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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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Collects space systems cost and related data (flight rate, payload, etc.) over time. Gathers only public data. Non-recurring and recurring. Minimal data processing. Graph, visualize, add context. Focus on US space systems competitiveness. Keep fresh update as data arises, launches occur, etc. Keep fresh focus on recent data, indicative of the future.
Combining computer vision techniques, high-speed cameras, pressure-sensitive paint and a transonic wind tunnel, NASA studies unsteady aerodynamic forces on the Space Launch System rocket with unprecedented temporal and spatial resolution. NASA’s most powerful supercomputer, Pleiades, enables parallel processing and real-time visualization to investigate buffet forces and aeroacoustic physics.
The NASA Kepler and K2 Missions have recently revealed a population of transiting giant planets orbiting moderately evolved, low-luminosity red giant branch stars. Here, we present radial velocity (RV) measurements of three of these systems, revealing significantly non-zero orbital eccentricities in each case. Comparing these systems with the known planet population suggests that close-in giant planets around evolved stars tend to have more eccentric orbits than those around main sequence stars. We interpret this as tentative evidence that the orbits of these planets pass through a transient, moderately eccentric phase where they shrink faster than they circularize due to tides raised on evolved host stars. Additional RV measurements of currently known systems, along with new systems discovered by the recently launched NASA Transiting Exoplanet Survey Satellite (TESS) mission, may constrain the timescale and mass dependence of this process.
In this presentation we evaluate ozone profile retrievals from Ozone Mapping and Profiling Suite Limb Profiler (OMPS-LP) launched in October 2011 on Suomi NPP satellite. The data from OMPS-LP were re-processed with the new retrieval algorithm that combines measurements from UV and VIS parts of the spectra to retrieve a single ozone profile between 12.5 km (or cloud tops) and 57.5 km. To improve ozone retrievals in the Upper Troposphere Lower Stratosphere (UTLS) region, a new aerosol correction with a gamma-function particle size distribution was implemented. The OMPS-LP altitude registration was also adjusted in the new version to ensure the stability of the ozone record. The continuous ozone profile observations with Suomi NPP OMPS LP exceeds 10 years. We will compare OMPS-LP ozone profiles with coincident stratospheric ozone measurements from NASA Aura MLS, ISS SAGE III, and OMPS NP and with the observations in UTLS and middle stratosphere from a ground-based network of ozone sonde stations to evaluate accuracy and precision of OMPS-LP retrievals and overall stability of the record. We will also investigate the quality of stratospheric ozone columns obtained from OMPS-LP and estimate the contribution of various errors to integrated stratospheric columns from OMPS-LP.
Mission integration is an iterative process by which a specific mission is formulated, refined, planned, and executed within the established vehicle(s), architecture, and ground systems design. Mission integration includes the people, vehicle(s) and ground hardware/software, products, processes, analyses, schedules, facilities, Certification of Flight Readiness, etc. The Artemis Mission Integration Task Team (MITT) developed a series of products and processes to support the complex mission integration across various Programs within the Artemis Mission Campaign (Orion, Space Launch Systems, Exploration Ground Systems, Gateway, Human Landing System, and Extravehicular Activity and Human Surface Mobility). The Moon to Mars (M2M) Program is referred to as ‘the enterprise’ as it includes both the M2M organization and the Programs supporting the Artemis Mission Campaign. Artemis Mission Integration has five phases: mission capability, mission definition, mission preparation, mission execution, and post-mission assessment. This paper focuses on one of the enterprise-level mission checkpoints as a kick-off to the Mission Preparation phase, the Mission Integration Review (MIR), which occurs 18-24 months prior to launch. The MIR helps to confirm the defined mission technical baseline is within the existing analyzed design envelope. Details are provided on the identification of dependencies, issues, or gaps for mission-specific objectives and requirements, as well as the definition of the analysis, training, mission execution products, facilities, and detailed supporting operations requirements. The MIR was held for both Artemis I and II and this paper aims to share with the aerospace community its value as we prepare for upcoming Artemis Missions.
