies due to lack of necessary knowledge during the analysis. Even so, CAPO provides an easy way for user to interact with the parallelization process. The OpenMP version was generated within a day after the analysis was completed. Due to sequential algorithms involved, code sections in TLNS3D and INS3D need to be restructured by hand to produce more efficient parallel codes. An included figure shows preliminary test results of the generated OVERFLOW with several test cases in single zone. The MPI data points for the small test case were taken from a handcoded MPI version. As we can see, CAPO's version has achieved 18 fold speed up on 32 nodes of the SGI O2K. For the small test case, it outperformed the MPI version. These results are very encouraging, but further work is needed. For example, although CAPO attempts to place directives on the outer- most parallel loops in an interprocedural framework, it does not insert directives based on the best manual strategy. In particular, it lacks the support of parallelization at the multi-zone level. Future work will emphasize on the development of methodology to work in a multi-zone level and with a hybrid approach. Development of tools to perform more complicated code transformation is also needed.

Jin, H.↗
NASA NTRS2001

First Materials Science Research Rack Capabilities and Design Features

The first Materials Science Research Rack (MSRR-1) will accommodate dual Experiment Modules (EM's) and provide simultaneous on-orbit processing operations capability. The first international Materials Science Experiment Module for the MSRR-1 is an international cooperative research activity between NASA's Marshall Space Flight Center (MSFC) and the European Space Agency's (ESA) European Space Research and Technology Center. (ESTEC). This International Standard Payload Rack (ISPR) will contain the Materials Science Laboratory (MSL) developed by ESA as an Experiment Module. The MSL Experiment Module will accommodate several on-orbit exchangeable experiment-specific Module Inserts. Module Inserts currently planned are a Quench Module Insert, Low Gradient Furnace, Solidification with Quench Furnace, and Diffusion Module Insert. The second Experiment Module for the MSRR-1 configuration is a commercial device supplied by MSFC's Space Products Department (SPD). It includes capabilities for vapor transport processes and liquid metal sintering. This Experiment Module will be replaced on-orbit with other NASA Materials Science EMs.

Schaefer, D.↗
NASA NTRS2003

Membrane-Based Functions in the Origin of Cellular Life

How simple membrane peptides performed such essential proto-cellular functions as transport of ions and organic matter across membranes separating the interior of the cell from the environment, capture and utilization of energy, and transduction of environmental signals, is a key question in protobiological evolution. On the basis of detailed, molecular-level computer simulations we investigate how these peptides insert into membranes, self-assemble into higher-order structures and acquire functions. We have studied the insertion of an a-helical peptide containing leucine (L) and serine (S) of the form (LSLLLSL)S into a model membrane. The transmembrane state is metastable, and approximately 15 kcal/mol is required to insert the peptide into the membrane. Investigations of dimers formed by (LSLLLSL)S and glycophorin A demonstrate how the favorable free energy of helix association can offset the unfavorable free energy of insertion, leading to self- assembly of peptide helices in the membrane. An example of a self-assembled structure is the tetrameric transmembrane pore of the influenza virus M2 protein, which is an efficient and selective voltage-gated proton channel. Our simulations explain the gating mechanism and provide guidelines how to reengineering the channel to act as a simple proton pump. In general, emergence of integral membrane proteins appears to be quite feasible and may be easier to envision than the emergence of water-soluble proteins.

