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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Supraoptimal carbon dioxide effects on growth of soybean [Glycine max (L.) Merr.]

In tightly closed environments used for human life support in space, carbon dioxide (CO2) partial pressures can reach 500 to 1000 Pa, which may be supraoptimal or toxic to plants used for life support. To study this, soybeans [Glycine max (L.) Merr. cvs. McCall and Pixie] were grown for 90 days at 50, 100, 200, and 500 Pa partial pressure CO2 (500, 1000, 2000, and 5000 ppm). Plants were grown using recirculating nutrient film technique with a 12-h photoperiod, a 26 degrees C/20 degrees C thermoperiod, and approximately 300 micromoles m-2 s-1 photosynthetic photon flux (PPF). Seed yield and total biomass were greatest at 100 Pa for cv. McCall, suggesting that higher CO2 levels were supraoptimal. Seed yield and total biomass for cv. Pixie showed little difference between CO2 treatments. Average stomatal conductance of upper canopy leaves at 50 Pa CO2 approximately 500 Pa > 200 Pa > 100 Pa. Total water use over 90 d for both cultivars (combined on one recirculating system) equalled 822 kg water for 100 Pa CO2, 845 kg for 50 Pa, 879 kg for 200 Pa, and 1194 kg for 500 Pa. Water use efficiences for both cultivars combined equalled 3.03 (g biomass kg-1 water) for 100 Pa CO2, 2.54 g kg-1 for 200 Pa, 2.42 g kg-1 for 50 Pa, and 1.91 g kg-1 for 500 Pa. The increased stomatal conductance and stand water use at the highest CO2 level (500 Pa) were unexpected and pose interesting considerations for managing plants in a tightly closed system where CO2 concentrations may reach high levels.

NASA Discipline Life Support Systems↗
NASA NTRS2005

Mars Aeronomy Explorer (MAX): Study Employing Distributed Micro-Spacecraft

An overview of a Mars Aeronomy Explorer (MAX) mission design study performed at NASA's Jet Propulsion Laboratory is presented herein. The mission design consists of ten micro-spacecraft orbiters launched on a Delta IV to Mars polar orbit to determine the spatial, diurnal and seasonal variation of the constituents of the Martian upper atmosphere and ionosphere over the course of one Martian year. The spacecraft are designed to allow penetration of the upper atmosphere to at least 90 km. This property coupled with orbit precession will yield knowledge of the nature of the solar wind interaction with Mars, the influence of the Mars crustal magnetic field on ionospheric processes, and the measurement of present thermal and nonthermal escape rates of atmospheric constituents. The mission design incorporates alternative design paradigms that are more appropriate for-and in some cases motivate-distributed micro-spacecraft. These design paradigms are not defined by a simple set of rules, but rather a way of thinking about the function of instruments, mission reliability/risk, and cost in a systemic framework.

Mars↗
NASA NTRS2010

The Proposed Mars Astrobiology Explorer - Cacher [MAX-C] Rover: First Step in a Potential Sample Return Campaign

Sample return from Mars has been advocated by numerous scientific advisory panels for over 30 years, most prominently beginning with the National Research Council s [1] strategy for the exploration of the inner solar system, and most recently by the Mars Exploration Program Analysis Group (MEPAG s) Next Decade Science Analysis Group [2]. Analysis of samples here on Earth would have enormous advantages over in situ analyses in producing the data quality needed to address many of the complex scientific questions the community has posed about Mars. Instead of a small, predetermined set of analytical techniques, state of the art preparative and instrumental resources of the entire scientific community could be applied to the samples. The analytical emphasis could shift as the meaning of each result becomes better appreciated. These arguments apply both to igneous rocks and to layered sedimentary materials, either of which could contain water and other volatile constituents. In 2009 MEPAG formed the Mid-Range Rover Science Analysis Group (MRR-SAG) to formulate a mission concept that would address two general objectives: (1) conduct high-priority in situ science and (2) make concrete steps towards the potential return of samples to Earth. This analysis resulted in a mission concept named the Mars Astrobiology Explorer-Cacher (MAX-C), which was envisioned for launch in the 2018 opportunity. After extensive discussion, this group concluded that by far the most definitive contribution to sample return by this mission would be to collect and cache, in an accessible location, a suite of compelling samples that could potentially be recovered and returned by a subsequent mission. This would have the effect of separating two of the essential functions of MSR, the acquisition of the sample collection and its delivery to martian orbit, into two missions.

