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NASA NTRS2004

Foreign Object Damage in Flexure Bars of Two Gas-Turbine Grade Silicon Nitrides

Foreign object damage (FOD) behavior of two commercial gas-turbine grade silicon nitrides, AS800 and SN282, was determined at ambient temperature through strength testing of flexure test specimens impacted by steel ball projectiles with a diameter of 1.59 mm in a velocity range from 220 to 440 m/s. AS800 silicon nitride exhibited a greater FOD resistance than SN282, primarily due to its greater value of fracture toughness (K(sub lc)). The use of an additional equiaxed, fine-grained silicon nitride (NC132) showed that fracture toughness was a key material parameter affecting FOD resistance. The observed damages generated by projectile impact were typically in the forms of well- or ill-developed ring and cone cracks with little presence of radial cracks.

Choi, Sung R.↗
NASA NTRS2005

Atmospheric Fragmentation of the Canyon Diablo Meteoroid

About 50 kyr ago the impact of an iron meteoroid excavated Meteor Crater, Arizona, the first terrestrial structure widely recognized as a meteorite impact crater. Recent studies of ballistically dispersed impact melts from Meteor Crater indicate a compositionally unusually heterogeneous impact melt with high SiO2 and exceptionally high (10 to 25% on average) levels of projectile contamination. These are observations that must be explained by any theoretical modeling of the impact event. Simple atmospheric entry models for an iron meteorite similar to Canyon Diablo indicate that the surface impact speed should have been around 12 km/s [Melosh, personal comm.], not the 15-20 km/s generally assumed in previous impact models. This may help explaining the unusual characteristics of the impact melt at Meteor Crater. We present alternative initial estimates of the motion in the atmosphere of an iron projectile similar to Canyon Diablo, to constraint the initial conditions of the impact event that generated Meteor Crater.

Pierazzo, E.↗
NASA NTRS2000

Effects of Various Heat Treatments on the Ballistic Impact Properties of Inconel 718 Investigated

Uncontained failures of aircraft engine fan blades are serious events that can cause equipment damage and loss of life. Federal Aviation Administration (FAA) certification requires that all engines demonstrate the ability to contain a released fan blade with the engine running at full power. However, increased protection generally comes at the expense of weight. Proper choice of materials is therefore imperative to an optimized design. The process of choosing a good casing material is done primarily through trial and error. This costly procedure could be minimized if there was a better understanding of the relationships among static material properties, impact properties, and failure mechanisms. This work is part of a program being conducted at the NASA Glenn Research Center at Lewis Field to study these relationships. Ballistic impact tests were conducted on flat, square sheets of Inconel 718 that had been subjected to different heat treatments. Two heat treatments and the as-received condition were studied. In addition, results were compared with those from an earlier study involving a fourth heat treatment. The heat treatments were selected on the basis of their effects on the static tensile properties of the material. The impact specimens used in this study were 17.8-cm square panels that were centered and clamped over a 15.2-cm square hole in a 1.27-cm-thick steel plate. Three nominal plate thickness dimensions were studied, 1.0, 1.8, and 2.0 mm. For each thickness, all the specimens were taken from the same sheet of material. The projectile was a Ti-6Al-4V cylinder with a length of 25.4 mm, a diameter of 12.7 mm, and a mass ranging from 14.05 to 14.20 g. The projectiles were accelerated toward the specimens at normal incidence using a gas gun with a 2-m-long, 12.7-mm inner-diameter barrel. The ballistic limit for each heat treatment condition and thickness was determined by conducting a number of impact tests that bracketed as closely as possible the velocity required to penetrate the specimen.

