Status of NASA Research on Projectile Shape Effects-CFRP Impact Experiments
No abstract available
Miller, J.↗
Engineering topics
Christiansen, E.
Publications and source records attributed to Christiansen, E..
Status of NASA Research on Projectile Shape Effects-Impact Simulations
Numerical evaluation of a Whipple shield has yielded a ballistic model for cylindrical projectiles. A numerical model of a Whipple shield covered with a thermal-blanket that is representative of operational shields has been developed using CTH: The shield model has been evaluated against existing test data; The projectile model has been developed from preexisting models for graphite-epoxy materials. The model has been used to identify projectile characteristics at the ballistic limit for spherical and cylindrical projectiles; Considered normal impacts of the shield at 7 kilometers per second; Varied the angle between the axis-of-symmetry and velocity vector; Varied the length to the diameter ratio over a broad range. The results have been consolidated into a generalized model that can be adapted to existing spherical models
Miller, J.↗
Highly Reliable 3-Dimensional Woven Thermal Protection System for Mars Sample Return
No abstract available
Peterson, Keith↗
Predicting the Consequences of MMOD Penetrations on the International Space Station
The threat from micrometeoroid and orbital debris (MMOD) impacts on space vehicles is often quantified in terms of the probability of no penetration (PNP). However, for large spacecraft, especially those with multiple compartments, a penetration may have a number of possible outcomes. The extent of the damage (diameter of hole, crack length or penetration depth), the location of the damage relative to critical equipment or crew, crew response, and even the time of day of the penetration are among the many factors that can affect the outcome. For the International Space Station (ISS), a Monte-Carlo style software code called Manned Spacecraft Crew Survivability (MSCSurv) is used to predict the probability of several outcomes of an MMOD penetration-broadly classified as loss of crew (LOC), crew evacuation (Evac), loss of escape vehicle (LEV), and nominal end of mission (NEOM). By generating large numbers of MMOD impacts (typically in the billions) and tracking the consequences, MSCSurv allows for the inclusion of a large number of parameters and models as well as enabling the consideration of uncertainties in the models and parameters. MSCSurv builds upon the results from NASA's Bumper software (which provides the probability of penetration and critical input data to MSCSurv) to allow analysts to estimate the probability of LOC, Evac, LEV, and NEOM. This paper briefly describes the overall methodology used by NASA to quantify LOC, Evac, LEV, and NEOM with particular emphasis on describing in broad terms how MSCSurv works and its capabilities and most significant models.
Hyde, James↗
Estimating Consequences of MMOD Penetrations on ISS
The threat from micrometeoroid and orbital debris (MMOD) impacts on space vehicles is often quantified in terms of the probability of no penetration (PNP). However, for large spacecraft, especially those with multiple compartments, a penetration may have a number of possible outcomes. The extent of the damage (diameter of hole, crack length or penetration depth), the location of the damage relative to critical equipment or crew, crew response, and even the time of day of the penetration are among the many factors that can affect the outcome. For the International Space Station (ISS), a Monte-Carlo style software code called Manned Spacecraft Crew Survivability (MSCSurv) is used to predict the probability of several outcomes of an MMOD penetration-broadly classified as loss of crew (LOC), crew evacuation (Evac), loss of escape vehicle (LEV), and nominal end of mission (NEOM). By generating large numbers of MMOD impacts (typically in the billions) and tracking the consequences, MSCSurv allows for the inclusion of a large number of parameters and models as well as enabling the consideration of uncertainties in the models and parameters. MSCSurv builds upon the results from NASA's Bumper software (which provides the probability of penetration and critical input data to MSCSurv) to allow analysts to estimate the probability of LOC, Evac, LEV, and NEOM. This paper briefly describes the overall methodology used by NASA to quantify LOC, Evac, LEV, and NEOM with particular emphasis on describing in broad terms how MSCSurv works and its capabilities and most significant models.
Evans, H.↗
Surveys of ISS Returned Hardware for MMOD Impacts
Since February 2001, the Hypervelocity Impact Technology (HVIT) group at the Johnson Space Center in Houston has performed 26 post-flight inspections on space exposed hardware that have been returned from the International Space Station. Data on 1,024 observations of MMOD damage have been collected from these inspections. Survey documentation typically includes impact feature location and size measurements as well as microscopic photography (25-200x). Sampling of impacts sites for projectile residue was performed for the largest features. Results of Scanning Electron Microscopy (SEM) analysis to discern impactor source is included in the database. This paper will summarize the post-flight MMOD inspections, and focus on two inspections in particular: (1) Pressurized Mating Adapter-2 (PMA-2) cover returned in 2015 after 1.6 years exposure with 26 observed damages, and (2) Airlock shield panels returned in 2010 after 8.7 years exposure with 58 MMOD damages. Feature sizes from the observed data are compared to predictions using the Bumper risk assessment code.
