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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

The ballistic range and aerothermodynamic testing

A renewed interest in hypervelocity vehicles requires an increase in our knowledge of aerothermodynamic phenomena. Tests conducted in our ground-based facilities can be used both to better understand the physics of hypervelocity flight, and to calibrate and validate computer codes designed to predict vehicle performance in the hypervelocity environment. This paper reviews the requirements for aerothermodynamic testing and discusses the ballistic range and its capabilities. Examples of the kinds of testing performed in typical high-performance ballistic ranges are described; we draw heavily on experience obtained in the ballistics facilities at NASA Ames. Prospects for improving the capabilities of the ballistic range by using advanced instrumentation are discussed. Finally, recent developments in gun technology and their application to extend the capability of the ballistic range are summarized.

Strawa, A. W.↗
NASA NTRS1980

Heatshield material selection for advanced ballistic reentry vehicles

The Performance of staple rayon fiber and AVTEX continuous rayon fiber was evaluated as precursor materials for heatshields. The materials studied were referenced to the IRC FM5055A heatshield materials flown during the past decade. Three different arc jet facilities were used to simulate portions of the reentry environment. The IRC FM5055A and the AVTEX FM5055G, both continuous rayon fiber woven materials having the phenolic impregnant filled with carbon particles were compared. The AVTEX continuous fiber, unfilled material FM5822A was also examined to a limited extent. Test results show that the AVTEX FM5055G material provided a close substitute for the IRC FM5055A material both in terms of thermal protection and roll torque performance.

Legendre, P. J.↗
NASA NTRS2010

Characteristics of Whipple Shield Performance in the Shatter Regime

Ballistic limit equations define the failure of metallic Whipple shields in three parts: low velocity, shatter, and hypervelocity. Failure limits in the shatter regime are based on a linear interpolation between the onset of projectile fragmentation, and impulsive rupture of the shield rear wall. A series of hypervelocity impact tests have been performed on aluminum alloy Whipple shields to investigate failure mechanisms and performance limits in the shatter regime. Test results demonstrated a more rapid increase in performance than predicted by the latest iteration of the JSC Whipple shield BLE following the onset of projectile fragmentation. This increase in performance was found to level out between 4.0-5.0 km/s, with a subsequent decrease in performance for velocities up to 6.0 km/s. For a detached spall failure criterion, the failure limit was found to continually decrease up to a velocity of 7.0 km/s, substantially varying from the BLE, while for perforation-based failure an increase in performance was observed. An existing phenomenological ballistic limit curve was found to provide a more accurate reproduction of shield behavior that the BLE, however a number of underlying assumptions such as the occurrence of complete projectile fragmentation and the effect on performance of incipient projectile melt were found to be inaccurate. A cratering relationship based on the largest residual fragment size has been derived for application at velocities between 3.0-4.0 km/s, and was shown to accurately reproduce the trends of the experimental data. Further investigation is required to allow a full analytical description of shatter regime performance for metallic Whipple shields.

