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This final report describes the results of a program for a crashworthy, 14.5mm API tolerant fuel cell construction developed and subjected to qualification testing.

The patent describes a process for fabricating a structure by vacuum brazing tantalum alloy elements in which a columbium-vanadium brazing foil is placed between surfaces of the elements to be joined. A coating of yttria powder is applied to surfaces of elements that will be in contact with a tungsten hearth plate and tungsten pellets; the structure is positioned on the tungsten hearth plate in a vacuum heating zone with tungsten pellets disposed on top of the portion of the upper element to be joined so as to exert pressure thereon and control the gap between the surfaces to be joined; the elements are heated in the vacuum heating zone to a temperature sufficient to melt the brazing foil. The vacuuum brazing process is particularly suitable for use in fabricating honeycomb sandwich panels, using tantalum alloy structural elements., Supersedes AD-164 555.

The patent describes apparatus for preventing frost from forming on cold components in a moist atmosphere. The apparatus includes: a housing to enclose the cold components on which frost is to be prevented from forming, with the housing having an inlet and an outlet; a honeycomb mesh fixedly positioned across the entire outlet; a source of cold dry nitrogen connected to the housing inlet; and valving means, between the source of nitrogen and the housing inlet, to control the flow of the cold dry nitrogen from the nitrogen source to the housing.

The program was undertaken to extend the technology base for highly loaded composite wing box structure through the selective application of the newer advanced composite material systems and minimization of the use of metallic structural elements. The design engineering and manufacturing technology portion of the program demonstrated the high compression strength of boron/aluminum through 350 F and the weight savings of boron/aluminum skinned honeycomb sandwich construction over a titanium baseline design. This volume reports the results of a study to further increase the structural efficiency of boron/aluminum to the 550 F temperature regime. The use and practicality of secondarily retort brazing boron/aluminum faced titanium honeycomb core sandwich construction in overcoming the limitations of its adhesive bonded counterpart was investigated. Various aspects of the basic brazing process including core pretreatment and braze alloy requirements were investigated. Optimal parameters were selected and a preliminary process specification written. The basic properties; the strength of codiffusion braze bonded overlap and integral stepped joints; and the strength of single shear, double row bolted joints of optimally brazed sandwich panels and plain specimens subjected to the selected braze temperature are reported. Pertinent conclusion and recommendations for future study are presented. (Author)

Engineering data appropriate to the use of selected graphite/epoxy composite materials in dimensionally critical space applications is presented. The thermal strain characteristics of a large number of material/layup configurations over the temperature range -320F to 250F were investigated. Additional data from the investigation of two systems, HM-S/X-30 and GY-70/X-904, is also included. The out-gassing characteristics in a hard vacuum at temperatures up to 150F were measured including identification of outgassing on optical reflectors. Thermal conductivity data over the temperature range of -320F to 250F is presented as is the effect of moisture and the effect of thermal cycling of composite substrate lightweight mirrors over selected temperature intervals between -320F and 250F. The fabrication techniques, including the effects of layup, resin contact, and modular cure are discussed as they relate to dimensional stability. Finally, the design and fabrication of a lightweight 12-inch diameter fully figured spherical mirror is discussed and the results of a test of the figure are presented. (Author-PL)

The objective of this program was to advance the state-of-the-art technology for low cost; low weight; thermal imaging systems for remote piloted vehicles. The principle areas of study are in automatic focus compensation, structures, optic system designs and electronic multiplexing methods. The report covers preliminary designs for three types of FLIR systems and predicts quantity cost of production for quantities of 100 and 500. This report also includes design data that allows the application of technology to the reduction of piece part costs by using plastic parts. In this manner, temperature compensation of refractive optical systems, and compensated reflective optical systems, can be fabricated with improved performance and reduced weight. Valuable fall-out of this study includes; techniques that can be used in manned aircraft systems to improve performance, make them simpler to operate, and reduce system power and weight requirements.

