Advanced Technology Composite Fuselage - Materials and Processes

NASA Technical Reports Server (NTRS)

Scholz, D. B.; Dost, E. F.; Flynn, B. W.; Ilcewicz, L. B.; Nelson, K. M.; Sawicki, A. J.; Walker, T. H.; Lakes, R. S.

1997-01-01

The goal of Boeing's Advanced Technology Composite Aircraft Structures (ATCAS) program was to develop the technology required for cost and weight efficient use of composite materials in transport fuselage structure. This contractor report describes results of material and process selection, development, and characterization activities. Carbon fiber reinforced epoxy was chosen for fuselage skins and stiffening elements and for passenger and cargo floor structures. The automated fiber placement (AFP) process was selected for fabrication of monolithic and sandwich skin panels. Circumferential frames and window frames were braided and resin transfer molded (RTM'd). Pultrusion was selected for fabrication of floor beams and constant section stiffening elements. Drape forming was chosen for stringers and other stiffening elements. Significant development efforts were expended on the AFP, braiding, and RTM processes. Sandwich core materials and core edge close-out design concepts were evaluated. Autoclave cure processes were developed for stiffened skin and sandwich structures. The stiffness, strength, notch sensitivity, and bearing/bypass properties of fiber-placed skin materials and braided/RTM'd circumferential frame materials were characterized. The strength and durability of cocured and cobonded joints were evaluated. Impact damage resistance of stiffened skin and sandwich structures typical of fuselage panels was investigated. Fluid penetration and migration mechanisms for sandwich panels were studied.

Characterization and manufacture of braided composites for large commercial aircraft structures

NASA Technical Reports Server (NTRS)

Fedro, Mark J.; Willden, Kurtis

1992-01-01

Braided composite materials, one of the advanced material forms which is under investigation in Boeing's ATCAS program, have been recognized as a potential cost-effective material form for fuselage structural elements. Consequently, there is a strong need for more knowledge in the design, manufacture, test, and analysis of textile structural composites. The overall objective of this work is to advance braided composite technology towards applications to a large commercial transport fuselage. This paper summarizes the mechanics of materials and manufacturing demonstration results which have been obtained in order to acquire an understanding of how braided composites can be applied to a commercial fuselage. Textile composites consisting of 1D, 2D triaxial, and 3D braid patterns with thermoplastic and two RTM resin systems were investigated. The structural performance of braided composites was evaluated through an extensive mechanical test program. Analytical methods were also developed and applied to predict the following: internal fiber architectures, stiffnesses, fiber stresses, failure mechanisms, notch effects, and the entire history of failure of the braided composites specimens. The applicability of braided composites to a commercial transport fuselage was further assessed through a manufacturing demonstration. Three foot fuselage circumferential hoop frames were manufactured to demonstrate the feasibility of consistently producing high quality braided/RTM composite primary structures. The manufacturing issues (tooling requirements, processing requirements, and process/quality control) addressed during the demonstration are summarized. The manufacturing demonstration in conjunction with the mechanical test results and developed analytical methods increased the confidence in the ATCAS approach to the design, manufacture, test, and analysis of braided composites.

Global cost and weight evaluation of fuselage keel design concepts

NASA Technical Reports Server (NTRS)

Flynn, B. W.; Morris, M. R.; Metschan, S. L.; Swanson, G. D.; Smith, P. J.; Griess, K. H.; Schramm, M. R.; Humphrey, R. J.

1993-01-01

The Boeing program entitled Advanced Technology Composite Aircraft Structure (ATCAS) is focused on the application of affordable composite technology to pressurized fuselage structure of future aircraft. As part of this effort, a design study was conducted on the keel section of the aft fuselage. A design build team (DBT) approach was used to identify and evaluate several design concepts which incorporated different material systems, fabrication processes, structural configurations, and subassembly details. The design concepts were developed in sufficient detail to accurately assess their potential for cost and weight savings as compared with a metal baseline representing current wide body technology. The cost and weight results, along with an appraisal of performance and producibility risks, are used to identify a globally optimized keel design; one which offers the most promising cost and weight advantages over metal construction. Lastly, an assessment is given of the potential for further cost and weight reductions of the selected keel design during local optimization.

Design and Analysis of a Stiffened Composite Fuselage Panel

NASA Technical Reports Server (NTRS)

Dickson, J. N.; Biggers, S. B.

1980-01-01

A stiffened composite panel has been designed that is representative of the fuselage structure of existing wide bodied aircraft. The panel is a minimum weight design, based on the current level of technology and realistic loads and criteria. Several different stiffener configurations were investigated in the optimization process. The final configuration is an all graphite epoxy J-stiffened design in which the skin between adjacent stiffeners is permitted to buckle under design loads. Fail-safe concepts typically employed in metallic fuselage structure have been incorporated in the design. A conservative approach has been used with regard to structural details such as skin frame and stringer frame attachments and other areas where sufficient design data was not available.

Hybrid-Wing-Body Vehicle Composite Fuselage Analysis and Case Study

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

2014-01-01

Recent progress in the structural analysis of a Hybrid Wing-Body (HWB) fuselage concept is presented with the objective of structural weight reduction under a set of critical design loads. This pressurized efficient HWB fuselage design is presently being investigated by the NASA Environmentally Responsible Aviation (ERA) project in collaboration with the Boeing Company, Huntington Beach. The Pultruded Rod-Stiffened Efficient Unitized Structure (PRSEUS) composite concept, developed at the Boeing Company, is approximately modeled for an analytical study and finite element analysis. Stiffened plate linear theories are employed for a parametric case study. Maximum deflection and stress levels are obtained with appropriate assumptions for a set of feasible stiffened panel configurations. An analytical parametric case study is presented to examine the effects of discrete stiffener spacing and skin thickness on structural weight, deflection and stress. A finite-element model (FEM) of an integrated fuselage section with bulkhead is developed for an independent assessment. Stress analysis and scenario based case studies are conducted for design improvement. The FEM model specific weight of the improved fuselage concept is computed and compared to previous studies, in order to assess the relative weight/strength advantages of this advanced composite airframe technology

Advanced technology composite aircraft structures

NASA Technical Reports Server (NTRS)

Ilcewicz, Larry B.; Walker, Thomas H.