Deterministic damage tolerance guidelines for highly reusable and efficiently designed launch systems can be challenging to meet and result in overly conservative assessments for tightly controlled manufacturing processes. Deterministic approaches may be unconservative when the structure is workmanship sensitive and there is a wide spread in fracture properties. Reliability-based damage tolerance assessments targeting a component reliability over the service life commensurate to mission risk posture is a promising alternative. Low production rates, short fleets, and lack of standards or guidance has slowed down widespread adoption of this method. Guidance for the robust application of reliability-based damage tolerance for space systems are proposed. These guidelines cover the treatment of uncertainty in damage tolerance analysis, the collection of data to develop probabilistic distributions, uncertainty propagation methods, and types of hardware.
A system, method, and computer program product for warming a cache for a task launch is described. The method includes the steps of receiving a task data structure that defines a processing task, extracting information stored in a cache warming field of the task data structure, and, prior to executing the processing task, generating a cache warming instruction that is configured to load one or more entries of a cache storage with data fetched from a memory.
The National Aeronautics and Space Administration (NASA) recognizes the tremendous potential that CubeSats (very small satellites) have to inexpensively demonstrate advanced technologies, collect scientific data, and enhance student engagement in Science, Technology, Engineering, and Mathematics (STEM). The CubeSat Launch Initiative (CSLI) was created to provide launch opportunities for CubeSats developed by academic institutions, non-profit entities, and NASA centers. This presentation will provide an overview of the CSLI, its benefits, and its results. This presentation will also provide high level CubeSat 101 information for prospective CubeSat developers, describing the development process from concept through mission operations while highlighting key points that developers need to be mindful of.
The Consort 3 is a commercial suborbital rocket that carried 12 microgravity experiments. It was launched on a Starfire rocket on May 16, 1990, from the Naval Ordnance Missile Test Station facilities at the U.S. Army's White Sands Missile Range (WSMR), NM. The videotape opens with approximately 2 minutes of a man speaking into a microphone but there is no sound. This is followed by a brief summary of the payload, and the expected trajectory, a view of the launch vehicle, the countdown and the launch. The videotape then shows a film clip from the University of Alabama, with Dr. Francis Wessling, project manager for the Consort 3 project, speaking about the mission goals in the materials sciences experimentation. The video shows footage of the payload being assembled. The next section is a discussion by Dr. Roy Hammustedt, of Pennsylvania State University, who reviews the Penn State Bio Module,and the goal of learning about the effects of gravity on physiology. This is followed by George Maybee, from McDonald Douglas, who spoke about the payload integration process while the video shows some of the construction. The last section of the videotape shows a press conference at the launch site. Ana Villamil answers questions from the press about the flight.
The Columbia Accident Investigation Board (CAIB) was clear in its assessment of the loss of the Space Shuttle Columbia on February 3, 2003. Foam liberated from the External Tank (ET) impacting the brittle wing leading edge (WLE) of the orbiter causing the vehicle to disintegrate upon re-entry. Naturally, the CAB pointed out numerous issues affecting this exact outcome in hopes of correcting systems of systems failures any one of which might have altered the outcome. However, Discovery s recent return to flight (RTF) illustrates the primacy of erosion of foam and the risk of future undesirable outcomes. It is obvious that the original RTF focused approach to this problem was not equal to a comprehensive foam debris reduction activity consistent with the high national value of the Space Shuttle assets. The root cause is really very simple when looking at the spray-on foam insulation for the entire ET as part of the structure (e.g., actual stresses > materials allowable) rather than as some sort of sizehime limited ablator. This step is paramount to accepting the CAB recommendation of eliminating debris or in meeting any level of requirements due to the fundamental processes ensuring structural materials maintain their integrity. Significant effort has been expended to identify root cause of the foam debris In-Flight Anomaly (FA) of STS-114. Absent verifiable location specific data pre-launch (T-0) and in-flight, only a most probable cause can be identified. Indeed, the literature researched corroborates NASNTM-2004-2 13238 disturbing description of ill defined materials characterization, variable supplier constituents and foam processing irregularities. Also, foam is sensitive to age and the exposed environment making baseline comparisons difficult without event driven data. Conventional engineering processes account for such naturally occurring variability by always maintaining positive margins. Success in a negative margin range is not consistently achieved. Looking at the ET S spray-on foam insulation as part of the structural system (e.g., glass half full mentality) will create an environment where ET debris levels as low as reasonably achievable (ALARA) can be realized. ALARA is a NASA requirements philosophy deployed for the complex, mission altering radiation exposure requirements for life safety of astronauts. In the Shuttle s case, reasonableness is established by exhaustive engineering rigor, allowable debris size/quantity, technology maturity and programmatic constraints. A more robust urethane foam thermal protection system (TPS) will enhance the hctionality of the new Ares I Crew Launch Vehicle (CLV) Upper Stage. This paper will outline the strategy for a comprehensive effort to reduce ET foam debris and outline steps leading to an improved foam TPS. The NASA must remain committed to such an approach no matter what becomes of the next flight s actual debris field lest we fall back into a false sense of security. This commitment along with full implementation of all the other CAB recommendations such as orbiter hardening will significantly improve the Shuttle system, the engineering workforce, future capabilities & alternate policy offramps, national human resource protection, high value national asset protection and increase the level of service to the overall NASA mission.
