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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Chemical nonequilibrium effects on flowfields for aeroassist orbital transfer vehicles

Acceptable altitude and velocity ranges for the perigee of aeroassist orbital transfer vehicles (AOTV) have been determined for various values of ballistic coefficient and lift-to-drag ratio. The chemical nonequilibrium effect on the flowfield of a simple but representatively high-drag, blunt configuration was investigated for these perigee conditions, ranging from 73 to 90 km in altitude and 8 to 9.3 km/s in velocity. The analysis was performed using a two-dimensional viscous-shock-layer code which had previously demonstrated good success in predicting nonequilibrium heating on the Space Shuttle. The results indicate lower ratio of nonequilibrium/equilibrium heating for the present AOTV perigee conditions than for the Shuttle flight conditions, due to higher velocity. The trend for the degree of nonequilibrium with the effect of varying wall catalycity and configuration was determined over the altitude range. Some implications to future design consideration are stated.

Shinn, J. L.↗
NASA NTRS1992

Physiologically constrained aerocapture for manned Mars missions

Aerobraking has been proposed as a critical technology for manned missions to Mars. The variety of mission architectures currently under consideration presents aerobrake designers with an enormous range of potential entry scenarios. Two of the most important considerations in the design of an aerobrake are the required control authority (lift-to-drag ratio) and the aerothermal environment which the vehicle will encounter. Therefore, this study examined the entry corridor width and stagnation-point heating rate and load for the entire range of probable entry velocities, lift-to-drag ratios, and ballistic coefficients for capture at both Earth and Mars. To accomplish this, a peak deceleration limit for the aerocapture maneuvers had to be established. Previous studies had used a variety of load limits without adequate proof of their validity. Existing physiological and space flight data were examined, and it was concluded that a deceleration limit of 5 G was appropriate. When this load limit was applied, numerical studies showed that an aerobrake with an L/D of 0.3 could provide an entry corridor width of at least 1 degree for all Mars aerocaptures considered with entry velocities up to 9 km/s. If 10 km/s entries are required, an L/D of 0.4 to 0.5 would be necessary to maintain a corridor width of at least 1 degree. For Earth return aerocapture, a vehicle with an L/D of 0.4 to 0.5 was found to provide a corridor width of 0.7 degree or more for all entry velocities up to 14.5 km/s. Aerodynamic convective heating calculations were performed assuming a fully catalytic, 'cold' wall; radiative heating was calculated assuming that the shock layer was in thermochemical equilibrium. Heating rates were low enough for selected entries at Mars that a radiatively cooled thermal protection system might be feasible, although an ablative material would be required for most scenarios. Earth return heating rates were generally more severe than those encountered by the Apollo vehicles, and would require ablative heat shields in all cases.

Lyne, James Evans↗
NASA NTRS1995

Environmentally compatible solid rocket propellants

Hercules' clean propellant development research is exploring three major types of clean propellant: (1) chloride-free formulations (no chlorine containing ingredients), being developed on the Clean Propellant Development and Demonstration (CPDD) contract sponsored by Phillips Laboratory, Edwards Air Force Base, CA; (2) low HCl scavenged formulations (HCl-scavenger added to propellant oxidized with ammonium perchlorate (AP)); and (3) low HCl formulations oxidized with a combination of AN and AP (with or without an HCl scavenger) to provide a significant reduction (relative to current solid rocket boosters) in exhaust HCl. These propellants provide performance approaching that of current systems, with less than 2 percent HCl in the exhaust, a significant reduction (greater than or equal to 70 percent) in exhaust HCl levels. Excellent processing, safety, and mechanical properties were achieved using only readily available, low cost ingredients. Two formulations, a sodium nitrate (NaNO3) scavenged HTPB and a chloride-free hydroxy terminated polyether (HTPE) propellant, were characterized for ballistic, mechanical, and rheological properties. In addition, the hazards properties were demonstrated to provide two families of class 1.3, 'zero-card' propellants. Further characterization is planned which includes demonstration of ballistic tailorability in subscale (one to 70 pound) motors over the range of burn rates required for retrofit into current Hercules space booster designs (Titan 4 SRMU and Delta 2 GEM).

