Search NASA⌕ Search

SEARCH · Search NASA

Results for “ballistic performance”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Aero-Assisted Pre-Stage for Ballistic and Aero-Assisted Launch Vehicles

A concept of an aero-assisted pre-stage is proposed, which enables launch of both ballistic and aero-assisted launch vehicles from conventional runways. The pre-stage can be implemented as a delta-wing with a suitable undercarriage, which is mated with the launch vehicle, so that their flight directions are coaligned. The ample wing area of the pre-stage combined with the thrust of the launch vehicle ensure prompt roll-out and take-off of the stack at airspeeds typical for a conventional jet airliner. The launch vehicle is separated from the pre-stage as soon as safe altitude is achieved, and the desired ascent trajectory is reached. Nominally, the pre-stage is non-powered. As an option, to save the propellant of the launch vehicle, the pre-stage may have its own short-burn propulsion system, whereas the propulsion system of the launch vehicle is activated at the separation point. A general non-dimensional analysis of performance of the pre-stage from roll-out to separation is carried out and applications to existing ballistic launch vehicle and hypothetical aero-assisted vehicles (spaceplanes) are considered.

access to space↗
NASA NTRS1977

Aerodynamic behavior of the Viking entry vehicle - Ground test and flight results

An extensive series of tests of the Viking entry vehicle flying in pure CO2 was conducted in a ballistic range at Ames Research Center. The primary purpose of these tests was to calibrate the aerodynamic lift and drag characteristics in order to allow the density, pressure, and temperature profiles of the Martian atmosphere to be determined from onboard instrumentation carried on Viking. Both the Viking 1 and Viking 2 entry vehicles performed flawlessly during entry and descent, and the atmosphere structure was deduced to an altitude of about 120 km. A description is given of the ballistic range tests and of the aerodynamic behavior of the full scale entry vehicles during entry into the Martian atmosphere. Some comparisons between ground test and flight results are shown.

Kirk, D. B.↗
NASA NTRS1972

Potential advantages of solar electric propulsion for outer planet orbiters.

Past studies of solar electric propulsion for outer planet orbiters have generally emphasized the advantages of flight time reduction and payload increases. However, several subtle advantages exist, which may become important in an environment of increasingly difficult requirements as ways to extend current technology are sought. These advantages accrue primarily because of the inherent capability, unique to electric propulsion, to efficiently shape a trajectory while enroute. Stressed in this paper are: the ability to meet orbital constraints due to assumed radiation belts, science flexibility in a dual launch program, increased numbers of observational passes, and the lengthening of launch periods. These are examined for years representative of relatively easy and difficult ballistic missions. The results indicate that an early investment in solar electric technology will provide a strong performance foundation for a long range outer planet exploration program which evolves from current spacecraft technology.

Sauer, C. G.↗
NASA NTRS2017

Spacecraft Re-Entry Impact Point Targeting Using Aerodynamic Drag

The ability to re-enter the atmosphere at a desired location is important for spacecraft containing components that may survive re-entry. While impact point targeting has traditionally been initiated through impulsive burns with chemical thrusters on large vehicles such as the Space Shuttle, and the Soyuz and Apollo capsules, many small spacecraft do not host thrusters and require an alternative means of impact point targeting to ensure that falling debris do not cause harm to persons or property. This paper discusses the use of solely aerodynamic drag force to perform this targeting. It is shown that by deploying and retracting a drag device to vary the ballistic coefficient of the spacecraft, any desired longitude and latitude on the ground can be targeted provided that the maneuvering begins early enough and the latitude is less than the inclination of the orbit. An analytical solution based on perturbations from a numerically propagated trajectory is developed to map the initial state and ballistic coefficient profile of a spacecraft to its impact point. This allows the ballistic coefficient profile necessary to reach a given target point to be rapidly calculated, making it feasible to generate the guidance for the decay trajectory onboard the spacecraft. The ability to target an impact point using aerodynamic drag will enhance the capabilities of small spacecraft and will enable larger space vehicles containing thrusters to save fuel by more effectively leveraging the available aerodynamic drag.

