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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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Open-loop pulsed transfer functions of simplified thrust vector control system for Saturn-type vehicle with unsampled input derived by several methods
Approaches used in the selection of lunar landing sites are discussed, giving attention to a survey of major constraints affecting site selection. The science objectives of lunar exploration are considered, taking into account exploration goals, the acquisition of required data, mission science planning, orbital science, lunar surface traverse capability, the mare basalts, lunar interior composition, regolith, breccias, lunar chronology, lunar surface processes, and the origin of the moon. Subjects examined in connection with Apollo system capabilities include the Apollo missions, development missions, lunar exploration missions, mission design requirements, navigation, launch vehicle considerations, methods of determining accessibility, lunar module descent considerations, landing site redesignation, and questions of landability.
The paper discusses a concept for a ground-based flight-test facility, a large aeroballistic range capable of launching models large enough to permit the installation of significant amounts of onboard instrumentation. The large model size would also provide thick shock/boundary layers, thus providing the opportunity to measure flow-field properties via advanced offboard diagnostics. Current fabrication technology should permit the construction of a two-stage light-gas gun capable of accelerating 20 to 25 cm diameter models to velocities on the order of 6 km/sec. An electromagnetic launcher or a ram accelerator is considered as a potential means for achieving larger model sizes and/or greater launch velocities. Possible methods for obtaining and recording data are discussed as are the test chamber and model deceleration section requirements.
The conceptual design of a rocket-powered, two-stage fully reusable launch vehicle has been performed as a part of NASA's Advanced Manned Launch System (AMLS) study. This paper summarizes the structural design and loads analysis of this two-stage fully reusable launch vehicle. The method used to determine the structural weights consists of generating a finite-element model for each vehicle, selecting a set of critical loading conditions, determining the loads on the model caused by those conditions, determining the model response and changing the sizes of individual elements to obtain a safe structure. The integrated loads on the two-stage vehicle were obtained from a three-degrees-of-freedom trajectory analysis.
Solar Electric Propulsion (SEP) when used for station keeping and final orbit insertion has been shown to increase a geostationary satellite's payload when launched by existing expendable launch vehicles. In the case of reusable launch vehicles or expendable launch vehicles where an upper stage is an expensive option, this methodology can be modified by using the existing on-board apogee chemical system to perform a perigee burn and then letting the electric propulsion system complete the transfer to geostationary orbit. The elimination of upper stages using on-board chemical and electric propulsion systems was thus examined for GEO spacecraft. Launch vehicle step-down from an Atlas IIAR to a Delta 7920 (no upper stage) was achieved using expanded on-board chemical tanks, 40 kW payload power for electric propulsion, and a 60 day elliptical to GEO SEP orbit insertion. Optimal combined chemical and electric trajectories were found using SEPSPOT. While Hall and ion thrusters provided launch vehicle step-down and even more payload for longer insertion times, NH3 arcjets had insufficient performance to allow launch vehicle step-down. Degradation levels were only 5% to 7% for launch step-down cases using advanced solar arrays. Results were parameterized to allow comparisons for future reusable launch vehicles. Results showed that for an 8 W/kg initial power/launch mass power density spacecraft, 50% to 100% more payload can be launched using this method.
This paper proposes a new method and transportation system to fly into space, to the Moon, Mars, and other planets. This transportation system uses a mechanical energy transfer and requires only minimal energy so that it provides a 'Free Trip' into space. The method uses the rotary and kinetic energy of planets, asteroids, moons, satellites and other natural space bodies. computations for the following projects: 1. Non-Rocket Method for free launch of payload in Space and to other planets. The low cost project will accommodate one hundred thousand tourists annually. 2. Free Trips to the Mars for two thousand annually. 3. Free Trips to the Moon for ten thousand people annually. The projects use artificial materials like nanotubes and whiskers that have a ratio of tensile strength to density equal 4 million meters. In the future, nanotubes will be produced that can reach a specific stress up 100 millions meter and will significantly improve the parameters of suggested projects. The author is prepared to discuss the problems with serious organizations that want to research and develop these inventions.
