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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 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

NASA's Precision Segmented Reflectors (PSR) project

The PSR project representing a step toward developing the technology base needed to support future advanced astrophysics missions is described. The project is focused on producing high precision lightweight panels, large structures with proven erectability and accuracy, and lightweight and low power controls capable of achieving a high degree control of accuracy. Panels of sizes from 0.4 to 0.9 meter are fabricated to verify various design concepts for improving panel surface accuracy; in particular, a 0.4 meter panel is fabricated and refinished to a high degree of accuracy.

Pawlik, E.↗
NASA NTRS1985

Abrasion behavior of aluminum and composite skin coupons, stiffened skins and stiffened panels representative of transport airplane structures

A three-phase investigation was conducted to compare the friction and wear response of aluminum and graphite-epoxy composite materials when subjected to loading conditions similar to those experienced by the skin panels on the underside of a transport airplane during an emergency belly landing on a runway surface. The first phase involved a laboratory test which used a standard belt sander to provide the sliding abrasive surface. Small skin-coupon test specimens were abraded over a range of pressures and velocities to determine the effects of these variables on the coefficient of friction and wear rate. The second phase involved abrading I-beam stiffened skins on actual runway surface over the same range of pressures and velocities used in the first phase. In the third phase, large stiffened panels which most closely resembled transport fuelage skin construction were abraded on a runway surface. This report presents results from each phase of the investigation and shows comparisons between the friction and wear behavior of the aluminum and graphite-epoxy composite materials.

Jackson, K. E.↗
NASA NTRS1972

Design of a convective cooling system for a Mach 6 hypersonic transport airframe.

The design of a convective cooling system for an aluminum alloy airframe was established to verify preliminary weight estimates and to define future development requirements. A water glycol coolant was circulated through a closed loop network of supply and return lines to absorb incident aerodynamic heating from structural skin panels with integral passages and to transfer this heat input to a heat exchanger where it is rejected to the hydrogen fuel. Partial shielding of the aluminum alloy structure reduced heat loads to levels compatible with engine fuel flow requirements.

Anthony, F. M.↗
NASA NTRS1980

A cellular glass substrate solar concentrator

The design of a second generation point focusing solar concentration is discussed. The design is based on reflective gores fabricated of thin glass mirror bonded continuously to a contoured substrate of cellular glass. The concentrator aperture and structural stiffness was optimized for minimum concentrator cost given the performance requirement of delivering 56 kWth to a 22 cm diameter receiver aperture with a direct normal insolation of 845 watts sq m and an operating wind of 50 kmph. The reflective panel, support structure, drives, foundation and instrumentation and control subsystem designs, optimized for minimum cost, are summarized. The use of cellular glass as a reflective panel substrate material is shown to offer significant weight and cost advantages compared to existing technology materials.

Bedard, R.↗
NASA NTRS2001

STOP Analysis and Optimization of a Very-Low-Distortion Space Instrument: HST WFC3 Case Study

