Trajectory simulation for axisymmetric planetary entry vehicles.
Computer program for parametric analysis and trajectory simulation for axisymmetric planetary entry vehicles
PLANETARY ENTRY↗
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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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Biosatellite 2 experiment environment
Reentry vehicle and experiment packages of Biosatellite 2 spacecraft
Look, B. C.↗
Processing and analysis of data from the University of Chicago experiments on the IMP 4 and 5 spacecraft
Data from the Charged Particle Experiments of the IMP-4 and 5 spacecraft were processed and analyzed from launch through reentry. Approximately 95 percent of all data transmitted was recovered, processed, and analyzed. The computer programming and operational procedures which were developed to perform the processing and analysis of the data are discussed. The scientific results obtained from the analysis of the data are summarized. A bibliography of documents based on these experiments is included.
Simpson, J. A.↗
Investigation of sonic boom for the Space Shuttle: Low cross-range orbiter
It is desired that the Space Shuttle Orbiter be capable of landing at airports equipped to handle present-day jet transports. Since the majority of such airports are located near heavily populated areas, an investigation has been undertaken to determine whether or not the sonic boom generated during reentry of Space Shuttle Orbiters is potentially a serious problem. The investigation was concerned with the low cross-range orbiter and reentry concept proposed by Faget of the Manned Spacecraft Center (MSC). This report describes the approach used and presents the results obtained to date.
Levy, Lionel L., Jr.↗
Managing Space Radiation Risk in the New Era of Space Exploration
Space exploration is a risky enterprise. Rockets launch astronauts at enormous speeds into a harsh, unforgiving environment. Spacecraft must withstand the bitter cold of space and the blistering heat of reentry. Their skin must be strong enough to keep the inside comfortably pressurized and tough enough to resist damage from micrometeoroids. Spacecraft meant for lunar or planetary landings must survive the jar of landing, tolerate dust, and be able to take off again. For astronauts, however, there is one danger in space that does not end when they step out of their spacecraft. The radiation that permeates space -- unattenuated by Earth s atmosphere and magnetosphere -- may damage or kill cells within astronauts bodies, resulting in cancer or other health consequences years after a mission ends. The National Aeronautics and Space Administration (NASA) has recently embarked on Project Constellation to implement the Vision for Space Exploration -- a program announced by President George W. Bush in 2004 with the goal of returning humans to the Moon and eventually transporting them to Mars. To adequately prepare for the safety of these future space explorers, NASA s Exploration Systems Mission Directorate requested that the Aeronautics and Space Engineering Board of the National Research Council establish a committee to evaluate the radiation shielding requirements for lunar missions and to recommend a strategic plan for developing the radiation mitigation capabilities needed to enable the planned lunar mission architecture
Source record↗
Genesis Preliminary Examination: Ellipsometry Overview
The Genesis spacecraft returned to Earth on September 8, 2004, experiencing a non-nominal reentry in which both the drogue and main parachutes failed to deploy causing the capsule to impact the surface of the UTTR desert at a speed of approximately 310 kph (193 mph). The impact caused severe damage to the capsule and a breach of the science canister in the field. The science canister was recovered and transported to the cleanroom at UTTR within approximately 8 hours of reentry. Although the ground water table did not rise to canister level before removal, damp soil and debris from the heat shield and other spacecraft components did enter the canister and contaminate some collector surfaces. The objective of preliminary examination of the Genesis collectors is to provide the science community with the information necessary to request the most useful samples for their analysis.
Stansbery, E. K.↗
Orbit, reentry, and landing attachment for globes
Navigational device, invented to aid recovery of spacecraft from any orbit, also illustrates motions of satellites relative to earth and their entry-ranging requirements. Device rapidly and accurately defines lateral range requirements for spacecraft returning to any desired site without manual or computerized calculation of orbital equations of motion.