Pereira, J. Michael↗
NASA NTRS2003

Superior Ballistic Impact Resistance Achieved by the Co-Base Alloy Haynes 25

The fan case in a jet engine is required to contain a fan blade in the rare event of a blade loss during operation. Because of its function, the fan case is the largest structural component in high-bypass-ratio turbofan engines used in commercial aircraft. Therefore, the use of lighter and stronger materials would be advantageous in most engines and is practically a necessity in the latest generation of high-bypass engines. Small panels, 7 in. wide by 7 in. long, of a number of metallic alloys were impact tested at room temperature with a 0.50-caliber blunt-nose titanium alloy projectile at the NASA Glenn Research Center (ref. 1). These metallic systems included several high-strength aluminum (Al) alloys, Al-based laminates, aluminum metal matrix composites (Al-MMCs), nickel-base superalloys (Inconel 718 and 625), several titanium (Ti) alloys in different heat treated conditions, 304L stainless steel, a stainless-steel-based laminate, and a high strength steel (Nitronic 60). It was determined that a simple Co-base alloy (Haynes 25) had the best impact resistance on an areal weight basis. Haynes 25 was at least 10 percent better than IMI 550, the best titanium alloy tested to date, and it was far superior to other metals, especially at higher impact velocities (greater than 1100 ft/sec). Because this material could be ideal for fan containment applications in supersonic aircraft as a replacement for titanium, impact tests were also conducted at 371 oC and compared with results from alloys tested at elevated temperature under previous programs (i.e., Inconel 718, Ti-6242, M-152, Timetal 21S, and Aeromet 100). Although cobalt-base alloys are used in some high-temperature engine applications, to our knowledge they are not used in any containment systems. Advantages of cobalt over titanium include lower cost, easier processing, better high-temperature strength, and no fire hazard if tip rub occurs. Future plans include testing of lightweight sandwich panels with Haynes 25 as a core material in the form of a foam or lattice block structure and scaling up the current tests by using blade-simulating projectiles impacting large plates and half rings.

Hebsur, Mohan G.↗
NASA NTRS2006

A Mesh Refinement Study on the Impact Response of a Shuttle Leading-Edge Panel Finite Element Simulation

A study was performed to examine the influence of varying mesh density on an LS-DYNA simulation of a rectangular-shaped foam projectile impacting the space shuttle leading edge Panel 6. The shuttle leading-edge panels are fabricated of reinforced carbon-carbon (RCC) material. During the study, nine cases were executed with all possible combinations of coarse, baseline, and fine meshes of the foam and panel. For each simulation, the same material properties and impact conditions were specified and only the mesh density was varied. In the baseline model, the shell elements representing the RCC panel are approximately 0.2-in. on edge, whereas the foam elements are about 0.5-in. on edge. The element nominal edge-length for the baseline panel was halved to create a fine panel (0.1-in. edge length) mesh and doubled to create a coarse panel (0.4-in. edge length) mesh. In addition, the element nominal edge-length of the baseline foam projectile was halved (0.25-in. edge length) to create a fine foam mesh and doubled (1.0-in. edge length) to create a coarse foam mesh. The initial impact velocity of the foam was 775 ft/s. The simulations were executed in LS-DYNA for 6 ms of simulation time. Contour plots of resultant panel displacement and effective stress in the foam were compared at four discrete time intervals. Also, time-history responses of internal and kinetic energy of the panel, kinetic and hourglass energy of the foam, and resultant contact force were plotted to determine the influence of mesh density.

Fasanella, Edwin L.↗
NASA NTRS2007

Analysis of Cometary Dust Impact Residues in the Aluminum Foil Craters of Stardust

In January 2006, the sample return capsule from NASA s Stardust spacecraft successfully returned to Earth after its seven year mission to comet Wild-2. While the principal capture medium for comet dust was low-density graded silica aerogel, the 1100 series aluminum foil (approximately 100 m thick) which wrapped around the T6064 aluminum frame of the sample tray assembly (STA) contains micro-craters that constitute an additional repository for Wild-2 dust. Previous studies of similar craters on spacecraft surfaces, e.g. the Long Duration Exposure Facility (LDEF), have shown that impactor material can be preserved for elemental and mineralogical characterization, although the quantity of impact residue in Stardust craters far exceeds previous missions. The degree of shock-induced alteration experienced by the Wild-2 particles impacting on foil will generally be greater than for those captured in the low-density aerogel. However, even some of the residues found in LDEF craters showed not only survival of crystalline silicates but even their solar flare tracks, which are extremely fragile structures and anneal at around 600 C. Laboratory hypervelocity experiments, using analogues of Wild-2 particles accelerated into flight-grade foils under conditions close to those of the actual encounter, showed retention of abundant projectile residues at the Stardust encounter velocity of 6.1 km/s. During the preliminary examination (PE) of the returned foils, using optical and electron microscopy studies, a diverse range in size and morphologies of micro-craters was identified. In this abstract we consider the state of residue preservation in a diverse range of craters with respect to their elemental composition and inferred mineralogy of the original projectiles.