Wilson, Michael A.↗
NASA NTRS2003

Membrane peptides and their role in protobiological evolution

How simple membrane peptides performed such essential protocellular functions as transport of ions and organic matter across membranes separating the interior of the cell from the environment, capture and utilization of energy, and transduction of environmental signals, is a key question in protobiological evolution. On the basis of detailed, molecular-level computer simulations we explain how these peptides fold at water-membrane interfaces, insert into membranes, self-assemble into higher-order structures and acquire functions. We have investigated the interfacial behavior and folding of several peptides built of leucine and glutamine residues and have demonstrated that many of them tend to adopt ordered structures. Further, we have studied the insertion of an alpha-helical peptide containing leucine (L) and serine (S) of the form (LSLLLSL)3 into a model membrane. The transmembrane state is metastable, and approximately 15 kcal mol(-1) is required to insert the peptide into the membrane. Investigations of dimers formed by (LSLLLSL)3 and glycophorin A demonstrate how the favorable free energy of helix association can offset the unfavorable free energy of insertion, leading to self-assembly of peptide helices in the membrane. An example of a self-assembled structure is the tetrameric transmembrane pore of the influenza virus M2 protein, which is an efficient and selective voltage-gated proton channel. Our simulations explain the gating mechanism and provide guidelines how to re-engineer the channel to act as a simple proton pump. In general, emergence of integral membrane proteins appears to be quite feasible and may be easier to envision than the emergence of water-soluble proteins.

Models, Chemical↗
NASA NTRS2000

Radio Frequency (RF) Micro-Electromechanical Systems (MEMS) Switches for Space Communications

Micro-electromechanical systems (MEMS) is an emerging technology for radio frequency (RF) systems because it has the potential to dramatically decrease loss and improve efficiency. In this paper, we address the design and fabrication of novel MEMS switches being developed at NASA Glenn Research Center. Two types of switches are being developed: a microstrip series single pole single throw (SPST) switch and a coplanar waveguide (CPW) series SPST and single pole double throw (SPDT) switches. These are being fabricated as an integral part of 50 Ohm microstrip and CPW RF integrated circuits using microfabrication techniques. The construction of the switch relies on a cantilever beam that is partially supported by a dielectric post. The cantilever beam is electro-magnetically actuated. To decrease stiction, a Si3N4 thin film is deposited over the contact area. Thus, when the switch is closed, the ON-state insertion loss is governed by the parallel plate capacitance formed by the two contacts. The isolation in the OFF-state is governed by the parasitic capacitance when the cantilever is in the up position. RF MEMS switches have been demonstrated with 80% lower insertion loss than conventional solid state devices (GaAs Metal Semiconductor Field Effect Transistors (MESFETs) and Silicon PIN diodes) based switches. For example, a conventional GaAs five-bit phase shifter which is required for beam steering in a phased array antenna has approximately 7 dB of insertion loss at 26.5 GHz where as a comparable MEMS based phase shifter is expected to have only 2 dB of insertion loss. This translates into 56% lower power dissipation and therefore decreases the thermal load on the spacecraft and also reduces the power amplifier requirements. These benefits will enable NASA to build the next generation of deep space science crafts and micro/nano satellites.

Simons, Rainee N.↗
NASA NTRS2004

Hollow Cathode and Keeper-region Plasma Measurements Using Ultra-fast Miniature Scanning Probes

In order to support the development of comprehensive performance and life models for future deep space missions that will utilize ion thrusters, we have undertaken a study of the plasma structure in hollow cathodes using an new pneumatic scanning probe diagnostic. This device is designed to insert a miniature probe directly into the hollow cathode orifice from either the upstream insert region in the interior of the hollow cathode, or from the downstream keeper-plasma region at the exit of the hollow cathode, to provide complete axial profiles of the discharge plasma parameters. Previous attempts to diagnose this region with probes was Limited by the melting of small probes in the intense discharge near the orifice, or caused significant perturbation of the plasma by probes large enough to survive. Our new probe is extremely compact, and when configured as a single Langmuir probe, the ceramic tube insulator is only 0.5mm in diameter and the current collecting conductor has a total area of 0.002 cm2. A series of current-voltage characteristics are obtained by applying a rapid sawtooth voltage waveform to the probe as it is scanned by the pneumatic actuator into and out of the plasma region, The bellow-sealed pneumatic drive scans the probe 4 cm in the cathode insert region and 10 cm in the anode/keeper plasmas region at average speeds of about 1 mm/msec, and the residence time at the end of the insertion stroke in the densest part of the plasma near the orifice is measured to be only 10 msec. Since the voltage sweep time is fast compared to the motion of the probe, axial profiles of the plasma density, temperature and potential with reasonable spatial resolution are obtained. Measurements of the internal cathode pressures and the axial plasma-parameter profiles for a hollow cathode operating at discharge currents of up to 35 A in xenon will be presented.