1991-01-01

Work performed during the 25th month on NAS1-18889, Advanced Technology Composite Aircraft Structures, is summarized. The main objective of this program is to develop an integrated technology and demonstrate a confidence level that permits the cost- and weight-effective use of advanced composite materials in primary structures of future aircraft with the emphasis on pressurized fuselages. The period from 1-31 May 1991 is covered.

Subcomponent tests for composite fuselage technology readiness

NASA Technical Reports Server (NTRS)

Madan, R. C.; Hawley, A. V.

1989-01-01

An account is given of a NASA research effort aimed at the development of an all-composite transport aircraft fuselage incorporating joints and cutouts which meets all design requirements. The design, construction, and analysis activities associated with the 30-ft-long fuselage section gave attention to critically important subcomponent specimens, including shear-tee pulloff specimens, stiffened and unstiffened cutout panels, longitudinal and transverse skin splices, longeron runouts, transverse skin-longerons, stiffened shear panels, and window belt panels. The analysis of large cutouts was conducted with coupling FEM analyses incorporating accurate failure criteria for tension and shear; the strategic application of S2 glass fiber plies around cutouts was demonstrated both analytically and experimentally to increase load capacity with virtually no weight penalty.

Design and analysis of a stiffened composite fuselage panel

NASA Technical Reports Server (NTRS)

Dickson, J. N.; Biggers, S. B.

1980-01-01

The design and analysis of stiffened composite panel that is representative of the fuselage structure of existing wide bodied aircraft is discussed. The panel is a minimum weight design, based on the current level of technology and realistic loads and criteria. Several different stiffener configurations were investigated in the optimization process. The final configuration is an all graphite/epoxy J-stiffened design in which the skin between adjacent stiffeners is permitted to buckle under design loads. Fail safe concepts typically employed in metallic fuselage structure have been incorporated in the design. A conservative approach has been used with regard to structural details such as skin/frame and stringer/frame attachments and other areas where sufficient design data was not available.

Textile technology development

NASA Technical Reports Server (NTRS)

Shah, Bharat M.

1995-01-01

The objectives of this report were to evaluate and select resin systems for Resin Transfer Molding (RTM) and Powder Towpreg Material, to develop and evaluate advanced textile processes by comparing 2-D and 3-D braiding for fuselage frame applications and develop window belt and side panel structural design concepts, to evaluate textile material properties, and to develop low cost manufacturing and tooling processes for the automated manufacturing of fuselage primary structures. This research was in support of the NASA and Langley Research Center (LaRc) Advanced Composite Structural Concepts and Materials Technologies for Primary Aircraft Structures program.

78 FR 66317 - Special Conditions: Learjet Inc. Model LJ-200-1A10; Airplane Fuselage Post-Crash Fire Survivability

Federal Register 2010, 2011, 2012, 2013, 2014

2013-11-05

... feature associated with advanced composite materials in the construction of its fuselage and wings. The... is the first airplane manufactured by Learjet Inc. to utilize advanced composite materials in the... composite materials in the construction of its fuselage and wings. In accordance with Sec. 21.16, fuselage...

Development of Textile Reinforced Composites for Aircraft Structures

NASA Technical Reports Server (NTRS)

Dexter, H. Benson

1998-01-01

NASA has been a leader in development of composite materials for aircraft applications during the past 25 years. In the early 1980's NASA and others conducted research to improve damage tolerance of composite structures through the use of toughened resins but these resins were not cost-effective. The aircraft industry wanted affordable, robust structures that could withstand the rigors of flight service with minimal damage. The cost and damage tolerance barriers of conventional laminated composites led NASA to focus on new concepts in composites which would incorporate the automated manufacturing methods of the textiles industry and which would incorporate through-the-thickness reinforcements. The NASA Advanced Composites Technology (ACT) Program provided the resources to extensively investigate the application of textile processes to next generation aircraft wing and fuselage structures. This paper discusses advanced textile material forms that have been developed, innovative machine concepts and key technology advancements required for future application of textile reinforced composites in commercial transport aircraft. Multiaxial warp knitting, triaxial braiding and through-the-thickness stitching are the three textile processes that have surfaced as the most promising for further development. Textile reinforced composite structural elements that have been developed in the NASA ACT Program are discussed. Included are braided fuselage frames and window-belt reinforcements, woven/stitched lower fuselage side panels, stitched multiaxial warp knit wing skins, and braided wing stiffeners. In addition, low-cost processing concepts such as resin transfer molding (RTM), resin film infusion (RFI), and vacuum-assisted resin transfer molding (VARTM) are discussed. Process modeling concepts to predict resin flow and cure in textile preforms are also discussed.

Test results for composite specimens and elements containing joints and cutouts

NASA Technical Reports Server (NTRS)

Sumida, P. T.; Madan, R. C.; Hawley, A. V.