Graham, G. A.↗
NASA NTRS2007

Investigation of MMOD Impact on Shuttle Payload Bay Door Radiator

Post flight inspections on the Space Shuttle Atlantis conducted after the STS-115 mission revealed a 0.11 inch (2.8 mm) hole in the outer facesheet of the starboard payload bay door radiator panel #4. This hole is the possible result of micrometeoroid/orbiting debris (MMOD) impact. The payload bay door radiators in this region are 0.5 inch (12.7 mm) thick aluminum honeycomb with 0.011 in (0.279 mm) thick aluminum facesheets topped with 0.005 in (0.127 mm) silver-Teflon tape. Inner facesheet damage included a 0.267 in (6.78 mm) long through crack with measurable deformation in the area of 0.2 in (5.1 mm). There was also a 0.031 in (0.787 mm) diameter hole in the rear facesheet. A large approximately 1 in (25 mm) diameter region of honeycomb was also destroyed. Since the radiators are located on the inside of the shuttle payload bay doors which are closed during ascent and reentry, the damage could only have occurred during the on-orbit portion of the mission. This paper will document the data collected from the impact site and will include results of the SEM/EDX analysis. Evidence will be presented that suggests a source of the impact as well as an analysis of the impact site features that indicate projectile directionality. Results of hypervelocity impact testing on representative samples in an attempt to simulate the impact event will be presented and discussed. Finally, the results of a study showing the regions of the orbiter vehicle that would be vulnerable to an equivalent projectile will be given.

Hyde, J. L.↗
NASA NTRS2007

Comparison of Integrated Radiation Transport Models with TEPC Measurements for the Average Quality Factors in Spaceflights

The purpose of this work is to test our theoretical model for the interpretation of radiation data measured in space. During the space missions astronauts are exposed to the complex field of radiation type and kinetic energies from galactic cosmic rays (GCR), trapped protons, and sometimes solar particle events (SPEs). The tissue equivalent proportional counter (TEPC) is a simple time-dependent approach for radiation monitoring for astronauts on board the International Space Station. Another and a newer approach to Microdosimetry is the use of silicon-on-insulator (SOI) technology launched on the MidSTAR-1 mission in low Earth orbit (LEO). In the radiation protection practice, the average quality factor of a radiation field is defined as a function of linear energy transfer (LET), Q(sub ave)(LET). However, TEPC measures the average quality factor as a function of the lineal energy y, Q(sub ave)(y), defined as the average energy deposition in a volume divided by the average chord length of the volume. Lineal energy, y, deviates from LET due to energy straggling, delta-ray escape or entry, and nuclear fragments produced in the detector volume. Monte Carlo track structure simulation was employed to obtain the response of a TEPC irradiated with charged particle for an equivalent site diameter of 1 micron of wall-less counter. The calculated data of the energy absorption in the wall-less counter were compiled for various y values for several ion types at various discrete projectile energy levels. For the simulation of TEPC response from the mixed radiation environments inside a spacecraft, such as, Space Shuttle and International Space Station, the complete microdosimetric TEPC response, f( y, E, Z), were calculated with the Monte Carlo theoretical results by using the first order Lagrangian interpolation for a monovariate function at a given y value (y = 0.1 keV/micron 5000 keV/micron) at any projectile energy level (E = 0.01 MeV/u to 50,000 MeV/u) of each specific radiation type (Z = 1 to 28). Because the anomalous response has been observed at large event sizes in the experiment due to the escape of energy out of sensitive volume by delta-rays and the entry of delta-rays from the high-density wall into the low-density gas-volume cavity, Monte Carlo simulation was also made for the response of a walled-TEPC with wall thickness 2 mm and density 1 g/cm(exp 3). The radius of cavity was set to 6.35 mm and a gas density 7.874 x 10(exp -5) g/cm(exp 3). The response of the walled- and the wall-less counters were compared. The average quality factor Q(sub ave)(y) for trapped protons on STS-89 demonstrated the good agreement between the model calculations and flight TEPC data as shown. Using an integrated space radiation model (this includes the transport codes HZETRN and BRYNTRN, the quantum nuclear interaction model QMSFRG) and the resultant response distribution functions of walled-TEPC from Monte-Carlo track simulations, we compared model calculations with walled-TEPC measurements from NASA missions in LEO and made predictions for the lunar and the Mars missions. The Q(sub ave)(y) values for the trapped or the solar protons ranged from 1.9-2.5. This over-estimates the Qave(LET) values which ranged from 1.4-1.6. Both quantities increase with shield thickness due to nuclear fragmentation. The Q(sub ave)(LET) for the complete GCR spectra was found to be 3.5-4.5, while flight TEPCs measured 2.9-3.4 for Q(sub ave)(y). The GCR values are decreasing with the shield thickness. Our analysis for a proper interpretation of data supports the use of TEPCs for monitoring space radiation environment.