Allen, Carlton C.↗
NASA NTRS2010

Design and Analysis of Outer Mold Line Close-outs for the Max Launch Abort System (MLAS) Flight Experiment

In 2007, the NASA Exploration Systems Mission Directorate (ESMD) chartered the NASA Engineering Safety Center (NESC) to demonstrate an alternate launch abort concept as risk mitigation for the Orion project's baseline "tower" design. On July 8, 2009, a full scale, passive aerodynamically stabilized Max Launch Abort System (MLAS) pad abort demonstrator was successfully launched from NASA Goddard Space Flight Center's Wallops Flight Facility. Aerodynamic close-outs were required to cover openings on the MLAS fairing to prevent aerodynamic flow-through and to maintain the MLAS OML surface shape. Two-ply duct tape covers were designed to meet these needs. The duct tape used was a high strength fiber reinforced duct tape with a rubberized adhesive that demonstrated 4.6 lb/in adhesion strength to the unpainted fiberglass fairing. Adhesion strength was observed to increase as a function of time. The covers were analyzed and experimentally tested to demonstrate their ability to maintain integrity under anticipated vehicle ascent pressure loads and to not impede firing of the drogue chute mortars. Testing included vacuum testing and a mortar fire test. Tape covers were layed-up on thin Teflon sheets to facilitate installation on the vehicle. Custom cut foam insulation board was used to fill mortar hole and separation joint cavities and provide support to the applied tape covers. Flight test results showed that the tape covers remained adhered during flight.

Woods-Vedeler, Jessica A.↗
NASA NTRS2010

GN and C Design Overview and Flight Test Results from NASA's Max Launch Abort System (MLAS)

The National Aeronautics and Space Administration (NASA) Engineering and Safety Center (NESC) designed, developed and flew the alternative Max Launch Abort System (MLAS) as risk mitigation for the baseline Orion spacecraft launch abort system (LAS) already in development. The NESC was tasked with both formulating a conceptual objective system (OS) design of this alternative MLAS as well as demonstrating this concept with a simulated pad abort flight test. The goal was to obtain sufficient flight test data to assess performance, validate models/tools, and to reduce the design and development risks for a MLAS OS. Less than 2 years after Project start the MLAS simulated pad abort flight test was successfully conducted from Wallops Island on July 8, 2009. The entire flight test duration was 88 seconds during which time multiple staging events were performed and nine separate critically timed parachute deployments occurred as scheduled. Overall, the as-flown flight performance was as predicted prior to launch. This paper provides an overview of the guidance navigation and control (GN&C) technical approaches employed on this rapid prototyping activity. This paper describes the methodology used to design the MLAS flight test vehicle (FTV). Lessons that were learned during this rapid prototyping project are also summarized.

Dennehy, Cornelius J.↗
NASA NTRS2010

Guidance, Navigation and Control (GN and C) Design Overview and Flight Test Results from NASA's Max Launch Abort System (MLAS)

The National Aeronautics and Space Administration Engineering and Safety Center designed, developed and flew the alternative Max Launch Abort System (MLAS) as risk mitigation for the baseline Orion spacecraft launch abort system already in development. The NESC was tasked with both formulating a conceptual objective system design of this alternative MLAS as well as demonstrating this concept with a simulated pad abort flight test. Less than 2 years after Project start the MLAS simulated pad abort flight test was successfully conducted from Wallops Island on July 8, 2009. The entire flight test duration was 88 seconds during which time multiple staging events were performed and nine separate critically timed parachute deployments occurred as scheduled. This paper provides an overview of the guidance navigation and control technical approaches employed on this rapid prototyping activity; describes the methodology used to design the MLAS flight test vehicle; and lessons that were learned during this rapid prototyping project are also summarized.