1988-01-01

A program was conducted to develop the technology for joints and cutouts in a composite fuselage that meets all design requirements of a large transport aircraft for the 1990s. An advanced trijet derivative of the DC-10 was selected as the baseline aircraft. Design and analysis of a 30-foot-long composite fuselage barrel provided a realistic basis for the test effort. The primary composite material was Hexcel F584 resin on 12 K IM6 fiber, in tape and broadgoods form. Fiberglass broadgoods were used in E-glass and S-glass fiber form in the cutout region of some panels. Additionally, injection-molded chopped graphite fiber/PEEK was used for longeron-to-frame shear clips. The test effort included four groups of test specimens, beginning with coupon specimens of mono-layer and cross-piled laminates, progressing through increasingly larger and more complex specimens, and ending with two 4- by 5-foot curved fuselage side panels. One of the side panels incorporated a transverse skin splice, while the second included two cabin window cutouts.

Nonlinear and progressive failure aspects of transport composite fuselage damage tolerance

NASA Technical Reports Server (NTRS)

Walker, Tom; Ilcewicz, L.; Murphy, Dan; Dopker, Bernhard

1993-01-01

The purpose is to provide an end-user's perspective on the state of the art in life prediction and failure analysis by focusing on subsonic transport fuselage issues being addressed in the NASA/Boeing Advanced Technology Composite Aircraft Structure (ATCAS) contract and a related task-order contract. First, some discrepancies between the ATCAS tension-fracture test database and classical prediction methods is discussed, followed by an overview of material modeling work aimed at explaining some of these discrepancies. Finally, analysis efforts associated with a pressure-box test fixture are addressed, as an illustration of modeling complexities required to model and interpret tests.

Weight Assessment for Fuselage Shielding on Aircraft With Open-Rotor Engines and Composite Blade Loss

NASA Technical Reports Server (NTRS)

Carney, Kelly; Pereira, Michael; Kohlman, Lee; Goldberg, Robert; Envia, Edmane; Lawrence, Charles; Roberts, Gary; Emmerling, William

2013-01-01

The Federal Aviation Administration (FAA) has been engaged in discussions with airframe and engine manufacturers concerning regulations that would apply to new technology fuel efficient "openrotor" engines. Existing regulations for the engines and airframe did not envision features of these engines that include eliminating the fan blade containment systems and including two rows of counter-rotating blades. Damage to the airframe from a failed blade could potentially be catastrophic. Therefore the feasibility of using aircraft fuselage shielding was investigated. In order to establish the feasibility of this shielding, a study was conducted to provide an estimate for the fuselage shielding weight required to provide protection from an open-rotor blade loss. This estimate was generated using a two-step procedure. First, a trajectory analysis was performed to determine the blade orientation and velocity at the point of impact with the fuselage. The trajectory analysis also showed that a blade dispersion angle of 3deg bounded the probable dispersion pattern and so was used for the weight estimate. Next, a finite element impact analysis was performed to determine the required shielding thickness to prevent fuselage penetration. The impact analysis was conducted using an FAA-provided composite blade geometry. The fuselage geometry was based on a medium-sized passenger composite airframe. In the analysis, both the blade and fuselage were assumed to be constructed from a T700S/PR520 triaxially-braided composite architecture. Sufficient test data on T700S/PR520 is available to enable reliable analysis, and also demonstrate its good impact resistance properties. This system was also used in modeling the surrogate blade. The estimated additional weight required for fuselage shielding for a wing- mounted counterrotating open-rotor blade is 236 lb per aircraft. This estimate is based on the shielding material serving the dual use of shielding and fuselage structure. If the shielding material is not used for dual purpose, and is only used for shielding, then the additional weight per aircraft is estimated to be 428 lb. This weight estimate is based upon a number of assumptions that would need to be revised when applying this concept to an actual airplane design. For example, the weight savings that will result when there is no fan blade containment system, manufacturing limitations which may increase the weight where variable thicknesses was assumed, engine placement on the wing versus aft fuselage, etc.

Advanced composite structural concepts and material technologies for primary aircraft structures

NASA Technical Reports Server (NTRS)

Jackson, Anthony

1991-01-01

Structural weight savings using advanced composites have been demonstrated for many years. Most military aircraft today use these materials extensively and Europe has taken the lead in their use in commercial aircraft primary structures. A major inhibiter to the use of advanced composites in the United States is cost. Material costs are high and will remain high relative to aluminum. The key therefore lies in the significant reduction in fabrication and assembly costs. The largest cost in most structures today is assembly. As part of the NASA Advanced Composite Technology Program, Lockheed Aeronautical Systems Company has a contract to explore and develop advanced structural and manufacturing concepts using advanced composites for transport aircraft. Wing and fuselage concepts and related trade studies are discussed. These concepts are intended to lower cost and weight through the use of innovative material forms, processes, structural configurations and minimization of parts. The approach to the trade studies and the downselect to the primary wing and fuselage concepts is detailed. The expectations for the development of these concepts is reviewed.

NASA/HAA Advanced Rotorcraft Technology and Tilt Rotor Workshops. Volume 3: Aerodynamics and Structures Session

NASA Technical Reports Server (NTRS)

1980-01-01

Advanced rotorcraft technology and tilt rotor aircraft were discussed. Rotorcraft performance, acoustics, and vibrations were discussed, as was the use of composite materials in rotorcraft structures. Rotorcraft aerodynamics, specifically the aerodynamic phenomena of a rotating and the aerodynamics of fuselages, was discussed.

Resin transfer molding for advanced composite primary wing and fuselage structures

NASA Technical Reports Server (NTRS)

Markus, Alan

1992-01-01

The stitching and resin transfer molding (RTM) processes developed at Douglas Aircraft Co. are successfully demonstrating significant cost reductions with good damage tolerance properties. These attributes were identified as critical to application of advanced composite materials to commercial aircraft primary structures. The RTM/stitching developments, cost analyses, and test results are discussed of the NASA Advanced Composites Technology program.