A launch vehicle ground-wind-loads program was conducted at the NASA Langley Transonic Dynamics Tunnel. The objectives were to quantify key aerodynamic and structural characteristics that impact the occurrence of large wind-induced oscillations of a launch vehicle when exposed to ground winds prior to launch. Of particular interest is the dynamic response of a launch vehicle when a von Kármán vortex street forms in the wake of the vehicle resulting in quasiperiodic lift and drag forces.Vehicle response to these quasiperiodic forces can become quite large when the frequency of vortex shedding nears that of a lowly-damped structural mode thereby exciting a resonant response. The study of ground-wind-loads presents unique challenges to quantify significant characteristics of the approaching wind and to relate the wind-tunnel acquired static, dynamic, and gravitational loads to full-scale vehicles. This paper will explain the correlation process between model-scale and full-scale wind characteristics and resulting structural loads. Characterization of the wind at a launch site requires knowledge of the vehicle dynamics and the anemometer performance. Vehicle dynamics dictate the frequency range of interest in characterization of the atmospheric turbulence, and in defining the required data window-length to model dynamic oscillation build-up during wind-gust analysis. Knowledge of the anemometer performance is required to infer peak wind magnitudes since most anemometers cannot directly measure instantaneous peak values. Following proper wind-speed correlation, static and dynamic wind-tunnel loads must be converted to full-scale equivalent values to derive design and operational guidance. In addition to the direct computation of full-scale equivalent loads using relevant scaling parameters, the corrections applied to dynamic loads accounting for differences in structural damping will also be discussed. Furthermore, a phenomenon that manifests in many slender and flexible launch vehicles is a load magnification due to gravity resulting from mass offset under deflection. A method for scaling this contribution from model-scale to full-scale is presented. Finally, a comparison of wind-tunnel derived loads using the described methodologies to launch vehicle measured loads is presented.
A launch vehicle ground-wind-loads program was conducted at the NASA Langley Transonic Dynamics Tunnel. The objectives were to quantify key aerodynamic and structural characteristics that impact the occurrence of large wind-induced oscillations of a launch vehicle when exposed to ground winds prior to launch. Of particular interest is the dynamic response of a launch vehicle when a von Kármán vortex street forms in the wake of the vehicle resulting in quasiperiodic lift and drag forces. Vehicle response to these quasiperiodic forces can become quite large when the frequency of vortex shedding nears that of a lowly-damped structural mode thereby exciting a resonant response. The study of ground-wind-loads presents unique challenges to quantify significant characteristics of the approaching wind and to relate the wind-tunnel acquired static, dynamic, and gravitational loads to full-scale vehicles. This paper will explain the correlation process between model-scale and full-scale wind characteristics and resulting structural loads. Characterization of the wind at a launch site requires knowledge of the vehicle dynamics and the anemometer performance. Vehicle dynamics dictate the frequency range of interest in characterization of the atmospheric turbulence, and in defining the required data window-length to model dynamic oscillation build-up during wind-gust analysis. Knowledge of the anemometer performance is required to infer peak wind magnitudes since most anemometers cannot directly measure instantaneous peak values. Following proper wind-speed correlation, static and dynamic wind-tunnel loads must be converted to full-scale equivalent values to derive design and operational guidance. In addition to the direct computation of full-scale equivalent loads using relevant scaling parameters, the corrections applied to dynamic loads accounting for differences in structural damping will also be discussed. Furthermore, a phenomenon that manifests in many slender and flexible launch vehicles is a load magnification due to gravity resulting from mass offset under deflection. A method for scaling this contribution from model-scale to full-scale is presented. Finally, a comparison of wind-tunnel derived loads using the described methodologies to launch vehicle measured loads is presented.