Hyde, James↗
Micrometeoroid and Orbital Debris Risk Assessment With Bumper 3
The Bumper 3 computer code is the primary tool used by NASA for micrometeoroid and orbital debris (MMOD) risk analysis. Bumper 3 (and its predecessors) have been used to analyze a variety of manned and unmanned spacecraft. The code uses NASA's latest micrometeoroid (MEM-R2) and orbital debris (ORDEM 3.0) environment definition models and is updated frequently with ballistic limit equations that describe the hypervelocity impact performance of spacecraft materials. The Bumper 3 program uses these inputs along with a finite element representation of spacecraft geometry to provide a deterministic calculation of the expected number of failures. The Bumper 3 software is configuration controlled by the NASA/JSC Hypervelocity Impact Technology (HVIT) Group. This paper will demonstrate MMOD risk assessment techniques with Bumper 3 used by NASA's HVIT Group. The Permanent Multipurpose Module (PMM) was added to the International Space Station in 2011. A Bumper 3 MMOD risk assessment of this module will show techniques used to create the input model and assign the property IDs. The methodology used to optimize the MMOD shielding for minimum mass while still meeting structural penetration requirements will also be demonstrated.
Hyde, J.↗
Development of the Space Debris Sensor
The Space Debris Sensor (SDS) is a NASA experiment scheduled to fly aboard the International Space Station (ISS) starting in 2017. The SDS is the first flight demonstration of the Debris Resistive/Acoustic Grid Orbital NASA-Navy Sensor (DRAGONS) developed and matured by the NASA Orbital Debris Program Office. The DRAGONS concept combines several technologies to characterize the size, speed, direction, and density of small impacting objects. With a minimum two-year operational lifetime, SDS is anticipated to collect statistically significant information on orbital debris ranging from 50 micron to 500 micron in size. This paper describes the SDS features and how data from the ISS mission may be used to update debris environment models. Results of hypervelocity impact testing during the development of SDS and the potential for improvement on future sensors at higher altitudes will be reviewed.
Hamilton, J.↗
Development of the Space Debris Sensor (SDS)
Debris Resistive/Acoustic Grid Orbital Navy-NASA Sensor (DRAGONS) is an impact sensor designed to detect and characterize collisions with small orbital debris: from 50 microns to greater than 1millimeter debris size detection; Characterizes debris size, speed, direction, and density. The Space Debris Sensor (SDS) is a flight demonstration of DRAGONS on the International Space Station: Approximately 1 square meter of detection area facing the ISS velocity vector; Minimum two year mission on Columbus External Payloads Facility (EPF); Minimal obstruction from ISS hardware; Development is nearing final checkout and integration with the ISS; Current launch schedule is SpaceX13, about September 2017, or SpaceX14, about Jan 2018.
Hamilton, Joe↗
Development of the Space Debris Sensor (SDS)
The Space Debris Sensor (SDS) is a NASA experiment scheduled to fly aboard the International Space Station (ISS) starting in 2018. The SDS is the first flight demonstration of the Debris Resistive/Acoustic Grid Orbital NASA-Navy Sensor (DRAGONS) developed and matured at NASA Johnson Space Center's Orbital Debris Program Office. The DRAGONS concept combines several technologies to characterize the size, speed, direction, and density of small impacting objects. With a minimum two-year operational lifetime, SDS is anticipated to collect statistically significant information on orbital debris ranging from 50 microns to 500 microns in size. This paper describes the features of SDS and how data from the ISS mission may be used to update debris environment models. Results of hypervelocity impact testing during the development of SDS and the potential for improvement on future sensors at higher altitudes will be reviewed.
Hamilton, J.↗
Development and Evaluation of the Next Generation of Meteoroid and Orbital Debris Shields
Recent events such as the Chinese anti-satellite missile test in January 2007 and the collision between a Russian Cosmos satellite and US Iridium satellite in February 2009 are responsible for a rapid increase in the population of orbital debris in Low Earth Orbit (LEO). Without active debris removal strategies the debris population in key orbits will continue to increase, requiring enhanced shielding capabilities to maintain allowable penetration risks. One of the more promising developments in recent years for meteoroid and orbital debris shielding (MMOD) is the application of open cell foams. Although shielding onboard the International Space Station is the most capable ever flown, the most proficient configuration (stuffed Whipple shield) requires an additional ~30% of the shielding mass for non-ballistic requirements (e.g. stiffeners, fasteners, etc.). Open cell foam structures provide similar mechanical performance to more traditional structural components such as honeycomb sandwich panels, as well as improved projectile fragmentation and melting as a result of repeated shocking by foam ligaments. In this paper, the preliminary results of an extensive hypervelocity impact test program on next generation MMOD shielding configurations incorporating open-cell metallic foams are reported.