Textile composite fuselage structures development

NASA Technical Reports Server (NTRS)

Jackson, Anthony C.; Barrie, Ronald E.; Chu, Robert L.

1993-01-01

Phase 2 of the NASA ACT Contract (NAS1-18888), Advanced Composite Structural Concepts and Materials Technology for Transport Aircraft Structures, focuses on textile technology, with resin transfer molding or powder coated tows. The use of textiles has the potential for improving damage tolerance, reducing cost and saving weight. This program investigates resin transfer molding (RTM), as a maturing technology for high fiber volume primary structures and powder coated tows as an emerging technology with a high potential for significant cost savings and superior structural properties. Powder coated tow technology has promise for significantly improving the processibility of high temperature resins such as polyimides.

Advanced resin systems and 3D textile preforms for low cost composite structures

NASA Technical Reports Server (NTRS)

Shukla, J. G.; Bayha, T. D.

1993-01-01

Advanced resin systems and 3D textile preforms are being evaluated at Lockheed Aeronautical Systems Company (LASC) under NASA's Advanced Composites Technology (ACT) Program. This work is aimed towards the development of low-cost, damage-tolerant composite fuselage structures. Resin systems for resin transfer molding and powder epoxy towpreg materials are being evaluated for processability, performance and cost. Three developmental epoxy resin systems for resin transfer molding (RTM) and three resin systems for powder towpregging are being investigated. Various 3D textile preform architectures using advanced weaving and braiding processes are also being evaluated. Trials are being conducted with powdered towpreg, in 2D weaving and 3D braiding processes for their textile processability and their potential for fabrication in 'net shape' fuselage structures. The progress in advanced resin screening and textile preform development is reviewed here.

Advanced Technology Composite Fuselage-Structural Performance

NASA Technical Reports Server (NTRS)

Walker, T. H.; Minguet, P. J.; Flynn, B. W.; Carbery, D. J.; Swanson, G. D.; Ilcewicz, L. B.

1997-01-01

Boeing is studying the technologies associated with the application of composite materials to commercial transport fuselage structure under the NASA-sponsored contracts for Advanced Technology Composite Aircraft Structures (ATCAS) and Materials Development Omnibus Contract (MDOC). This report addresses the program activities related to structural performance of the selected concepts, including both the design development and subsequent detailed evaluation. Design criteria were developed to ensure compliance with regulatory requirements and typical company objectives. Accurate analysis methods were selected and/or developed where practical, and conservative approaches were used where significant approximations were necessary. Design sizing activities supported subsequent development by providing representative design configurations for structural evaluation and by identifying the critical performance issues. Significant program efforts were directed towards assessing structural performance predictive capability. The structural database collected to perform this assessment was intimately linked to the manufacturing scale-up activities to ensure inclusion of manufacturing-induced performance traits. Mechanical tests were conducted to support the development and critical evaluation of analysis methods addressing internal loads, stability, ultimate strength, attachment and splice strength, and damage tolerance. Unresolved aspects of these performance issues were identified as part of the assessments, providing direction for future development.

Design considerations for composite fuselage structure of commercial transport aircraft

NASA Technical Reports Server (NTRS)

Davis, G. W.; Sakata, I. F.

1981-01-01

The structural, manufacturing, and service and environmental considerations that could impact the design of composite fuselage structure for commercial transport aircraft application were explored. The severity of these considerations was assessed and the principal design drivers delineated. Technical issues and potential problem areas which must be resolved before sufficient confidence is established to commit to composite materials were defined. The key issues considered are: definition of composite fuselage design specifications, damage tolerance, and crashworthiness.

Detailed design of a lattice composite fuselage structure by a mixed optimization method

NASA Astrophysics Data System (ADS)

Liu, D.; Lohse-Busch, H.; Toropov, V.; Hühne, C.; Armani, U.

2016-10-01

In this article, a procedure for designing a lattice fuselage barrel is developed. It comprises three stages: first, topology optimization of an aircraft fuselage barrel is performed with respect to weight and structural performance to obtain the conceptual design. The interpretation of the optimal result is given to demonstrate the development of this new lattice airframe concept for the fuselage barrel. Subsequently, parametric optimization of the lattice aircraft fuselage barrel is carried out using genetic algorithms on metamodels generated with genetic programming from a 101-point optimal Latin hypercube design of experiments. The optimal design is achieved in terms of weight savings subject to stability, global stiffness and strain requirements, and then verified by the fine mesh finite element simulation of the lattice fuselage barrel. Finally, a practical design of the composite skin complying with the aircraft industry lay-up rules is presented. It is concluded that the mixed optimization method, combining topology optimization with the global metamodel-based approach, allows the problem to be solved with sufficient accuracy and provides the designers with a wealth of information on the structural behaviour of the novel anisogrid composite fuselage design.

Damage criticality and inspection concerns of composite-metallic aircraft structures under blunt impact

NASA Astrophysics Data System (ADS)

Zou, D.; Haack, C.; Bishop, P.; Bezabeh, A.