The objectives of this program were establishment of a technology to fabricate helicopter titanium main rotor blade spars cut from sheet material and joined by continuous seam diffusion bonding (CSDB) and the eventual application of this technology in the manufacturing of various components at reduced overall cost. The effects of different thermal treatments after bonding were determined by tensile and fatigue testing of small test specimens fabricated from titanium sheet flat panels joined by CSDB. Three 9-foot long spar sections were then fabricated, and two were tested in fatigue. The results indicate that continuous seam diffusion bonding has fatigue strength characteristics as good as, or better than, present fusion welded main rotor blade spar specimens and should be further evaluated for this application. (Modified author abstract)

The beta backscatter technique was determined to be feasible for the measurement of contaminants and surface preparation coatings. Titanium and aluminum honeycomb panels with bonded and disbonded areas were examined to identify (1) contaminants or surface preparation properties which result in disbonding and (2) surface preparation coatings which result in acceptable bonds. The poor bonding of the titanium panels was between the metal skin and the primer coating. Electron microscopy showed that the surface of the titanium skin was too smooth to favor physical bonding. Further examination showed there were no bonded areas suitable for analysis. The skins of the aluminum panels showed the bonded areas exhibited a microetchpit structure which was not present in the poorly bonded areas. It was concluded that the lack of bond of the aluminum panels was not the result of surface contamination but rather a function of the pickling parameters and perhaps the annealing of the aluminum. (Author)

The purpose of this research is to develop techniques for rapidly repair-ing bomb damaged runways. Structural concepts for filling and/or spanning the crater are studied. Spanning systems, such as clear span and intermediate support systems, structural fill systems including inter- locking fills, and waterbag systems are considered as possible solutions. Only one concept, a honeycomb structural fill technique, fulfills the requirements of simplicity, speed, and dependability to merit further testing. This technique consists of filling a large portion of the crater with several modules constructed of interconnected, vertically placed polyvinyl chloride (PVC) pipes. Laboratory tests are conducted on bonding agents, capacity of individual pipe, several different size modules, and various module arrangements. The F-111 aircraft wheel loads are considered as the design loading. Results of these tests lead to a PVC module design for full-sclae field tests. A small preliminary field test indicates that the repair technique is feasible. (Author-PL)

The patent relates to a module construction system which can be used, and can be adapted to be used, to meet varying structural requirements for different construction applications, including use as an aircraft runway and as other surfaces for aircraft, including taxiways and ramps. The system includes a plurality of identical load bearing structural modules. Each identical load bearing structural module is shaped in the form of a rectangular solid, and comprises an aluminum honeycomb core bonded to and between parallel flat skins of epoxy-fiberglass having a polyurethane coating.

The U.S. Supersonic Transport (SST) program resulted in many significant developments in titanium technology such as metallurgy, brazing, welding, and aircraft structural analysis. A two-phase follow-on program was funded by DOT/FAA to document areas of significant development completed during the SST effort and to expand selected technological areas most applicable to future American aerospace and other industry technology. This report summarizes the work conducted on titanium structure during the two-phase DOT/SST program and provides a concise reference to the more detailed reports available., Report on SST Technology Follow-on Program. Continuation of Contract DOT-FA-SS-71-12. See also AD0913365 and ADB000286.

Problem of reducing radiator size by placing fins inside of tubes is discussed. Calculations of radiator performance are compared with that of conventional radiator. Results show that internally finned radiators can dissipate the same heat as conventional radiators with same frontal area, but require 55% less volume and 18% less power expended. If radiators have same volume, the new designed requires 20% less pressure drop.

Honeycomb sandwich panels of various types fabricated of aluminum and graphite/epoxy composite facings bonded to aluminum and phenolic fiberglass (HRD) cores were subjected to 30-day accelerated aging exposures to high humidity (95% RH) and salt spray environments. Sandwiches were tested for flatwise tensile and flexural strength before and after environmental aging. In addition, weight loss determinations and visual examination of corrosion effects were performed. Severe galvanic corrosion was found on graphite/epoxy-aluminum core sandwiches during and after the salt spray exposure, and mechanical strength retention was drastically reduced. Those sandwiches made of graphite/ epoxy-aluminum core incorporating a 5-mil-thick glass fabric interliner in the core-to-face bonds (between the facings and the modified epoxy adhesive) withstood salt spray exposure extremely well, i.e., no evidence of galvanic corrosion or significant reduction of mechanical strength properties occurred. None of the sandwich types tested were affected by the 30-day 95% relative humidity exposure at 120 degrees F.