propagation↗
NASA NTRS2007

Space Shuttle MMOD Threat Mitigation Techniques

Prior to each shuttle mission, threat assessments are performed to determine the risk of critical penetration, payload bay door radiator tube leak and crew module window replacement from Micrometeoroid and Orbital Debris (MMOD). Mission parameters, such as vehicle attitude, exposure time and altitude are used as inputs for the analysis. Ballistic limit equations, based on hypervelocity impact testing of shuttle materials are used to estimate the critical particle diameters of the outer surfaces of the vehicle. The assessments are performed using the BUMPER computer code at the NASA/JSC Hypervelocity Impact Technology Facility (HITF). The most critical involves the calculation of Loss of Crew and Vehicle (LOCV) risk. In recent years, NASA has implemented several techniques to reduce the risk to the Shuttle from MMOD impacts. This paper will describe on-orbit inspection of the reinforced carbon-carbon (RCC) regions and the methods used discern hypervelocity impact damage. Impact damage contingency plans and on-orbit repair techniques will also be discussed. The wing leading edge impact detection system (WLEIDS) and it's role in the reduction of on-orbit risk reduction will be presented. Finally, an analysis of the effectivity of alternative shuttle flight attitudes on MMOD risk will be demonstrated.

Hyde, Justin L.↗
NASA NTRS2008

Space Shuttle MMOD Threat Mitigation Techniques

Prior to each shuttle mission, threat assessments are performed to determine the risk of critical penetration, payload bay door radiator tube leak and crew module window replacement from Micrometeoroid and Orbital Debris (MMOD). Mission parameters, such as vehicle attitude, exposure time and altitude are used as inputs for the analysis. Ballistic limit equations, based on hypervelocity impact testing of shuttle materials are used to estimate the critical particle diameters of the outer surfaces of the vehicle. The assessments are performed using the BUMPER computer code at the NASA/JSC Hypervelocity Impact Technology Facility (HITF). The most critical involves the calculation of Loss of Crew and Vehicle (LOCV) risk. An overview of significant MMOD impacts on the Payload Bay Door radiators, wing leading edge reinforced carbon-carbon (RCC) panels and crew module windows will be presented, along with a discussion of the techniques NASA has implemented to reduce the risk from MMOD impacts. This paper will describe on-orbit inspection of the RCC regions and the methods used discern hypervelocity impact damage. Impact damage contingency plans and on-orbit repair techniques will also be discussed. The wing leading edge impact detection system (WLEIDS) and it s role in the reduction of on-orbit risk reduction will be presented. Finally, an analysis of alternative shuttle flight attitudes on MMOD risk will be demonstrated.

Hyde, J. L.↗
NASA NTRS2013

Atmospheric Entry Studies for Venus Missions: 45 Sphere-Cone Rigid Aeroshells and Ballistic Entries

The present study considers direct ballistic entries into the atmosphere of Venus using a 45deg sphere-cone rigid aeroshell, a legacy shape that has been used successfully in the past in the Pioneer Venus Multiprobe Mission. For a number of entry mass and heatshield diameter combinations (i.e., various ballistic coefficients) and entry velocities, the trajectory space in terms of entry flight path angles between skip out and -30deg is explored with a 3DoF trajectory code, TRAJ. From these trajectories, the viable entry flight path angle space is determined through the use of mechanical and thermal performance limits on the thermal protection material and science payload; the thermal protection material of choice is entry-grade carbon phenolic, for which a material thermal response model is available. For mechanical performance, a 200 g limit is placed on the peak deceleration load experienced by the science instruments, and 10 bar is assumed as the pressure limit for entry-grade carbon-phenolic material. For thermal performance, inflection points in the total heat load distribution are used as cut off criteria. Analysis of the results shows the existence of a range of critical ballistic coefficients beyond which the steepest possible entries are determined by the pressure limit of the material rather than the deceleration load limit.

Prabhu, Dinesh K.↗
NASA NTRS2012

Investigation of Transonic Wake Dynamics for Mechanically Deployable Entry Systems

A numerical investigation of transonic flow around a mechanically deployable entry system being considered for a robotic mission to Venus has been performed, and preliminary results are reported. The flow around a conceptual representation of the vehicle geometry was simulated at discrete points along a ballistic trajectory using Detached Eddy Simulation (DES). The trajectory points selected span the low supersonic to transonic regimes with freestream Mach numbers from 1:5 to 0:8, and freestream Reynolds numbers (based on diameter) between 2:09 x 10(exp 6) and 2:93 x 10(exp 6). Additionally, the Mach 0:8 case was simulated at angles of attack between 0 and 5 . Static aerodynamic coefficients obtained from the data show qualitative agreement with data from 70deg sphere-cone wind tunnel tests performed for the Viking program. Finally, the effect of choices of models and numerical algorithms is addressed by comparing the DES results to those using a Reynolds Averaged Navier-Stokes (RANS) model, as well as to results using a more dissipative numerical scheme.

