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NASA NTRS2002

Life Prediction/Reliability Data of Glass-Ceramic Material Determined for Radome Applications

Brittle materials, ceramics, are candidate materials for a variety of structural applications for a wide range of temperatures. However, the process of slow crack growth, occurring in any loading configuration, limits the service life of structural components. Therefore, it is important to accurately determine the slow crack growth parameters required for component life prediction using an appropriate test methodology. This test methodology also should be useful in determining the influence of component processing and composition variables on the slow crack growth behavior of newly developed or existing materials, thereby allowing the component processing and composition to be tailored and optimized to specific needs. Through the American Society for Testing and Materials (ASTM), the authors recently developed two test methods to determine the life prediction parameters of ceramics. The two test standards, ASTM 1368 for room temperature and ASTM C 1465 for elevated temperatures, were published in the 2001 Annual Book of ASTM Standards, Vol. 15.01. Briefly, the test method employs constant stress-rate (or dynamic fatigue) testing to determine flexural strengths as a function of the applied stress rate. The merit of this test method lies in its simplicity: strengths are measured in a routine manner in flexure at four or more applied stress rates with an appropriate number of test specimens at each applied stress rate. The slow crack growth parameters necessary for life prediction are then determined from a simple relationship between the strength and the applied stress rate. Extensive life prediction testing was conducted at the NASA Glenn Research Center using the developed ASTM C 1368 test method to determine the life prediction parameters of a glass-ceramic material that the Navy will use for radome applications.

Choi, Sung R.↗

NASA NTRS2003

Characterization of a Ultra-high Temperature Ceramic Composite

Ultra-high temperature ceramics (UHTC) are of interest for hypersonic vehicle leading edge applications. Monolithic UHTCs are of concern because of their low fracture toughness and brittle behavior. UHTC composites (UHTCC) are being investigated as a possible approach to overcome these deficiencies. In this study a small sample of a UHTCC was evaluated by limited mechanical property tests, furnace oxidation exposures, and oxidation exposures in a flowing environment. The composite was prepared from a carbon fiber perform using ceramic particulates and a preceramic polymer. The as-received composite plate was non-uniform from front to back surface. Plate dimensions were 150 x 150 x 6 mm. The back surface had a fibrous, uniform appearance; XRD analysis revealed the presence of Sic and C. The front surface was smooth and non-uniform in appearance with evidence of a coarse grain structure produced by a liquid phase; XRD analysis revealed the presence of HfB2. Microcracks were present throughout the thickness as one might expect from a carbon fiber reinforced composite with attendant large thermal expansion mismatch between the matrix phases and the fibers. The HfB2 phase on the front surface was comparable in thickness to a fiber ply or about 0.6 mm, and surface microcracks were evident. Limited four point flexural tests were carried out at span to depth ratios of approximately 14 and 16 with markedly different results. Tests were run with the front or the back surface in tension. At the shorter span to depth failures occurred under a loading pin for both orientations. At a span to depth of 16 failures occurred in the center of the span with fracture clearly initiating from a tensile failure. Ultimate flexural strength, strain at ultimate stress, stress and strain at deviation from linear elastic behavior are reported. Strains at ultimate stress ranged from about 0.6 to 0.7 % for the back surface in tension, and 0.4 to 0.6 for the front surface in tension. At constant span to depth the strain at ultimate stress was about 0.2% greater for the back surface in tension and the ultimate strength was also higher. Strengths were in line with predictions from theory. Furnace oxidation studies were carried out at 1627 and 1927OC in a static furnace environment using ten minute cycles and one, five, and ten cycles. Limited oxidation studies were also carried out in a flowing oxyacetylene torch environment. Specimens were photographed, and weight and dimensional changes were determined. XRD and SEM characterizations were performed. Weight losses were attributed primarily to carbon fiber oxidation. The composite survived the torch test with little visible distress. Further details will be determined once metallographic studies are completed.