The present invention is a space launch system and method to propel a payload bearing craft into earth orbit. The invention has two, or preferably, three stages. The upper stage has rocket engines capable of carrying a payload to orbit and provides the capability of releasably attaching to the lower, or preferably, middle stage. Similar to the lower stage, the middle stage is a reusable booster stage that employs all air breathing engines, is recoverable, and can be turned-around in a short time between missions.
Develop Survivability vs Time Model as a decision-evaluation tool to assess various emergency egress methods used at Launch Complex 39B (LC 39B) and in the Vehicle Assembly Building (VAB) on NASAs Kennedy Space Center. For each hazard scenario, develop probability distributions to address statistical uncertainty resulting in survivability plots over time and composite survivability plots encompassing multiple hazard scenarios.
Computational aeroacoustic simulations using the structured overset grid approach and higher-order finite difference methods within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are presented for slat noise predictions. The simulations are part of a collaborative study comparing noise generation mechanisms between a conventional slat and a Krueger leading edge flap. Simulation results are compared with experimental data acquired during an aeroacoustic test in the NASA Langley Quiet Flow Facility. Details of the structured overset grid, numerical discretization, and turbulence model are provided.
Hybrid RANS/LES simulations using the structured overset grid approach and a low dissipation finite-difference method within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are presented for jet noise prediction. The simulations are part of a validation effort to demonstrate jet noise prediction capability, to assess noise characteristics of next generation quiet supersonic commercial jets. Results are compared with experimental data acquired in the Small Hot Jet Acoustic Rig in the Aeroacoustic Propulsion Laboratory at NASA Glenn Research Center. Details of the structured overset grids, numerical discretization, and turbulence model are provided. Near-field comparisons to PIV data and far-field comparisons to microphone data are discussed. Excellent agreement of time-averaged mean quantities along the jet center-line and lip-line are obtained. Good agreement with root-mean-squared (RMS) quantities along the jet center-line are also obtained. An over prediction of lip-line RMS at the nozzle exit is observed leading to a small over-prediction of side-line far-field noise (less than 2 dB). Good agreement with far-field noise in the Mach wave radiation directions is achieved.
Lattice Boltzmann (LB) based Large Eddy Simulation (LES), Reynolds-averaged Navier-Stokes (RANS) as well as hybrid RANS/LES methods within the Launch Ascent and Vehicle Aerodynamics (LAVA) solver framework are applied to NASA's wall-mounted hump. Computational results are compared with experiments performed by Greenblatt et al. A detailed comparison between the accuracy and resolution requirements of the two approaches for turbulence resolving simulations, as well as the suitability of different grid paradigms (body-fitted curvilinear and block structured Cartesian) are presented. This test case is part of NASA's Revolutionary Computational Aerosciences (RCA) sub-project which addresses the technical challenge of predicting flow separation and reattachment accurately. Improvements in predictive accuracy by as much as 90% are demonstrated using LB as well as hybrid RANS/LES approaches compared to state-of-the-art steady state RANS simulations.
Exposed window surfaces on-orbit are vulnerable to a variety of damage including impacts. These impacts impart stress into the panes and reduce the remaining life of the window. Spacecraft windows are evaluated before and after flight to determine the minimum strength of the glass prior to launch. Currently, no method exists to accurately evaluate the windows on-orbit. An on-orbit optical system is under development to evaluate spacecraft windowpane material for remaining flight life by directly measuring the stress in the glass. This optical system builds on the industry-standard practice of calculating the amount of impact-induced stress through birefringence measurement. In industry, the birefringence is imaged through the pane with a detector and light source on opposite sides. On-orbit, there is only access to one side of the pane. This system has been designed to image birefringence by reflecting the light off the pane to a detector on the same side of the glass. Previous studies have shown a correlation between material stress distribution and the associated K-value. This system is designed to determine low strength areas which can be caused by defects due to impact damage. Glass samples will be imaged and broken in order to correlate impact-induced stress to the fracture strength of the material. Once this correlation is determined, a quantitative assessment of the stress areas in spacecraft windows will occur prior to and during mission operations to reliably determine the overall strength of the glass.
