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NASA NTRS1966

Momentum Transfer and Cratering Phenomena Associated With the Impact of Aluminum Spheres Into Thick Aluminum Targets at Velocities to 24,000 Feet Per Second

Spheres of 2017-T4 aluminum with diameters of, 1/16, 1/8, and ½ inch were launched into semi-infinite targets of 2024-T4 and 1100-0 aluminum. The momentum transfer during impact was measured by means of a simple ballistic pendulum to which the target was attached. The ratio of target momentum to projectile momentum increased with increasing projectile velocity and with projectile diameter. Mass loss due to cratering, penetration, and crater diameter were also obtained for these experiments. It was noted that the variation of these parameters was also a function of the projectile diameter. No conclusive evidence was uncovered to explain this apparent scale effect. However, close examination of the various craters suggests that the cratering process is, in fact, different for the different size projectiles. The cause for this difference has yet to be resolved.

MOMENTUM TRANSFER↗

NASA NTRS1992

Applications of the ram accelerator to hypervelocity aerothermodynamic testing

A ram accelerator used as a hypervelocity launcher for large-scale aeroballistic range applications in hypersonics and aerodynamics research is presented. It is an in-bore ramjet device in which a projectile shaped like the centerbody of a supersonic ramjet is propelled down a stationary tube filled with a tailored combustible gas mixture. Ram accelerator operation has been demonstrated at 39 mm and 90 mm bores, supporting the proposition that this launcher concept can be scaled up to very large bore diameters of the order of 30-60 cm. It is concluded that high quality data obtained from the tube wall and projectile during the aceleration process itself are very useful for understanding aerothermodynamics of hypersonic flow in general, and for providing important CFD validation benchmarks.

Bruckner, A. P.↗

NASA NTRS1992

Investigation of advanced propulsion technologies: The RAM accelerator and the flowing gas radiation heater

The two principal areas of advanced propulsion investigated are the ram accelerator and the flowing gas radiation heater. The concept of the ram accelerator is presented as a hypervelocity launcher for large-scale aeroballistic range applications in hypersonics and aerothermodynamics research. The ram accelerator is an in-bore ramjet device in which a projectile shaped like the centerbody of a supersonic ramjet is propelled in a stationary tube filled with a tailored combustible gas mixture. Combustion on and behind the projectile generates thrust which accelerates it to very high velocities. The acceleration can be tailored for the 'soft launch' of instrumented models. The distinctive reacting flow phenomena that have been observed in the ram accelerator are relevant to the aerothermodynamic processes in airbreathing hypersonic propulsion systems and are useful for validating sophisticated CFD codes. The recently demonstrated scalability of the device and the ability to control the rate of acceleration offer unique opportunities for the use of the ram accelerator as a large-scale hypersonic ground test facility. The flowing gas radiation receiver is a novel concept for using solar energy to heat a working fluid for space power or propulsion. Focused solar radiation is absorbed directly in a working gas, rather than by heat transfer through a solid surface. Previous theoretical analysis had demonstrated that radiation trapping reduces energy loss compared to that of blackbody receivers, and enables higher efficiencies and higher peak temperatures. An experiment was carried out to measure the temperature profile of an infrared-active gas and demonstrate the effect of radiation trapping. The success of this effort validates analytical models of heat transfer in this receiver, and confirms the potential of this approach for achieving high efficiency space power and propulsion.

Bruckner, A. P.↗

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 NTRS2004

Hypervelocity Impact Energy Loss and Track Shape in Aerogels: Theory and Experiment

We have developed and tested a theoretical model of impact cratering in aerogels that derives the impact crater dimensions directly from the energy and momentum deposition of the projectile. In this paper, our work is outlined.