Levine, Stanley R.↗

NASA NTRS2003

Polymer/Silicate Nanocomposites Developed for Improved Strength and Thermal Stability

Over the past decade, polymer-silicate nanocomposites have been attracting considerable attention as a method of enhancing polymer properties. The nanometer dimensions of the dispersed silicate reinforcement can greatly improve the mechanical, thermal, and gas barrier properties of a polymer matrix. In a study at the NASA Glenn Research Center, the dispersion of small amounts (less than 5 wt%) of an organically modified layered silicate (OLS) into the polymer matrix of a carbon-fiber-reinforced composite has improved the thermal stability of the composite. The enhanced barrier properties of the polymer-clay hybrid are believed to slow the diffusion of oxygen into the bulk polymer, thereby slowing oxidative degradation of the polymer. Electron-backscattering images show cracking of a nanocomposite matrix composite in comparison to a neat resin matrix composite. The images show that dispersion of an OLS into the matrix resin reduces polymer oxidation during aging and reduces the amount of cracking in the matrix significantly. Improvements in composite flexural strength, flexural modulus, and interlaminar shear strength were also obtained with the addition of OLS. An increase of up to 15 percent in these mechanical properties was observed in composites tested at room temperature and 288 C. The best properties were seen with low silicate levels, 1 to 3 wt%, because of the better dispersion of the silicate in the polymer matrix.

Campbell, Sandi G.↗

NASA NTRS2006

Slow Crack Growth and Fracture Toughness of Sapphire for the International Space Station Fluids and Combustion Facility

The fracture toughness, inert flexural strength, and slow crack growth parameters of the r- and a-planes of sapphire grown by the Heat Exchange Method were measured to qualify sapphire for structural use in the International Space Station. The fracture toughness in dry nitrogen, K(sub Ipb), was 2.31 +/- 0.12 MPa(square root of)m and 2.47 +/- 0.15 MPa(squre root of)m for the a- and r-planes, respectively. Fracture toughness measured in water via the operational procedure in ASTM C1421 was significantly lower, K(sub Ivb) = 1.95+/- 0.03 MPa(square root of)m, 1.94 +/- 0.07 and 1.77 +/- 0.13 MPa(square root of)m for the a- , m- and r-planes, respectively. The mean inert flexural strength in dry nitrogen was 1085 +/- 127 MPa for the r-plane and 1255 +/- 547 MPa for the a-plane. The power law slow crack growth exponent for testing in water was n = 21 +/- 4 for the r-plane and n (greater than or equal to) 31 for the a-plane. The power law slow crack growth coefficient was A = 2.81 x 10(exp -14) m/s x (MPa(squre root of)m)/n for the r-plane and A (approx. equals)2.06 x 10(exp -15) m/s x (MPa(square root of)m)/n for the a-plane. The r- and a-planes of sapphire are relatively susceptible to stress corrosion induced slow crack growth in water. However, failure occurs by competing modes of slow crack growth at long failure times and twinning for short failure time and inert environments. Slow crack growth testing needs to be performed at low failure stress levels and long failure times so that twinning does not affect the results. Some difficulty was encountered in measuring the slow crack growth parameters for the a-plane due to a short finish (i.e., insufficient material removal for elimination of the damage generated in the early grinding stages). A consistent preparation method that increases the Weibull modulus of sapphire test specimens and components is needed. This would impart higher component reliability, even if higher Weibull modulus is gained at the sacrifice of absolute strength of the component. The current specification frequently used for the preparation of sapphire test specimens and components (e.g., a "60/40" scratch-dig finish) is inadequate to avoid a short finish.

Salem, Jonathan A.↗

NASA NTRS2007

E-Textile Antennas for Space Environments

The ability to integrate antennas and other radio frequency (RF) devices into wearable systems is increasingly important as wireless voice, video, and data sources become ubiquitous. Consumer applications including mobile computing, communications, and entertainment, as well as military and space applications for integration of biotelemetry, detailed tracking information and status of handheld tools, devices and on-body inventories are driving forces for research into wearable antennas and other e-textile devices. Operational conditions for military and space applications of wireless systems are often such that antennas are a limiting factor in wireless performance. The changing antenna platform, i.e. the dynamic wearer, can detune and alter the radiation characteristics of e-textile antennas, making antenna element selection and design challenging. Antenna designs and systems that offer moderate bandwidth, perform well with flexure, and are electronically reconfigurable are ideally suited to wearable applications. Several antennas, shown in Figure 1, have been created using a NASA-developed process for e-textiles that show promise in being integrated into a robust wireless system for space-based applications. Preliminary characterization of the antennas with flexure indicates that antenna performance can be maintained, and that a combination of antenna design and placement are useful in creating robust designs. Additionally, through utilization of modern smart antenna techniques, even greater flexibility can be achieved since antenna performance can be adjusted in real-time to compensate for the antenna s changing environment.