Christiansen, E.↗
STS-118 Radiator Impact Damage
During the August 2007 STS-118 mission to the International Space Station, a micro-meteoroid or orbital debris (MMOD) particle impacted and completely penetrated one of shuttle Endeavour s radiator panels and the underlying thermal control system (TCS) blanket, leaving deposits on (but no damage to) the payload bay door. While it is not unusual for shuttle orbiters to be impacted by small MMOD particles, the damage from this impact is larger than any previously seen on the shuttle radiator panels. A close-up photograph of the radiator impact entry hole is shown in Figure 1, and the location of the impact on Endeavour s left-side aft-most radiator panel is shown in Figure 2. The aft radiator panel is 0.5-inches thick and consists of 0.011 inch thick aluminum facesheets on the front and back of an aluminum honeycomb core. The front facesheet is additionally covered by a 0.005 inch thick layer of silver-Teflon thermal tape. The entry hole in the silver-Teflon tape measured 8.1 mm by 6.4 mm (0.32 inches by 0.25 inches). The entry hole in the outer facesheet measured 7.4 mm by 5.3 mm (0.29 inches by 0.21 inches) (0.23 inches). The impactor also perforated an existing 0.012 inch doubler that had been bonded over the facesheet to repair previous impact damage (an example that lightning can strike the same place twice, even for MMOD impact). The peeled-back edge around the entry hole, or lip , is a characteristic of many hypervelocity impacts. High velocity impact with the front facesheet fragmented the impacting particle and caused it to spread out into a debris cloud. The debris cloud caused considerable damage to the internal honeycomb core with 23 honeycomb cells over a region of 28 mm by 26 mm (1.1 inches by 1.0 inches) having either been completely destroyed or partially damaged. Figure 3 is a view of the exit hole in the rear facesheet, and partially shows the extent of the honeycomb core damage and clearly shows the jagged petaled exit hole through the backside facesheet. The rear facesheet exit hole damage including cracks in the facesheet measures 14 mm by 14 mm (0.55 inches by 0.55 inches). The remnants of the impacting particle and radiator panel material blown through the rear facesheet hole also created two penetrations in the TCS blanket 115 mm (4.5 inches) behind the rear facesheet. Figure 4 shows these two impacts, which are located 75 mm (3 inches) apart. Some deposits of material were found on the payload bay door beneath the TCS blanket, but no additional damage occurred to the door. Figure 5 illustrates the relationship of the facesheet entry hole to the TCS blanket damage, which may indicate the direction of the impacting particle. The image on the left side of Figure 5 shows an overhead view of the damaged radiator after the facesheet holes were cored out of the panel. The entry hole location and the two underlying TCS blanket damage sites are annotated on the image. Section A-A, running through the entry hole and TCS blanket damage locations, describes a 25 angle from the longitudinal axis of the shuttle. The 2nd impact angle can be seen in section A-A on the right side of Figure 5. An average 17 angle of impact to the surface normal was derived by measuring the angles of the two damage sites in TCS blanket to the entry hole.
Lear, Dana M.↗
Investigation of MMOD Impact on STS-115 Shuttle Payload Bay Door Radiator
The Orbiter radiator system consists of eight individual 4.6 m x 3.2 m panels located with four on each payload bay door. Forward panels #1 and #2 are 2.3 cm thick while the aft panels #3 and #4 have a smaller overall thickness of 1.3 cm. The honeycomb radiator panels consist of 0.028 cm thick Aluminum 2024-T81 facesheets and Al5056-H39 cores. The face-sheets are topped with 0.005 in. (0.127 mm) silver-Teflon tape. The radiators are located on the inside of the shuttle payload bay doors, which are closed during ascent and reentry, limiting damage to the on-orbit portion of the mission. Post-flight inspections at the Kennedy Space Center (KSC) following the STS-115 mission revealed a large micrometeoroid/orbital debris (MMOD) impact near the hinge line on the #4 starboard payload bay door radiator panel. The features of this impact make it the largest ever recorded on an orbiter payload bay door radiator. The general location of the damage site and the adjacent radiator panels can be seen in Figure 2. Initial measurements of the defect indicated that the hole in the facesheet was 0.108 in. (2.74 mm) in diameter. Figure 3 shows an image of the front side damage. Subsequent observations revealed exit damage on the rear facesheet. Impact damage features on the rear facesheet included a 0.03 in. diameter hole (0.76 mm), a approx.0.05 in. tall bulge (approx.1.3 mm), and a larger approx.0.2 in. tall bulge (approx.5.1 mm) that exhibited a crack over 0.27 in. (6.8 mm) long. A large approx.1 in. (25 mm) diameter region