Dominguez, G.↗

NASA NTRSDate not supplied

An Efficient and Effective Light Gas Gun Design for Millimeter-Scale Hypervelocity Testing

This paper serves to document a design for a light gas gun that was used for 3 decades at NASA’s Langley Research Center. By adapting readily-available equipment and supplies, this apparatus was capable of accelerating projectiles at up to 9 kilometers per second. The lack of a projectile-carrying sabot sacrifices some adaptability in projectile size, but results in a design that can be cleaned and cycled with far less labor input. Overall limitation on the projectile impact energies attainable effectively limit the applications to assessing the relatively thin shields employed for robotic spacecraft (less useful for those used for manned missions). The original gun has been decommissioned, but it is believed that this design has application for robotic spacecraft shield testing in the modern world, allowing cost-effective generation of larger sample sizes than with conventional sabot-based light gas guns.

Scott Hull↗

NASA NTRSDate not supplied

An Efficient and Effective Light Gas Gun Design for Millimeter-scale Hypervelocity Testing

This paper serves to document a design for a light gas gun that was used for 3 decades at NASA’s Langley Research Center. By adapting readily-available equipment and supplies, this apparatus was capable of accelerating projectiles at up to 9 kilometers per second (when new). The lack of a projectile-carrying sabot sacrifices some adaptability in projectile size, but results in a design that can be cleaned and cycled with far less labor input. Overall limitation on the projectile impact energies attainable effectively limit the applications to assessing the relatively thin shields employed for robotic spacecraft (less useful for those used for manned missions). The original gun has been decommissioned, but it is believed that this design has application for robotic spacecraft shield testing in the modern world, allowing cost-effective generation of larger sample sizes than with conventional sabot-based light gas guns.

Scott Hull↗

NASA NTRS1971

An analytical and experimental study of the behavior of semi-infinite metal targets under hypervelocity impact

The material strength and strain rate effects associated with the hypervelocity impact problem were considered. A yield criterion involving the second and third invariants of the stress deviator and a strain rate sensitive constitutive equation were developed. The part of total deformation which represents change in shape is attributable to the stress deviator. Constitutive equation is a means for analytically describing the mechanical response of a continuum under study. The accuracy of the yield criterion was verified utilizing the published two and three dimensional experimental data. The constants associated with the constitutive equation were determined from one dimensional quasistatic and dynamic experiments. Hypervelocity impact experiments were conducted on semi-infinite targets of 1100 aluminum, 6061 aluminum alloy, mild steel, and commercially pure lead using spherically shaped and normally incident pyrex projectiles.

Chakrapani, B.↗

NASA NTRS1972

Appraisal of UTIAS implosion-driven hypervelocity launchers and shock tubes.

A critical appraisal is made of the design, research, development, and operation of the novel UTIAS implosion-driven hypervelocity launchers and shock tubes. Explosively driven (PbN6-lead azide, PETN-pentaerythritetetranitrate) implosions in detonating stoichiometric hydrogen-oxygen mixtures have been successfully developed as drivers for hypervelocity launchers and shock tubes in a safe and reusable facility. Intense loadings at very high calculated pressures, densities, and temperatures, at the implosion center, cause severe problems with projectile integrity. Misalignment of the focal point can occur and add to the difficulty in using small caliber projectiles. In addition, the extreme driving conditions cause barrel expansion, erosion, and possible gas leakage from the base to the head of the projectile which cut the predicted muzzle velocities to half or a third of the lossless calculated values. However, in the case of a shock-tube operation these difficulties are minimized or eliminated and the possibilities of approaching Jovian reentry velocities are encouraging.

Glass, I. I.↗

NASA NTRS2015

Orbital Debris Assesment Tesing in the AEDC Range G

The space environment presents many hazards for satellites and spacecraft. One of the major hazards is hypervelocity impacts from uncontrolled man-made space debris. Arnold Engineering Development Complex (AEDC), The National Aeronautics and Space Administration (NASA), The United States Air Force Space and Missile Systems Center (SMC), the University of Florida, and The Aerospace Corporation configured a large ballistic range to perform a series of hypervelocity destructive impact tests in order to better understand the effects of space collisions. The test utilized AEDC's Range G light gas launcher, which is capable of firing projectiles up to 7 km/s. A non-functional full-scale representation of a modern satellite called the DebriSat was destroyed in the enclosed range enviroment. Several modifications to the range facility were made to ensure quality data was obtained from the impact events. The facility modifcations were intended to provide a high impact energy to target mass ratio (>200 J/g), a non-damaging method of debris collection, and an instrumentation suite capable of providing information on the physics of the entire imapct event.