Jacox, James L.↗
NASA NTRS2016

Engineering Polymer Blends for Impact Damage Mitigation

Structures containing polymers such as DuPont's Surlyn® 8940, demonstrate puncture healing when impacted by a 9 millimeter projectile traveling from speeds near 300 meters per second (1,100 feet per second) to hypervelocity impacts in the micrometeoroid velocity range of 5 kilometers per second (16,000 feet per second). Surlyn® 8940 puncture heals over a temperature range of minus 30 degrees Centigrade to plus 70 degrees Centigrade and shows potential for use in pressurized vessels subject to impact damage. However, such polymers are difficult to process and limited in applicability due to their low thermal stability, poor chemical resistance and overall poor mechanical properties. In this work, several puncture healing engineered melt formulations were developed. Moldings of melt blend formulations were impacted with a 5.56 millimeter projectile with a nominal velocity of 945 meters per second (3,100 feet per second) at about 25 degrees Centigrade, 50 degrees Centigrade and 100 degrees Centigrade, depending upon the specific blend being investigated. Self-healing tendencies were determined using surface vacuum pressure tests and tensile tests after penetration using tensile dog-bone specimens (ASTM D 638-10). For the characterization of tensile properties both pristine and impacted specimens were tested to obtain tensile modulus, yield stress and tensile strength, where possible. Experimental results demonstrate a range of new puncture healing blends which mitigate damage in the ballistic velocity regime.

Gordon, Keith L.↗
NASA NTRS2009

Performance of Whipple Shields at Impact Velocities above 9 km/s

Whipple shields were first proposed as a means of protecting spacecraft from the impact of micrometeoroids in 1947 [1] and are currently in use as micrometeoroid and orbital debris shields on modern spacecraft. In the intervening years, the function of the thin bumper used to shatter or melt threatening particles has been augmented and enhanced by the use of various types and configurations of intermediate layers of various materials. All shield designs serve to minimize the threat of a spall failure or perforation of the main wall of the spacecraft as a result of the impact of the fragments. With increasing use of Whipple shields, various ballistic limit equations (BLEs) for guiding the design and estimating the performance of shield systems have been developed. Perhaps the best known and most used are the "new" modified Cour-Palais (Christiansen) equations [2]. These equations address the three phases of impact: (1) ballistic (<3 km/s), where the projectile is moving too slowly to fragment and essentially penetrates as an intact projectile; (2) shatter (3 to 7 km/s), where the projectile fragments at impact and forms an expanding cloud of debris fragments; and (3) melt/vaporization (>7 km/s), where the projectile melts or vaporizes at impact. The performance of Whipple shields and the adequacy of the BLEs have been examined for the first two phases using the results of impact tests obtained from two-stage, light-gas gun test firings. Shield performance and the adequacy of the BLEs has not been evaluated in the melt/vaporization phase until now because of the limitations of launchers used to accelerate projectiles with controlled properties to velocities above 7.5 km/s. A three-stage, light-gas gun, developed at the University of Dayton Research Institute (UDRI) [3], is capable of launching small, aluminum spheres to velocities above 9 km/s. This launcher was used to evaluate the ballistic performance of two Whipple shield systems, various thermal protection system materials, and other spacecraft-related materials to the impact of 1.6-mm- to 2.6-mm-diameter, 2017-T4 aluminum spheres at impact velocities ranging from 8.91 km/s to 9.28 km/s. Test results, details of the shield systems, and nominal ballistic limits for the two Whipple shields are shown in Figures 1 and 2.

Christiansen, Eric L.↗
NASA NTRSDate not supplied

Impact of Magnetohydrodynamic (MHD) Induced Drag on High Ballistic Coefficient Aerocapture at Neptune

This paper presents analyses illustrating the benefits of magnetohydrodynamic (MHD) drag force control for a capsule performing aerocapture at Neptune relative to the same capsule utilizing conventional aerodynamic-only control across a range of design and flight conditions. MHD technology utilizes an onboard mag-net to create a magnetic field that interacts with the ionized plasma surrounding the capsule at hypersonic velocities, thus creating an electron current and inducing a Lorentz force used to increase capsule drag. The study assumes a 4.57-meterMars Science Laboratory (MSL) 70-degree sphere-cone capsule aerodynamic shape arrives hyperbolically at Neptune and aerocaptures into a synchronous Tri-ton orbit (3986 km x 430,000 km orbit inclined 157 degrees) that offers repeated flyby opportunities. Hyperbolic arrival velocity and the capsule entry mass (i.e., ballistic coefficient) are varied over a wide range of values to study the impact of these variables on key flight performance and sizing parameters. Results show reduced heat rates, heat loads, and thermal protection system mass for MHD cases at all conditions, with the benefit of MHD relative to conventional aerodynamic-only options increasing as ballistic coefficient increases, suggesting the payoff for MHD is significantly better for high-mass entry system design concepts with high ballistic coefficients.