Kim, Myung-Hee Y.↗
NASA NTRS2007

Measurement of Primary Ejecta From Normal Incident Hypervelocity Impact on Lunar Regolith Simulant

The National Aeronautics and Space Administration (NASA) continues to make progress toward long-term lunar habitation. Critical to the design of a lunar habitat is an understanding of the lunar surface environment. A subject for further definition is the lunar primary ejecta environment. The document NASA SP-8013 was developed for the Apollo program and is the latest definition of the primary ejecta environment. There is concern that NASA SP-8013 may over-estimate the lunar primary ejecta environment. NASA's Meteoroid Environment Office (MEO) has initiated several tasks to improve the accuracy of our understanding of the lunar surface primary ejecta environment. This paper reports the results of experiments on projectile impact into pumice targets, simulating lunar regolith. The Ames Vertical Gun Range (AVGR) was used to accelerate spherical Pyrex projectiles of 0.29g to velocities ranging between 2.5 km/s and 5.18 km/s. Impact on the pumice target occurred at normal incidence. The ejected particles were detected by thin aluminum foil targets placed around the pumice target in a 0.5 Torr vacuum. A simplistic technique to characterize the ejected particles was formulated. Improvements to this technique will be discussed for implementation in future tests.

Edwards, David L.↗
NASA NTRS2007

Reinforced Carbon-Carbon Subcomponent Flat Plate Impact Testing for Space Shuttle Orbiter Return to Flight

Following the tragedy of the Space Shuttle Columbia on February 1, 2003, a major effort commenced to develop a better understanding of debris impacts and their effect on the Space Shuttle subsystems. An initiative to develop and validate physics-based computer models to predict damage from such impacts was a fundamental component of this effort. To develop the models it was necessary to physically characterize Reinforced Carbon-Carbon (RCC) and various debris materials which could potentially shed on ascent and impact the Orbiter RCC leading edges. The validated models enabled the launch system community to use the impact analysis software LS DYNA to predict damage by potential and actual impact events on the Orbiter leading edge and nose cap thermal protection systems. Validation of the material models was done through a three-level approach: fundamental tests to obtain independent static and dynamic material model properties of materials of interest, sub-component impact tests to provide highly controlled impact test data for the correlation and validation of the models, and full-scale impact tests to establish the final level of confidence for the analysis methodology. This paper discusses the second level subcomponent test program in detail and its application to the LS DYNA model validation process. The level two testing consisted of over one hundred impact tests in the NASA Glenn Research Center Ballistic Impact Lab on 6 by 6 in. and 6 by 12 in. flat plates of RCC and evaluated three types of debris projectiles: BX 265 External Tank foam, ice, and PDL 1034 External Tank foam. These impact tests helped determine the level of damage generated in the RCC flat plates by each projectile. The information obtained from this testing validated the LS DYNA damage prediction models and provided a certain level of confidence to begin performing analysis for full-size RCC test articles for returning NASA to flight with STS 114 and beyond.