Stern, Eric↗
NASA NTRS2006

Scaling Equations for Ballistic Modeling of Solid Rocket Motor Case Breach

This paper explores the development of a series of scaling equations that can take a known nominal motor performance and scale it for small and growing case failures. This model was developed for the Malfunction-Turn Study as part of Return to Flight activities for the Space Shuttle program. To verify the model, data from the Challenger accident (STS- 51L) were used. The model is able to predict the motor performance beyond the last recorded Challenger data and show how the failed right hand booster would have performed if the vehicle had remained intact.

McMillin, Joshua E.↗
NASA NTRS2015

OSIRIS-REx, Returning the Asteroid Sample

This paper addresses the technical aspects of the sample return system for the upcoming Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission. The overall mission design and current implementation are presented as an overview to establish a context for the technical description of the reentry and landing segment of the mission.The prime objective of the OSIRIS-REx mission is to sample a primitive, carbonaceous asteroid and to return that sample to Earth in pristine condition for detailed laboratory analysis. Targeting the near-Earth asteroid Bennu, the mission launches in September 2016 with an Earth reentry date of September 24, 2023.OSIRIS-REx will thoroughly characterize asteroid Bennu providing knowledge of the nature of near-Earth asteroids that is fundamental to understanding planet formation and the origin of life. The return to Earth of pristine samples with known geologic context will enable precise analyses that cannot be duplicated by spacecraft-based instruments, revolutionizing our understanding of the early Solar System. Bennu is both the most accessible carbonaceous asteroid and one of the most potentially Earth-hazardous asteroids known. Study of Bennu addresses multiple NASA objectives to understand the origin of the Solar System and the origin of life and will provide a greater understanding of both the hazards and resources in near-Earth space, serving as a precursor to future human missions to asteroids.This paper focuses on the technical aspects of the Sample Return Capsule (SRC) design and concept of operations, including trajectory design and reentry retrieval. Highlights of the mission are included below.The OSIRIS-REx spacecraft provides the essential functions for an asteroid characterization and sample return mission: attitude control propulsion power thermal control telecommunications command and data handling structural support to ensure successful rendezvous with Bennu characterization of Bennus properties delivery of the sampler to the surface, and return of the spacecraft to the vicinity of the Earth sample collection, performed by the Touch-and-Go Sample Acquisition Mechanism (TAGSAM), to acquire a regolith sample from the surface Earth re-entry and SRC recovery. Following sample collection, OSIRIS-REx drifts away from Bennu until the Asteroid Departure Maneuver is commanded on March 4, 2021, sending OSIRIS-REx on a ballistic return cruise to Earth. No additional large deterministic maneuvers are required to return the SRC to Earth. During the cruise, tracking and trajectory correction maneuvers (TCMs) are performed as necessary to precisely target the entry corridor. As OSIRIS-REx approaches Earth, the reentry plans are reviewed starting about a year before arrival, and preparations begin. The spacecraft is targeted away from the Earth until 7 days before entry. The final two trajectory correction maneuvers bring the spacecraft on target toward the Utah Test and Training Range (UTTR), with sufficient time for contingency resolution. The SRC releases 4 hours prior to atmospheric entry interface and, using the Stardust capsule heritage design, employs a traditional drogue and main parachute descent system for a soft touchdown.

reentry↗
NASA NTRS2016

Bumper 3 Update for IADC Protection Manual

The Bumper code has been the standard in use by NASA and contractors to perform meteoroid/debris risk assessments since 1990. It has undergone extensive revisions and updates [NASA JSC HITF website; Christiansen et al., 1992, 1997]. NASA Johnson Space Center (JSC) has applied BUMPER to risk assessments for Space Station, Shuttle, Mir, Extravehicular Mobility Units (EMU) space suits, and other spacecraft (e.g., LDEF, Iridium, TDRS, and Hubble Space Telescope). Bumper continues to be updated with changes in the ballistic limit equations describing failure threshold of various spacecraft components, as well as changes in the meteoroid and debris environment models. Significant efforts are expended to validate Bumper and benchmark it to other meteoroid/debris risk assessment codes. Bumper 3 is a refactored version of Bumper II. The structure of the code was extensively modified to improve maintenance, performance and flexibility. The architecture was changed to separate the frequently updated ballistic limit equations from the relatively stable common core functions of the program. These updates allow NASA to produce specific editions of the Bumper 3 that are tailored for specific customer requirements. The core consists of common code necessary to process the Micrometeoroid and Orbital Debris (MMOD) environment models, assess shadowing and calculate MMOD risk. The library of target response subroutines includes a board range of different types of MMOD shield ballistic limit equations as well as equations describing damage to various spacecraft subsystems or hardware (thermal protection materials, windows, radiators, solar arrays, cables, etc.). The core and library of ballistic response subroutines are maintained under configuration control. A change in the core will affect all editions of the code, whereas a change in one or more of the response subroutines will affect all editions of the code that contain the particular response subroutines which are modified. Note that the Bumper II program is no longer maintained or distributed by NASA.