2015-04-01

Composite aircraft structures such as fuselage and wings are subject to impact from many sources. Ground service equipment (GSE) vehicles are regarded as realistic sources of blunt impact damage, where the protective soft rubber is used. With the use of composite materials, blunt impact damage is of special interest, since potential significant structural damage may be barely visible or invisible on the structure's outer surface. Such impact can result in local or non-local damage, in terms of internal delamination in skin, interfacial delamination between stiffeners and skin, and fracture of internal reinforced component such as stringers and frames. The consequences of these events result in aircraft damage, delays, and financial cost to the industry. Therefore, it is necessary to understand the criticality of damage under this impact and provide reliable recommendations for safety and inspection technologies. This investigation concerns a composite-metallic 4-hat-stiffened and 5-frame panel, designed to represent a fuselage structure panel generic to the new generation of composite aircraft. The test fixtures were developed based on the correlation between finite element analyses of the panel model and the barrel model. Three static tests at certain amount of impact energy were performed, in order to improve the understanding of the influence of the variation in shear ties, and the added rotational stiffness. The results of this research demonstrated low velocity high mass impacts on composite aircraft fuselages beyond 82.1 kN of impact load, which may cause extensive internal structural damage without clear visual detectability on the external skin surface.

Global Cost and Weight Evaluation of Fuselage Side Panel Design Concepts

NASA Technical Reports Server (NTRS)

Polland, D. R.; Finn, S. R.; Griess, K. H.; Hafenrichter, J. L.; Hanson, C. T.; Ilcewicz, L. B.; Metschan, S. L.; Scholz, D. B.; Smith, P. J.

1997-01-01

This report documents preliminary design trades conducted under NASA contracts NAS1 18889 (Advanced Technology Composite Aircraft Structures, ATCAS) and NAS1-19349 (Task 3, Pathfinder Shell Design) for a subsonic wide body commercial aircraft fuselage side panel section utilizing composite materials. Included in this effort were (1) development of two complete design concepts, (2) generation of cost and weight estimates, (3) identification of technical issues and potential design enhancements, and (4) selection of a single design to be further developed. The first design concept featured an open-section stringer stiffened skin configuration while the second was based on honeycomb core sandwich construction. The trade study cost and weight results were generated from comprehensive assessment of each structural component comprising the fuselage side panel section from detail fabrication through airplane final assembly. Results were obtained in three phases: (1) for the baseline designs, (2) after global optimization of the designs, and (3) the results anticipated after detailed design optimization. A critical assessment of both designs was performed to determine the risk associated with each concept, that is the relative probability of achieving the cost and weight projections. Seven critical technical issues were identified as the first step towards side panel detailed design optimization.

Application of Ultrasonic Phased Array Technology to the Detection of Defect in Composite Stiffened-structures

NASA Astrophysics Data System (ADS)

Zhou, Yuan-Qi; Zhan, Li-Hua

2016-05-01

Composite stiffened-structure consists of the skin and stringer has been widely used in aircraft fuselage and wings. The main purpose of the article is to detect the composite material reinforced structure accurately and explore the relationship between defect formation and structural elements or curing process. Based on ultrasonic phased array inspection technology, the regularity of defects in the manufacture of composite materials are obtained, the correlation model between actual defects and nondestructive testing are established. The article find that the forming quality of deltoid area in T-stiffened structure is obviously improved by pre-curing, the defects of hat-stiffened structure are affected by the mandrel. The results show that the ultrasonic phased array inspection technology can be an effectively way for the detection of composite stiffened-structures, which become an important means to control the defects of composite and improve the quality of the product.

The Department of Defense Statement on the Science and Technology Program by Mr. H. Mark Grove, Assistant Deputy Under Secretary of Defense for Research and Advanced Technology Before the Defense Subcommittee of the Committee on Appropriations of the United States House of Representatives, 97th Congress, Second Session,

DTIC Science & Technology

1982-06-16

Technology Intergration (AFTI) Pro- gram, (2) a nonmetallic composite helicopter fuselage, and (3) a new initiative to develop Short Take-Off and...range of passive sonobuoys is being extended and the performance of a high gain extended life deployed horizontal line array is being investigated

Mechanical and analytical screening of braided composites for transport fuselage applications

NASA Technical Reports Server (NTRS)

Fedro, Mark J.; Gunther, Christian; Ko, Frank K.

1991-01-01

The mechanics of materials progress in support of the goal of understanding the application of braided composites in a transport aircraft fuselage are summarized. Composites consisting of both 2-D and 3-D braid patterns are investigated. Both consolidation of commingled graphite/PEEK and resin transfer molding of graphite-epoxy braided composite processes are studied. Mechanical tests were used to examine unnotched tension, open hole tension, compression, compression after impact, in-plane shear, out-of-plane tension, bearing, and crippling. Analytical methods are also developed and applied to predict the stiffness and strengths of test specimens. A preliminary study using the test data and analytical results is performed to assess the applicability of braided composites to a commercial aircraft fuselage.

NASA technology program for future civil air transports

NASA Technical Reports Server (NTRS)

Wright, H. T.

1983-01-01

An assessment is undertaken of the development status of technology, applicable to future civil air transport design, which is currently undergoing conceptual study or testing at NASA facilities. The NASA civil air transport effort emphasizes advanced aerodynamic computational capabilities, fuel-efficient engines, advanced turboprops, composite primary structure materials, advanced aerodynamic concepts in boundary layer laminarization and aircraft configuration, refined control, guidance and flight management systems, and the integration of all these design elements into optimal systems. Attention is given to such novel transport aircraft design concepts as forward swept wings, twin fuselages, sandwich composite structures, and swept blade propfans.