Ryan, S.↗
NASA NTRS2009

Recent Advances in Thermoplastic Puncture-Healing Polymers

Self-healing materials provide a route for enhanced damage tolerance in materials for aerospace applications. In particular, puncture-healing upon impact has the potential to mitigate significant damage caused by high velocity micrometeoroid impacts. This type of material also has the potential to improve damage tolerance in load bearing structures to enhance vehicle health and aircraft durability. The materials being studied are those capable of instantaneous puncture healing, providing a mechanism for mechanical property retention in lightweight structures. These systems have demonstrated healing capability following penetration of fast moving projectiles -- velocities that range from 9 mm bullets shot from a gun (approx.330 m/sec) to close to micrometeoroid debris velocities of 4800 m/sec. In this presentation, we report on a suite of polymeric materials possessing this characteristic. Figure 1 illustrates the puncture healing concept. Puncture healing in these materials is dependent upon how the combination of a polymer's viscoelastic properties responds to the energy input resulting from the puncture event. Projectile penetration increases the temperature in the vicinity of the impact. Self-healing behavior occurs following puncture, whereby energy must be transferred to the material during impact both elastically and inelastically, thus establishing two requirements for puncture healing to occur: a.) The need for the puncture event to produce a local melt state in the polymer material and b.) The molten material has to have sufficient melt elasticity to snap back and close the hole. 1,2 Previous ballistic testing studies revealed that Surlyn materials warmed up to a temperature approx.98 C during projectile puncture (3 C higher than it s melting temperature). 1,2 The temperature increase produces a localized flow state and the melt elasticity to snap back thus sealing the hole. Table 1 lists the commercially polymers studied here, together with their physical properties. The polymers were selected based on chemical structure, tensile strengths, tensile moduli, glass transition temperature, melting temperatures, and impact strength. The thermal properties of the polymers were characterized by Differential Scanning Calorimetry (DSC) and Dynamic Mechanical Analysis (DMA). Mechanical properties were assessed by a Sintech 2W instron according to ASTM D1708 or D638 at crosshead speeds of 5.08 cm/min. 7.6 cm x 7.6 cm panels of the different materials were prepared and ballistic testing was performed at various temperatures. The panels were shot with a .223 caliber semiautomatic rifle from a distance of 23 meters at various temperatures. Chronographs were used to measure initial and final bullet velocity. Temperatures at the site of impact were measured using a FLIR ThermaCAM S60 thermal camera. A Vision Research model Phantom 9 high speed video camera was used to capture high speed video footage of ballistics testing.

Gordon, K. L.↗
NASA NTRS1992

Damage tolerance of a composite sandwich with interleaved foam core

A composite sandwich panel consisting of carbon fiber-reinforced plastic (CFRP) skins and a syntactic foam core was selected as an appropriate structural concept for the design of wind tunnel compressor blades. Interleaving of the core with tough interlayers was done to prevent core cracking and to improve damage tolerance of the sandwich. Simply supported sandwich beam specimens were subjected to low-velocity drop-weight impacts as well as high velocity ballistic impacts. The performance of the interleaved core sandwich panels was characterized by localized skin damage and minor cracking of the core. Residual compressive strength (RCS) of the skin, which was derived from flexural test, shows the expected trend of decreasing with increasing size of the damage, impact energy, and velocity. In the case of skin damage, RCS values of around 50 percent of the virgin interleaved reference were obtained at the upper impact energy range. Based on the similarity between low-velocity and ballistic-impact effects, it was concluded that impact energy is the main variable controlling damage and residual strength, where as velocity plays a minor role.

Ishai, Ori↗
NASA NTRS2009

Deployable Aeroshell Flexible Thermal Protection System Testing

Deployable aeroshells offer the promise of achieving larger aeroshell surface areas for entry vehicles than otherwise attainable without deployment. With the larger surface area comes the ability to decelerate high-mass entry vehicles at relatively low ballistic coefficients. However, for an aeroshell to perform even at the low ballistic coefficients attainable with deployable aeroshells, a flexible thermal protection system (TPS) is required that is capable of surviving reasonably high heat flux and durable enough to survive the rigors of construction handling, high density packing, deployment, aerodynamic loading and aerothermal heating. The Program for the Advancement of Inflatable Decelerators for Atmospheric Entry (PAIDAE) is tasked with developing the technologies required to increase the technology readiness level (TRL) of inflatable deployable aeroshells, and one of several of the technologies PAIDAE is developing for use on inflatable aeroshells is flexible TPS. Several flexible TPS layups were designed, based on commercially available materials, and tested in NASA Langley Research Center's 8 Foot High Temperature Tunnel (8ft HTT). The TPS layups were designed for, and tested at three different conditions that are representative of conditions seen in entry simulation analyses of inflatable aeroshell concepts. Two conditions were produced in a single run with a sting-mounted dual wedge test fixture. The dual wedge test fixture had one row of sample mounting locations (forward) at about half the running length of the top surface of the wedge. At about two thirds of the running length of the wedge, a second test surface drafted up at five degrees relative to the first test surface established the remaining running length of the wedge test fixture. A second row of sample mounting locations (aft) was positioned in the middle of the running length of the second test surface. Once the desired flow conditions were established in the test section the dual wedge test fixture, oriented at 5 degrees angle of attack down, was injected into the flow. In this configuration the aft sample mounting location was subjected to roughly twice the heat flux and surface pressure of the forward mounting location. The tunnel was run at two different conditions for the test series: 1) 'Low Pressure', and 2) 'High Pressure'. At 'Low Pressure' conditions the TPS layups were tested at 6W/cm2 and 11W/cm2 while at 'High Pressure' conditions the TPS layups were tested at 11W/cm2 and 20W/cm2. This paper details the test configuration of the TPS samples in the 8Ft HTT, the sample holder assembly, TPS sample layup construction, sample instrumentation, results from this testing, as well as lessons learned.

