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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 307 records · Page 17
All sourcesNASA NTRSDOE OSTI
NASA NTRS1976

Planetary/DOD entry technology flight experiments. Volume 3: Planetary entry flight experiments handbook

The environments produced by entry into Jupiter and Saturn atmospheres are summarized. Worst case design environments are identified and the effect of entry angle, type of atmosphere and ballistic coefficient variations are presented. The range of environments experienced during earth entry is parametrically described as a function of initial entry conditions. The sensitivity of these environments to vehicle ballistic coefficient and nose radius is also shown. An elliptical deorbit maneuver strategy is defined in terms of the velocity increment required versus initial entry conditions and apoapsis altitude. Mission time, ground track, and out of plane velocity penalties are also presented. Performance capabilities of typical shuttle launched boosters are described including the initial entry conditions attainable as a function of paylaod mass and apoapsis altitude.

Christensen, H. E.↗
NASA NTRS1992

Space debris characterization in support of a satellite breakup model

The Space Kinetic Impact and Debris Branch began an ambitious program to construct a fully analytical model of the breakup of a satellite under hypervelocity impact. In order to provide empirical data with which to substantiate the model, debris from hypervelocity experiments conducted in a controlled laboratory environment were characterized to provide information of its mass, velocity, and ballistic coefficient distributions. Data on the debris were collected in one master data file, and a simple FORTRAN program allows users to describe the debris from any subset of these experiments that may be of interest to them. A statistical analysis was performed, allowing users to determine the precision of the velocity measurements for the data. Attempts are being made to include and correlate other laboratory data, as well as those data obtained from the explosion or collision of spacecraft in low earth orbit.

Fortson, Bryan H.↗
NASA NTRS1955

A comparative analysis of the performance of long-range hypervelocity vehicles

Long-range hypervelocity vehicles are studied in terms of their motion in powered flight, and their motion and aerodynamic heating in unpowered flight. Powered flight is analyzed for an idealized propulsion system which rather closely approaches present-day rocket motors. Unpowered flight is characterized by a return to earth along a ballistic, skip, or glide trajectory. Only those trajectories are treated which yield the maximum range for a given velocity at the end of powered flight. Aerodynamic heating is treated in a manner similar to that employed previously by the senior authors is studying ballistic missiles (NACA RM A53D28), with the exception that radiant as well as convective heat transfer is considered in connection with glide and skip vehicles. The ballistic vehicle is found to be the least efficient of the several types studied in the sense that it generally requires the highest velocity at the end of powered flight in order to attain a given range. This disadvantage may be offset, however, by reducing convective heat transfer to the re-entry body through the artifice of increasing pressure drag in relation to friction drag - that is, by using a blunt body. Thus the kinetic energy required by the vehicle at the end of powered flight may be reduced by minimizing the mass of coolant material involved.

Eggers, Alfred J , Jr↗
NASA NTRSDate not supplied

Ballistic Lunar Return Trajectories for Sustainable Cargo Return and Entry System Technology Development

As part of the sustained approach for the lunar Gateway and Artemis program as a whole, NASA is extending the logistics supply chain beyond low Earth orbit and to the Moon. This supply chain includes the possibility of lunar sample and cargo return. To enable these future return missions, the possibility for incorporating demonstration payloads including various entry, descent, and landing technologies is examined. Possible implementations include deployable entry vehicles, high speed sample return capsules, aeroassist technologies, and control technologies for guided hypersonic flight. The mission concepts utilize the secondary payload capabilities provided on a relatively low-cost logistic module. The logistic module may perform close Earth flyby, pointing, and release of return systems prior to disposal, with reentry velocities for payloads on the order of 11.5 km/sec. In this paper, we study the various return mission architectures available during the Artemis program to define the range of mission possibilities. Potential options include a reusable logistics module using a hypersonic inflatable aerodynamic decelerator, externally mounted entry system vehicle technologies, and a reusable sample return vehicle using a similar inflatable architecture with a feathered configuration with applied aerodynamic control. In these cases, the use of a ballistic lunar return trajectory is assumed, and a sensitivity analysis of midcourse corrections and the possibility of a lunar gravity assist for refining Earth entry interface points is provided. Furthermore, a novel controller for controlling a vehicle during reentry imposing heating limitations is introduced.