New generation optical instruments with very demanding stability requirements are being proposed and developed for space applications. STOP (Structural-Thermal-Optical Performance) analysis and optimization is crucial in meeting the very tight distortion budgets of these instruments. This presentation outlines STOP analysis and optimization approach in the context of WFC3 (Wide-Field Camera 3), which is a radial instrument designed to replace the Wide-Field Planetary Camera 2 (WFPC2) of the Hubble Space Telescope (HST). WFC3 houses two separate channels, UVIS and IR, and will have greater throughput and sensitivity than WFPC2. WFC3 line-of-sight alignment budget for the UVIS and IR channels are as small as 10 and 20 milli-arcsec, respectively. Its optical bench is the most critical subsystem effecting the optical stability of WFC3 hence our effort concentrates on the design and analysis of the bench and its interfaces. Structural analysis has accompanied the mechanical design of the bench since the initial concept study. A high fidelity structural Finite Element Model (FEM) of the bench has been developed and used for minimizing its thermally induced distortions as well as sizing it to meet the stiffness and strength requirements of a Shuttle launch. The bench is a composite honeycomb panel box structure with a very low planar Coefficient of Thermal Expansion (CTE) of approximately 0.1 ppm/C. Optic components are mounted to super-INVAR inserts bonded into the panels. The bench is kinematically supported on three HST latches via interface struts, which are tailored to exhibit negative CTE to cancel out the thermal motions of the latches. The interface struts also incorporate flexure elements to minimize the mechanical distortions coming into the bench from its enclosure. Bench FEM is coupled with the enclosure FEM to quantify these effects. Short term or on-orbit STOP analysis includes distortion due to the temperature variations of the bench, the struts, and the enclosure. Long term or ground-to-orbit STOP analysis includes distortional effects of gravity release, desorption, and assembly in addition to the ground-to-orbit temperature variations. A rigorous testing program has been implemented for verifying the material properties and the analysis predictions. STOP analysis results demonstrate that both the short-term and the long-term alignment budgets will be met. Presentation will cover design and analysis details that are critical to a successful implementation of the STOP analysis and optimization process.

Kunt, Cengiz↗
NASA NTRS1978

Computerized structural sizing at NASA Langley Research Center

Programs at the NASA Langley Research Center associated with the development of computerized structural sizing technology are reviewed. Particular attention is given to (1) lightweight columns for space structure applications, (2) stiffened composite panels for aerospace structures, (3) thermal structures for high-speed aircraft and space vehicles, (4) structural sizing methodology for finite-element structural models, (5) the sizing of large complex structural systems in multidisciplinary environments. Improvements to computational efficiency are noted with reference to a reduced number of sizing variables, a reduced number of constraints, and improved sizing algorithms.

Stroud, W. J.↗
NASA NTRS1981

Heat pipes cool probe and sandwich panel

Two concepts integrate heat-pipe technology. Probe with heat-pipe cooled jacket is self-contained, passive, and has no moving parts, unlike conventional air and water cooled probes. It is used in hostile, high temperature environments like wind tunnels and powerplants or on high-speed research and hypersonic cruise vehicles. Heat-pipe sandwich panel combines structural efficiency of sandwich with thermal efficiency of heat-pipe. It is used to eliminate thermal gradients and stresses, minimize thermal distortions, and transfer heat from one face of panel to other.

Camarda, C. J.↗
NASA NTRS1980

Emergency in-flight egress opening for general aviation aircraft

In support of a stall/spin research program, an emergency in-flight egress system is being installed in a light general aviation airplane. To avoid a major structural redesign for a mechanical door, an add-on 11.2 kg pyrotechnic-actuated system was developed to create an opening in the existing structure. The airplane skin will be explosively severed around the side window, across a central stringer, and down to the floor, creating an opening of approximately 76 by 76 cm. The severed panel will be jettisoned at an initial velocity of approximately 13.7 m/sec. System development included a total of 68 explosive severance tests on aluminum material using small samples, small and full scale flat panel aircraft structural mock-ups, and an actual aircraft fuselage. These tests proved explosive sizing/severance margins, explosive initiation, explosive product containment, and system dynamics.

Bement, L. J.↗
NASA NTRS1980

Emergency in-flight egress opening for general aviation aircraft

An emergency in-flight egress system was installed in a light general aviation airplane. The airplane had no provision for egress on the left side. To avoid a major structural redesign for a mechanical door, an add on 11.2 kg (24.6 lb) pyrotechnic-actuated system was developed to create an opening in the existing structure. The skin of the airplane was explosively severed around the side window, across a central stringer, and down to the floor, creating an opening of approximately 76 by 76 cm. The severed panel was jettisoned at an initial velocity of approximately 13.7 m/sec. System development included a total of 68 explosive severance tests on aluminum material using small samples, small and full scale flat panel aircraft structural mockups, and an actual aircraft fuselage. These tests proved explosive sizing/severance margins, explosive initiation, explosive product containment, and system dynamics. This technology is applicable to any aircraft of similar construction.