Entry Descent and Landing↗
NASA NTRSDate not supplied

The Adaptable, Deployable, Entry and Placement Technology (ADEPT) Enabling Advanced Entry, Descent, and Landing Capabilities for SmallSat Missions

The Adaptable, Deployable Entry and Placement Technology (ADEPT) is a mechanically deployable low-ballistic coefficient aeroshell entry system which performs entry, descent, and landing (EDL) functions for a broad range of planetary destinations including Mars, Venus, Titan and Earth. The deployable system allows mission planners to develop an aeroshell design that is stowed like a folded umbrella, and yet prior to the EDL mission segment, transforms into a deployed, near-rigid low ballistic coefficient configuration. The ADEPT architecture is scalable from a sub-1m class to 6m diameter entry vehicles for Venus and and it has been studied at up to 15m – 20m scale for human exploration class missions. There has been growing interest in performing interplanetary missions with small spacecraft as secondary payloads and ADEPT offers small satellite mission designers the ability to consider Aeroassist technologies to expand mission applications. The ADEPT architecture introduces a new approach for entry vehicle design. It deploys a flexible 3D woven carbon fabric as the primary drag surface. It is capable of surviving high aerothermal heating during entry, while maintaining a robust impermeable membrane that provides high drag for vehicle deceleration. The ADEPT project team has advanced this decelerator technology via systems-level testing at the one meter diameter (nano-ADEPT) scale. A subsonic aeroloads test (May 2015) and a pathfinder arc-jet aeroheating test (Sept 2015) and a successful sounding rocket flight (SR-1) launch of a 0.7 meter deployed diameter ADEPT (September 2018). The SR-1 flight experiment demonstrated most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in zero-g, exo-atmospheric conditions, and passive ballistic re-entry of a 70 degree half-angle cone geometry. The ADEPT SR-1 sounding rocket flight experiment had to address challenges such as stowing within the constrained payload envelope, fabric folding, and packaging concerns.

Paul Wercinski↗
NASA NTRS1976

The broadening of the calcium resonance line in a high-temperature helium atmosphere

The paper describes experimental measurements of the shape of the Ca I resonance line at 4227 A in a high-temperature helium atmosphere. A ballistic piston compressor was used to produce hot helium in the temperature range from 3000 to 7000 K and the number-density range from 2 to 4 by 10 to the 20th power per cu cm, which conditions approximate those postulated for the atmospheres of certain cool white-dwarf stars. The Boltzmann temperature of the calcium atoms was measured by the brightness-emissivity method, the absorption line shape was measured with a twelve-channel polychromator, and the helium kinetic temperature was calculated from the equation of an ideal gas. The observed deviation from thermodynamic equilibrium is illustrated. The value of the ratio of the damping constant to the helium number density at 5000 K is found to be 45 (+ or - 4) by 10 to the -22nd power A cu cm, or 4.7 (+ or - 0.4) by 10 to the -9th power rad/s cu cm. It is noted that no existing line-broadening calculation for the calcium-helium system is consistent with the present results. Recent studies of the 4227-A line in spectra of the sun and the white dwarf van Maanen 2 are discussed.

