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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 199 records · Page 11
All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

Space Shuttle Crawler Transporter Vibration Analysis in Support of Rollout Fatigue Load Spectra Verification Program

The Crawler Transporter (CT), designed and built for the Apollo Program in the 1960's and surpassing its initial operational life, has become an integral part of the Space Shuttle Program (SSP). The CT transports the Space Shuttle Vehicle (SSV) stack, atop the Mobile Launch Platform (MLP), from the Vehicle Assembly Building (VAB) to the launch pad. This support structure provides hydraulic jacking, leveling and load equalization for the 12 million pound stack on its 3.5-5.0 mile rollout to the launch pad. Major elements of the SSV, consisting of the orbiter, solid rocket boosters (SRB) and external tank (ET) have required fatigue analyses as part of the mission life certification. Compared to rollout vibration, the SSV sees relatively high vibration loads during launch, ascent, descent and landing phases of the mission. Although preliminary measured SRB vibration levels during rollout were of low amplitude and frequency, the duration of the rollout phase is typically high, from 5-6 hours. As part of an expanded mission life assessment, additional certification effort was initiated to define fatigue load spectra for rollout. This study addresses the CT vibration analyses in support of the rollout fatigue study. Structural models developed for modal and vibration analyses were used to identify unique CT, CT/MLP and CT/MLP/SRB vibration characteristics for comparison to instrumented rollout tests. Whereas the main structural and vibration characteristics of the SSV are well defined, minimum analytical and vibration test data on the Crawler Transporter were available. Unique vibration characteristics of the CT are attributable to the drive mechanism, hydraulic jacking system, structural framing and the CT-to-MLP support pad restraints. Initial tests performed on the CT/MLP/SRB configuration showed reasonable correlation with predicted mode shapes and frequencies.

Margasahayam, Ravi N.↗
NASA NTRS1992

Coupled loads analysis for Space Shuttle payloads

Described here is a method for determining the transient response of, and the resultant loads in, a system exposed to predicted external forces. In this case, the system consists of four racks mounted on the inside of a space station resource node module (SSRNMO) which is mounted in the payload bay of the space shuttle. The predicted external forces are forcing functions which envelope worst case forces applied to the shuttle during liftoff and landing. This analysis, called a coupled loads analysis, is used to couple the payload and shuttle models together, determine the transient response of the system, and then recover payload loads, payload accelerations, and payload to shuttle interface forces.

Eldridge, J.↗
NASA NTRS1997

A Model for Space Shuttle Orbiter Tire Side Forces Based on NASA Landing Systems Research Aircraft Test Results

Forces generated by the Space Shuttle orbiter tire under varying vertical load, slip angle, speed, and surface conditions were measured using the Landing System Research Aircraft (LSRA). Resulting data were used to calculate a mathematical model for predicting tire forces in orbiter simulations. Tire side and drag forces experienced by an orbiter tire are cataloged as a function of vertical load and slip angle. The mathematical model is compared to existing tire force models for the Space Shuttle orbiter. This report describes the LSRA and a typical test sequence. Testing methods, data reduction, and error analysis are presented. The LSRA testing was conducted on concrete and lakebed runways at the Edwards Air Force Flight Test Center and on concrete runways at the Kennedy Space Center (KSC). Wet runway tire force tests were performed on test strips made at the KSC using different surfacing techniques. Data were corrected for ply steer forces and conicity.

Carter, John F.↗
NASA NTRS2007

Notch Fatigue Strength of a PM Disk Superalloy

New powder metallurgy (PM) disk superalloys, such as ME3, LSHR, and Alloy 10, have been developed in recent years which enable rim temperatures in turbine disk applications to approach 1300 F. Before these alloys can be utilized at 1300 F their long term durability must be ensured. One of the key requirements for disk rims is notch fatigue strength. This issue is extremely important and is a direct result of the blade attachment geometry employed at the disk rim. Further, the imposition of a dwell at maximum load, associated with take off and landing, can also affect notch fatigue strength. For these reasons a study has been undertaken to assess the notch dwell fatigue strength of a modern PM disk alloy through spin pit evaluation of a prototypical disk. The first element of this program involves screening potential heat treatments with respect to notch fatigue strength at 1300 F utilizing a conventional notch fatigue specimen with a stress concentration factor (K(sub t)) of 2 and a 90 sec dwell at peak load. The results of this effort are reported in this paper including the downselect of an optimal heat treatment, from a notch fatigue standpoint.

