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At least 307 records · Page 17

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NASA NTRS2009

Hypervelocity Impact Performance of Open Cell Foam Core Sandwich Panel Structures

Metallic foams are a relatively new class of materials with low density and novel physical, mechanical, thermal, electrical and acoustic properties. Although incompletely characterized, they offer comparable mechanical performance to traditional spacecraft structural materials (i.e. honeycomb sandwich panels) without detrimental through-thickness channeling cells. There are two competing types of metallic foams: open cell and closed cell. Open cell foams are considered the more promising technology due to their lower weight and higher degree of homogeneity. Leading micrometeoroid and orbital debris shields (MMOD) incorporate thin plates separated by a void space (i.e. Whipple shield). Inclusion of intermediate fabric layers, or multiple bumper plates have led to significant performance enhancements, yet these shields require additional non-ballistic mass for installation (fasteners, supports, etc.) that can consume up to 35% of the total shield weight [1]. Structural panels, such as open cell foam core sandwich panels, that are also capable of providing sufficient MMOD protection, represent a significant potential for increased efficiency in hypervelocity impact shielding from a systems perspective through a reduction in required non-ballistic mass. In this paper, the results of an extensive impact test program on aluminum foam core sandwich panels are reported. The effect of pore density, and core thickness on shielding performance have been evaluated over impact velocities ranging from 2.2 - 9.3 km/s at various angles. A number of additional tests on alternate sandwich panel configurations of comparable-weight have also been performed, including aluminum honeycomb sandwich panels (see Figure 1), Nomex honeycomb core sandwich panels, and 3D aluminum honeycomb sandwich panels. A total of 70 hypervelocity impact tests are reported, from which an empirical ballistic limit equation (BLE) has been derived. The BLE is in the standard form suitable for implementation in risk analysis software, and includes the effect of panel thickness, core density, and facesheet material properties. A comparison between the shielding performance of foam core sandwich panel structures and common MMOD shielding configurations is made for both conservative (additional 35% non-ballistic mass) and optimistic (additional mass equal to 30% of bumper mass) considerations. Suggestions to improve the shielding performance of foam core sandwich panels are made, including the use of outer mesh layers, intermediate fabric/composite layers, and varying pore density.

Ryan, Shannon↗

NASA NTRS2000

Lead-Free Propellant for Propellant Actuated Devices

Naval Surface Warfare Center, Indian Head Division's CAD/PAD Department has been working to remove toxic compounds from our products for about a decade. In 1992, we embarked on an effort to develop a lead-free double base propellant to replace that of a foreign sole source. At the time there were availability concerns. In 1995, the department developed a strategic proposal to include a wider range of products. Efforts included such efforts as removing lead sheathing from linear explosives and replacing lead azide and lead styphnate compounds. This paper will discuss efforts specifically related to developing non-leaded double base propellant for use in various Propellant Actuated Devices (PADs) for aircrew escape systems. The propellants can replace their leaded counterparts, mitigating lead handling, processing, or toxic exposure to the environment and personnel. This work eliminates the use of leaded compounds, replacing them with a more environmentally benign metal-organic salt. Historically double-base propellants have held an advantage over other families of energetic materials through their relative insensitivity of the burning rate to changes in temperature and pressure. This desirable ballistic effect has been obtained with the use of a lead-organic salt alone or in a physical mixture with a copper-organic salt, or more recently with a lead-copper complex. These ballistic modifiers are typically added to the double-base 'paste' prior to gelatinization on heated calendars or one type or another. The effect of constant burning rate over a pressure range is called a 'plateau' while an even more beneficial effect of decreasing burning rate with increasing pressure is termed a 'mesa.' The latter effect results in very low temperature sensitivity of the propellant burning rate. Propellants with such effects are ideal tactical rocket motor propellants. The use of lead compounds poses a concern for the environment and personnel safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacturing, rocket firings, and disposal. Two separate programs are underway at Indian Head to replace the leaded ballistic modifiers in NOSIH-AA-2 and KU propellants. Although the system requirements were very different, the programs gained valuable information from one another throughout their course.