Bement, L. J.↗
NASA NTRS2010

ACT Payload Shroud Structural Concept Analysis and Optimization

Aerospace structural applications demand a weight efficient design to perform in a cost effective manner. This is particularly true for launch vehicle structures, where weight is the dominant design driver. The design process typically requires many iterations to ensure that a satisfactory minimum weight has been obtained. Although metallic structures can be weight efficient, composite structures can provide additional weight savings due to their lower density and additional design flexibility. This work presents structural analysis and weight optimization of a composite payload shroud for NASA s Ares V heavy lift vehicle. Two concepts, which were previously determined to be efficient for such a structure are evaluated: a hat stiffened/corrugated panel and a fiber reinforced foam sandwich panel. A composite structural optimization code, HyperSizer, is used to optimize the panel geometry, composite material ply orientations, and sandwich core material. HyperSizer enables an efficient evaluation of thousands of potential designs versus multiple strength and stability-based failure criteria across multiple load cases. HyperSizer sizing process uses a global finite element model to obtain element forces, which are statistically processed to arrive at panel-level design-to loads. These loads are then used to analyze each candidate panel design. A near optimum design is selected as the one with the lowest weight that also provides all positive margins of safety. The stiffness of each newly sized panel or beam component is taken into account in the subsequent finite element analysis. Iteration of analysis/optimization is performed to ensure a converged design. Sizing results for the hat stiffened panel concept and the fiber reinforced foam sandwich concept are presented.

Zalewski, Bart B.↗
NASA NTRS2012

Improvements in Cold-Plate Fabrication

Five improvements are reported in cold-plate fabrication. This cold plate is part of a thermal control system designed to serve on space missions. The first improvement is the merging of the end sheets of the cold plate with the face sheets of the structural honeycomb panel. The cold plate, which can be a brazed assembly, uses the honeycomb face sheet as its end sheet. Thus, when the honeycomb panel is fabricated, the face sheet that is used is already part of the cold plate. In addition to reducing weight, costs, and steps, the main benefit of this invention is that it creates a more structurally sound assembly. The second improvement involves incorporation of the header into the closure bar to pass the fluid to a lower layer. Conventional designs have used a separate header, which increases the geometry of the system. The improvement reduces the geometry, thus allowing the cold plate to fit into smaller area. The third improvement eliminates the need of hose, tube, or manifold to supply the cooling fluid externally. The external arrangement can be easily damaged and is vulnerable to leakage. The new arrangement incorporates an internal fluid transfer tube. This allows the fluid to pass from one cold plate to the other without any exposed external features. The fourth improvement eliminates separate fabrication of cold plate(s) and structural members followed by a process of attaching them to each other. Here, the structural member is made of material that can be brazed just as that of the cold plate. Now the structural member and the cold plate can be brazed at the same time, creating a monolithic unit, and thus a more structurally sound assembly. Finally, the fifth improvement is the elimination of an additional welding step that can damage the braze joints. A tube section, which is usually welded on after the braze process, is replaced with a more structurally sound configuration that can be brazed at the same time as the rest of the cold plate.

Zaffetti, Mark A.↗
NASA NTRS1983

Design, fabrication, installation and flight service evaluation of a composite cargo ramp skin on a model CH-53 helicopter

The installation of a composite skin panel on the cargo ramp of a CH-530 marine helicopter is discussed. The composite material is of Kevlar/Epoxy (K/E) which replaces aluminum outer skins on the aft two bays of the ramp. The cargo ramp aft region was selected as being a helicopter airframe surface subjected to possible significant field damage and would permit an evaluation of the long term durability of the composite skin panel. A structural analysis was performed and the skin shears determined. Single lap joints of K/E riveted to aluminum were statically tested. The joint tests were used to determine bearing allowables and the required K/E skin gage. The K/E skin panels riveted to aluminum edge members were tested in a shear fixture to confirm the allowable shear and bearing strengths. Impact tests were conducted on aluminum skin panels to determine energy level and damage relationship. The K/E skin panels of various ply orientations and laminate thicknesses were then impacted at similar energy levels. The results of the analysis and tests were used to determine the required K/E skin gages in each of the end two bays of the ramp.