Dennehy, Cornelius J.↗
NASA NTRS2011

Testing Strategies and Methodologies for the Max Launch Abort System

The National Aeronautics and Space Administration (NASA) Engineering and Safety Center (NESC) was tasked to develop an alternate, tower-less launch abort system (LAS) as risk mitigation for the Orion Project. The successful pad abort flight demonstration test in July 2009 of the "Max" launch abort system (MLAS) provided data critical to the design of future LASs, while demonstrating the Agency s ability to rapidly design, build and fly full-scale hardware at minimal cost in a "virtual" work environment. Limited funding and an aggressive schedule presented a challenge for testing of the complex MLAS system. The successful pad abort flight demonstration test was attributed to the project s systems engineering and integration process, which included: a concise definition of, and an adherence to, flight test objectives; a solid operational concept; well defined performance requirements, and a test program tailored to reducing the highest flight test risks. The testing ranged from wind tunnel validation of computational fluid dynamic simulations to component ground tests of the highest risk subsystems. This paper provides an overview of the testing/risk management approach and methodologies used to understand and reduce the areas of highest risk - resulting in a successful flight demonstration test.

Schaible, Dawn M.↗
NASA NTRS2011

Modeling, Analysis and Simulation Approaches Used in Development of the National Aeronautics and Space Administration Max Launch Abort System

The National Aeronautics and Space Administration (NASA) Engineering and Safety Center was chartered to develop an alternate launch abort system (LAS) as risk mitigation for the Orion Project. Its successful flight test provided data for the design of future LAS vehicles. Design of the flight test vehicle (FTV) and pad abort trajectory relied heavily on modeling and simulation including computational fluid dynamics for vehicle aero modeling, 6-degree-of-freedom kinematics models for flight trajectory modeling, and 3-degree-of-freedom kinematics models for parachute force modeling. This paper highlights the simulation techniques and the interaction between the aerodynamics, flight mechanics, and aerodynamic decelerator disciplines during development of the Max Launch Abort System FTV.

Yuchnovicz, Daniel E.↗
NASA NTRS2011

Max Launch Abort System (MLAS) Landing Parachute Demonstrator (LPD) Drop Test

The Landing Parachute Demonstrator (LPD) was conceived as a low-cost, rapidly-developed means of providing soft landing for the Max Launch Abort System (MLAS) crew module (CM). Its experimental main parachute cluster deployment technique and off-the-shelf hardware necessitated a full-scale drop test prior to the MLAS mission in order to reduce overall mission risk. This test was successfully conducted at Wallops Flight Facility on March 6, 2009, with all vehicle and parachute systems functioning as planned. The results of the drop test successfully qualified the LPD system for the MLAS flight test. This document captures the design, concept of operations and results of the drop test.

Shreves, Christopher M.↗
NASA NTRS2011

Max Launch Abort System (MLAS) Pad Abort Test Vehicle (PATV) II Attitude Control System (ACS) Integration and Pressurization Subsystem Dynamic Random Vibration Analysis

In order to mitigate catastrophic failures on future generation space vehicles, engineers at the National Aeronautics and Space Administration have begun to integrate a novel crew abort systems that could pull a crew module away in case of an emergency at the launch pad or during ascent. The Max Launch Abort System (MLAS) is a recent test vehicle that was designed as an alternative to the baseline Orion Launch Abort System (LAS) to demonstrate the performance of a "tower-less" LAS configuration under abort conditions. The MLAS II test vehicle will execute a propulsive coast stabilization maneuver during abort to control the vehicles trajectory and thrust. To accomplish this, the spacecraft will integrate an Attitude Control System (ACS) with eight hypergolic monomethyl hydrazine liquid propulsion engines that are capable of operating in a quick pulsing mode. Two main elements of the ACS include a propellant distribution subsystem and a pressurization subsystem to regulate the flow of pressurized gas to the propellant tanks and the engines. The CAD assembly of the Attitude Control System (ACS) was configured and integrated into the Launch Abort Vehicle (LAV) design. A dynamic random vibration analysis was conducted on the Main Propulsion System (MPS) helium pressurization panels to assess the response of the panel and its components under increased gravitational acceleration loads during flight. The results indicated that the panels fundamental and natural frequencies were farther from the maximum Acceleration Spectral Density (ASD) vibrations which were in the range of 150-300 Hz. These values will direct how the components will be packaged in the vehicle to reduce the effects high gravitational loads.