Aerospace Plane Technology, Research and Development Efforts in Europe

DTIC Science & Technology

1991-07-25

to conventional titanium alloys. Moreover, ti- aluminide has one-half the weight of the material previously used at these high temperatures. Real Gas...for the engine’s blades , turbine blisk (a turbine disk with integral blades ), Page 44 GAO/NSIAID-91-194 Aerospace Plane Technology Chapter 3 European...X-30 fuselage section from silicon carbide-reinforced titanium and manufactured an X-30 fuel tank from a graphite-polyamide composite. Although

A global range military transport: The ostrich

NASA Technical Reports Server (NTRS)

Aguiar, John; Booker, Cecilia; Hoffman, Eric; Kramar, James; Manahan, Orlando; Serranzana, Ray; Taylor, Mike

1993-01-01

Studies have shown that there is an increasing need for a global range transport capable of carrying large numbers of troops and equipment to potential trouble spots throughout the world. The Ostrich is a solution to this problem. The Ostrich is capable of carrying 800,000 pounds 6,500 n.m. and returning with 15 percent payload, without refueling. With a technology availability date in 2010 and an initial operating capability of 2015, the aircraft incorporates many advanced technologies including laminar flow control, composite primary structures, and a unique multibody design. By utilizing current technology, such as using McDonnell Douglas C-17 fuselage for the outer fuselages on the Ostrich, the cost for the aircraft was reduced. The cost of the Ostrich per aircraft is $1.2 billion with a direct operating cost of $56,000 per flight hour. The Ostrich will provide a valuable service as a logistical transport capable of rapidly projecting a significant military force or humanitarian aid anywhere in the world.

Evaluation of a metal fuselage panel selectively reinforced with filamentary composites for space shuttle application

NASA Technical Reports Server (NTRS)

Wennhold, W. F.

1974-01-01

The use of high strength and modulus of advanced filamentary composites to reduce the structural weight of aerospace vehicles was investigated. Application of the technology to space shuttle components was the primary consideration. The mechanical properties for the boron/epoxy, graphite/epoxy, and polyimide data are presented. Structural testing of two compression panel components was conducted in a simulated space shuttle thermal environment. Results of the tests are analyzed.

Impact Testing and Simulation of a Crashworthy Composite Fuselage Section with Energy-Absorbing Seats and Dummies

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.

2002-01-01

A 25-ft/s vertical drop test of a composite fuselage section was conducted with two energy-absorbing seats occupied by anthropomorphic dummies to evaluate the crashworthy features of the fuselage section and to determine its interaction with the seats and dummies. The 5-ft. diameter fuselage section consists of a stiff structural floor and an energy-absorbing subfloor constructed of Rohacel foam blocks. The experimental data from this test were analyzed and correlated with predictions from a crash simulation developed using the nonlinear, explicit transient dynamic computer code, MSC.Dytran. The anthropomorphic dummies were simulated using the Articulated Total Body (ATB) code, which is integrated into MSC.Dytran.

Impact Testing and Simulation of a Crashworthy Composite Fuselage Section with Energy-Absorbing Seats and Dummies

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.

2002-01-01

A 25-ft/s vertical drop test of a composite fuselage section was conducted with two energy-absorbing seats occupied by anthropomorphic dummies to evaluate the crashworthy features of the fuselage section and to determine its interaction with the seats and dummies. The 5-ft diameter fuselage section consists of a stiff structural floor and an energy-absorbing subfloor constructed of Rohacel foam blocks. The experimental data from this test were analyzed and correlated with predictions from a crash simulation developed using the nonlinear, explicit transient dynamic computer code, MSC.Dytran. The anthropomorphic dummies were simulated using the Articulated Total Body (ATB) code, which is integrated into MSC.Dytran.

Assessment of Damage Containment Features of a Full-Scale PRSEUS Fuselage Panel Through Test and Teardown

NASA Technical Reports Server (NTRS)

Bergan, Andrew C.; Bakuckas, John G., Jr.; Lovejoy, Andrew E.; Jegley, Dawn C.; Awerbuch, Jonathan; Tan, Tein-Min

2012-01-01

An area that shows promise in enhancing structural integrity of aircraft and aerospace structures is the integrally stitched composite technology. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept developed by Boeing Research and Technology and the National Aeronautics and Space Administration. A joint test program on the assessment of damage containment capabilities of the PRSEUS concept for curved fuselage structures was conducted recently at the Federal Aviation Administration William J. Hughes Technical Center. The panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure load conditions up to fracture, with a through-the-thickness, two-bay notch severing the central stiffener. For the purpose of future progressive failure analysis development and verification, extensive post failure nondestructive and teardown inspections were conducted. Detailed inspections were performed directly ahead of the notch tip where stable damage progression was observed. These examinations showed: 1) extensive delaminations developed ahead of the notch tip, 2) the extent and location of damage, 3) the typical damage mechanisms observed in composites, and 4) the role of stitching and warp-knitting in the failure mechanisms. The objective of this paper is to provide a summary of results from these posttest inspections.

Application study of filamentary composites in a commercial jet aircraft fuselage

NASA Technical Reports Server (NTRS)

Johnson, R. W.; June, R. R.

1972-01-01

A study of applications of filamentary composite materials to aircraft fuselage structure was performed. General design criteria were established and material studies conducted using the 727-200 forebody as the primary structural component. Three design approaches to the use of composites were investigated: uniaxial reinforcement of metal structure, uniaxial and biaxial reinforcement of metal structure, and an all-composite design. Materials application studies for all three concepts were conducted on fuselage shell panels, keel beam, floor beams, floor panels, body frames, fail-safe straps, and window frames. Cost benefit studies were conducted and developmental program costs estimated. On the basis of weight savings, cost effectiveness, developmental program costs, and potential for early application on commercial aircraft, the unaxial design is recommended for a 5-year flight service evaluation program.

Crashworthy Evaluation of a 1/5-Scale Model Composite Fuselage Concept

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.