Polk, Marshall↗

NASA NTRS1994

Cratering and penetration experiments in teflon targets at velocities from 1 to 7 km/s

Approximately 20 sq m of protective thermal blankets, largely composed of Teflon, were retrieved from the Long Duration Exposure Facility after the spacecraft spent approximately 5.7 years in space. Examination of these blankets revealed that they contained thousands of hypervelocity impact features ranging from micron-sized craters to penetration holes several millimeters in diameter. We conducted impact experiments to reproduce such features and to understand the relationships between projectile size and the resulting crater or penetration hole diameter over a wide range of impact velocities. Such relationships are needed to derive the size and mass frequency distribution and flux of natural and man-made particles in low-earth orbit. Powder propellant and light-gas guns were used to launch soda-lime glass spheres into pure Teflon targets at velocities ranging from 1 to 7 km/s. Target thickness varied over more than three orders of magnitude from finite halfspace targets to very thin films. Cratering and penetration of massive Teflon targets is dominated by brittle failure and the development of extensive spall zones at the target's front and, if penetrated, the target's rear side. Mass removal by spallation at the back side of Teflon targets may be so severe that the absolute penetration hole diameter can become larger than that of a standard crater. The crater diameter in infinite halfspace Teflon targets increases, at otherwise constant impact conditions, with encounter velocity by a factor of V (exp 0.44). In contrast, the penetration hole size in very thin foils is essentially unaffected by impact velocity. Penetrations at target thicknesses intermediate to these extremes will scale with variable exponents of V. Our experimental matrix is sufficiently systematic and complete, up to 7 km/s, to make reasonable recommendations for velocity-scaling of Teflon craters and penetrations. We specifically suggest that cratering behavior and associated equations apply to all impacts in which the shock-pulse duration of the projectile is shorter than that assigned a unique projectile size, provided an impact velocity is known or assumed. This calibration seems superior to the traditional ballistic-limit approach.

Horz, Friedrich↗

NASA NTRS1983

Impact cratering and spall failure at gabbro

Both hypervelocity impact and dynamic spall experiments were carried out on a series of well-indurated samples of gabbro. The impact experiments carried out with 0.04 to 0.2g, 5-6 km/sec projectiles produced deci-centimeter-sized craters and demonstrated crater efficiencies of 6/10 to the - 9 g/erg, and order of magnitude greater than in metal and some two to three times that of previous experiments on less strong igneous rocks. Most of the crater volume (some 60 to 80%) is due to spall failure. Distribution of cumulative fragment number, as a function of mass of fragments with masses greater than 0.1 gram yield values of b = d(log10N sub f)dlog10(m)of -0.5 to -0.6, where N sub f is the cumulate number of fragments and m is the mass of fragments. These values are in agreement or slightly higher than those obtained for less strong rocks and indicate that a large fraction of the ejectra resides in a few large fragments.

Lange, M. A.↗

NASA NTRS1984

Impact cratering and spall failure of gabbro

Both hypervelocity impact and dynamic spall experiments were carried out on a series of well-indurated samples of gabbro. The impact experiments carried out with 0.04 to 0.2 g, 5-6 km/sec projectiles produced deci-centimeter-sized craters and demonstrated crater efficiencies of 6/10 to the -9 g/erg, and order of magnitude greater than in metal and some two to three times that of previous experiments on less strong igneous rocks. Most of the crater volume (some 60 to 80 percent) is due to spall failure. Distribution of cumulative fragment number, as a function of mass of fragments with masses greater than 0.1 gram yield values of b = d(log10N sub f)dlog10(m) of -0.5 to -0.6, where N sub f is the cumulate number of fragments and m is the mass of fragments. These values are in agreement or slightly higher than those obtained for less strong rocks and indicate that a large fraction of the ejecta resides in a few large fragments.