Gayda, John↗
NASA NTRS2016

Green Propellant Landing Demonstration at U.S. Range

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Propellant↗
NASA NTRS2010

Orion Entry Performance-Based Center-of-Gravity Box

The Orion capsule is designed both for Low Earth Orbit missions to the ISS and for missions to the moon. For ISS class missions, the capsule will use an Apollo-style direct entry. For lunar return missions, depending on the timing of the mission, the capsule could perform a direct entry or a skip entry of up to 4800 n.mi. in order to land in the coastal waters of California. The physics of atmospheric re-entry determine the capability of the Orion vehicle. For a given vehicle mass and shape, physics tells us that the driving parameters for an entry vehicle are the hypersonic lift-to-drag ratio (L/D) and the flight path angle at entry interface (gamma(sub EI)). The design of the Orion atmospheric re-entry must meet constraints during both nominal and dispersed flight conditions on landing accuracy, heating rate, total heat load, sensed acceleration, and proper disposal of the Service Module. These constraints define an entry corridor in the space of L/D-gamma(sub EI); if the vehicle falls within this corridor, then all constraints are met. The gamma(sub EI) dimension of the corridor can be further constrained by the gloads experienced during emergency entries. Thus, the entry performance for the Orion vehicle can be described completely by the L/D. Bounds on the hypersonic L/D necessary to achieve all the mission requirements can be defined for the given entry corridor. Landing accuracy performance drives the lower limit on L/D. In order to achieve the desired landing accuracy, a minimum L/D must be ensured. The design of the Thermal Protection System (TPS) drives the upper limit on L/D. A higher L/D can drive mass into the design of the TPS. Conversely, once the TPS is designed, the L/D must be ensured to stay below a certain limit in order for the TPS to stay within its design envelop. The L/D must stay within its upper and lower bounds during dispersed flight conditions. L/D is a function of both the aerodynamics and the center-of-gravity (CG) of the vehicle. The aerodynamics of the vehicle are determined by Computational Fluid Mechanics (CFD) and wind tunnel tests. However, the aerodynamics are not known precisely. Instead, an aerodynamic database has been developed where the aerodynamic coefficients are known to fall within a probabilistic band defined by upper and lower bounds. It is expected that the probabilistic band will shrink after the first missions are flown and real-world data is collected. Until that time, the Orion must be designed to the current aerodynamic database. Thus, for a given aerodynamic database with given uncertainties, the allowable range in L/D can be mapped to an allowable box for the CG location. The CG box is used to set requirements on the dispersions allowed for vehicle packaging and cargo storage. As the aerodynamic uncertainties decrease, the size of the CG box can increase. This paper discusses the technique used to map the minimum and maximum L/D bounds set by the entry performance requirements to the allowable dispersions in CG while accounting for aerodynamic uncertainties. The L/D is defined as the ratio of the lift force to the drag force. It is equivalent to the ratio of lift coefficient (C(sub L)) over drag coefficient (C(sub D)). C(sub L) and C(sub D) are functions of Mach number (M) and angle of attack (alpha). A Mach number of 25 is used as a measuring point of the hypersonic L/D. Variations in C(sub L), C(sub D) and alpha cause variations in L/D. Equation (1) shows the three contributions to the variation in L/D.

Rea, Jeremy R.↗
NASA NTRS2005

Gas Wave Bearings: A Stable Alternative to Journal Bearings for High-Speed Oil-Free Machines

To run both smoothly and efficiently, high-speed machines need stable, low-friction bearings to support their rotors. In addition, an oil-free bearing system is a common requirement in today's designs. Therefore, self-acting gas film bearings are becoming the bearing of choice in high-performance rotating machinery, including that used in the machine tool industry. Although plain journal bearings carry more load and have superior lift and land characteristics, they suffer from instability problems. Since 1992, a new type of fluid film bearing, the wave bearing, has been under development at the NASA Lewis Research Center in Cleveland, Ohio, by Dr. Florin Dimofte, a Senior Research Associate of the University of Toledo. One unique characteristic of the waved journal bearing that gives it improved capabilities over conventional journal bearings is the low-amplitude waves of its inner diameter surface. The radial clearance is on the order of one thousandth of the shaft radius, and the wave amplitude is nominally up to one-half the clearance. This bearing concept offers a load capacity which is very close to that of a plain journal bearing, but it runs more stably at nominal speeds.

Dimofte, Florin↗
NASA NTRS1977

Detection by pilots of system failures during instrument landings

The effects of gust disturbances and the pilot's participation mode in the control task on his work-load level and failure detection performance during a simulated low visibility landing approach are examined. We found that the participation mode had a strong effect on the pilot's work load, the induced work load being lowest when the pilot acted as a monitoring element during a coupled approach and highest when the pilot was an active element in the control loop. The effects of differential work load and participation mode on failure detection were separated. The participation mode was shown to have a dominant effect on the failure detection performance, with a failure in a monitored (coupled) axis being detected significantly faster than a comparable failure in a manually controlled axis.