Goodwin, John L.↗

NASA NTRS1984

The aerodynamics of some guided projectiles

Some characteristic projectile shapes are considered with various added components intended to provide lift, stability, and control. The intent of the additions is to provide some means for altering the normal ballistic flight path of a projectile for various purposes such as: achieving greater accuracy at the impact point, selecting alternate impact points, extending range, improved evasion, and so on. The configurations presented illustrate the effects of a flare, wings, and tails for providing stability and lift, and the effects of aft-tails, a close-coupled flap, and all-moving forward wings for control. The relative merits of the various configurations, all of which provided for flight path alterations, are discussed.

Spearman, M. L.↗

NASA NTRS1984

The aerodynamics of some guided projectiles

Spearman, M. L.↗

NASA NTRS1992

Robotic planetary mission benefits from nuclear electric propulsion

Several interesting planetary missions are either enabled or significantly enhanced by nuclear electric propulsion (NEP) in the 50 to 100 kW power range. These missions include a Pluto Orbiter/Probe with an 11-year flight time and several years of operational life in orbit versus a ballistic very fast (13 km/s) flyby which would take longer to get to Pluto and would have a very short time to observe the planet. (A ballistic orbiter would take about 40 years to get to Pluto). Other missions include a Neptune Orbiter/Probe, a Jupiter Grand Tour orbiting each of the major moons in order, an Uranus Orbiter/Probe, a Multiple Mainbelt Asteroid Rendezvous orbiting six selected asteroids, and a Comet Nucleus Sample Return. This paper discusses potential missions and compares the nuclear electric propulsion option to the conventional ballistic approach on a parametric basis.

Kelley, James H.↗

NASA NTRS1991

A parametric study of manned aerocapture at Mars

The results of a parametric study of corridor width and stagnation point heating are presented for a range of a probable mission designs and vehicle configurations. Entry velocities were varied from 6 to 10 km/s, L/Ds from 0.1 to 1.0, and ballistic coefficients from 100 to 500 kg/sq m. It is found that vehicles with an L/D of 0.4 to 0.5 provide an entry corridor width of at least 1 deg for velocities up to 10 km/s. For entry velocities below approximately 7 km/s, radiative cooling may be possible for the thermal protection system. It is recommended that, at higher entry speeds, ablative heat shields be used. Maximum integrated stagnation point heat loads were equivalent to or less than those experienced by the Space Shuttle on a typical reentry.

Tauber, M. E.↗

NASA NTRS2004

Atmospheric Models for Aerocapture Systems Studies

Aerocapture uses atmospheric drag to decelerate into captured orbit from interplanetary transfer orbit. This includes capture into Earth orbit from, for example, Lunar-return or Mars-return orbit. Eight Solar System destinations have sufficient atmosphere for aerocapture to be applicable - three of the rocky planets (Venus, Earth, and Mars), four gas giants (Jupiter, Saturn, Uranus, and Neptune), and Saturn's moon Titan. These destinations fall into two groups: (1) The rocky planets, which have warm surface temperatures (approx. 200 to 750 K) and rapid decrease of density with altitude, and (2) the gas giants and Titan, which have cold temperatures (approx. 70 to 170 K) at the surface or 1-bar pressure level, and slow rate of decrease of density with altitude. The height variation of average density with altitude above 1-bar pressure level for the gas giant planets is shown. The periapsis density required for aerocapture of spacecraft having typical values of ballistic coefficient (a measure of mass per unit cross-sectional area) is also shown. The aerocapture altitudes at the gas giants would typically range from approx. 150 to 300 km. Density profiles are compared for the rocky planets with those for Titan and Neptune. Aerocapture at the rocky planets would occur at heights of approx. 50 to 100 km. For comparison, typical density and altitudes for aerobraking operations (circularizing a highly elliptical capture orbit, using multiple atmospheric passes) are also indicated.

Justus, C. G.↗

NASA NTRS2016

On the Theoretical Framework of Magnetized Outflows from Stellar-Mass Black Holes and Related Observations