This paper presents the updated ablative analysis methods used to determine Artemis 1 launch load environment on the flame deflector for the design based on Artemis I Assessments. The flame trench under Pad 39B at Kennedy Space Center contains a flame deflector to safely divert the exhaust plume from the SLS rocket during launch. During launch of Artemis I the refurbished flame trench and the new flame deflector experienced peak temperatures of over 2,000 degrees Fahrenheit (over 1,000 degrees Celsius) for several seconds. These extreme conditions caused ablation (material removal) from the steel plates. This paper contains flame deflector heat flux values for design based on Artemis I assessment. Post flight images of the flame deflector are shown along with “Pre vs Post” Flame Deflector scan image to measure the post flight deviations. Using COMSOL, a method to calculate a heat flux value due to measured deviations in the deflector plates caused by ablation is shown.
The IUE spacecraft was launched with prime and redundant mechanical Panoramic Attitude Sensors (PAS) to determine coarse spacecraft pointing. Attitude determination typically took at least 24 hours. After launch both systems failed. A new method was developed which required pointing the spacecraft at the antisolar position. After the failure of the 4th IUE gyro, it was no longer possible to point in the antisolar direction. A second method was developed which utilizes IUE's ability to track the sun with a solid state two-dimensional sun sensor. Attitude determination can now be completed in several hours. An hour is required for coarse position measurement and several more hours are needed, using a small 15 arc minute square finder camera, for final attitude confirmation. These methods should be of use for other spacecraft where weight is critical or there is a desire to avoid mechanical devices.
In order to perform post mortem analysis on materials launched in gas guns and powder guns, various methods have been explored to soft catch these materials. The majority of the deformation profile must be from the initial impact during the experiment, not from the subsequent impact with the energy absorption material in the catch tank. This process becomes increasingly more difficult with increased velocities. This report explores a method to soft catch materials launched at velocities 3000 $\frac{m}{s}$.
Ozone observations from ozonesondes, the DIAL and AROTEL lidars aboard the DC-8, in situ ozone measurements from the ER-2 and satellite ozone measurements from POAM were used to assess ozone loss during the SOLVE 1999-2000 campaign. We compare three different methods of computing the ozone loss. The first method simply compares the time sequence of ozonesondes taken at the same station inside the vortex from December through the end of March. In the second method, ozonesondes from a variety of stations are compared using a variant on the Match technique. This method uses short (approx. 5-10 day) forward diabatic trajectories to connect various sonde launches. In the third method, the measurements are simply injected into a diabatic trajectory model and carried forward in time from December 1 to March 16. Over 60,000 individual measurements were used in the last calculation. Again, ozone loss is estimated by comparing vortex interior measurements made early in the campaign with those made later in the campaign. The diabatic nature of the second and third methods calculation presumably corrects for the normal increase in ozone within the vortex due to downward advection. The three methods agree that the largest ozone loss occurs between 400 and 460 K potential temperatures (approx. 16-20 km) with slightly over 1.5 ppmv lost over the winter period. Between 460 K and 500 K (approx. 22 km) net ozone loss is less than 0.8 ppmv. From 500K to 600K (26 km) net loss is less than 0.5 ppmv.
The Environmental Test Facility (ETF), located at NASA-Marshall Space Flight Center, Huntsville, Alabama, has provided thermal vacuum testing for several major programs since the 1960's. The ETF consists of over 13 thermal vacuum chambers sized and configured to handle the majority of test payloads. The majority of tests require a hard vacuum with heating and cryogenics. NASA's Return-to-Flight program requested testing to simulate a launch from the ground to flight using vacuum, heating and cryogenics. This paper describes an effective method for simulating a launch.
Rocket launches generate infrasound signatures that have been detected at great distances. Due to the sparsity of the networks that have made these detections, however, most signals are detected tens of minutes to hours after the rocket launch. In this work, a method of near-real-time detection of rocket launches using data from a network of smartphones located 10–70 km from launch sites is presented. A machine learning model is trained and tested on the open-access Aggregated Smartphone Timeseries of Rocket-generated Acoustics (ASTRA), Smartphone High-explosive Audio Recordings Dataset (SHAReD), and ESC-50 datasets, resulting in a final accuracy of 97% and a false positive rate of <1%. The performance and behavior of the model are summarized, and its suitability for persistent monitoring applications is discussed.