Ephrath, A. R.↗
NASA NTRS2015

Overview of the 2nd Gen 3.7m HIAD Static Load Test

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for human class payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). In October of 2014, a 3.7m HIAD inflatable structure with an integrated flexible thermal protection sys-tem (F-TPS) was subjected to a static load test series to verify the designs structural performance. The 3.7m HIAD structure was constructed in a 70 deg sphere-cone stacked-toroid configuration using eight inflatable tori, which were joined together using adhesives and high strength textile webbing to help distribute the loads throughout the inflatable structure. The inflatable structure was fabricated using 2nd generation structural materials that permit an increase in use temperature to 400 C+ as compared to the 250 C limitation of the 1st generation materials. In addition to the temperature benefit, these materials also offer a 40 reduction in structure mass. The 3.7m F-TPS was fabricated using high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. The F-TPS was constructed of 2nd generation TPS materials increasing its heating capability from 35W sq cm to over 100W sq cm. This test article is the first stacked-torus HIAD to be fabricated and tested with a 70 deg sphere-cone. All previous stacked-torus HIADs have employed a 60o sphere-cone. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for dis-placement of the inflatable structure as loads were applied. The tub rim was attached to the floor to provide an airtight seal. The center body of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to characterize deformed shape, shoulder deflection, strap loads, and cord loads as a function of structural configuration and applied static load. In this overview, the 3.7m HIAD static load test series will be discussed in detail, including the 3.7m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally results and conclusions from the test series.

HIAD↗
NASA NTRS1975

A Monte Carlo analysis of the Viking lander dynamics at touchdown

The performance of the Viking lander has been evaluated by using a Monte Carlo simulation, and all results are presented in statistical form. The primary objectives of this analysis were as follows: (1) to determine the three sigma design values of maximum rigid body accelerations and the minimum clearance of the lander body during landing; (2) to determine the probability of an unstable landing; and (3) to determine the probability of the lander body striking a rock. Two configurations were analyzed with the only difference being in the ability of the primary landing gear struts to carry tension loads.

Muraca, R. J.↗
NASA NTRS1981

Contributions to a thermodynamic model of Earth systems on rivers

A model for the chemical (ground water) erosion and physical (bed load, including sedimentation) erosion of the land was developed. The rudiments of the relation between a regulated sea level (for the past 2500 million years) and the episodic rise and erosion of continents was examined to obtain some notion of the process scalings. Major process scales of about 200 years, 100,000 years, 3 My, 40 My, 300 My were estimated. It was suggested that a program targeted at ecological management would have to become familiar with processes at the first four scales (i.e., from glaciation to the horizontal movement of continents). The study returns to the initial premise. In order to understand and manage Earth biology (life, and modern man), it is necessary minimally to pursue systems' biogeology at a considerable number of process space and time scales via their irreversible thermodynamic couplings.

Iberall, A. S.↗
NASA NTRS1981

Structural dynamics payload loads estimates

The development of a full scale payload integration method which reduces the cost of a load cycle and is capable of land long very large systems is discussed. The method solves the coupled booster/payload system equations and does not involve additional approximations as compared to the standard transient analysis. The method is cost effective and directly applicable to the shuttle payload design case.

Engels, R. C.↗
NASA NTRS1923

Speed limits of aircraft

This paper is restricted to the question of attainable speed limits and attacks the problem from different angles. Theoretical limits due to air resistance are presented along with design factors which may affect speed such as wing loads, wing areas, wing section shifting, landing speeds, drag-lift ratios, and power coefficients.

Everling, E↗
NASA NTRS1994

Multi-partitioning for ADI-schemes on message passing architectures

A kind of discrete-operator splitting called Alternating Direction Implicit (ADI) has been found to be useful in simulating fluid flow problems. In particular, it is being used to study the effects of hot exhaust jets from high performance aircraft on landing surfaces. Decomposition techniques that minimize load imbalance and message-passing frequency are described. Three strategies that are investigated for implementing the NAS Scalar Penta-diagonal Parallel Benchmark (SP) are transposition, pipelined Gaussian elimination, and multipartitioning. The multipartitioning strategy, which was used on Ethernet, was found to be the most efficient, although it was considered only a moderate success because of Ethernet's limited communication properties. The efficiency derived largely from the coarse granularity of the strategy, which reduced latencies and allowed overlap of communication and computation.