Pritchard, E. B.↗
Development of a Composite Material Aerodynamic Demise Model for the Object Reentry Survival Analysis Tool (ORSAT)
The reentry demise of fiber-reinforced polymer (FRP) composites is an increasing concern for modern spacecraft at end of life. Unlike traditional materials such as metals, shredding of the material by aerodynamic forces appears to be a major component of the reentry demise mechanism. This paper will describe a new mechanical, strength-based, material demise model for the Object Reentry Survival Analysis Tool (ORSAT) Version 7.1. The model is based on laboratory in-situ and residual strength tests of several FRP materials performed during the NASA Orbital Debris Program Office’s Phase II Composite Material Demise test campaign. To test the residual strength of partially charred FRP materials, the authors tested 107 rectangular shaped samples of different thicknesses that were previously exposed to high-enthalpy flow at the University of Texas at Austin’s Inductively Coupled Plasma Torch facility. Tests were performed under normal atmospheric conditions at NASA Johnson Space Center’s Experimental Impact Laboratory using a Chatillon TCD1000 tensile test machine configured with a three-point bending jig. Mean fracture load, residual strength, delamination, and flexural modulus were measured for each test. An engineering model of aeromechanical demise for each material was developed by correlating the measured residual strength of the samples with the duration and magnitude of the applied heat flux and the mass loss and char progression. This model has been implemented in ORSAT 7.1 for the built-in, charring carbon fiber/epoxy and charring glass fiber/epoxy material models. The new model was verified by calculating the demise of hundreds of FRP composite fragments of various shapes and sizes and checking the calculations for consistency, with specific focus on material characterization in one-dimension. Further tests are planned in hypersonic flow facilities to validate the assumptions used in the model.
Priscilla A. Mendoza↗
Development of a Composite Material Aerodynamic Demise Model for the Object Reentry Survival Analysis Tool (ORSAT)
The reentry demise of fiber-reinforced polymer (FRP) composites is an increasing concern for modern spacecraft at end of life. Unlike traditional materials such as metals, shredding of the material by aerodynamic forces appears to be a major component of the reentry demise mechanism. This paper will describe a new mechanical, strength-based, material demise model for the Object Reentry Survival Analysis Tool (ORSAT) Version 7.1. The model is based on laboratory in-situ and residual strength tests of several FRP materials performed during the NASA Orbital Debris Program Office’s Phase II Composite Material Demise test campaign. To test the residual strength of partially charred FRP materials, the authors tested 107 rectangular shaped samples of different thicknesses that were previously exposed to high-enthalpy flow at the University of Texas at Austin’s Inductively Coupled Plasma Torch facility. Tests were performed under normal atmospheric conditions at NASA Johnson Space Center’s Experimental Impact Laboratory using a Chatillon TCD1000 tensile test machine configured with a three-point bending jig. Mean fracture load, residual strength, delamination, and flexural modulus were measured for each test. An engineering model of aeromechanical demise for each material was developed by correlating the measured residual strength of the samples with the duration and magnitude of the applied heat flux and the mass loss and char progression. This model has been implemented in ORSAT 7.1 for the built-in, charring carbon fiber/epoxy and charring glass fiber/epoxy material models. The new model was verified by calculating the demise of hundreds of FRP composite fragments of various shapes and sizes and checking the calculations for consistency, with specific focus on material characterization in one-dimension. Further tests are planned in hypersonic flow facilities to validate the assumptions used in the model.