For brazing titanium under 1100 F, binary information formed the basis of selecting two promising ternary basis alloys for development, (1) Ag- Cu-Ge and (2) Ag-Cu-Sn. Melt and flow behavior was favorable with Ag-Cu-Sn base alloys, but work was discontinued because of excessive joint brittleness related to intermetallic films formed on the substrate. Alloys based on the Ag-Cu-Ge ternary also formed intermetallic films, but the joints had intermediate strength which justified evaluation. Average shear strength of lap joints varied from 10,750 psi for aged specimens to 11,900 psi for specimens after brazing at 1100 F. These marginal properties would justify only limited application of the Ag-Cu Ge based alloys., Report on Metallic Materials.

Design graphs, formulas and procedures for the design and evaluation of the electrical performance of metal space frame radomes is presented covering specifically transmission loss, boresight error, and side-lobe variation. The relative merits of metal and dielectric elements is analyzed. The study consists of developing a straight forward simple procedure for the optimum sizing of solid beam elements for large space frame radomes. Equations were derived expressing the relationship between the element dimensions necessary to satisfy a prescribed electrical transmission loss and the structural integrity of the space frame. These equations, considered as design constraints, were utilized in the selection of the element size which satisfied a defined optimum criterion. Two optimization criteria were adopted: first, the minimization of the transmission loss subject to a constraint of structural integrity and, second, the reduction of the total structural weight of the elements which simultaneously satisfy the transmission loss and structural constraint. A sample design computation was conducted as an example of the procedure. The second optimization criterion was applied to a 150 foot diameter radome which was required to sustain 150 mph winds. A procedure for implementing the first optimization criterion employing the method of Lagrangian multipliers was presented resulting in four equations to be solved simultaneously.

The exploration of space will undoubtedly require a wide variety of large structures. Several approaches to producing those large structures become immediately apparent. Small modules or components of the overall structure could be transported into the desired position and assembled into the overall structure. Another approach is to develop an expandable structure which has a small packaging volume and can be transported into the desired position in its fully assembled configuration. Of equal importance to the small packaging volume is the strength-to-weight ratio of the finished structure. Excess weight transported into the space environment means larger rockets and highly increased costs of launch. GCA Viron Division, in conjunction with several subcontractors, have developed a concept which produces a structure with a low packaging volume and provides high strength with a low weight penalty. This paper will discuss the development phases, materials research, fabrication techniques, experimental development, and actual model space structures which have been fabricated based on the expandable honeycomb concept., Presented at the Conference on Aerospace Expandable Structures, held in Minneapolis, MN, on 25-27 May 1965. Published in CONTRAILS, the Transactions of the Conference on Aerospace Expandable Structures, p50-73, 1965. Prepared in collaboration with Experimental Fabrication Division, Wright-Patterson AFB, OH.

The effects of absorbed moisture on the energy-dissipating characteristics of paper honeycomb were first studied in the Structural Mechanics Research Laboratory at the University of Texas at Austin and reported in 1959. In this study, results of the former study are reexamined and most of the earlier experimental work is repeated using improved techniques. Results indicate that moisture content has no significant effect on average crushing strength, or the energy-dissipating capacity of paper honeycomb until the moisture content exceeds 14% of the dry weight of the sample. Taking into consideration the slow rate at which paper honeycomb absorbs moisture from the air, it is concluded that moisture content is not likely, under ordinary circumstances, to be a significant consideration in the use of paper honeycomb as a cushioning material. This conclusion is essentially the same as the one reached as a result of the earlier study.