Melis, Matthew E.↗
NASA NTRS2008

Comparison of Ejecta Distributions from Normal Incident Hypervelocity Impact on Lunar Regolith Simulant

The National Aeronautics and Space Administration (NASA) is progressing toward long-term lunar habitation. Critical to the design of a lunar habitat is an understanding of the lunar surface environment; of specific importance is the primary meteoroid and subsequent ejecta environment. The document, NASA SP-8013, was developed for the Apollo program and is the latest definition of the ejecta environment. There is concern that NASA SP-8013 may over-estimate the lunar ejecta environment. NASA's Meteoroid Environment Office (MEO) has initiated several tasks to improve the accuracy of our understanding of the lunar surface ejecta environment. This paper reports the results of experiments on projectile impact into powered pumice and unconsolidated JSC-1A Lunar Mare Regolith stimulant (JSC-1A) targets. The Ames Vertical Gun Range (AVGR) was used to accelerate projectiles to velocities in excess of 5 km/s and impact the targets at normal incidence. The ejected particles were detected by thin aluminum foil targets placed around the impact site and angular distributions were determined for ejecta. Comparison of ejecta angular distribution with previous works will be presented. A simplistic technique to characterize the ejected particles was formulated and improvements to this technique will be discussed for implementation in future tests.

Edwards, David L.↗
NASA NTRS2008

High Velocity Linear Induction Launcher with Exit-Edge Compensation for Testing of Aerospace Components

Advances in ultra high speed linear induction electromagnetic launchers over the past decade have focused on magnetic compensation of the exit and entry-edge transient flux wave to produce efficient and compact linear electric machinery. The paper discusses two approaches to edge compensation in long-stator induction catapults with typical end speeds of 150 to 1,500 m/s. In classical linear induction machines, the exit-edge effect is manifest as two auxiliary traveling waves that produce a magnetic drag on the projectile and a loss of magnetic flux over the main surface of the machine. In the new design for the Stator Compensated Induction Machine (SCIM) high velocity launcher, the exit-edge effect is nulled by a dual wavelength machine or alternately the airgap flux is peaked at a location prior to the exit edge. A four (4) stage LIM catapult is presently being constructed for 180 m/s end speed operation using double-sided longitudinal flux machines. Advanced exit and entry edge compensation is being used to maximize system efficiency, and minimize stray heating of the reaction armature. Each stage will output approximately 60 kN of force and produce over 500 G s of acceleration on the armature. The advantage of this design is there is no ablation to the projectile and no sliding contacts, allowing repeated firing of the launcher without maintenance of any sort. The paper shows results of a parametric study for 500 m/s and 1,500 m/s linear induction launchers incorporating two of the latest compensation techniques for an air-core stator primary and an iron-core primary winding. Typical thrust densities for these machines are in the range of 150 kN/sq.m. to 225 kN/sq.m. and these compete favorably with permanent magnet linear synchronous machines. The operational advantages of the high speed SCIM launcher are shown by eliminating the need for pole-angle position sensors as would be required by synchronous systems. The stator power factor is also improved.