of the honeycomb core was also damaged. Refer to Figure 4 for an image of the backside damage to the panel. No damage was found on thermal blankets or payload bay door structure under the radiator panel. Figure 5 shows the front facesheet with the thermal tape removed. Ultrasound examination indicated a maximum facesheet debond extent of approximately 1 in. (25 mm) from the entry hole. X-ray examinations revealed damage to an estimated 31 honeycomb cells with an extent of 0.85 in. x 1.1 in. (21.6 x 27.9 mm). Pieces of the radiator at and surrounding the impact site were recovered during the repair procedures at KSC. They included the thermal tape, front facesheet, honeycomb core, and rear facesheet. These articles were examined at JSC using a scanning electron microscope (SEM) with an energy dispersive x-ray spectrometer (EDS). Figure 6 shows SEM images of the entry hole in the facesheet. The asymmetric height of the lip may be attributed to projectile shape and impact angle. Numerous instances of a glass-fiber organic matrix composite were observed in the facesheet tape sample. The fibers were approximately 10 micrometers in diameter and variable lengths. EDS analysis indicated a composition of Mg, Ca, Al, Si, and O. Figures 7 and 8 present images of the fiber bundles, which were believed to be circuit board material based on similarity in fiber diameter, orientation, consistency, and composition. A test program was initiated in an attempt to simulate the observed damage to the radiator facesheet and honeycomb. Twelve test shots were performed using projectiles cut from a 1.6 mm thick fiberglass circuit board substrate panel. Results from test HITF07017, shown in figures 9 and 10, correlates with the observed impact features reasonably well. The test was performed at 4.14 km/sec with an impact angle of 45 degrees using a cylindrical projectile with a diameter and length of 1.25 mm. The fiberglass circuit board material had a density of 1.65 g/cu cm, giving a projectile mass of 2.53 mg. An analysis was performed using the Bumper code to estimate the probability of impact to the shuttle from a 1.25 mm diameter particle. Table 1 shows a 1.6% chance (impact odds = 1 in 62) of a 1.25 mm or larger MMOD impact on the radiators of the vehicle during a typical ISS mission. There is a 0.4% chance (impact odds = 1 in 260) that a 1.25 mm or larger MMOD particle would impact the RCC wing leading edge and nose cap during a typical miion. Figure 11 illustrates the vulnerable areas of the wing leading edge reinforced carbon-carbon (RCC), an area of the vehicle that is very sensitive to impact damage. The highlighted red, orange, yellow, and light green areas would be expected to experience critical damage if impacted by an OD particle such as the one that hit the RH4 radiator panel on STS-115.
Hyde, J.↗
STS-114 Micrometeoroid/Orbital Debris (MMOD) Post-Flight Assessment
NASA Johnson Space Center (JSC) personnel assisted Kennedy Space Center (KSC) inspection teams in the identification of 41 micrometeoroid/orbital debris (MMOD) impact sites on the OV-103 vehicle (Discovery) during STS-114 postflight inspections. There were 14 MMOD impacts reported on the crew module windows (Figure 1). The largest impact feature, a 6.6 mm x 5.8 mm crater on window #4, was caused by a particle with an estimated diameter of 0.22 mm (Figure 2). This impact was among the largest ever recorded on a crew module window. The window was removed and replaced. Scanning Electron Microscope/Energy Dispersive X-ray (SEM/EDX) analysis of dental mold samples from the impact site to determine particle origin was inconclusive, possibly due to contamination picked up on the ferry flight from Edwards Air Force Base to KSC. The radiators on the inside of the payload bay doors sustained 19 impacts (Figure 3) with one of the impacts causing a face sheet perforation. The 0.61 mm diameter hole was produced by a particle with an estimated diameter of 0.4 mm, which approaches the 0.5-mm critical particle diameter of the wing leading edge reinforced carbon-carbon (RCC) panel high-temperature regions (Zone 3, Figure 4) that was established during Return to Flight testing of the RCC panels. An inspection of the payload bay door exterior insulation (FRSI) revealed a 5.8 mm x 4.5 mm defect that was caused by an MMOD particle with unknown composition, as the sample obtained was contaminated. Figure 5 provides a summary of the exterior surface survey that was conducted following the STS-114 mission. Two windows were removed and replaced due to hypervelocity impact. Nineteen impacts were recorded on the payload bay door radiators, with one face sheet penetration. Three impact sites were identified on the FRSI. There were four hypervelocity impact sites detected on the wing leading edge RCC panels. One impact was detected on the top cover of the TPS sample box (TSB) payload that was mounted on a carrier in the aft portion of the payload bay.
Hyde, J.↗