Hughes, Stephen J.↗
NASA NTRS1977

A measurement of the width and shift of the Fe I 3719.94 A line broadened by helium

Results are reported for measurements of the widths and shifts of the overlapping helium-broadened profiles of the Fe I absorption lines at 3719.94 and 3722.56 A, which were performed in a ballistic piston compressor using a saturated solution of FeCl3 in ethanol. It is found that at 4000 K, the 3719.94-A line has a gamma/n (FWHM) value of approximately 5.2 billionths rad/sec per cu cm, the blue shift of this line corresponds to a beta/n value of about 0.49 billionth rad/sec per cu cm, and the 3722.56-A line has the same values. These results are compared with previous experimental determinations and with theoretical calculations of neutral-line broadening for the Fe I lines. The significance of these measurements with respect to the solar spectrum and cool DG white dwarfs is discussed.

Driver, R. D.↗
NASA NTRS1981

Development of energy-absorbing reaction-sintered Si3N4 surface layers on hot-pressed Si3N4

Energy-absorbing Si3N4 surface layers on dense Si3N4 substrates were formed by in-place nitridation of fine-grained silicon powder. Ballistic impact tests performed on samples with 1-mm thick layers at room temperature and 1370 C showed up to an eightfold increase in the energy necessary to fracture the substrate. For maximum impact resistance, a small amount (about 20 vol %) of residual Si must be present in the reaction-sintered Si3N4 surface layer. Thermal cycling to 1370 C did not affect impact resistance, even though a considerable amount of SiO2 formed within the reaction-sintered Si3N4 layer during cycling. Erosion testing of samples in a Mach 0.8 burner rig at 1370 C resulted in minimal surface recession of the surface layer. Chemically vapor-deposited SiC-coated material similarly tested exhibited no surface recession.

Brennan, J. J.↗
NASA NTRS1989

A code calibration program in support of the Aeroassist Flight Experiment

The code calibration program for the Langley Aerothermodynamic Upwind Relaxation Algorithm to be used as support for the Aeroassist Flight Experiment (AFE) is discussed. Comparisons between experimental data and numerical simulations are made which focus on perfect-gas tests involving a scale model of the AFE. Aspects of the thermochemical nonequilibrium model are called into question by the results of ground tests performed in a ballistic range and in a shock tunnel.