Matthew M. Wittal↗
NASA NTRS1996

Design of a Subscale Propellant Slag Evaluation Motor Using Two-Phase Fluid Dynamic Analysis

Small pressure perturbations in the Space Shuttle Reusable Solid Rocket Motor (RSRM) are caused by the periodic expulsion of molten aluminum oxide slag from a pool that collects in the aft end of the motor around the submerged nozzle nose during the last half of motor operation. It is suspected that some motors produce more slag than others due to differences in aluminum oxide agglomerate particle sizes that may relate to subtle differences in propellant ingredient characteristics such as particle size distributions or processing variations. A subscale motor experiment was designed to determine the effect of propellant ingredient characteristics on the propensity for slag production. An existing 5 inch ballistic test motor was selected as the basic test vehicle. The standard converging/diverging nozzle was replaced with a submerged nose nozzle design to provide a positive trap for the slag that would increase the measured slag weights. Two-phase fluid dynamic analyses were performed to develop a nozzle nose design that maintained similitude in major flow field features with the full scale RSRM. The 5 inch motor was spun about its longitudinal axis to further enhance slag collection and retention. Two-phase flow analysis was used to select an appropriate spin rate along with other considerations, such as avoiding bum rate increases due to radial acceleration effects. Aluminum oxide particle distributions used in the flow analyses were measured in a quench bomb for RSRM type propellants with minor variations in ingredient characteristics. Detailed predictions for slag accumulation weights during motor bum compared favorably with slag weight data taken from defined zones in the subscale motor and nozzle. The use of two-phase flow analysis proved successful in gauging the viability of the experimental program during the planning phase and in guiding the design of the critical submerged nose nozzle.

Whitesides, R. Harold↗
NASA NTRSDate not supplied

Multiscale Modeling of Woven Ablative Thermal Protection System Materials

The NASA Entry Systems Modeling project maintains a portfolio of computational model and tool development activities focused on reducing performance uncertainties in ablative Thermal Protection System (TPS) materials for NASA missions. The development activities span material scale and strive to allow microstructural characterization of material structure and properties, mesoscale analyses of damage, and macroscale evaluation of heatshield performance and recession in a given aerothermodynamic environment. This talk will detail the application of developed capabilities at all three scales to the woven TPS material that the Agency has selected as the heatshield for the Mars Sample Return Earth Entry System (MSR-EES) mission – 3D Mid-Density Carbon Phenolic (3MDCP). Each of the applications focuses on driving down uncertainties in material performance and thus risk for MSR-EES and other future missions that may leverage woven TPS. At the microscale, machine learning techniques are used to characterize images from destructive microscopy and inform structural variability. At the mesoscale, Lagrangian techniques are used to simulate ballistic impact and interpret damage modes noted in experiments. At the macroscale, coupled flow-material response techniques are validated by Arc Jet testing to enable heatshield design for missions with massive ablation.

Justin B Haskins↗
NASA NTRS2011

Capabilities of the Impact Testing Facility at Marshall Space Flight Center

The test and analysis capabilities of the Impact Testing Facility at NASA's Marshall Space Flight Center are described. Nine different gun systems accommodate a wide range of projectile and target sizes and shapes at velocities from subsonic through hypersonic, to accomplish a broad range of ballistic and hypervelocity impact tests. These gun systems include ballistic and microballistic gas and powder guns, a two-stage light gas gun, and specialty guns for weather encounter studies. The ITF "rain gun" is the only hydrometeor impact gun known to be in existence in the United States that can provide single impact performance data with known raindrop sizes. Simulation of high velocity impact is available using the Smooth Particle Hydrodynamic Code. The Impact Testing Facility provides testing, custom test configuration design and fabrication, and analytical services for NASA, the Department of Defense, academic institutions, international space agencies, and private industry in a secure facility located at Marshall Space Flight Center, on the US Army's Redstone Arsenal in Huntsville, Alabama. This facility performs tests that are subject to International Traffic in Arms Regulations (ITAR) and DoD secret classified restrictions as well as proprietary and unrestricted tests for civil space agencies, academic institutions, and commercial aerospace and defense companies and their suppliers.

Finchum, Andy↗
NASA NTRS1991

Single wall penetration equations

Five single plate penetration equations are compared for accuracy and effectiveness. These five equations are two well-known equations (Fish-Summers and Schmidt-Holsapple), two equations developed by the Apollo project (Rockwell and Johnson Space Center (JSC), and one recently revised from JSC (Cour-Palais). They were derived from test results, with velocities ranging up to 8 km/s. Microsoft Excel software was used to construct a spreadsheet to calculate the diameters and masses of projectiles for various velocities, varying the material properties of both projectile and target for the five single plate penetration equations. The results were plotted on diameter versus velocity graphs for ballistic and spallation limits using Cricket Graph software, for velocities ranging from 2 to 15 km/s defined for the orbital debris. First, these equations were compared to each other, then each equation was compared with various aluminum projectile densities. Finally, these equations were compared with test results performed at JSC for the Marshall Space Flight Center. These equations predict a wide variety of projectile diameters at a given velocity. Thus, it is very difficult to choose the 'right' prediction equation. The thickness of a single plate could have a large variation by choosing a different penetration equation. Even though all five equations are empirically developed with various materials, especially for aluminum alloys, one cannot be confident in the shield design with the predictions obtained by the penetration equations without verifying by tests.