1999-01-01

A 1/5-scale model composite fuselage concept for light aircraft and rotorcraft has been developed to satisfy structural and flight loads requirements and to satisfy design goals for improved crashworthiness. The 1/5-scale model fuselage consists of a relatively rigid upper section which forms the passenger cabin, a stiff structural floor, and an energy absorbing subfloor which is designed to limit impact forces during a crash event. The focus of the present paper is to describe the crashworthy evaluation of the fuselage concept through impact testing and finite element simulation using the nonlinear, explicit transient dynamic code, MSC/DYTRAN. The energy absorption behavior of two different subfloor configurations was determined through quasi-static crushing tests. For the dynamic evaluation, each subfloor configuration was incorporated into a 1/5-scale model fuselage section, which was impacted at 31 ft/s vertical velocity onto a rigid surface. The experimental data demonstrate that the fuselage section with a foam-filled subfloor configuration satisfied the impact design requirement. In addition, the fuselage section maintained excellent energy absorption behavior for a 31 ft/s vertical drop test with a 15 deg-roll impact attitude. Good correlation was obtained between the experimental data and analytical results for both impact conditions.

Development of High Toughness Sheet and Extruded Products for Airplane Fuselage Structures

NASA Astrophysics Data System (ADS)

Magnusen, P. E.; Mooy, D. C.; Yocum, L. A.; Rioja, R. J.

High specific ultimate strength and high plane stress fracture toughness are primary requirements of aircraft fuselage skins. The performance of alloys/products used in high performance fuselage applications is first reviewed. The specific fracture toughness for products such as 2017-T3, 2024-T3, 2524-T3 and 6013-T6, is discussed as a function of their composition and microstructure. Then the performance of modern Al-Li alloys/products such as 2199 and 2060 sheet and 2099 and 2055 extrusions is examined. It is concluded that the performance of Li containing alloys/products offer significant improvements over non-Li containing conventional fuselage products because of the optimization of strengthening precipitates and grain microstructures. The role of chemical composition on resulting microstructures is discussed.

Comparison of ALE and SPH Simulations of Vertical Drop Tests of a Composite Fuselage Section into Water

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fuchs, Yvonne T.

2008-01-01

Simulation of multi-terrain impact has been identified as an important research area for improved prediction of rotorcraft crashworthiness within the NASA Subsonic Rotary Wing Aeronautics Program on Rotorcraft Crashworthiness. As part of this effort, two vertical drop tests were conducted of a 5-ft-diameter composite fuselage section into water. For the first test, the fuselage section was impacted in a baseline configuration without energy absorbers. For the second test, the fuselage section was retrofitted with a composite honeycomb energy absorber. Both tests were conducted at a nominal velocity of 25-ft/s. A detailed finite element model was developed to represent each test article and water impact was simulated using both Arbitrary Lagrangian Eulerian (ALE) and Smooth Particle Hydrodynamics (SPH) approaches in LS-DYNA, a nonlinear, explicit transient dynamic finite element code. Analytical predictions were correlated with experimental data for both test configurations. In addition, studies were performed to evaluate the influence of mesh density on test-analysis correlation.

Thermal Inspection of a Composite Fuselage Section Using a Fixed Eigenvector Principal Component Analysis Method

NASA Technical Reports Server (NTRS)

Zalameda, Joseph N.; Bolduc, Sean; Harman, Rebecca

2017-01-01

A composite fuselage aircraft forward section was inspected with flash thermography. The fuselage section is 24 feet long and approximately 8 feet in diameter. The structure is primarily configured with a composite sandwich structure of carbon fiber face sheets with a Nomex(Trademark) honeycomb core. The outer surface area was inspected. The thermal data consisted of 477 data sets totaling in size of over 227 Gigabytes. Principal component analysis (PCA) was used to process the data sets for substructure and defect detection. A fixed eigenvector approach using a global covariance matrix was used and compared to a varying eigenvector approach. The fixed eigenvector approach was demonstrated to be a practical analysis method for the detection and interpretation of various defects such as paint thickness variation, possible water intrusion damage, and delamination damage. In addition, inspection considerations are discussed including coordinate system layout, manipulation of the fuselage section, and the manual scanning technique used for full coverage.

Comparison of Hard Surface and Soft Soil Impact Performance of a Crashworthy Composite Fuselage Concept

NASA Technical Reports Server (NTRS)

Sareen, Ashish K.; Sparks, Chad; Mullins, B. R., Jr.; Fasanella, Edwin; Jackson, Karen

2002-01-01

A comparison of the soft soil and hard surface impact performance of a crashworthy composite fuselage concept has been performed. Specifically, comparisons of the peak acceleration values, pulse duration, and onset rate at specific locations on the fuselage were evaluated. In a prior research program, the composite fuselage section was impacted at 25 feet per second onto concrete at the Impact Dynamics Research Facility (IDRF) at NASA Langley Research Center. A soft soil test was conducted at the same impact velocity as a part of the NRTC/RITA Crashworthy and Energy Absorbing Structures project. In addition to comparisons of soft soil and hard surface test results, an MSC. Dytran dynamic finite element model was developed to evaluate the test analysis correlation. In addition, modeling parameters and techniques affecting test analysis correlation are discussed. Once correlated, the analytical methodology will be used in follow-on work to evaluate the specific energy absorption of various subfloor concepts for improved crash protection during hard surface and soft soil impacts.

An overview of the crash dynamics failure behavior of metal and composite aircraft structures

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Boitnott, Richard L.; Fasanella, Edwin L.; Jones, Lisa E.