plasma generation↗
NASA NTRS2008

Experimental and Numerical Characterization of High Heat Fluxes During Transient Blackbody Calibrations

High heat fluxes are encountered in numerous applications, such as hypersonic vehicles in flight, fires, and engines, Calibration of heat flux gages may be performed in a dual cavity cylindrical blackbody resulting in a transient calibration environment. To characterize the transient heat fluxes. experiments were performed on a dual cavity cylindrical blackbody at nominal temperatures varying from 800 C to 1900 C in increments of 100 C. Based on experiments, the optimum heat flux sensor insertion location as measured from the center partition was determined. The pre-insertion steady state axial temperature profile is compared experimentally, numerically, and analytically. The effect of convection in the blackbody cavity during the insertion is calculated and found to be less than 2 per cent. Also, an empirical correlation for predicting the emissivity of the blackbody is included. Detailed transient thermal models have been developed to simulate the heat flux calibration process at two extreme fluxes. The high (1MW/sq m) and relatively low (70 kw/sq m) fluxes are reported in this article. The transient models show the effect of inserting a heat flux gage at room temperature on the thermal equilibrium of the blackbody at 1800 C and 800 C nominal temperatures, respectively. Also, heat flux sensor outputs are derived from computed sensor temperature distributions and compared to experimental results.

Abdelmessih, Amanie N.↗
NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO) Thruster Control Mode Design and Flight Experience

National Aeronautics and Space Administration s (NASA) Goddard Space Flight Center (GSFC) in Greenbelt, MD, designed, built, tested, and launched the Lunar Reconnaissance Orbiter (LRO) from Cape Canaveral Air Force Station on June 18, 2009. The LRO spacecraft is the first operational spacecraft designed to support NASA s return to the Moon, as part of the Vision for Space Exploration. LRO was launched aboard an Atlas V 401 launch vehicle into a direct insertion trajectory to the Moon. Twenty-four hours after separation the propulsion system was used to perform a mid-course correction maneuver. Four days after the mid-course correction a series of propulsion maneuvers were executed to insert LRO into its commissioning orbit. The commission period lasted eighty days and this followed by a second set of thruster maneuvers that inserted LRO into its mission orbit. To date, the spacecraft has been gathering invaluable data in support of human s future return to the moon. The LRO Attitude Control Systems (ACS) contains two thruster based control modes: Delta-H and Delta-V. The design of the two controllers are similar in that they are both used for 3-axis control of the spacecraft with the Delta-H controller used for momentum management and the Delta-V controller used for orbit adjust and maintenance maneuvers. In addition to the nominal purpose of the thruster modes, the Delta-H controller also has the added capability of performing a large angle slew maneuver. A suite of ACS components are used by the thruster based control modes, for both initialization and control. For initialization purposes, a star tracker or the Kalman Filter solution is used for providing attitude knowledge and upon entrance into the thruster based control modes attitude knowledge is provided via rate propagation using a inertial reference unit (IRU). Rate information for the controller is also supplied by the IRU. Three-axis control of the spacecraft in the thruster modes is provided by eight 5-lbf class attitude control thrusters configured in two sets of four thrusters for redundancy purposes. Four additional 20-lbf class thrusters configured in two sets of two thrusters are used for Lunar Orbit Insertion maneuvers. The propulsion system is one the few systems on-board the LRO spacecraft that has built in redundancy. The Delta-H controller consists of a Proportional-Derivative (PD) controller with a structural filter on the thrusters and a Proportional controller on the reaction wheels. The PD control that employs the thrusters is used for attitude and rate control. The Proportional controller on the reaction wheels is used for commanding the wheels to a new momentum state. The ground commands used for the Delta-H controller are the system momentum vector, reaction wheel momentum, maximum expected command time, and which set of attitude control thrusters to use. The ability to command both the system momentum vector and reaction wheel momentum in the Delta-H controller provides both a capability and an additional source of operator error. Large angle slews via the Delta-H controller is achievable via this commands because these commands are used for the exit mode criteria. Setting these commands to non-consistent values prevents the mode from exiting nominally.