Kuznetsov, Stephen↗
NASA NTRS2006

Laboratory Simulation of Impacts upon Aluminum Foils of the Stardust Spacecraft: Calibration of Dust Particle Size from Comet Wild 2

Metallic aluminium alloy foils exposed on the forward, comet-facing surface of the aerogel tray on the Stardust spacecraft are likely to have been impacted by the same cometary particle population as the dedicated impact sensors and the aerogel collector. The ability of soft aluminium alloy to record hypervelocity impacts as bowl-shaped craters offers an opportunistic substrate for recognition of impacts by particles of a wide potential size range. In contrast to impact surveys conducted on samples from low Earth orbit, the simple encounter geometry for Stardust and Wild 2, with a known and constant spacecraft-particle relative velocity and effective surface-perpendicular impact trajectories, permits closely comparable simulation in laboratory experiments. For a detailed calibration programme we have selected a suite of spherical glass projectiles of uniform density and hardness characteristics, with well-documented particle size range from 10 microns to nearly 100 microns. Light gas gun buckshot firings of these particles at approximately 6km s)exp -1) onto samples of the same foil as employed on Stardust have yielded large numbers of craters. Scanning electron microscopy of both projectiles and impact features has allowed construction of a calibration plot, showing a linear relationship between impacting particle size and impact crater diameter. The close match between our experimental conditions and the Stardust mission encounter parameters should provide another opportunity to measure particle size distributions and fluxes close to the nucleus of Wild 2, independent of the active impact detector instruments aboard the Stardust spacecraft.

Kearsley, A. T.↗
NASA NTRS2008

Analysis of Regolith Simulant Ejecta Distributions from Normal Incident Hypervelocity Impact

The National Aeronautics and Space Administration (NASA) has established the Constellation Program. The Constellation Program has defined one of its many goals as long-term lunar habitation. Critical to the design of a lunar habitat is an understanding of the lunar surface environment; of specific importance is the primary meteoroid and subsequent ejecta environment. The document, NASA SP-8013 'Meteoroid Environment Model Near Earth to Lunar Surface', was developed for the Apollo program in 1969 and contains the latest definition of the lunar ejecta environment. There is concern that NASA SP-8013 may over-estimate the lunar ejecta environment. NASA's Meteoroid Environment Office (MEO) has initiated several tasks to improve the accuracy of our understanding of the lunar surface ejecta environment. This paper reports the results of experiments on projectile impact into powdered pumice and unconsolidated JSC-1A Lunar Mare Regolith simulant targets. Projectiles were accelerated to velocities between 2.45 and 5.18 km/s at normal incidence using the Ames Vertical Gun Range (AVGR). The ejected particles were detected by thin aluminum foil targets strategically placed around the impact site and angular ejecta distributions were determined. Assumptions were made to support the analysis which include; assuming ejecta spherical symmetry resulting from normal impact and all ejecta particles were of mean target particle size. This analysis produces a hemispherical flux density distribution of ejecta with sufficient velocity to penetrate the aluminum foil detectors.

Edwards, David L.↗
NASA NTRS2009

Micro-Satellite Impact Tests to Investigate Multi-Layer Insulation Fragments

This paper summarizes two satellite impact experiments completed in 2008. The objective of the experiments is to investigate the physical properties of satellite fragments, including those originated from Multi-Layer Insulation (MLI) and solar panels. The ultimate goal is to use the results to improve the NASA Standard Breakup Model. The targets were two cubic micro-satellites, 20 cm by 20 cm by 20 cm in size, and approximately 1,500 g in mass. The main structure of each micro-satellite was composed of five layers; the top and bottom layers and three internal layers parallel to the top and bottom layers, plus four side panels. The top layer was equipped with solar cells that was mounted to an aluminum honeycomb sandwich panel with CFRP face sheets. The four side panels and the bottom layer are all covered with MLI. The two satellite impact experiments were conducted using the two-stage light gas gun at the Kyushu Institute of Technology in Kitakyusyu, Japan. For the first experiment (labeled Shot F), the satellite was oriented in such a way that the solar panel was facing the incoming projectile, a 39.3 g aluminum alloy solid sphere. For the second experiment (labeled Shot R), the satellite was oriented so that the solar panel was on the opposite side of the impact surface. The projectile used in the second shot was a 39.2 g aluminum alloy solid sphere. The impact speeds of Shot F and Shot R were 1.74 km/s and 1.78 km/s, respectively. The ratio of the impact kinetic energy to satellite mass for the two experiments was about 40 J/g. Both target satellites were completely fragmented, although there were noticeable differences in the characteristics of the fragments. Approximately 1,800 fragments were collected from Shot F but only 1,000 fragments were collected from Shot R. This difference primarily comes from the number of needle-like CFRP and MLI fragments. The difference in CFRP pieces depends on how the CFRP panels were fragmented. Regarding the MLI pieces, a significant difference in size and number can be observed. The largest MLI pieces in Shot F are almost of the same size as the side panels, whereas those in Shot R are larger by about a factor of two. The collected fragments and MLI pieces will be measured and analyzed using the same method as described in the NASA Standard Breakup Model. This paper will present: (1) the area-to-mass ratio, size, and mass distributions of the fragments, and (2) the differences in fragment properties between Shot F and Shot R.