Christiansen, Eric L.↗
DOE OSTI2023

A study of convective core overshooting as a function of stellar mass based on two-dimensional hydrodynamical simulations

ABSTRACT We perform two-dimensional (2D) numerical simulations of core convection for zero-age main-sequence stars covering a mass range from 3 to 20 M⊙. The simulations are performed with the fully compressible time-implicit code music. We study the efficiency of overshooting, which describes the ballistic process of convective flows crossing a convective boundary, as a function of stellar mass and luminosity. We also study the impact of artificially increasing the stellar luminosity for 3 M⊙ models. The simulations cover hundreds to thousands of convective turnover time-scales. Applying the framework of extreme plume events previously developed for convective envelopes, we derive overshooting lengths as a function of stellar masses. We find that the overshooting distance (dov) scales with the stellar luminosity (L) and the convective core radius (rconv). We derive a scaling law $d_{\rm ov} \propto L^{1/3} r_{\rm conv}^{1/2}$, which is implemented in a one-dimensional stellar evolution code and the resulting stellar models are compared to observations. The scaling predicts values for the overshooting distance that significantly increase with stellar mass, in qualitative agreement with observations. Quantitatively, however, the predicted values are underestimated for masses ≳10 M⊙. Our 2D simulations show the formation of a nearly adiabatic layer just above the Schwarzschild boundary of the convective core, as exhibited in recent three-dimensional simulations of convection. The most luminous models show a growth in size with time of the nearly adiabatic layer. This growth seems to slow down as the upper edge of the nearly adiabatic layer gets closer to the maximum overshooting length and as the simulation time exceeds the typical thermal diffusive time-scale in the overshooting layer.

79 ASTRONOMY AND ASTROPHYSICS↗
NASA NTRS1991

Propellant development for the Advanced Solid Rocket Motor

The properties of a propellant developed for the NASA Advanced Solid Rocket Motor (ASRM) are described in terms of its composition, performance, and compliance to NASA specifications. The class 1.3 HTPB/AP/A1 propellant employs an ester plasticizer and the content of ballistic solids is set at 88 percent. Ammonia evolution is prevented by the utilization of a neutral bonding agent which allows continuous mixing. The propellant also comprises a bimodal AP blend with one ground fraction, ground AP of at least 20 microns, and ferric oxide to control the burning rate. The propellant's characteristics are discussed in terms of tradeoffs in AP particle size and the types of Al powder, bonding agent, and HTPB polymer. The size and shape of the ballistic solids affect the processability, ballistic properties, and structural properties of the propellant. The revised baseline composition is based on maximizing the robustness of in-process viscosity, structural integrity, and burning-rate tailoring range.

Landers, L. C.↗
NASA NTRS1990

Flight motor set 360L001 (STS-26R). Volume 1: System overview, revision A

The NASA space shuttle flight STS-26R, launched at 11:37.00.009 am, EDT on 29 Sep. 1988, used the redesigned solid rocket motors (RSRM) 360L001A and 360L001B. Evaluation of the ground environment instrumentation (GEI) data recorded prior to flight showed no launch commit criteria violations; that the field joint heater and aft skirt thermal conditioning systems performed adequately; and that the GEI data showed good agreement with thermal model predictions. Evaluation of the developmental flight instrumentation (DFI) revealed excellent agreement with both the predicted and required ballistic specifications. All parameters were well within the GEI specification requirements including propellant burn rates, specific impulse values, and thrust imbalance. Recorded strain values also indicated satisfactory radial growth and stress levels, as well as verification of adequate safety factors. Postflight inspection of the insulation, seals, case, and nozzles showed overall excellent performance. Some thermal DFI protective cork was missing, and inoperative field joint vent valves on the thermal protection cork allowed water entry into the field joints upon splashdown. Evaluation of these anomalies, as well as complete evaluation of all Redesigned Solid Rocket Motor components, is contained.