Priscilla A. Mendoza↗
Simulation for Prediction of Entry Article Demise (SPEAD): an Analysis Tool for Spacecraft Safety Analysis and Ascent/Reentry Risk Assessment
For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of For the purpose of performing safety analysis and risk assessment for a probable offnominal suborbital/orbital atmospheric reentry resulting in vehicle breakup, a synthesis of trajectory propagation coupled with thermal analysis and the evaluation of node failure is required to predict the sequence of events, the timeline, and the progressive demise of spacecraft components. To provide this capability, the Simulation for Prediction of Entry Article Demise (SPEAD) analysis tool was developed. This report discusses the capabilities, modeling, and validation of the SPEAD analysis tool. SPEAD is applicable for Earth or Mars, with the option for 3 or 6 degrees-of-freedom (DOF) trajectory propagation. The atmosphere and aerodynamics data are supplied in tables, for linear interpolation of up to 4 independent variables. The gravitation model can include up to 20 zonal harmonic coefficients. The modeling of a single motor is available and can be adapted to multiple motors. For thermal analysis, the aerodynamic radiative and free-molecular/continuum convective heating, black-body radiative cooling, conductive heat transfer between adjacent nodes, and node ablation are modeled. In a 6- DOF simulation, the local convective heating on a node is a function of Mach, angle-ofattack, and sideslip angle, and is dependent on 1) the location of the node in the spacecraft and its orientation to the flow modeled by an exposure factor, and 2) the geometries of the spacecraft and the node modeled by a heating factor and convective area. Node failure is evaluated using criteria based on melting temperature, reference heat load, g-load, or a combination of the above. The failure of a liquid propellant tank is evaluated based on burnout flux from nucleate boiling or excess internal pressure. Following a component failure, updates are made as needed to the spacecraft mass and aerodynamic properties, nodal exposure and heating factors, and nodal convective and conductive areas. This allows the trajectory to be propagated seamlessly in a single run, inclusive of the trajectories of components that have separated from the spacecraft. The node ablation simulates the decreasing mass and convective/reference areas, and variable heating factor. A built-in database provides the thermo-mechanical properties of
Ling, Lisa↗
Summary of EOS flight dynamics analysis
From a flight dynamics perspective, the Earth Observing System (EOS) spacecraft present a number of challenges to mission designers. The Flight Dynamics Support Branch of NASA GSFC has examined a number of these challenges, including managing the EOS constellation, disposing of the spacecraft at the end-of-life (EOL), and achieving the appropriate mission orbit given launch vehicle and ascent propulsion constraints. The EOS program consists of a number of spacecraft including EOS-AM, an ascending node spacecraft, EOS-PM, a descending node spacecraft, the EOS Chemistry mission (EOS-CHEM), the EOS Altimetry Laser (EOS-LALT), and the EOS-Altimetry Radar (EOS-RALT). The orbit characteristics of these missions are presented. In order to assure that downlinking data from each spacecraft will be possible without interference between any two spacecraft, a careful examination of the relationships between spacecraft and how to maintain the spacecraft in a configuration which would minimize these communications problems must be made. The FDSB has performed various analyses to determine whether the spacecraft will be in a position to interfere with each other, how the orbit dynamics will change the relative positioning of the spacecraft over their lifetimes, and how maintenance maneuvers could be performed, if needed, to minimize communications problems. Prompted by an activity at NASA HQ to set guidelines for spacecraft regarding their end-of-life dispositions, much analysis has also been performed to determine the spacecraft lifetime of EOS-AM1 under various conditions, and to make suggestions regarding the spacecraft disposal. In performing this analysis, some general trends have been observed in lifetime calculations. The paper will present the EOS-AM1 lifetime results, comment on general reentry conclusions, and discuss how these analyses reflect on the HQ NMI. Placing the EOS spacecraft into their respective mission orbits involves some intricate maneuver planning to assure that all mission orbit requirements are met, given the initial conditions supplied by the launch vehicle at injection. The FDSB has developed an ascent scenario to meet the mission requirements. This paper presents results of the ascent analysis.
Newman, Lauri Kraft↗
Radiative properties of advanced spacecraft heat shield materials
Experimental results are presented to show the effects of simulated reentry exposure by convective heating and by radiant heating on spectral and total emittance of statically oxidized Inconel 617 and Haynes HS188 superalloys to 1260 K and a silicide coatea (R512E) columbium 752 alloy to 1590 K. Convective heating exposures were conducted in a supersonic arc plasma wind tunnel using a wedge-shaped specimen configuration. Radiant tests were conducted at a pressure of .003 atmospheres of dry air at a flow velocity of several meters per second. Convective heating specimens were subjected to 8, 20, and 38 15-min heating cycles, and radiant heating specimens were tested for 10, 20, 50, and 100 30-min heating cycles. Changes in radiative properties are explained in terms of changes in composition resulting from simulated reentry tests. The methods used to evaluate morphological, compositional and crystallographic changes include: Auger electron spectroscopy; scanning electron microscopy; X-ray diffraction analysis; and electron microprobe analysis.