A program was conducted to evaluate the performance of liquid-nitrogen-cooled condensers in pumping the exhaust of mercury ion thrustors. Tests were conducted with two different condensers. One of the condensers was a smooth cylinder having a total surface area of 3. 8 square meters, and the other condenser was of honeycomb type of construction and had a total surface area of 62. 9 square meters. Two targets, having apex angles of 320 and 1480, respectively, were used with each condenser. Ion beam current ranged from 10 to 200 milliamperes and accelerating potential from 2. 0 to 6. 0 kilovolts. The results were evaluated on the basis of a performance factor F* which is a function of the condenser accommodation and sticking coefficients. A mean value of F* of 0. 75 was determined in this study. A visual observation of mercury distribution patterns showed that the ratio of effective pumping area to actual geometric area was 0. 4 for the smooth condenser and 0. 1 for the honeycomb condenser. This study showed no significant difference in the effect of the two targets.

This report describes the work accomplished on Research and Development of S-1C Heat Shield Panels utilizing brazed stainless steel honeycomb sandwich construction for the period January 29, 1963, to January 29, 1964, inclusive. The principal effort in this program was the stress, thermal and design analysis of two S-1C heat shield panel designs which differed by virtue of the panel mounting system. As a consequence of this analytical study and the NASA S-1C Heat Shield Panel Tests conducted at Wylie Laboratories, either panel design appears to satisfy the S-1C requirements provided the M-31 insulation is retained on the panel by deformation of the honeycomb insulation reinforcement. Additional items investigated include: (1) Application of Beryllium Sheet to S-1C Heat Shield panels. (2) Deflection Characteristics of M-31 Insulation with Deformed Stainless Steel Honeycomb Reinforcement. (3) Analysis of Braze Defects, Braze Quality Standards and Repair Methods. (3) Analysis of Holes in Heat Shield Panels and Experimental Measurements of Stresses at Hole Boundaries. (5) Design Recommendations for Heat Shield Panels.

The Honeycomb Aerodynamic Structure (SKIN) of the LANCE XM234 Warhead was damaged during the engineer test/expanded service test (ET/EST) program at White Sands Missile Range (WSMR), New Mexico. This damage ride it necessary to the warhead support cradles on the LANCE system tracked vehicles. The first modification consisted of changing the station 71.5 saddle to provide a larger bearing area (65 sq in.) and better control of hold down strap tension. The second modification consisted of moving the forward saddle from station 71.5 to station 61.54 in order to support the warhead section where structural joint rings exist. In addition to cradle modifications shock absorbers were in stalled on the two rear roadwheels of the LANCE tracked vehicle to help reduce the road shock and vibration environment for the missiles.

Two approximate analytical models are defined for a landing configuration in which a spherical payload can sometimes penetrate into its crushable casing. Results for both models are found to agree reasonably well with two previous experimental measurements. Design examples are presented for an impact velocity of 300 ft per sec. These are based on choices of zero or 'perfect' payload bonding, and of either a balsa-like or honeycomb-like class of crushable material. The greatest difference between the models for these examples is a 29-percent discrepancy in the required maximum crushing stress. A particular pair of examples gives the unexpected result that penetration can provide a decrease in crushable material weight by a factor greater than 4 when the honeycomb-like class of material is required without penetration, but the more efficient balsa-like class is feasible with penetration.

Significant reports covering the period since 1962 were selected for this study, with reference to relevant earlier works. Recent research in honeycomb sandwich structural materials concerns: (1) the investigation of manufacturing methods and design; (2) testing of panels of various constructions for the effects of elevated temperatures, edgewise compression, thermal response, energy absorption, vibration, and strength as affected by surface waviness and core cell size; (3) testing the joining (welding and brazing) properties of various alloys.