Ekrami, Yasamin↗
NASA NTRS2013

Flight Performance Feasibility Studies for the Max Launch Abort System

In 2007, the NASA Engineering and Safety Center (NESC) initiated the Max Launch Abort System Project to explore crew escape system concepts designed to be fully encapsulated within an aerodynamic fairing and smoothly integrated onto a launch vehicle. One objective of this design was to develop a more compact launch escape vehicle that eliminated the need for an escape tower, as was used in the Mercury and Apollo escape systems and what is planned for the Orion Multi-Purpose Crew Vehicle (MPCV). The benefits for the launch vehicle of eliminating a tower from the escape vehicle design include lower structural weights, reduced bending moments during atmospheric flight, and a decrease in induced aero-acoustic loads. This paper discusses the development of encapsulated, towerless launch escape vehicle concepts, especially as it pertains to the flight performance and systems analysis trade studies conducted to establish mission feasibility and assess system-level performance. Two different towerless escape vehicle designs are discussed in depth: one with allpropulsive control using liquid attitude control thrusters, and a second employing deployable aft swept grid fins to provide passive stability during coast. Simulation results are presented for a range of nominal and off-nominal escape conditions.

Tarabini, Paul V.↗
NASA NTRS2014

The Max Launch Abort System - Concept, Flight Test, and Evolution

The NASA Engineering and Safety Center (NESC) is an independent engineering analysis and test organization providing support across the range of NASA programs. In 2007 NASA was developing the launch escape system for the Orion spacecraft that was evolved from the traditional tower-configuration escape systems used for the historic Mercury and Apollo spacecraft. The NESC was tasked, as a programmatic risk-reduction effort to develop and flight test an alternative to the Orion baseline escape system concept. This project became known as the Max Launch Abort System (MLAS), named in honor of Maxime Faget, the developer of the original Mercury escape system. Over the course of approximately two years the NESC performed conceptual and tradeoff analyses, designed and built full-scale flight test hardware, and conducted a flight test demonstration in July 2009. Since the flight test, the NESC has continued to further develop and refine the MLAS concept.

Gilbert, Michael G.↗
NASA NTRS2016

A Proposed Ascent Abort Flight Test for the Max Launch Abort System

The NASA Engineering and Safety Center initiated the Max Launch Abort System (MLAS) Project to investigate alternate crew escape system concepts that eliminate the conventional launch escape tower by integrating the escape system into an aerodynamic fairing that fully encapsulates the crew capsule and smoothly integrates with the launch vehicle. This paper proposes an ascent abort flight test for an all-propulsive towerless escape system concept that is actively controlled and sized to accommodate the Orion Crew Module. The goal of the flight test is to demonstrate a high dynamic pressure escape and to characterize jet interaction effects during operation of the attitude control thrusters at transonic and supersonic conditions. The flight-test vehicle is delivered to the required test conditions by a booster configuration selected to meet cost, manufacturability, and operability objectives. Data return is augmented through judicious design of the boost trajectory, which is optimized to obtain data at a range of relevant points, rather than just a single flight condition. Secondary flight objectives are included after the escape to obtain aerodynamic damping data for the crew module and to perform a high-altitude contingency deployment of the drogue parachutes. Both 3- and 6-degree-of-freedom trajectory simulation results are presented that establish concept feasibility, and a Monte Carlo uncertainty assessment is performed to provide confidence that test objectives can be met.