Kennedy, Timothy F.↗

NASA NTRS1990

Submillimeter Wave Astronomy Satellite Conceptual Design Review

The SWAS instrument consists of the moving telescope assembly, the interface baseplate assembly, and the thermal control housing. The movin g telescope structure holds the primary mirror, secondary mirror, and receiver front end in precise alignment. This structure also carries the receiver cold plate radiators, which view cold space both directl y and through reflection off the primary mirror. The moving telescope assembly attaches to the interface baseplate with an open pivot frame through which flex leads from the receiver front end also pass. Two flexure-mounted linear actuators tilt the moving telescope assembly u p to +/-3 deg in two axes. The interface plate serves as a mounting for these actuators as well as the acousto-optic spectrometer, the inst rument control electronics, the star tracker, and the balance of the receiver components. The thermal control housing attaches to the inte rface baseplate around its edge. The top cylindrical part of this hou sing serves to shade the receiver cold plate radiators from Sun and Earth influence. The bottom "D" shaped portion of the housing is therma lly isolated from the top and forms the electronlcs radiator. Heat is conducted from the Interface baseplate into this lower portion of the shell. The radiator geometry and conduction paths create a very stab le thermal environment in the center of the interface baseplate where the AOS and receiver IF amplifiers are heat sinked. The outer surface of the thermal control housing is covered with fused silica second s urface mirrors that reflect visible light and radiate infrared energy . The instrument is mounted to the spacecraft through four thermally- isolating titanium flexures an the underside of the interface basepla te.

Source record↗

NASA NTRS1986

Rotor blade construction for circulation control aircraft

A circulation control aircraft rotor blade having a spanwise Coanda surface 16 and a plurality of spanwise extending flexible composite material panels 18 cooperating with the surface to define slots for the discharge of compressed air from within the blade with each panel having first flexure means 60 associated with screw adjustments 36 for establishing a slot opening preload and second flexure means 62 associated with screw adjustments 38 for establishing a slot maximum opening.

Carter, Sr., Donald R.↗

NASA NTRS2006

Isolated resonator gyroscope with a drive and sense plate

The present invention discloses a resonator gyroscope comprising a vibrationally isolated resonator including a proof mass, a counterbalancing plate having an extensive planar region, and one or more flexures interconnecting the proof mass and counterbalancing plate. A baseplate is affixed to the resonator by the one or more flexures and sense and drive electrodes are affixed to the baseplate proximate to the extensive planar region of the counterbalancing plate for exciting the resonator and sensing movement of the gyroscope. The isolated resonator transfers substantially no net momentum to the baseplate when the resonator is excited.

Challoner, A. Dorian↗

NASA NTRS2006

Isolated resonator gyroscope with a drive and sense plate

Challoner, A. Dorian↗

NASA NTRS2006

Isolated resonator gyroscope with isolation trimming using a secondary element

The present invention discloses a resonator gyroscope including an isolated resonator. One or more flexures support the isolated resonator and a baseplate is affixed to the resonator by the flexures. Drive and sense elements are affixed to the baseplate and used to excite the resonator and sense movement of the gyroscope. In addition, at least one secondary element (e.g., another electrode) is affixed to the baseplate and used for trimming isolation of the resonator. The resonator operates such that it transfers substantially no net momentum to the baseplate when the resonator is excited. Typically, the isolated resonator comprises a proof mass and a counterbalancing plate.

Challoner, A. Dorian↗

NASA NTRS1989

Composites with improved fiber-resin interfacial adhesion

The adhesion of fiber reinforcement such as high modulus graphite to a matrix resin such as polycarbonate is greatly enhanced by applying a very thin layer, suitably from 50 Angstroms to below 1000 Angstroms, to the surface of the fiber such as by immersing the fiber in a dilute solution of the matrix resin in a volatile solvent followed by draining to remove excess solution and air drying to remove the solvent. The thin layer wets the fiber surface. The very dilute solution of matrix resin is able to impregnate multifilament fibers and the solution evenly flows onto the surface of the fibers. A thin uniform layer is formed on the surface of the fiber after removal of the solvent. The matrix resin coated fiber is completely wetted by the matrix resin during formation of the composite. Increased adhesion of the resin to the fibers is observed at fracture. At least 65 percent of the surface of the graphite fiber is covered with polycarbonate resin at fracture whereas uncoated fibers have very little matrix resin adhering to their surfaces at fracture and epoxy sized graphite fibers exhibit only slightly higher coverage with matrix resin at fracture. Flexural modulus of the composite containing matrix resin coated fibers is increased by 50 percent and flexural strength by 37 percent as compared to composites made with unsized fibers.