An outline is given of the development history of Royal Aircraft Establishment(RAE) carbon fibre reinforced plastics. Initially, development was inhibited by poor interlaminar shear strength. However, composite properties have improved since the advent of various surface treatments derived by the fibre producers. Present work indicates certain criteria by which resins are selected for use as a matrix with the two available types of RAE fibre. These are more critical for filament wound cylindrical components than for pressed flat laminates, as shown by levels of fibre compaction achieved on hoop wound NOL rings and helically wound tubes. Currently, the strength data fall well below that available if the full strength of the single fibre or bundle is utilised. However, the majority of the fibre stiffness has been achieved in unidirectional composites. Illustrations are shown of a carbon fibre reinforced plastic and honeycomb sandwich satellite structure which has been constructed to investigate the efficacy of methods selected for laminating polygonal cylindrical shapes, the subsequent adhesive bonding and the final machining to complete the structural assembly

An experimental investigation has been conducted to develop highly efficient energy absorbing structural elements from radiofrequency transparent materials. The materials were evaluated using two types of energy absorption processes; the fragmenting tube and the crushing of honeycomb and foamed materials. Plastic reinforced glass fabric tubes, employed in the fragmenting tube process, yielded specific energies of 28,400 and 12,300 ft-lb/lb for fabric orientation in the longitudinal and hoop directions respectively. The crushing of nylon phenolic honeycomb yielded a value of specific energy of l4,400 ft-lb/lb for a 3/16 inch cell size, 9.0 lb/ft3 density, and with a usable stroke of 80% of initial length. Radio frequency transmission losses were well within acceptable limits for the specified 100 to 2000 megacycle frequency range investigated.

Specimens representative of metal aircraft structural components reinforced with boron filamentary composites were manufactured and tested under cyclic loading, cyclic temperature, or continuously applied loading to evaluate some of the factors that affect structural integrity under cyclic conditions. Bonded, stepped joints were used throughout to provide composite-to-metal transition regions at load introduction points. Honeycomb panels with titanium or aluminum faces reinforced with unidirectional boron composite were fatigue tested at constant amplitude under completely reversed loading. Results indicated that the matrix material was the most fatigue-sensitive part of the design, with debonding initiating in the stepped joints. However, comparisons with equal weight all-metal specimens show a 10 to 50 times improved fatigue life. Fatigue crack propagation and residual strength were studied for several different stiffened panel concepts, and were found to vary considerably depending on the configuration. Weight savings up to 30 percent may be realized with the better concepts when compared to all-metal structure. Composite-reinforced metal specimens were also subjected to creep and thermal cycling tests. The creep tests at 50 percent of tensile ultimate load were inconclusive due to large scatter in the limited tests. Thermal cycling of stepped joint tensile specimens resulted in a ten percent decrease in residual strength after 4000 cycles.

The purpose of this document is to summarize the mechanical properties of Hexcel heat resistant reinforced plastic honeycomb. All the honeycomb types evaluated were manufactured on production equipment. Most of them are standard production materials and any honeycomb considered as experimental is identified by the Prefix X in the designation. The following table provides a description of these materials., Supersedes rept. dated 25 Sep 58.

This paper treats the instability of a honeycomb sandwich cylinder under axial compression, torsion and external lateral pressure. The simply supported honeycomb-sandwich cylinder considered here has face sheets of equal thickness and an isotropic honeycomb core. Curves are presented for the cases of axial compression and lateral pressure only, as well as torsion only. The combined case is handled by means of interaction equations based on the curves shown.

The patent provides a lenticulated rear projection screen comprising a multitude of tiny aspherical lenses, preferably arranged in a honeycomb pattern. It provides a method of producing a lenticulated rear projection screen comprising a multitude of tiny aspherical lenses arranged in a honeycomb pattern.

The patent describes an improved acoustic window or lens comprising a metallic core element and a thermoplastic filler which is shaped to conform to complex curves. A method of manufacture is also disclosed, wherein voids in said core are filled with particles of thermoplastic and then heated to melt the particles into a fluid which hardens to fill the voids.

An investigation was conducted to determine the vibrational characteristics of sandwich panels over a wide range of environmental pressure or density. Natural frequencies and damping of panels having aluminum honeycomb and polyurethane foam cores were determined for several modes of vibration over a pressure range from 760 torr (1 atmosphere) to 10-6 torr. Results indicate that no significant changes in either natural frequency or damping occur between a pressure of 1 torr and 10-6 torr. A decrease in pressure from 1 atmosphere to 1 torr produces an increase in frequency from 2 to 10 percent due to "apparent mass" effects and possible increases in panel thickness and shear rigidity. Furthermore, a corresponding decrease in pressure results in a linear decrease in damping due to a reduction in acoustic radiation. The change in frequency and damping with pressure is dependent upon the mode shape of the panel.