Lowry, D. W.↗
NASA NTRS1999

Thermal and Mechanical Performance of a Carbon/Carbon Composite Spacecraft Radiator

Carbon-carbon composite materials offer greater thermal efficiency, stiffness to weight ratio, tailorability, and dimensional stability than aluminum. These lightweight thermal materials could significantly reduce the overall costs associated with satellite thermal control and weight. However, the high cost and long lead-time for carbon-carbon manufacture have limited their widespread usage. Consequently, an informal partnership between government and industrial personnel called the Carbon-Carbon Spacecraft Radiator Partnership (CSRP) was created to foster carbon-carbon composite use for thermally and structurally demanding space radiator applications. The first CSRP flight opportunity is on the New Millennium Program (NMP) Earth Orbiter-1 (EO-1) spacecraft, scheduled for launch in late 1999. For EO-1, the CSRP designed and fabricated a Carbon-Carbon Radiator (CCR) with carbon-carbon facesheets and aluminum honeycomb core, which will also serve as a structural shear panel. While carbon-carbon is an ideal thermal candidate for spacecraft radiators, in practice there are technical challenges that may compromise performance. In this work, the thermal and mechanical performance of the EO-1 CCR is assessed by analysis and testing. Both then-nal and mechanical analyses were conducted to predict the radiator response to anticipated launch and on-orbit loads. The thermal model developed was based on thermal balance test conditions. The thermal analysis was performed using SINDA version 4.0. Structural finite element modeling and analysis were performed using SDRC/1-DEAS and UAI/NASTRAN, respectively. In addition, the CCR was subjected to flight qualification thermal/vacuum and vibration tests. The panel meets or exceeds the requirements for space flight and demonstrates promise for future satellite missions.

Kuhn, Jonathan↗
NASA NTRS2015

High Altitude Venus Operational Concept (HAVOC): Proofs of Concept

The atmosphere of Venus is an exciting destination for both further scientific study and future human exploration. A recent internal NASA study of a High Altitude Venus Operational Concept (HAVOC) led to the development of an evolutionary program for the exploration of Venus, with focus on the mission architecture and vehicle concept for a 30-day crewed mission into Venus's atmosphere at 50 kilometers. Key technical challenges for the mission include performing the aerocapture maneuvers at Venus and Earth, inserting and inflating the airship at Venus during the entry sequence, and protecting the solar panels and structure from the sulfuric acid in the atmosphere. Two proofs of concept were identified that would aid in addressing some of the key technical challenges. To mitigate the threat posed by the sulfuric acid ambient in the atmosphere of Venus, a material was needed that could protect the systems while being lightweight and not inhibiting the performance of the solar panels. The first proof of concept identified candidate materials and evaluated them, finding FEP-Teflon (Fluorinated Ethylene Propylene-Teflon) to maintain 90 percent transmittance to relevant spectra even after 30 days of immersion in concentrated sulfuric acid. The second proof of concept developed and verified a packaging algorithm for the airship envelope to inform the entry, descent, and inflation analysis.

Jones, Christopher A.↗
NASA NTRS2002

Structural Acoustic Response of a Shape Memory Alloy Hybrid Composite Panel (Lessons Learned)

This study presents results from an effort to fabricate a shape memory alloy hybrid composite (SMAHC) panel specimen and test the structure for dynamic response and noise transmission characteristics under the action of thermal and random acoustic loads. A method for fabricating a SMAHC laminate with bi-directional SMA reinforcement is described. Glass-epoxy unidirectional prepreg tape and Nitinol ribbon comprise the material system. Thermal activation of the Nitinol actuators was achieved through resistive heating. The experimental hardware required for mechanical support of the panel/actuators and for establishing convenient electrical connectivity to the actuators is presented. Other experimental apparatus necessary for controlling the panel temperature and acquiring structural acoustic data are also described. Deficiency in the thermal control system was discovered in the process of performing the elevated temperature tests. Discussion of the experimental results focuses on determining the causes for the deficiency and establishing means for rectifying the problem.