Hayashida, K. B.↗
NASA NTRS2010

Multi-Mission Earth Entry Vehicle: Aerodynamic and Aerothermal Analysis of Trajectory Environments

Multi-mission Earth Entry Vehicle (MMEEV) is designed to deliver small payloads from space to Earth's surface by flying an uncontrolled ballistic entry, which ends with ground impact. The included range of entry velocities is from 10 to 16 km/s. The range of ballistic coefficients is from 41.94 to 128.74 kg/m2, which insures a low subsonic terminal velocity on the order of 50 m/sec. The range of entry flight path angles, considered in this analysis is from -5 to -25 degrees. The assessment and parametric characterization of aeroheating and aerodynamic performance of the capsule during entry is the subject of this paper.

Trumble, Kerry↗
NASA NTRS1979

Feasibility of SiC composite structures for 1644 deg gas turbine seal applications

The feasibility of silicon carbide composite structures was evaluated for 1644 K gas turbine seal applications. The silicon carbide composites evaluated consisted of Si/SiC Silcomp (Trademark) - and sintered silicon carbide as substrates, both with attached surface layers containing BN as an additive. A total of twenty-eight candidates with variations in substrate type and density, and layer chemistry, density, microstructure, and thickness were evaluated for abradability, cold particle erosion resistance, static oxidation resistance, ballistic impact resistance, and fabricability. The BN-free layers with variations in density and pore size were later added for evaluation. The most promising candidates were evaluated for Mach 1.0 gas oxidation/erosion resistance from 1477 K to 1644 K. The as-fabricated rub layers did not perform satisfactorily in the gas oxidation/erosion tests. However, preoxidation was found to be beneficial in improving the hot gas erosion resistance. Overall, the laboratory and rig test evaluations show that material properties are suitable for 1477 K gas turbine seal applications.

Darolia, R.↗
NASA NTRS2022

Impact Simulation of Composite Panels for Aerospace Applications

The aerospace research community has been working towards the development of a comprehensive framework for the design of composite structures at the component, subassembly, assembly, and system levels. Such a framework requires accurate, efficient and easy-to-use mathematical and computational models that are in sync with physical reality. The research work in this paper is designed to help an aerospace structural engineer move towards this goal by simulating dynamic impact to composite panels. First, the existing orthotropic visco-elastic visco-plastic material model in LS-DYNA commercial finite element program, MAT_COMPOSITE_TABULATED_PLASTICITY_DAMAGE (MAT_213) is enhanced with the development and use of rate-dependent damage data. Using state-of-the-art experimental procedures, impact validation test data are created. Second, the versatility of the constitutive model is examined by its ability to simulate a low velocity (27 ft/s, 8 m/s) impact test, and two high velocity impact tests (385–417 ft/s, 117–127 m/s) that straddle the ballistic limit of the projectile and the composite panels. Third, these impact events are modeled using three different finite element types – under-integrated and fully-integrated solid elements, and fully-integrated thin shell elements. Insight into accuracy versus numerical efficiency is gained by comparing the performance of solid versus thin shell elements. The numerical results show that accurate predictions of these impact events can be obtained with reasonable effort and care.

Simulations↗
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 NTRS1974

Explosive propulsion applications

The feasibility and application of an explosive propulsion concept capable of supporting future unmanned missions in the post-1980 era were examined and recommendations made for advanced technology development tasks. The Venus large lander mission was selected as the first in which the explosive propulsion concept can find application. A conceptual design was generated and its performance, weight, costs, and interaction effects determined. Comparisons were made with conventional propulsion alternatives. The feasibility of the explosive propulsion system was verified for planetology experiments within the dense atmosphere of Venus as well as the outer planets. Additionally, it was determined that the Venus large lander mission could be augmented ballistically with a significant delivery margin.

Nakamura, Y.↗
NASA NTRS2006

Capabilities of NASA/Marshall Space Flight Center's Impact Testing Facility

The Impact Testing Facility (ITF) serves as an important installation for materials science at Marshall Space Flight Center (MSFC). With an array of air, powder, and two-stage light gas guns, a variety of projectile and target types and sizes can be accommodated. The ITF allows for simulation of impactors from rain to micrometeoroids and orbital debris on materials being investigated for space, atmospheric, and ground use. Expendable, relatively simple launch assemblies are used to obtain well-documented results for impact conditions comparable to those from ballistic and rocket sled ranges at considerably lower cost. In addition, for applications requiring study of impacts at speeds in excess of those attainable by gun launches, hydrocode simulations, validated by test data, can be used to extend the velocity range. In addition to serving various NASA directorates, the ITF has performed testing on behalf of the European and Russian space agencies, as well as the Department of Defense, and academic institutions. The m s contributions not only enable safer space flight for NASA s astronauts, but can help design materials and structures to protect soldiers and civilians on Earth, through advances in body armor, aircraft survivability, and a variety of other applications.