Vanderwijngaart, Rob F.↗
NASA NTRSDate not supplied

Mars Sample Return, Sample Retrieval Lander, Reaction Control System Jet Interaction Supersonic Wind Tunnel Test Overview with CFD Predictions

NASA's Mars Sample Return campaign will be launching several missions over the next decade that will work together to return rock samples from Mars back to Earth. The Sample Retrieval Lander (SRL) will deliver the Mars Ascent Vehicle and fetch rover to the surface of Mars in 2006. Rock samples collected by the Mars 2020 Perseverance rover, landing in early 2021, will be loaded on the the ascent vehicle to be launched into Mars orbit for retrieval by yet another spacecraft. The Sample Retrieval Lander will be a blunt entry capsule similar to past Mars entry vehicles like Mars Science Laboratory and Viking. The vehicle will fly a guided entry, using a small lift vector produced by a non-zero trim angle of attack to eliminate downrange and crossrange position errors at the point of parachute deploy. This energy and heading management is achieved with a reaction control system (RCS) that directs the bank angle of the vehicle and also minimizes unwanted capsule dynamics. The reaction control system and control design is based on the Mars Science Laboratory and Mars 2020 RCS systems. However, due to packaging constraints, the backshell of this new entry vehicle has a different geometry than those earlier designs. To certify the RCS system for flight the project must characterize the jet plume interactions with the capsule backshell that could impair or significantly augment the RCS control authority. This characterization will be done through a combination of computational fluid dynamics (CFD) analysis and wind tunnel test. Two candidate arrangements of the RCS jets have been identified for the SRL vehicle and are currently under evaluation before final selection. The aero/RCS plume interactions of these candidate configurations have been measured in a supersonic wind tunnel test in NASA Langley's Unitary Plan Wind Tunnel. The test was conducted in the fall of 2020 and data is currently being reduced. An overview of the candidate RCS configurations are presented here with an overview of the wind tunnel model design, jet scaling and scaled nozzle design, and the test matrix. Preliminary CFD runs are presented with an assessment of the predicted plumes and their interaction with the wake flow of the vehicle. The predicted effects of the model sting is provided as well. This high fidelity wind tunnel test is being conducted much earlier in the SRL project than would normally be done. The test was funded as part of a CFD evaluation task funded by NASA's Aerosciences Evaluation and Test Capabilities Project. The objective of the evaluation task was to compare the ability of CFD to predict complex flows with data that can be measured in the Langley Unitary Plan Wind Tunnel. RCS Jet interactions were selected as a type of complex flow that is important to NASA missions. In addition to providing useful data to the SRL project, there was added emphasis on quantifying the accuracy of the CFD predictions and wind tunnel test data. An overview of the uncertainty quantification methodologies for computational and experimental portions of this test is presented.

blunt body↗
NASA NTRS1972

Analysis of legged landers for the survivable soft landing of instrument payloads.

Two methods of analysis have been developed for legged planetary landers. The first of these, the Large Displacement Gear Analysis, is a design tool for statically determining the large displacement stroking behavior, energy absorption characteristics, and internal load distributions in a single gear. The second, the Landing Dynamics Analysis, predicts the spatial landing dynamics of a legged lander. Effects of structural flexibility, elastic-plastic gear load characteristics, and soil properties on the loads, motions, and stability of the lander may be determined. Validation of the analytical techniques is accomplished through comparison of predicted results and experimental data obtained during a model test program.

Laurenson, R. M.↗
NASA NTRS1991

Mars entry-to-landing trajectory optimization and closed loop guidance

The guidance strategy of the Mars Rover Sample Return mission is presented in detail. Aeromaneuver versus aerobrake trades are examined, and an aerobrake analysis is presented which takes into account targeting, guidance, flight control, trajectory profile, delivery accuracy. An aeromaneuver analysis is given which includes the entry corridor, maneuver footprint, guidance, preentry phase, constant drag phase, equilibrium guide phase, variable drag phase, influence of trajectory profile on the entry flight loads, parachute deployment conditions and strategies, and landing accuracy. The Mars terminal descent phase is analyzed.

Ilgen, Marc R.↗
NASA NTRS2014

Overview of the 6 Meter HIAD Inflatable Structure and Flexible TPS Static Load Test Series

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). At the beginning of 2014, a 6m HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify the designs structural performance. The 6m HIAD structure was constructed in a stacked toroid configuration using nine inflatable torus segments composed of fiber reinforced thin films, which were joined together using adhesives and high strength textile woven structural straps to help distribute the loads throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for displacement of the inflatable structure as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The centerbody of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to provide many data sets including: deformed shape, shoulder deflection, strap loads, cord loads, inflation pressures, and applied static load.In this overview, the 6m HIAD static load test series will be discussed in detail, including the 6m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally initial results and conclusions from the test series.

inflatable aerodynamic decelerator↗