Driver, R. D.↗
NASA NTRS1993

A hypersonic vehicle approach to planetary exploration

An enhanced Mars network class mission using a lifting hypersonic entry vehicle is proposed. The basic vehicle, derived from a mature hypersonic flight system called SWERVE, offers several advantages over more conventional low L/D or ballistic entry systems. The proposed vehicle has greatly improved lateral and cross range capability (e.g., it is capable of reaching the polar regions during less than optimal mission opportunities), is not limited to surface target areas of low elevation, and is less susceptible to problems caused by Martian dust storms. Further, the integrated vehicle has attractive deployment features and allows for a much improved evolutionary path to larger vehicles with greater science capability. Analysis of the vehicle is aided by the development of a Mars Hypersonic Flight Simulator from which flight trajectories are obtained. Atmospheric entry performance of the baseline vehicle is improved by a deceleration skirt and transpiration cooling system which significantly reduce TPS (Thermal Protection System) and flight battery mass. The use of the vehicle is also attractive in that the maturity of the flight systems make it cost-competitive with the development of a conventional low L/D entry system. Finally, the potential application of similar vehicles to other planetary missions is discussed.

Murbach, Marcus S.↗
NASA NTRS2011

A Survey of Ballistic Transfers to the Lunar Surface

In this study techniques are developed which allow an analysis of a range of different types of transfer trajectories from the Earth to the lunar surface. Trajectories ranging from those obtained using the invariant manifolds of unstable orbits to those derived from collision orbits are analyzed. These techniques allow the computation of trajectories encompassing low-energy trajectories as well as more direct transfers. The range of possible trajectory options is summarized, and a broad range of trajectories that exist as a result of the Sun's influence are computed and analyzed. The results are then classified by type, and trades between different measures of cost are discussed.

Earth-Moon trajectories↗
DOE OSTI2022

Dominant Energy Carrier Transitions and Thermal Anisotropy in Epitaxial Iridium Thin Films

High aspect ratio metal nanostructures are commonly found in a broad range of applications such as electronic compute structures and sensing. The self-heating and elevated temperatures in these structures, however, pose a significant bottleneck to both the reliability and clock frequencies of modern electronic devices. Any notable progress in energy efficiency and speed requires fundamental and tunable thermal transport mechanisms in nanostructured metals. Here, in this work, time-domain thermoreflectance is used to expose cross-plane quasi-ballistic transport in epitaxially grown metallic Ir(001) interposed between Al and MgO(001). Thermal conductivities ranges from roughly 65 (96 in-plane) to 119 (122 in-plane) W m -1 K -1 for 25.5–133.0 nm films, respectively. Further, low defects afforded by epitaxial growth are suspected to allow the observation of electron–phonon coupling effects in sub-20 nm metals with traditionally electron-mediated thermal transport. Via combined electro-thermal measurements and phenomenological modeling, the transition is revealed between three modes of cross-plane heat conduction across different thicknesses and an interplay among them: electron dominant, phonon dominant, and electron–phonon energy conversion dominant. The results substantiate unexplored modes of heat transport in nanostructured metals, the insights of which can be used to develop electro-thermal solutions for a host of modern microelectronic devices and sensing structures.

36 MATERIALS SCIENCE↗
NASA NTRS1987

Momentum transfer from oblique impacts

A completely satisfactory experiment would be in a low gravity environment where the effect of momentum imparted by ejecta impacting the surface can be removed or controlled from momentum transfer during impact. Preliminary estimates can be made using a ballistic pendulum. Such experiments were initiated at the NASA-Ames Vertical Gun Range in order to examine momentum transfer due to impact vaporization for oblique impacts. The preliminary results indicate that momentum from oblique impacts is very inefficient: decreasing with increasing impact velocity and perhaps size; increasing with decreasing density; and increasing with increasing impact angle. At face value, such results minimize the effect of momentum transfer by grazing impact; the more probable impact angles of 30 deg would have a greater effect, contrary to the commonly held impression.

Schultz, Peter H.↗
NASA NTRS1997

NASA/JSC orbital debris models

The orbital debris models developed within the framework of the NASA Johnson Space Center's (JSC's) orbital debris program, are categorized as environment definition and risk assessment models. The EVOLVE, CHAIN and the orbital debris engineering model 1996 (ORDEM96) computer programs determine the past, present and future orbital particulate environment, while the BUMPER and debris assessment software (DAS) computer programs provide a means for evaluating the risks of specific space missions. These models are presented. To support these models and to conduct specialized analyses, NASA/JSC employs a range of auxiliary models, including explosion and collision satellite breakup models, hypervelocity impact ballistic limit models, orbit propagation and decay models, space traffic models and solid rocket motor effluent models.