Hovater, Mary↗
NASA NTRSDate not supplied

The NASA JSC Hypervelocity Impact Technology (HVIT) Office

The Hypervelocity Impact Technology (HVIT) office at the Johnson Space Center (JSC) is dedicated to support NASA flight programs in meeting their meteoroid and orbital debris (MMOD) protection requirements both efficiently and effectively, with a minimum of shielding mass, volume, and cost. This paper provides an overview of HVIT capabilities in providing hypervelocity tests, analyses, MMOD risk assessments and shielding expertise. HVIT personnel work closely with NASA Program personnel to develop program-specific MMOD protection requirements which are achievable based on initial risk assessments. Then HVIT proceeds with hypervelocity impact testing and hydrocode assessments to update and verify ballistic limit equations used in the MMOD risk assessment. The final MMOD risk assessment prior to flight will show compliance with MMOD requirements. This final assessment will also incorporate changes in the MMOD environment from meteoroid showers and orbital debris breakups. After the mission is flown, post-flight inspections are performed by HVIT to document and trend MMOD damage to returned spacecraft surfaces and hardware. A majority of the hypervelocity impact tests are performed at the NASA WSTF Remote Hypervelocity Test Laboratory (RHTL), although other hypervelocity facilities that provide complementary test capabilities are called on when necessary to fill data gaps. HVIT accomplishments in NASA mission support are described. In addition, current HVIT activities to aid on-going NASA crewed missions and science missions are presented.

Hypervelocity↗
NASA NTRS1962

Optimum Interplanetary Rendezvous Trajectories With Powerlimited Vehicles

The optimum-thrust equations for both variable and constant thrust are presented. These thrust programs are used to generate rendezvous trajectories from the Earth to Mars for various flight times and launch dates during the years 1968-71. The manner in which the propulsion requirements vary with flight time and launch date are considered, and a comparison of vehicle performance using the variable- and constant-thrust programs is presented. The optimization of the pro- pulsion system parameters is discussed, and the existence of optimum launch dates is interpreted in terms of certain transversality conditions derivable from the calculus of variations. A brief comparison of the advanced propulsion vehicle and the ballistic vehicle propulsion requirements is made for Earth-Mars rendezvous trajectories. An appendix considering the analytical basis for this work is included.

INTERPLANETARY TRAJECTORY↗
NASA NTRS1989

Space Shuttle propulsion performance reconstruction from flight data

The aplication of extended Kalman filtering to estimating Space Shuttle Solid Rocket Booster (SRB) performance, specific impulse, from flight data in a post-flight processing computer program. The flight data used includes inertial platform acceleration, SRB head pressure, and ground based radar tracking data. The key feature in this application is the model used for the SRBs, which represents a reference quasi-static internal ballistics model normalized to the propellant burn depth. Dynamic states of mass overboard and propellant burn depth are included in the filter model to account for real-time deviations from the reference model used. Aerodynamic, plume, wind and main engine uncertainties are included.

Rogers, Robert M.↗
NASA NTRS1990

Preliminary subsystem designs for the Assured Crew Return Vehicle (ACRV), volume 1

A series of design studies is presented concerning the Assured Crew Return Vehicle (ACRV) for Space Station Freedom. Four alternate designs are presented for the ACRV braking and landing system. Options presented include: ballistic and lifting body reentries; the use of high-lift, high-payload aerodynamic decelerators, as well as conventional parachutes; landing systems designed for water landings, land landings, or both; and an aerial recovery system. All four design options presented combine some or all of the above attributes, and all meet performance requirements established by the ACRV Program Office. Two studies of ACRV growth options are also presented. Use of the ACRV or a similarly designed vehicle in several roles for possible future space missions is discussed, along with the required changes to a basic ACRV to allow it to perform these missions optimally. The outcome of these studies is a set of recommendations to the ACRV Program Office describing the vehicle characteristics of the basic ACRV which lend themselves most readily to be adapted for use in other missions. Finally, the impacts on the design of the ACRV due to its role as a medical emergency vehicle were studied and are presented. The use of the ACRV in this manner will impact its shape, internal configuration, and equipment.

Source record↗