Garecht, Diane M.↗
NASA NTRS1994

Power Electronics Development for the SPT-100 Thruster

Russian electric propulsion technologies have recently become available on the world market. Of significant interest is the Stationary Plasma Thruster (SPT) which has a significant flight heritage in the former Soviet space program. The SPT has performance levels of up to 1600 seconds of specific impulse at a thrust efficiency of 0.50. Studies have shown that this level of performance is well suited for stationkeeping applications, and the SPT-100, with a 1.35 kW input power level, is presently being evaluated for use on Western commercial satellites. Under a program sponsored by the Innovative Science and Technology Division of the Ballistic Missile Defense Organization, a team of U.S. electric propulsion specialists observed the operation of the SPT-100 in Russia. Under this same program, power electronics were developed to operate the SPT-100 to characterize thruster performance and operation in the U.S. The power electronics consisted of a discharge, cathode heater, and pulse igniter power supplies to operate the thruster with manual flow control. A Russian designed matching network was incorporated in the discharge supply to ensure proper operation with the thruster. The cathode heater power supply and igniter were derived from ongoing development projects. No attempts were made to augment thruster electromagnet current in this effort. The power electronics successfully started and operated the SPT-100 thruster in performance tests at NASA Lewis, with minimal oscillations in the discharge current. The efficiency of the main discharge supply was measured at 0.92, and straightforward modifications were identified which could increase the efficiency to 0.94.

Hamley, John A.↗
NASA NTRS1988

Flight motor set 360L001 (STS-26R), volume 1

The NASA space shuttle flight STS-26R, launched at 11:37.00.009 a.m. EDT on 29 Sep. 1988, used the redesigned solid rocket motors (RSRM) 360LOO1A and 360LOO1B. Evaluation of the ground environment instrumentation (GEI) data recorded prior to flight showed: (1) no launch commit criteria violations, (2) that the field joint heater and aft skirt thermal conditioning systems performed adequately, and (3) that the GEI data showed good agreement with thermal model predictions. Evaluation of the developmental flight instrumentation (DFI) revealed excellent agreement with both the predicted and required ballistic specifications. All parameters were well within the CEI specification requirements including propellant burn rates, specific impulse values, and thrust imbalance. Recorded strain values also indicated satisfactory radial growth and stress levels, as well as verification of adequate safety factors. Postflight inspection of the insulation, seals, case, and nozzles showed overall excellent performance. Some thermal DFI protective cork was missing, and inoperative field joint vent valves on the thermal protection cork allowed water entry into the field joints upon splashdown. Evaluation of these anomalies, as well as complete evaluation of all RSRM components, is presented.

Ricks, Glen A.↗
NASA NTRS1981

The influence of ship motion of manual control skills

The effects of ship motion on a range of typical manual control skills were examined on the Warren Spring ship motion simulator driven in heave, pitch, and roll by signals taken from the frigate HMS Avenger at 13 m/s (25 knots) into a force 4 wind. The motion produced a vertical r.m.s. acceleration of 0.024g, mostly between 0.1 and 0.3 Hz, with comparatively little pitch or roll. A task involving unsupported arm movements was seriously affected by the motion; a pursuit tracking task showed a reliable decrement although it was still performed reasonably well (pressure and free moving tracking controls were affected equally by the motion); a digit keying task requiring ballistic hand movements was unaffected. There was no evidence that these effects were caused by sea sickness. The differing response to motion of the different tasks, from virtual destruction to no effect, suggests that a major benefit could come from an attempt to design the man/control interface onboard ship around motion resistant tasks.

Mcleod, P.↗
NASA NTRS1988

A comparison of experimental and computational results for 5 and 10 degree cones at high Mach numbers

Experiments were performed on 5 and 10 deg slender cones at a velocity of approximately 5 km/sec in the NASA-Ames ballistic ranges. The flowfields for the cones were computed using ideal-gas and chemical nonequilibrium-air parabolized Navier-Stokes codes. Experimentally determined drag coefficients and shock shapes are compared with the results of the computer codes. Both the flight-data analysis methods and the computational codes are examined to achieve the most meaningful comparison. Under the conditions of the experiments, skin-friction drag makes up approximately 50 percent of the total drag for the 5 deg cone and 30 percent of the total drag for the 10 deg cone. Computed drag coefficients of the 10 deg cone agree well with the experimental values; predictions fall below the experimental values for the 5 deg cone.

Strawa, A. W.↗