Cizmecioglu, Muzaffer↗

NASA NTRS2006

Finite Element Development of Honeycomb Panel Configurations with Improved Transmission Loss

The higher stiffness-to-mass ratio of a honeycomb panel compared to a homogeneous panel results in a lower acoustic critical frequency. Above the critical frequency the panel flexural wave speed is acoustically fast and the structure becomes a more efficient radiator with associated lower sound transmission loss. Finite element models of honeycomb sandwich structures are presented featuring areas where the core is removed from the radiating face sheet disrupting the supersonic flexural and shear wave speeds that exist in the baseline honeycomb panel. These modified honeycomb panel structures exhibit improved transmission loss for a pre-defined diffuse field sound excitation. The models were validated by the sound transmission loss of honeycomb panels measured in the Structural Acoustic Loads and Transmission (SALT) facility at the NASA Langley Research Center. A honeycomb core panel configuration is presented exhibiting a transmission loss improvement of 3-11 dB compared to a honeycomb baseline panel over a frequency range from 170 Hz to 1000 Hz. The improved transmission loss panel configuration had a 5.1% increase in mass over the baseline honeycomb panel, and approximately twice the deflection when excited by a static force.

Grosveld, Ferdinand W.↗

NASA NTRS2008

Advanced Active-Magnetic-Bearing Thrust-Measurement System

An advanced thrust-measurement system utilizes active magnetic bearings to both (1) levitate a floating frame in all six degrees of freedom and (2) measure the levitation forces between the floating frame and a grounded frame. This system was developed for original use in measuring the thrust exerted by a rocket engine mounted on the floating frame, but can just as well be used in other force-measurement applications. This system offers several advantages over prior thrust-measurement systems based on mechanical support by flexures and/or load cells: The system includes multiple active magnetic bearings for each degree of freedom, so that by selective use of one, some, or all of these bearings, it is possible to test a given article over a wide force range in the same fixture, eliminating the need to transfer the article to different test fixtures to obtain the benefit of full-scale accuracy of different force-measurement devices for different force ranges. Like other active magnetic bearings, the active magnetic bearings of this system include closed-loop control subsystems, through which the stiffness and damping characteristics of the magnetic bearings can be modified electronically. The design of the system minimizes or eliminates cross-axis force-measurement errors. The active magnetic bearings are configured to provide support against movement along all three orthogonal Cartesian axes, and such that the support along a given axis does not produce force along any other axis. Moreover, by eliminating the need for such mechanical connections as flexures used in prior thrust-measurement systems, magnetic levitation of the floating frame eliminates what would otherwise be major sources of cross-axis forces and the associated measurement errors. Overall, relative to prior mechanical-support thrust-measurement systems, this system offers greater versatility for adaptation to a variety of test conditions and requirements. The basic idea of most prior active-magnetic-bearing force-measurement systems is to calculate levitation forces on the basis of simple proportionalities between changes in those forces and changes in feedback-controlled currents applied to levitating electromagnetic coils. In the prior systems, the effects of gap lengths on fringing magnetic fields and the concomitant effects on magnetic forces were neglected. In the present system, the control subsystems of the active magnetic bearings are coupled with a computer-based automatic calibration system running special-purpose software wherein gap-length-dependent fringing factors are applied to current and magnetic-flux-based force equations and combined with a multipoint calibration method to obtain greater accuracy.

Imlach, Joseph↗

NASA NTRS2008

Elastic Distribution of Microshutters, Measurements Obtainable on James Web Space Telescope

Spectrographic astronomy measurements in the near-infrared region will be done by functional two-dimensional microshutter arrays that are being fabricated at the NASA Goddard Space Flight Center for the James Webb Space Telescope (JWST). These micro-shutter arrays will represent the first mission-critical MEMS devices to be flown in space. JWST will use microshutter arrays to select focal plane object. 2-D programmable aperture masks of more than 200,000 elements select such space object. The use of silicon wafer material promises high efficiency and high contrast. Microshutter operation temperature is around 35K. Microshutter arrays are fabricated as close-packed silicon nitride membranes with a unit cell size of 105 x 204 micrometers. A layer of magnetic material is deposited onto each shutter. Individual shutters are equipped with a torsion flexure. Reactive ion etching (RIE) releases the shutters so they can open up to 90 degrees using the torsion flexure. Shutter rotation is initiated into a silicon support structure via an external magnetic field. Two electrically independent aluminum electrodes are deposited, one onto each shutter and another onto the support structure side-wall, permitting electrostatic latching and 2-D addressing to hold specific shutters open via external electronics.