Lange, M. A.↗

NASA NTRS2003

Hypervelocity Impact Test Results for a Metallic Thermal Protection System

Hypervelocity impact tests have been performed on specimens representing metallic thermal protection systems (TPS) developed at NASA Langley Research Center for use on next-generation reusable launch vehicles (RLV). The majority of the specimens tested consists of a foil gauge exterior honeycomb panel, composed of either Inconel 617 or Ti-6Al-4V, backed with 2.0 in. of fibrous insulation and a final Ti-6Al-4V foil layer. Other tested specimens include titanium multi-wall sandwich coupons as well as TPS using a second honeycomb sandwich in place of the foil backing. Hypervelocity impact tests were performed at the NASA Marshall Space Flight Center Orbital Debris Simulation Facility. An improved test fixture was designed and fabricated to hold specimens firmly in place during impact. Projectile diameter, honeycomb sandwich material, honeycomb sandwich facesheet thickness, and honeycomb core cell size were examined to determine the influence of TPS configuration on the level of protection provided to the substructure (crew, cabin, fuel tank, etc.) against micrometeoroid or orbit debris impacts. Pictures and descriptions of the damage to each specimen are included.

Karr, Katherine L.↗

NASA NTRS2017

Hypervelocity Impact Testing of Materials for Additive Construction: Applications on Earth, the Moon, and Mars

Additive Construction is the process of building infrastructure such as habitats, garages, roads, berms, etcetera layer by layer (3D printing). The National Aeronautics and Space Administration (NASA) and the United States Army Corps of Engineers (USACE) are pursuing additive construction to build structures using resources available in-situ. Using materials available in-situ reduces the cost of planetary missions and operations in theater. The NASA team is investigating multiple binders that can be produced on planetary surfaces, including the magnesium oxide-based Sorel cement; the components required to make Ordinary Portland Cement (OPC), the common cement used on Earth, have been found on Mars. The availability of OPC-based concrete on Earth drove the USACE to pursue additive construction for base housing and barriers for military operations. Planetary and military base structures must be capable of resisting micrometeoroid impacts with velocities ranging from 11 to 72km/s for particle sizes 200 micrometers or more (depending on protection requirements) as well as bullets and shrapnel with a velocity of 1.036km/s with projectiles 5.66mm diameter and 57.40mm in length, respectively.

Hypervelocity Impact↗

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↗

NASA NTRS2009

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

deGroh, Henry C., III↗

NASA NTRS1994

The use of silica aerogel to collect interplanetary dust in space

We report on impact experiments to evaluate the suitability of silica aerogel as a capture medium for interplanetary dust particles in space. The structure of silica aerogel is such that even micrometer-sized hypervelocity particles may sense it as a target of very low density; consequently shock stresses and temperatures due to impact should be minimized. We performed impact experiments in a light gas gun that employed projectiles manufactured from forsterite, pyrrhotite, and calcite. From these experiments we have developed techniques that will permit the successful extrication of interplanetary dust particles from aerogel. We note a rough correlation between increasing particle penetration lengths and decreasing aerogel density, and there is a poor correlation of track lengths with impact velocity at laboratory-attainable velocities of 5-7 km/s. We conclude that aerogel track length should not be used even as (crude) velocity indicators. Transmission electron microscope studies show the re covered particulate residues are encased in melted silica aerogel, which possibly protects the residue from some impact damage. Individual minerals within interplanetary dust are expected to exhibit differential survival, and may be vaporized, melted, dehydrated, or structurally reordered. However, some mineral grains should successfully be captured in an unmelted state at encounter velocities less than 8 km/s.

Zolensky, Michael E.↗