Liou, J.C.↗
NASA NTRS2009

Investigation of Orbital Debris Impacts on Shuttle Radiator Panels

This paper documents the data collected from two hypervelocity micro-meteoroid orbital debris (MMOD) impact events where the shuttle payload bay door radiator sandwich panel was completely perforated. Scanning Electron Microscope/Energy-Dispersive x-ray Spectroscopy (SEM/EDS) analysis of impact residue provided evidence to identify the source of each impact. Impact site features that indicate projectile directionality are discussed, along with hypervelocity impact testing on representative samples conducted to simulate the impact event. The paper provides results of a study of impact risks for the size of particles that caused the MMOD damage and the regions of the orbiter vehicle that would be vulnerable to an equivalent projectile

Hyde, James L.↗
NASA NTRS2009

Design and Testing of Braided Composite Fan Case Materials and Components

Triaxial braid composite materials are beginning to be used in fan cases for commercial gas turbine engines. The primary benefit for the use of composite materials is reduced weight and the associated reduction in fuel consumption. However, there are also cost benefits in some applications. This paper presents a description of the braided composite materials and discusses aspects of the braiding process that can be utilized for efficient fabrication of composite cases. The paper also presents an approach that was developed for evaluating the braided composite materials and composite fan cases in a ballistic impact laboratory. Impact of composite panels with a soft projectile is used for materials evaluation. Impact of composite fan cases with fan blades or blade-like projectiles is used to evaluate containment capability. A post-impact structural load test is used to evaluate the capability of the impacted fan case to survive dynamic loads during engine spool down. Validation of these new test methods is demonstrated by comparison with results of engine blade-out tests.

Roberts, Gary D.↗
NASA NTRS2009

Meteoroid and Orbital Debris Threats to NASA's Docking Seals: Initial Assessment and Methodology

The Crew Exploration Vehicle (CEV) will be exposed to the Micrometeoroid Orbital Debris (MMOD) environment in Low Earth Orbit (LEO) during missions to the International Space Station (ISS) and to the micrometeoroid environment during lunar missions. The CEV will be equipped with a docking system which enables it to connect to ISS and the lunar module known as Altair; this docking system includes a hatch that opens so crew and supplies can pass between the spacecrafts. This docking system is known as the Low Impact Docking System (LIDS) and uses a silicone rubber seal to seal in cabin air. The rubber seal on LIDS presses against a metal flange on ISS (or Altair). All of these mating surfaces are exposed to the space environment prior to docking. The effects of atomic oxygen, ultraviolet and ionizing radiation, and MMOD have been estimated using ground based facilities. This work presents an initial methodology to predict meteoroid and orbital debris threats to candidate docking seals being considered for LIDS. The methodology integrates the results of ground based hypervelocity impacts on silicone rubber seals and aluminum sheets, risk assessments of the MMOD environment for a variety of mission scenarios, and candidate failure criteria. The experimental effort that addressed the effects of projectile incidence angle, speed, mass, and density, relations between projectile size and resulting crater size, and relations between crater size and the leak rate of candidate seals has culminated in a definition of the seal/flange failure criteria. The risk assessment performed with the BUMPER code used the failure criteria to determine the probability of failure of the seal/flange system and compared the risk to the allotted risk dictated by NASA's program requirements.

deGroh, Henry C., III↗