Kletetschka, Gunther↗

NASA NTRS2009

Performance of "Waterless Concrete"

Waterless concrete consists of molten elementary sulfur and aggregate. The aggregates in a lunar environment will be lunar rocks and soil. Sulfur is present on the Moon in Troilite soil (FeS) and, by oxidation of the soil, iron and sulfur can be produced. Sulfur concrete specimens were cycled between liquid nitrogen (approx.]91 C) and room temperature (^21 C) to simulate exposure to a lunar environment. Cycled and control specimens were subsequently tested in compression at room temperatures (^21 C) and ^-101 C. Test results showed that due to temperature cycling, the compressive strength of cycled specimens was 20% of those non-cycled. This reduction in strength can be attributed to the large differences in thermal coefficients of expansion of the materials constituting the concrete which promoted cracking. Similar sulfur concrete mixtures were strengthened with short and long glass fibres. The lunar regolith simulant was melted in a 25 cc Pt- Rh crucible in a Sybron Thermoline high temperature MoSi2 furnace at melting temperatures of 1450 to 1600 C for times of 30 min to i hour. Glass fibres and small rods were pulled from the melt. The glass fibres were used to reinforce sulfur concrete plated to improve the flexural strength of the sulfur concrete. Beams strengthened with glass fibres showed to exhibit an increase in the flexural strength by as much as 45%.

Toutanji, H. A.↗

NASA NTRS2010

An Approach to Designing Passive Self-Leveling Landing Gear with Application to the Lunar Lander

Once the lunar lander has touched down on the moon problems can occur if the crew module is not level. To mitigate, compliant landing gear provide a solution that would allow the module to be leveled once it has landed on some ground slope. The work presented here uses compliant joints, or flexures, for each leg of the module and optimizes the mechanics of these flexures such that the module can be passively leveled over a range of landing slopes. Preliminary results suggest that for landing on a slope of up to 12 deg the effective slope of the module can be reduced to a maximum of 2.5 deg.

Rippere, Troy B.↗

NASA NTRS2010

An Approach to Designing Passive Self-Leveling Landing Gear with Application to the Lunar Lander

Rippere, Troy B.↗

NASA NTRS2010

Miniature Piezoelectric Shaker for Distribution of Unconsolidated Samples to Instrument Cells

The planned Mars Science Laboratory mission requires inlet funnels for channeling unconsolidated powdered samples from the sampling and sieving mechanisms into instrument test cells, which are required to reduce cross-contamination of the samples and to minimize residue left in the funnels after each sample transport. To these ends, a solid-state shaking mechanism has been created that requires low power and is lightweight, but is sturdy enough to survive launch vibration. The funnel mechanism is driven by asymmetrically mounted, piezoelectric flexure actuators that are out of the load path so that they do not support the funnel mass. Each actuator is a titanium, flextensional piezoelectric device driven by a piezoelectric stack. The stack has Invar endcaps with a half-spherical recess. The Invar is used to counteract the change in stress as the actuators are cooled to Mars ambient temperatures. A ball screw is threaded through the actuator frame into the recess to apply pre-stress, and to trap the piezoelectric stack and endcaps in flexure. During the vibration cycle of the flextensional actuator frame, the compression in the piezoelectric stack may decrease to the point that it is unstressed; however, because the ball joint cannot pull, tension in the piezoelectric stack cannot be produced. The actuators are offset at 120 . In this flight design, redundancy is required, so three actuators are used though only one is needed to assist in the movement. The funnel is supported at three contact points offset to the hexapod support contacts. The actuator surface that does not contact the ring is free to expand. Two other configurations can be used to mechanically tune the vibration. The free end can be designed to drive a fixed mass, or can be used to drive a free mass to excite impacts (see figure). Tests on this funnel mechanism show a high density of resonance modes between 1 and 20 kHz. A subset of these between 9 and 12 kHz was used to drive the CheMin actuators at 7 V peak to peak. These actuators could be driven by a single resonance, or swept through a frequency range to decrease the possibility that a portion of the funnel surface was not coincident with a nodal line (line of no displacement). The frequency of actuation can be electrically controlled and monitored and can also be mechanically tuned by the addition of tuning mass on the free end of the actuator. The devices are solid-state and can be designed with no macroscopically moving parts. This design has been tested in a vacuum at both Mars and Earth ambient temperatures ranging from 30 to 25 C

Sherrit, Stewart↗