Cunnington, G. R.↗
Navigation accuracy for infrared equipped Apollo aircraft for reentry tracking
Navigational accuracy of infrared acquisition and tracking system for Apollo reentry aircraft
INFRARED TRACKING↗
Short and long periodic atmospheric variations between 25 and 200 km
Previously collected data on atmospheric pressure, density, temperature and winds between 25 and 200 km from sources including Meteorological Rocket Network data, ROBIN falling sphere data, grenade release and pitot tube data, meteor winds, chemical release winds, satellite data, and others were analyzed by a daily difference method and results on the distribution statistics, magnitude, and spatial structure of gravity wave and planetary wave atmospheric variations are presented. Time structure of the gravity wave variations were determined by the analysis of residuals from harmonic analysis of time series data. Planetary wave contributions in the 25-85 km range were discovered and found to have significant height and latitudinal variation. Long period planetary waves, and seasonal variations were also computed by harmonic analysis. Revised height variations of the gravity wave contributions in the 25 to 85 km height range were computed. An engineering method and design values for gravity wave magnitudes and wave lengths are given to be used for such tasks as evaluating the effects on the dynamical heating, stability and control of spacecraft such as the space shuttle vehicle in launch or reentry trajectories.
Justus, C. G.↗
Thermal insulation protection means
A system for providing thermal insulation for portions of a spacecraft which do not exceed 900 F during ascent or reentry relative to the earth's atmosphere is described. The thermal insulation is formed of relatively large flexible sheets of needled Nomex felt having a flexible waterproof coating. The thickness of the felt is sized to protect against projected temperatures and is attached to the structure by a resin adhesive. Vent holes in the sheets allow ventilation while maintaining waterproofing. The system is heat treated to provide thermal stability.
Dotts, R. L.↗
Gemini 8, This is Houston Flight
The historic first docking in space with the Agena is completed. Camaras record the harrowing experiences of the astronauts as Gemini VIII wildly gyrates through space following a malfunction. The spacecraft is separated from the Agena, brought under control and reentry is achieved.
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Fracture in Phenolic Impregnated Carbon Ablator
The thermal protection materials used for spacecraft heat shields are subjected to various thermal-mechanical loads during an atmospheric reentry which can threaten the structural integrity of the system. This paper discusses the development of a novel technique to understand the failure mechanisms inside the thermal protection material, Phenolic Impregnated Carbon Ablator (PICA). PICA has successfully flown on the Stardust spacecraft and was the TPS material chosen for the Mars Science Laboratory (MSL), that will fly in 2011. Although PICA has good thermal properties, structurally, it is a weak material. To thoroughly understand failure in PICA, experiments were performed using FiberForm(Registered TradeMark) (precursor of PICA), virgin and furnace-charred PICA. Several small samples were tested inside an electron microscope to investigate the failure mechanisms. Micrographs were obtained before and after the failure in order to study crack initiation and growth. Videos were obtained to capture failure mechanisms in real time. Stress-strain data was obtained simultaneously for all the samples with the help of a data acquisition system, integrated to the mechanical stages. It was found that cracks initiated and grew in the FiberForm when a critical stress limit was reached such that the carbon fibers separated from the binder. However, both for virgin and charred PICA, crack initiation and growth occurred in the matrix (phenolic) phase. Both virgin and charred PICA showed greater strength values compared to FiberForm coupons, confirming that the presence of the porous matrix helps in absorbing the fracture energy.
Agrawal, Parul↗
Integral Launch and Reentry Vehicle System. Volume 1, Book 1 - Configuration Design and Subsystems
Two stage to orbit reusable spacecraft system for use in transporting cargo and passengers to and from earth orbital space stations
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