Tartabini, Paul V.↗
NASA NTRS2017

Unusual Oxidative Limitations for Al-MAX Phases

Alumina-forming MAX phases are well-known for their excellent oxidation resistance, rivaling many metallic NiAl, NiCrAl, and FeCrAl counterparts and with upper temperature capability possible to approximately1400C. However a number of limitations have been emerging that need to be acknowledged to permit robust performance in demanding applications. Ti2AlC and Ti3AlC2 possess excellent scale adhesion, cyclic oxidation/moisture/volatility resistance, and TBC compatibility. However they are very sensitive to Al content and flux in order to maintain an exclusive Al2O3 scale without runaway oxidation of ubiquitous TiO2 transient scales. Accelerated oxidation has been shown to occur for Al-depleted, damaged, or roughened surfaces at temperatures less than 1200C. Conversely, Cr2AlC is less sensitive to transients, but exhibits volatile losses at 1200C or above if common Cr7C3 impurity phases are present. Poor scale adhesion is exhibited after oxidation at 1150C or above, where spallation occurs at the Cr7C3 (depletion zone) interface. Delayed spallation is significant and suggests a moisture-induced phenomenon similar to non-adherent metallic systems. Re-oxidation of this surface does not reproduce the initial pure Al2O3 behavior, but initiates a less-protective scale. Cr2AlC has also been shown to have good long term bonding with superalloys at 800C, but exhibits significant Beta-NiAl + Cr7C3 diffusion zones at 1100C and above. This may set limits on Cr2AlC as a high temperature TBC bond coat on Ni-based superalloys, while improving corrosion resistance in lower temperature applications.

oxidation resistance↗
DOE OSTI2023

On the Approximability of Random-Hypergraph MAX-3-XORSAT Problems with Quantum Algorithms

Constraint satisfaction problems are an important area of computer science. Many of these problems are in the complexity class NP which is exponentially hard for all known methods, both for worst cases and often typical. Fundamentally, the lack of any guided local minimum escape method ensures the hardness of both exact and approximate optimization classically, but the intuitive mechanism for approximation hardness in quantum algorithms based on Hamiltonian time evolution is poorly understood. We explore this question using the prototypically hard MAX-3-XORSAT problem class. We conclude that the mechanisms for quantum exact and approximation hardness are fundamentally distinct. We qualitatively identify why traditional methods such as quantum adiabatic optimization are not good approximation algorithms. We propose a new spectral folding optimization method that does not suffer from these issues and study it analytically and numerically. We consider random rank-3 hypergraphs including extremal planted solution instances, where the ground state satisfies an anomalously high fraction of constraints compared to truly random problems. We show that, if we define the energy to be $E = N_{unsat}-N_{sat}$, then spectrally folded quantum optimization will return states with energy $E \leq A E_{GS}$ (where $E_{GS}$ is the ground state energy) in polynomial time, where conservatively, $A \simeq 0.6$. We thoroughly benchmark variations of spectrally folded quantum optimization for random classically approximation-hard (planted solution) instances in simulation, and find performance consistent with this prediction. We do not claim that this approximation guarantee holds for all possible hypergraphs, though our algorithm's mechanism can likely generalize widely. These results suggest that quantum computers are more powerful for approximate optimization than had been previously assumed.

Kapit, Eliot↗
DOE OSTI2025

Thermophysical Properties of Ti3SiC2 MAX Phase Composites with SiC Reinforcement

In the present work, dense (∼100%) Ti 3 SiC 2 composites (TSC) are processed along with 20 vol% of SiC reinforcement (TSC20) via spark plasma sintering at 1400°C, 40 MPa, 15 min, and dynamic vacuum environment. Thermal expansion of both the composites increases from RT to 1273 K and linear fitting of data yields coefficient of thermal expansion (CTE) of 9.4 × 10 −6 K −1 for TSC which decreases to 8.3 × 10 −6 K −1 for TSC20. With increase in temperature from RT to 773 K, specific heat for both TSC and TSC20 composites is observed to increase from 598-850 J.kg −1 .K −1 , whereas thermal diffusivity and thermal conductivity values decrease with testing temperature. SiC reinforcement in Ti 3 SiC 2 resulted in improved thermal diffusivity from 12.7 to 18.7 mm 2 .s −1 and thermal conductivity from ∼57 to ∼79 W.m −1 .K −1 at RT. However, with increase in temperature (773 K), thermal diffusivity and conductivity decrease, and values get closer for both TSC and TSC20 composites. Further extrapolation of thermal conductivity data showed cross-over at ∼973 K due to domination of phonon-phonon scattering and thus lower values of thermal conductivity for TSC20 than TSC. Therefore, reduced CTE and higher thermal conductivity of TSC20 make it a viable choice for applications in high temperatures.

36 MATERIALS SCIENCE↗