Gnoffo, Peter A.↗
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 NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗
NASA NTRSDate not supplied

Orion Artemis I As Flown MMOD Analysis

Introduction The Lockheed Martin Orion spacecraft conducted the Artemis I flight around the Moon from November 16 through December 11, 2022. After the flight, an engineer from the NASA Johnson Space Center (JSC) Hypervelocity Impact Technology (HVIT) Group performed a Micrometeoroid and Orbital Debris MMOD analysis using the Bumper 3 risk assessment tool to predict the number of small impacts that would likely have occurred during the mission. Separately, a team from the same group inspected the Orion capsule for hypervelocity impact damage features. The results of the inspection were compared to those of the analysis to aid in improving the analysis, including the environment models. Scope of Work The spacecraft geometry model was created by Lockheed Martin during construction of the Artemis I Orion vehicle based on Computer Aided Design (CAD) models of the vehicle. New hypervelocity impact testing was performed to verify Ballistic Limit Equations (BLEs) used in the analysis to link impactor size and damage to the Thermal Protection System (TPS). The exact trajectory flown was recorded during the flight, including vehicle attitude. This data was used in conjunction with the ORDEM 3.2 and MEM 3 environment modeling tools to create models of particle flux impacting the spacecraft throughout the mission. Meteoroid shower forecast information was also included to account for additional particle flux associated with meteoroid showers. Inspection of the Orion capsule included the Backshell thermal tiles and the tape covering it, windows, fabric thermal materials, and small areas of other materials. Potential MMOD damage found was characterized using various techniques, including optical microscopy, computed tomography scanning, and X-ray spectroscopy. Findings The number of craters found in the Backshell tile, and their size distribution, matches well with the Bumper analysis prediction. Tape, window, and other material impacts recorded similarly align to Bumper analysis predictions. Conclusions and Recommendations This comparison of analysis with inspection of the hardware provides valuable insight into the MMOD environment and how accurately analysis tools assess the impact risk to spacecraft. As this was the first large, non-ablative returned surface from a lunar mission, this analysis extends the MMOD community’s insight beyond low Earth orbit into cis-lunar space.

MMOD↗
NASA NTRS2018

NASA ACC High Energy Dynamic Impact Methodology and Outcomes

For High Energy Dynamic Impact (HEDI) events, testing to evaluate the structural response of primary aircraft structure for design and certification is both expensive and time consuming. This paper discusses current work seeking to assess, develop, and validate appropriate analytical models that accurately predict physical response, damage, and failure modes for large scale composite structures in HEDI events. Four state-of-the-art Progressive Damage Analysis (PDA) methods were employed for this phased project: LS-DYNA MAT162, LS-DYNA MAT261, Smoothed Particle Galerkin (SPG), and EMU Peridynamics. Multiple material systems were considered, namely T700/5208 textile-infusion triaxial braid, T800/AMD-825 textile-infusion triaxial braid, IM7/8552 uni-directional tape, and SPG 196-PW/8552 plain-weave fabric. Extensive ballistic impact testing was performed to support this activity and measured results were compared to predictive models for assessment using panel delamination, panel displacement, force at the load cells, and threshold velocity as measures. Ultimately, the work under this activity provided significant progress in advancing the state-of-the-art in the use of PDA for HEDI events. Each material model had favorable performance comparing to test in some parameters and needed improvement in others. With the lessons learned from this activity, significant progress was made in the ability to predict panel behavior for a more general case beyond the flat panel in a ballistic impact event. Subsequent Phase II of the NASA ACC HEDI effort will continue to build on the coupon testing, flat panel ballistic impact testing, and analysis performed to-date with application of the PDA methods for intended material selections to test articles with greater complexity of configuration, curvature, and scale. It is not the intention of this paper to present a full set of data, but rather to give an overview of the NASA HEDI effort and show a small representative subset of the test and analysis results.

Hunziker, Kenneth J.↗
NASA NTRS2017

CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD) Ballistic Range Tests

A series of ballistic range tests were performed on a scaled model of the Supersonic Flight Demonstration Test (SFDT) intended to test the Supersonic Inflatable Aerodynamic Decelerator (SIAD) geometry. The purpose of these experiments were to provide aerodynamic coefficients of the vehicle to aid in mission and vehicle design. The experimental data spans the moderate Mach number range, $3.8-2.0$, with a total angle of attack ($alpha_T$) range, $10o-20o$. These conditions are intended to span the Mach-$alpha$ space for the majority of the SFDT experiment. In an effort to validate the predictive capabilities of Computational Fluid Dynamics (CFD) for free-flight aerodynamic behavior, numerical simulations of the ballistic range experiment are performed using the unstructured finite volume Navier-Stokes solver, US3D. Comparisons to raw vehicle attitude, and post-processed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data gives further insights into dynamic flow coupling leading to a potential mechanism for dynamic instability.