Attachment of honeycomb sandwich structure has been an important design consideration in the aerospace field since the early 50's when bonded sandwich was first used extensively as covering for the wings and tail of the supersonic B-58. Since then, virtually all military and an increasing number of commercial aircraft and space vehicles utilize honeycomb sandwich as primary structure. This paper illustrates the basic approaches to sandwich panel attachment and the variety of typical solutions. Successful honeycomb fastening design concepts from B-58. B-57F, and F-111 production programs at General Dynamics are illustrated. Sealing of sandwich joints for use as integral fuel tanks and application of specialty fasteners for honeycomb sandwich are discussed. Structural efficiency of 'fastened' stepped-down edge concepts versus full-depth panel edges is compared. A full-depth panel edge design with a stud and plug with spliceplate is currently an attractive design for fastening both metallic and composite-faced sandwich.

An experimental investigation was made to determine the resistance to heat penetration and deformations or buckling of three ablation materials during simulated exposure to lifting vehicle afterbody heating conditions. The ablation materials, which were bonded to inconel cones, were subjected to convective heating at cold-wall heating rates ranging from 4 to 40 Btu/sq ft-sec (45 to 454 kW/sq m) in arc jet streams of air and nitrogen. The ablation materials tested were a molded epoxy-based composite and a silicone elastomeric with and without honeycomb reinforcement. For the test conditions of the investigation, models with the epoxy-based composite and the unreinforced silicone elastomeric developed buckles in the ablation material; two models of the epoxy-based composite failed catastrophically. The honeycomb reinforcement in the silicone elastomeric was beneficial in restraining thermal expansion and maintaining char integrity. Except for tests in nitrogen, the silicone elastomerics provided the best resistance to heat penetration.

During the continuation of the first year's effort on forming shaped and hollow glass filaments, the process was refined. Precision windings were made to obtain close packing from rectangular microtape, both solid and hollow, while hollow hexagonal fiber was wound into a micro-honeycomb. Substantial mechanical properties data were developed on 2.3 in. diameter tubes precision wound from solid microtape into 90% by volume glass structures. These displayed no room temperature permeability to water and modulus of elasticity values the same as the glass; namely, 1,000,000 psi. Axial strengths of the tubes with only circumferential windings were as high as 25,000 psi. (MM)

This program analytically and experimentally investigates the design and thermostructural performance of integrated ablative/radiative thermal protection systems in two phases: CONCEPT DEVELOPMENT AND FULL-SCALE PANEL DEVELOPMENT. In Phase I, the program includes the investigation, analysis, and test of sub-scale specimens and various design concepts to determine design values for the full-scale panels and leading edges. Phase II covers the design and fabrication of large-scale lower surface ablative panels and leading edges for integration on the AFFDL Hypersonic Aerospace Test Structure (HATS). A vehicle configuration, the FDL-5A, is furnished by the Air Force to serve as a basepoint for the analytical and experimental phases of the program. (Author, modified-PL), Prepared in cooperation with HITCO, Gardena, Calif. Defense Products Div. and Acurex Corp., Mountain View, Calif. Aerotherm Div.

The performance of diffusion-bonded titanium sandwich constructions was determined by evaluating flatwise shear, flatwise tensile, and compressive properties at room temperature. Shear and flatwise tension fatigue data were also obtained at room temperature. Edgewise compression tests were conducted on sandwich with three different facing thicknesses at temperatures up to 1200F. Fatigue loading of edgewise compression specimens was done at room temperature. Shear fatigue strength approached about 10 percent of static shear strength and flatwise tensile fatigue strength about 5 percent of static strength. Edgewise compression fatigue endurance limit was from 60 to 70 percent of static strength values. Effects of increasing temperature dropped static edgewise compressive strength, especially at 1200F. (Author)