The spins of stellar-mass black holes (BHs) and the power outputs of their jets are measurable quantities. Unfortunately, the currently employed methods do not agree and the results are controversial. Two major issues concern the measurements of BH spin and beam (jet) power. The former issue can be resolved by future observations. But the latter issue can be resolved now, if we pay attention to what is expected from theoretical considerations. The question of whether a correlation has been found between the power outputs of few objects and the spins of their BHs is moot because BH beam power does not scale with the square of the spin of the BH. We show that the theoretical BH beam power is a strongly nonlinear function of spin that cannot be approximated by a quadratic relation, as is generally stated when the influence of the magnetic field is not accounted for in the Blandford & Znajek model. The BH beam power of ballistic jets should scale a lot more steeply with BH spin irrespective of the magnetic field assumed to thread the horizon and the spin range considered. This behavior may already be visible in the analyses of radio observations by Narayan & McClintock and Russell et al. In agreement with previous studies, we also find that the power output that originates in the inner regions of the surrounding accretion disks is higher than that from the BHs and it cannot be ignored in investigations of continuous compact jets from these systems.

accretion↗

NASA NTRS1977

Mercury orbiter transport study

A data base and comparative performance analyses of alternative flight mode options for delivering a range of payload masses to Mercury orbit are provided. Launch opportunities over the period 1980-2000 are considered. Extensive data trades are developed for the ballistic flight mode option utilizing one or more swingbys of Venus. Advanced transport options studied include solar electric propulsion and solar sailing. Results show the significant performance tradeoffs among such key parameters as trip time, payload mass, propulsion system mass, orbit size, launch year sensitivity and relative cost-effectiveness. Handbook-type presentation formats, particularly in the case of ballistic mode data, provide planetary program planners with an easily used source of reference information essential in the preliminary steps of mission selection and planning.

Friedlander, A. L.↗

NASA NTRS2000

Evaluation of Thermo-Mechanical Stability of COTS Dual-Axis MEMS Accelerometers for Space Applications

Microelectromechanical systems in MEMS is one of the fastest growing technologies in microelectronics, and is of great interest for military and aerospace applications. Accelerometers are the earliest and most developed representatives of MEMS. First demonstrated in 1979, micromachined accelerometers were used in automobile industry for air bag crash- sensing applications since 1990. In 1999, N4EMS accelerometers were used in NASA-JPL Mars Microprobe. The most developed accelerometers for airbag crash- sensing are rated for a full range of +/- 50 G. The range of sensitivity for accelerometers required for military or aerospace applications is much larger, varying from 20,000 G (to measure acceleration during gun and ballistic munition launches), and to 10(exp -6) G, when used as guidance sensors (to measure attitude and position of a spacecraft). The presence of moving parts on the surface of chip is specific to MEMS, and particularly, to accelerometers. This characteristic brings new reliability issues to micromachined accelerometers, including cyclic fatigue cracking of polysilicon cantilevers and springs, mechanical stresses that are caused by packaging and contamination in the internal cavity of the package. Studies of fatigue cracks initiation and growth in polysilicon showed that the fatigue damage may influence MEMS device performance, and the presence of water vapor significantly enhances crack initiation and growth. Environmentally induced failures, particularly, failures due to thermal cycling and mechanical shock are considered as one of major reliability concerns in MEMS. These environmental conditions are also critical for space applications of the parts. For example, the Mars pathfinder mission had experienced 80 mechanical shock events during the pyrotechnic separation processes.

Sharma, Ashok K.↗

NASA NTRS2000

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

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

Pereira, J. Michael↗

NASA NTRS1984

Separation processes during binary monotectic alloy production

Observation of microgravity solidification processes indicates that outside of sedimentation, at least two other important effects can separate the phases: critical-point wetting and spreading; and thermal migration of second-phase droplets due to interfacial tension gradients. It is difficult to study these surface tension effects while in a unit gravity field. In order to investigate the processes occurring over a temperature range, i.e., between a consolute point and the monotectic temperature, it is necessary to use a low-gravity environment. The MSFC drop tube (and tower), the ballistic trajectory KC-135 airplane, and the Space Shuttle are ideal facilities to aid formation and testing of hypotheses. Much of the early work in this area focuses on transparent materials so that process dynamics may be studied by optical techniques such as photography for viewing macro-processes; holography for studying diffusional growth; spinodal decomposition and coalescence; ellipsometry for surface wetting and spreading effects; and interferometry and spectroscopy for small-scale spatial resolution of concentration profiles.

Frazier, D. O.↗

NASA NTRS1991

Earth return aerocapture for manned Mars missions

Lift to drag ratio (L/D) requirements and stagnation point heating have been examined for a variety of probable entry conditions and vehicle configurations. It is found that vehicles with an L/D of 0.5 or more provide a corridor width of at least 0.7 degrees while keeping the peak deceleration load below 5 g for approach velocities up to 14.5 km/s. It is shown that stagnation point peak heating rates and integrated heat load critically depend on both entry velocity and ballistic coefficient. For the most severe cases under consideration, peak heating and integrated heat load are five times greater than those encountered by Apollo but within the range of experience for unmanned vehicles.