CFD modeling↗
NASA NTRS2016

Dynamic CFD Simulations of the MEADS II Ballistic Range Test Model

Dynamic, viscous, free-to-oscillate simulations of the Mars Entry Atmospheric Data System (MEADS) ballistic range model are performed using two different ow solvers, OVERFLOW and US3D. At the time of publication, data from the ballistic range experiment was not yet available, so the current work serves as a code-to-code exercise. Results from the analysis show good agreement between the predicted static aerodynamic coefficients for each solver. Both codes predict damped pitch oscillations for Mach 3:0 with initial amplitudes of 5⁰ and 30⁰, as well as for Mach 1:5 with initial amplitude of 30⁰. The two solvers predict undamped pitch oscillations for Mach 1:5 with initial amplitude of 5⁰. For most cases, US3D predicts less damping than OVERFLOW. The difference is attributed to higher pressures in the separated region of the wake, and the resultant effect on the backshell contribution to the pitching moment.

Stern, Eric C.↗
NASA NTRS1976

Impact ejecta on the moon

The response of a lunar-sized object to the impact of meteoroids no more than about 100 km in radius is studied by means of a numerical model. The partitioning of impact energy into the kinetic and internal energy of the ejecta is obtained by using the conservation of mass, momentum, and energy conservation equations in finite-difference form within an Eulerian framework with approximate equations of state. The calculations are performed for a 15 km/sec impact of an iron object 5 cm in radius on a gabbroic anorthosite surface. Ejecta ballistic analysis is then performed. Most of the material lost escaping the moon is lunar crust material. Only 0.2% of the meteoroid escapes, all in the vapor phase.

Okeefe, J. D.↗
NASA NTRS1972

Space Shuttle: Reentry stability and performance characteristics in the transonic and supersonic flight regimes of the Boeing ballistic recoverable booster

Experimental aerodynamic investigations were made in transonic and supersonic wind tunnels on a .008899 scale model of the Boeing model 979-145 Ballistic Recoverable Booster. The purpose of the tests was to define the stability and performance characteristics of the BRB at re-entry attitudes in the transonic and supersonic flight regimes. Data were obtained over a Mach number range from 0.6 to 4.0 at angles of attack between 50 deg and 85 deg at zero sideslip and at angles of sideslip between -17.5 deg and +15 deg at angles of attack between 50 deg and 85 deg.

Houser, J.↗
NASA NTRS2001

Impact Testing of Composites for Aircraft Engine Fan Cases

Before composite materials can be considered for use in the fan case of a commercial jet engine, the performance of a composite structure under blade-out loads needs to be demonstrated. The objective of this program is to develop an efficient test and analysis method for evaluating potential composite case concepts. Ballistic impact tests were performed on laminated glass/epoxy composites in order to identify potential failure modes and to provide data for analysis. Flat 7x7 in. panels were impacted with cylindrical titanium projectiles, and 15 in. diameter half-rings were impacted with wedge-shaped titanium projectiles. Composite failure involved local fiber fracture as well as tearing and delamination on a larger scale. A 36 in. diameter full-ring subcomponent was proposed for larger scale testing. Explicit, transient, finite element analyses were used to evaluate impact dynamics and subsequent global deformation for the proposed full-ring subcomponent test. Analyses on half-ring and quarter ring configurations indicated that less expensive smaller scale tests could be used to screen potential composite concepts when evaluation of local impact damage is the primary concern.

Roberts, Gary D.↗