Turner, Travis L.↗
NASA NTRS2004

Lightweight Deployable Mirrors with Tensegrity Supports

The upper part of Figure 1 shows a small-scale prototype of a developmental class of lightweight, deployable structures that would support panels in precise alignments. In this case, the panel is hexagonal and supports disks that represent segments of a primary mirror of a large telescope. The lower part of Figure 1 shows a complete conceptual structure containing multiple hexagonal panels that hold mirror segments. The structures of this class are of the tensegrity type, which was invented five decades ago by artist Kenneth Snelson. A tensegrity structure consists of momentfree compression members (struts) and tension members (cables). The structures of this particular developmental class are intended primarily as means to erect large segmented primary mirrors of astronomical telescopes or large radio antennas in outer space. Other classes of tensegrity structures could also be designed for terrestrial use as towers, masts, and supports for general structural panels. An important product of the present development effort is the engineering practice of building a lightweight, deployable structure as an assembly of tensegrity modules like the one shown in Figure 2. This module comprises two octahedral tensegrity subunits that are mirror images of each other joined at their plane of mirror symmetry. In this case, the plane of mirror symmetry is both the upper plane of the lower subunit and the lower plane of the upper subunit, and is delineated by the midheight triangle in Figure 2. In the configuration assumed by the module to balance static forces under mild loading, the upper and lower planes of each sub-unit are rotated about 30 , relative to each other, about the long (vertical) axis of the structure. Larger structures can be assembled by joining multiple modules like this one at their sides or ends. When the module is compressed axially (vertically), the first-order effect is an increase in the rotation angle, but by virtue of the mirror arrangement, the net first-order rotation between the uppermost and lowermost planes is zero. The need to have zero net rotation between these planes under all loading conditions in a typical practical structure is what prompts the use of the mirror configuration. Force and moment loadings other than simple axial compression produce only second-order deformations through strains in the struts and cables.

Zeiders, Glenn W.↗
DOE OSTI2025

Numerical analysis of gas exhaust in Wendelstein 7-X using the direct simulation Monte Carlo method

The present work is focused on a 3D numerical assessment of the Wendelstein 7-X (W7-X) particle exhaust. For all the numerical simulations the direct simulation Monte Carlo solver of the DIVGAS workflow, has been employed. The complex 3D geometry of the sub-divertor region includes the pumping gap panel, supporting structures, cooling pipes as well as the cryo-vacuum pump. All the considered flow simulations correspond to the Standard magnetic configuration of W7-X. The main conclusions, which can be extracted from the present numerical analysis could be summarized as follows; The coupling between EMC3-EIRENE and DIVGAS, which considers the fact that the incoming neutral particle flux at the sub-divertor is based on realistic plasma background, has been demonstrated. Three plasma scenarios have been considered, for which is clearly seen that by increasing the heating power, the neutral pressure as well as the resulting pumping efficiency is increased. The obtained numerical results of the neutral pressure in the sub-divertor lie within a more general scan matrix, which assumes a wider range of incoming particle flux, namely 10 19 –10 24 (s −1 ). It has been observed that, the sub-divertor neutral pressure is proportional to the incoming neutral particle flux, with the effective pumping speed to be a constant of proportionality. The influence of switching off the cryo-vacuum pump on the sub-divertor pressure is rather modest and a weak increase of the neutral pressure in the sub-divertor is expected. Correlations of the sub-divertor pressure with the total incoming particle flux as well as the individual pumped flux at each of the AEH and AEP sections have been deduced. Moreover, it has been demonstrated that the influence of the incoming neutral particle flux on the albedo coefficient at the AEH and AEP pumping gaps is rather weak. All the above numerical findings will actively support the optimization of the W7-X particle exhaust, in view of future experimental campaigns.

DSMC method↗