Tauber, M. E.↗

NASA NTRSDate not supplied

Parametric Studies of Human Mars Entry, Descent, and Landing Systems

This paper will present a parametric analysis for entry, descent, and landing (EDL) concepts, enabling rapid systems assessment and tradespace exploration. The entry system uses a hypersonic inflatable aerodynamic decelerator (HIAD) technology. The baseline system includes elements for Mars aerocapture (AC) and EDL segments of the mission. The mission concept of operations (ConOps) begins at Mars arrival in a polar inclination. After performing an aerocapture (AC) maneuver into a pre-defined Mars parking orbit, the AC HIAD is jettisoned. The rest of thevehicle stays in the parking orbit for up to one year. The EDL sequence starts with a deorbit burn at the apoapsis of the parking orbit. After hypersonic entry with the EDL HIAD, the entry system uses a supersonic retropropulsion maneuver to slow the vehicle for the descent and landing segments of the mission. The vehicle will maintain a constant velocity of 2.5 m/s for 5 seconds prior to landing.The system includes a Mars Ascent Vehicle (MAV), Mars Descent Module (MDM), and two HIADs. The MDM includes a primary structure, tanks, engines, and radiators. The primary structure is an aluminum-lithium (Al-Li)cruciform design similar to the structural design of the Apollo Lunar Module. The cruciform planform layout results in four outer bays, with adequate volume in the corners between outer bays to package four landing gears. The central bay is reserved for packaging the MAV and the recessed MAV engines. Two of the outer bays accommodate main propellant tanks, with one LOX and one CH4 tank in each bay. The two remaining outer bays each house four rocket engine systems and associated support structure. Each HIAD comprises an inflatable structure, flexible thermal protection system, gas, and gas generators. The HIAD design used in this study is a stacked-toroid concept with pairing loop straps and radial/chevron straps. The baseline system lands a 20-t payload on the Mars surface. It is assumed the vehicle arrives at 6.2 km/s relative velocity at 90° inclination and is captured to a 1-Sol parking orbit.The baseline design includes many assumptions such as margins, arrival state, ConOps options, parking orbit, physical dimensions, propellant options, and technology concepts. The impact of these parameters are quantified through systems-level sensitivity analyses, which capture the global impact—not at a component level—but at the systems level. The systems-level sensitivities expose major design drivers and importance of each assumption for a design.Through tradespace exploration, a wide range of systems parameters are examined and compared for several feasible design options. Studies have been completed for the following input parameters: payload mass, propellant options, AC/EDL ballistic coefficient, lander thrust to weight ratio (T/W) (surrogate for the maximum EDL g’s),engine specific impulse (Isp), parking orbit, and inclination.The final paper will present and discuss the parametric approach used in the study. It will also include the results of recent systems analyses, sensitivity analyses, and tradespace exploration