Hsu, Oscar C.↗
NASA NTRS2011

Monitored separation device

A device for separating and purifying useful quantities of particles comprises: a. an anolyte reservoir connected to an anode, the anolyte reservoir containing an electrophoresis buffer; b. a catholyte reservoir connected to a cathode, the catholyte reservoir also containing the electrophoresis buffer; c. a power supply connected to the anode and to the cathode; d. a column having a first end inserted into the anolyte reservoir, a second end inserted into the catholyte reservoir, and containing a separation medium; e. a light source; f. a first optical fiber having a first fiber end inserted into the separation medium, and having a second fiber end connected to the light source; g. a photo detector; h. a second optical fiber having a third fiber end inserted into the separation medium, and having a fourth fiber end connected to the photo detector; and i. an ion-exchange membrane in the anolyte reservoir.

Jackson, George William↗
NASA NTRS2011

Improved Rare-Earth Emitter Hollow Cathode

An improvement has been made to the design of the hollow cathode geometry that was created for the rare-earth electron emitter described in Compact Rare Earth Emitter Hollow Cathode (NPO-44923), NASA Tech Briefs, Vol. 34, No. 3 (March 2010), p. 52. The original interior assembly was made entirely of graphite in order to be compatible with the LaB6 material, which cannot be touched by metals during operation due to boron diffusion causing embrittlement issues in high-temperature refractory materials. Also, the graphite tube was difficult to machine and was subject to vibration-induced fracturing. This innovation replaces the graphite tube with one made out of refractory metal that is relatively easy to manufacture. The cathode support tube is made of molybdenum or molybdenum-rhenium. This material is easily gun-bored to near the tolerances required, and finish machined with steps at each end that capture the orifice plate and the mounting flange. This provides the manufacturability and robustness needed for flight applications, and eliminates the need for expensive e-beam welding used in prior cathodes. The LaB6 insert is protected from direct contact with the refractory metal tube by thin, graphite sleeves in a cup-arrangement around the ends of the insert. The sleeves, insert, and orifice plate are held in place by a ceramic spacer and tungsten spring inserted inside the tube. To heat the cathode, an insulating tube is slipped around the refractory metal hollow tube, which can be made of high-temperature materials like boron nitride or aluminum nitride. A screw-shaped slot, or series of slots, is machined in the outside of the ceramic tube to constrain a refractory metal wire wound inside the slot that is used as the heater. The screw slot can hold a single heater wire that is then connected to the front of the cathode tube by tack-welding to complete the electrical circuit, or it can be a double slot that takes a bifilar wound heater with both leads coming out the back. This configuration replaces the previous sheathed heater design that limited the cycling-life of the cathode.

Goebel, Dan M.↗
NASA NTRS2005

Fastener Starter

The Fastener Starter is a creative solution to prevent the loss of small fasteners during their installation. This is the only currently available tool that can firmly grip and hold a single screw, bolt, nut, washer, spacer, or any combination of these parts. Other commercially available fastener starters are unable to accommodate a variety of parts simultaneously. The Fastener Starter is a more capable and easier tool to use than prior tools. Its compact size allows it to be used effectively in cramped, difficult-to-see locations. Its design also allows it to be used with or without handles and extenders in other difficult-to-reach locations. It provides better protection against cross threading and loss of fasteners and associated parts. The Fastener Starter is non-magnetic and does not off-gas, thus meeting flight hardware requirements. The Fastener Starter incorporates a combination of features of several commercially available tools, providing an improved means of installing small fasteners. The Fastener Starter includes a custom molded insert that can be removed easily and replaced with a conventional tool bit (e.g., a screwdriver or hex-driver bit). When used with the insert, the Fastener Starter prevents cross threading and damage to internal threaded holes. This is achieved by allowing the fastener to slip within the tool insert when used without a conventional tool bit. Alternatively, without the insert and with a tool bit, the Fastener Starter can torque a fastener. The Fastener Starter has a square recess hole that accepts a conventional square drive handle or extension to accommodate a variety of applications by providing flexibility in handle style and length.

Chandler, Faith↗