The method and terms used to design the target parameters that will be used to insert the Orbiting Carbon Observatory 2 (OCO-2) into the A-train are defined. Comparisons are made to the plan that would have been used had OCO-1 successfully achieved its Injection Orbit. Major differences arise from the fact that OCO-1 was launched on a Taurus XL with a target 65 km below the A-Train while OCO-2 will be launched on a Delta II with a target only 15 km below the A-Train. The new plan is similar to the one used for the CloudSat/CALIPSO missions, but reformulated into a one-step iteration process that is easier to understand.
NASA-Goddard is responsible for the development of a ground system for the Upper Atmosphere Research Satellite (UARS) observatory, whose launch is scheduled for 1991. This ground system encompasses a dedicated Central Data Handling Facility (CDHF); attention is presently given to the management of software systems design and implementation phases for CDHF by the UARS organization. Also noted are integration and testing activities performed following software deliveries to the CDHF. The UARS project has an obvious requirement for a powerful and flexible data base management system; an off-the-shelf commercial system has been incorporated.
The role of simulations in the design of the HL-20, the crew-carrying unpowered lifting-body component of the NASA Personnel Launch System, is reviewed and illustrated with drawings and diagrams. Detailed consideration is given to the overall implementation of a real-time simulation of the HL-20 approach and landing phase, the baseline and experimental control laws used in the flight-control system, autoland guidance and control laws (vertical and lateral steering), the control-surface mixer and actuator model, and simulation results. The simulations allowed identification and correction of design problems with respect to the position of the landing gear and the original maximum L/D ratio of 3.2.
The Visible Imaging System (VIS) on Polar spacecraft of the NASA Goddard Space Flight Center was launched into orbit about Earth on February 24, 1996. Since shortly after launch, the Earth Camera subsystem of the VIS has been operated nearly continuously to acquire far ultraviolet, global images of Earth and its northern and southern auroral ovals. The only exceptions to this continuous imaging occurred for approximately 10 days at the times of the Polar spacecraft re-orientation maneuvers in October, 1996 and April, 1997. Since launch, approximately 525,000 images have been acquired with the VIS Earth Camera. The VIS instrument operational health continues to be excellent. Since launch, all systems have operated nominally with all voltages, currents, and temperatures remaining at nominal values. In addition, the sensitivity of the Earth Camera to ultraviolet light has remained constant throughout the operation period. Revised flight software was uploaded to the VIS in order to compensate for the spacecraft wobble. This is accomplished by electronic shuttering of the sensor in synchronization with the 6-second period of the wobble, thus recovering the original spatial resolution obtainable with the VIS Earth Camera. In addition, software patches were uploaded to make the VIS immune to signal dropouts that occur in the sliprings of the despun platform mechanism. These changes have worked very well. The VIS and in particular the VIS Earth Camera is fully operational and will continue to acquire global auroral images as the sun progresses toward solar maximum conditions after the turn of the century.
The Autonomous Sciencecraft Experiment (ASE) will fly onboard the Air Force TechSat 21 constellation of three spacecraft scheduled for launch in 2006. ASE uses onboard continuous planning, robust task and goal-based execution, model-based mode identification and reconfiguration, and onboard machine learning and pattem recognition to radically increase science retum by enabling intelligent downlink selection and autonomous retargeting. Demonstration of these capabilities in a flight environment will open up tremendous new opportunities in planetary science, space physics, and earth science that would be unreachable without this technology.
The Autonomous Sciencecraft Experiment (ASE) will fly onboard the Air Force TechSat 21 constellation of three spacecraft scheduled for launch in 2006. ASE uses onboard continuous planning, robust task and goal-based execution, model-based mode identification and reconfiguration, and onboard machine learning and pattern recognition to radically increase science return by enabling intelligent downlink selection and autonomous retargeting. In this paper we discuss how these AI technologies are synergistically integrated in a hybrid multi-layer control architecture to enable a virtual spacecruft science agent. Demonstration of these capabilities in a flight environment will open up tremendous new opportunities in planetary science, space physics, and earth science that would be unreachable without this technology.