Jamshid Samareh↗

NASA NTRS2004

The Wallops Flight Facility Rapid Response Range Operations Initiative

While the dominant focus on short response missions has appropriately centered on the launch vehicle and spacecraft, often overlooked or afterthought phases of these missions have been launch site operations and the activities of launch range organizations. Throughout the history of organized spaceflight, launch ranges have been the bane of flight programs as the source of expense, schedule delays, and seemingly endless requirements. Launch Ranges provide three basic functions: (1) provide an appropriate geographical location to meet orbital other mission trajectory requirements, (2) provide project services such as processing facilities, launch complexes, tracking and data services, and expendable products, and (3) assure safety and property protection to participating personnel and third-parties. The challenge with which launch site authorities continuously struggle, is the inherent conflict arising from projects whose singular concern is execution of their mission, and the range s need to support numerous simultaneous customers. So, while tasks carried out by a launch range committed to a single mission pale in comparison to efforts of a launch vehicle or spacecraft provider and could normally be carried out in a matter of weeks, major launch sites have dozens of active projects separate sponsoring organizations. Accommodating the numerous tasks associated with each mission, when hardware failures, weather, maintenance requirements, and other factors constantly conspire against the range resource schedulers, make the launch range as significant an impediment to responsive missions as launch vehicles and their cargo. The obvious solution to the launch site challenge was implemented years ago when the Department of Defense simply established dedicated infrastructure and personnel to dedicated missions, namely an Inter Continental Ballistic Missile. This however proves to be prohibitively expensive for all but the most urgent of applications. So the challenge becomes how can a launch site provide acceptably responsive mission services to a particular customer without dedicating extensive resources and while continuing to serve other projects? NASA's Wallops Flight Facility (WFF) is pursuing solutions to exactly this challenge. NASA, in partnership with the Virginia Commercial Space Flight Authority, has initiated the Rapid Response Range Operations Initiative (R3Ops). R3Ops is a multi-phased effort to incrementally establish and demonstrate increasingly responsive launch operations, with an ultimate goal of providing ELV-class services in a maximum of 7-10 days from initial notification routinely, and shorter schedules possible with committed resources. This target will be pursued within the reality of simultaneous concurrent programs, and ideally, largely independent of specialized flight system configurations. WFF has recently completed Phase 1 of R3Ops, an in-depth collection (through extensive expert interviews) and software modeling of individual steps by various range disciplines. This modeling is now being used to identify existing inefficiencies in current procedures, to identify bottlenecks, and show interdependencies. Existing practices are being tracked to provide a baseline to benchmark against as new procedures are implemented. This paper will describe in detail the philosophies behind WFF's R3Ops, the data collected and modeled in Phase 1, and strategies for meeting responsive launch requirements in a multi-user range environment planned for subsequent phases of this initiative.

Underwood, Bruce E.↗

NASA NTRS1987

Simulation of Airplane and Rocket Trajectories

Simulation and Optimization of Rocket Trajectories program (SORT) contains comprehensive mathematical models for simulating aircraft dynamics, freely falling objects, and many types of ballistic trajectories. Provides high-fidelity, three-degrees-of-freedom simulation for atmospheric and exoatmospheric flight. It numerically models vehicle subsystems and vehicle environment. Used for wide range of simulations. Written in machine-independent FORTRAN 77.

Wahbah, Magdy M.↗

NASA NTRS1991

RSRM-11 (36OW011) ballistics mass properties (STS-35)

The propulsion performance and reconstructed mass properties data from Thiolol's RSRM-11 motors which were assigned to the STS-35 launch are contained. The Thiokol manufacturing designations for the motors were 360W011A/360W011B, which are referred to as RSRM-11A and RSRM-B, respectively. The launch of STS-35 occurred on 2 December 1990 at the Eastern Test Range (ETR). The data contained herein was input to the STS-35 Flight Evaluation Report. The SRM propellant, TP-H1148, is a composite type solid propellants, formulated of polybutediene acrylic acid, acryonitrile terpolymer binder, epoxy curing agent, ammonium perchlorate oxidizer, and aluminum powder fuel. A small amount of burning rate catalyst (iron oxide) was added to achieve the desired propellant burn rate. The propellant evaluation and raw material information for the RSRM-11 are included. The ballistic performance presented was based on the Operational Flight Instrumentation (OFI) 12.5 sample per second pressure data for the steady state and tail off portion of the pressure trace. Recent studies have shown that the transducers are affected by the measuring system at KSC and temperature gradients created by the igniter heaters. Therefore, an adjustment to the data from each transducer is made to make the initial reading match the atmospheric pressure at the time of launch.

Hutchinson, B. J.↗

NASA NTRS1975

Outer planet entry nonequilibrium heating

An analytical tool has been developed which enables the impact of nonequilibrium effects on outer planet entry heating to be estimated. The analysis combines recent shock tube experiments, flow field calculations, and planetary entry trajectory analysis. The thickness of the nonequilibrium layer and its variation around the entry body have been correlated by a reaction flow parameter over a wide range of entry conditions. The influence of nonequilibrium effects on heating during entry into Saturn and Jupiter model atmospheres has been studied and the effect of vehicle size and ballistic coefficient determined. A nonequilibrium layer of significant thickness was found to exist during portions of entry into Saturn and Jupiter warm atmospheres. However, the nonequilibrium layer was found to be thin during the peak heating portions of the trajectory and resulted in reductions in total probe heating of less than 15 per cent.

Leibowitz, L. P.↗