Johnson, Nicholas↗
DOE OSTI2023

Insights into radiation resistance of titanium alloys from displacement cascade simulations

Radiation damage in beam window materials limits the use of high power proton beams in high energy physics research. The alloy Ti-6Al-4V is presently used as a beam window material but a prospective alternative, Ti-15V-3Cr-3Sn-3Al has been proposed. Since both these alloys contain dual phases at room temperature, we compare the radiation damage in the α and ß-phases of these two materials via primary knock-on atom (PKA) cascade simulations in the 10-40keV energy range. At PKA energies 30 and 40keV, the number of Frenkel pairs in the ballistic stage is higher in the ß-phase of Ti-6Al-4V than that in the ß-phase of Ti-15V-3Cr-3Sn-3Al almost by a factor of 2. The α-phase, of both these alloys, by far outperforms the ß-phases of the two alloys, both in terms of damage during the ballistic stage and in terms of the surviving defects. The average displacement threshold energy (E d ) in the α-phase of both alloys was found to be 66eV while that in the ß-phase of Ti-15-3 was 55 eV and in the ß-phase of Ti-6-4 was 46. Further, while the number of surviving defects is almost equal in both alloys, the vacancy and interstitial clustering mechanisms differ notably, which can impact the degree of radiation hardening and loss of ductility. Our simulations show larger vacancy and interstitial clusters form in Ti-6Al-4V as compared to that in Ti-15-3-3-3 alloy. These results indicate that Ti-15-3-3-3 alloys may be a promising candidate for next generation beam window material with a higher radiation tolerance than the existing Ti-6-4 alloy.

36 MATERIALS SCIENCE↗
NASA NTRS1988

Dynamics of exploding magma chambers: Implications for K-T volcanism and mass extinctions

Although it is well known that unconfined chemical explosives may yield pressures to several megabars on detonation in air, the explosive literature has yet to be accessed by some contributors to the volcanological literature who've indicated that pressures in excess of the overburden and/or tensile cannot be obtained. Idealized ballistic assessments of pressures internal to volcanoes yield pressures in the hundreds of kilobar range upon correction by addition of friction, etc. Previous assessments of exploding magma chamber pressure have been made from the characteristics of the Mt. St. Helens explosion. A variety of methods yield pressures of similar value: at least hundreds of kilobars. Such results are consistent with free energy requirements for quench supersaturation explosion, a process occurring in solidifying industrial melts. Several reviews of geochemical literature emphasize the carbon event at the Cretaceous-Tertiary (K-T) boundary as being an indicator of a massive dump of CO2 derived from the mantle and entering the atmosphere by extensive global volcanism. Oxygen isotope data indicates extreme warming at the end of the Cretaceous which is consistent with a greenhouse effect attending the CO2 event. Reaction rate equations for the quench supersaturation explosion mechanism indicated, are consistent with the rise in pressure to 30 kbar on solidification of magmatic melts, these pressures limited by the strength of the experimental apparatus.

Rice, A. R.↗
NASA NTRS1959

A Limited Study of a Hypothetical Winged Anti-ICBM Point-Defense Missile

A preliminary investigation was conducted to determine whether a warhead stage of an antimissile missile could be placed within an arbitrary 2-nautical-mile-radius maneuver cylinder around an intercontinental-ballistic-missile (ICBM) flight path above an altitude of 140,000 feet, a horizontal range of 40 nautical miles, at a flight-path angle of approximately 20 deg, and within 50 seconds after take-off using only aerodynamic forces to turn the antimissile missile. The preliminary investigation indicated that an antimissile missile using aerodynamic forces for turning was capable of intercepting the ICBM for the stated conditions of this study although the turning must be completed below an altitude of approximately 70,000 feet to insure that the antimissile missile will be at the desired flight-path angle. Trim lift coefficients on the order of 2 to 3 and a maximum normal-acceleration force of from 25g to 35g were necessary to place the warhead stage in intercept position. The preliminary investigation indicated that for the two boosters investigated the booster having a burning time of 10 seconds gave greater range up the ICBM flight path than did the booster having a burning time of 15 seconds for the same trim lift coefficient and required the least trim lift coefficient for the same range.

Brown, Clarence A., Jr.↗