1991-01-01

An overview of failure behavior results is presented from some of the crash dynamics research conducted with concepts of aircraft elements and substructure not necessarily designed or optimized for energy absorption or crash loading considerations. Experimental and analytical data are presented that indicate some general trends in the failure behavior of a class of composite structures that includes fuselage panels, individual fuselage sections, fuselage frames, skeleton subfloors with stringers and floor beams without skin covering, and subfloors with skin added to the frame stringer structure. Although the behavior is complex, a strong similarity in the static/dynamic failure behavior among these structures is illustrated through photographs of the experimental results and through analytical data of generic composite structural models.

National Aerospace Plane Integrated Fuselage/Cryotank Risk Reduction program

NASA Astrophysics Data System (ADS)

Dayton, K. E.

1993-06-01

The principal objectives and results of the National Aerospace Plane (NASP) Integrated Risk Reduction program are briefly reviewed. The program demonstrated the feasibility of manufacturing lightweight advanced composite materials for single-stage-to-orbit hypersonic flight vehicle applications. A series of combined load simulation tests (thermal, mechanical, and cryogenic) demonstrated proof of concept performance for an all unlined composite cryogenic fuel tank with flat end bulkheads and a high-temperature thin-shell advanced composite fuselage. Temperatures of the fuselage were as high as 1300 F, with 100 percent bending and shear loads applied to the tank while filled with 850 gallons of cryogenic fluid hydrogen (-425 F). Leak rates measured on and around the cryotank shell and bulkheads were well below acceptable levels.

Critical joints in large composite primary aircraft structures. Volume 2: Technology demonstration test report

NASA Technical Reports Server (NTRS)

Bunin, Bruce L.

1985-01-01

A program was conducted to develop the technology for critical structural joints in composite wing structure that meets all the design requirements of a 1990 commercial transport aircraft. The results of four large composite multirow bolted joint tests are presented. The tests were conducted to demonstrate the technology for critical joints in highly loaded composite structure and to verify the analytical methods that were developed throughout the program. The test consisted of a wing skin-stringer transition specimen representing a stringer runout and skin splice on the wing lower surface at the side of the fuselage attachment. All tests were static tension tests. The composite material was Toray T-300 fiber with Ciba-Geigy 914 resin in 10 mil tape form. The splice members were metallic, using combinations of aluminum and titanium. Discussions are given of the test article, instrumentation, test setup, test procedures, and test results for each of the four specimens. Some of the analytical predictions are also included.

Simulation and analysis of airborne antenna radiation patterns

NASA Technical Reports Server (NTRS)

Kim, J. J.; Burnside, Walter D.

1984-01-01

The objective is to develop an accurate and efficient analytic solution for predicting high frequency radiation patterns of fuselage-mounted airborne antennas. This is an analytic study of airborne antenna patterns using the Uniform Geometrical Theory of Diffraction (UTD). The aircraft is modeled in its most basic form so that the solution is applicable to general-type aircraft. The fuselage is modeled as a perfectly conducting composite ellipsoid; whereas, the wings, stabilizers, nose, fuel tanks, and engines, are simulated as perfectly conducting flat plates that can be attached to the fuselage and/or to each other. The composite-ellipsoid fuselage model is necessary to successfully simulate the wide variety of real world fuselage shapes. Since the antenna is mounted on the fuselage, it has a dominant effect on the resulting radiation pattern so it must be simulated accurately, especially near the antenna. Various radiation patterns are calculated for commercial, private, and military aircraft, and the Space Shuttle Orbiter. The application of this solution to numerous practical airborne antenna problems illustrates its versatility and design capability. In most cases, the solution accuracy is verified by the comparisons between the calculated and measured data.

Structural Configuration Systems Analysis for Advanced Aircraft Fuselage Concepts

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek; Welstead, Jason R.; Quinlan, Jesse R.; Guynn, Mark D.

2016-01-01

Structural configuration analysis of an advanced aircraft fuselage concept is investigated. This concept is characterized by a double-bubble section fuselage with rear mounted engines. Based on lessons learned from structural systems analysis of unconventional aircraft, high-fidelity finite-element models (FEM) are developed for evaluating structural performance of three double-bubble section configurations. Structural sizing and stress analysis are applied for design improvement and weight reduction. Among the three double-bubble configurations, the double-D cross-section fuselage design was found to have a relatively lower structural weight. The structural FEM weights of these three double-bubble fuselage section concepts are also compared with several cylindrical fuselage models. Since these fuselage concepts are different in size, shape and material, the fuselage structural FEM weights are normalized by the corresponding passenger floor area for a relative comparison. This structural systems analysis indicates that an advanced composite double-D section fuselage may have a relative structural weight ratio advantage over a conventional aluminum fuselage. Ten commercial and conceptual aircraft fuselage structural weight estimates, which are empirically derived from the corresponding maximum takeoff gross weight, are also presented and compared with the FEM- based estimates for possible correlation. A conceptual full vehicle FEM model with a double-D fuselage is also developed for preliminary structural analysis and weight estimation.

Preparation for Testing a Multi-Bay Box Subjected to Combined Loads

NASA Technical Reports Server (NTRS)

Rouse, Marshall; Jegley, Dawn

2015-01-01

The COmbined Loads Test System (COLTS) facility at NASA Langley Research Center provides a test capability to help develop validated structures technologies. The test machine was design to accommodate a range of fuselage structures and wing sections and subject them to both quasistatic and cyclic loading conditions. The COLTS facility is capable of testing fuselage barrels up to 4.6 m in diameter and 13.7 m long with combined mechanical, internal pressure, and thermal loads. The COLTS facility is currently being prepared to conduct a combined mechanical and pressure loading for a multi-bay pressure box to experimentally verify the structural performance of a composite structure which is 9.1 meters long and representative of a section of a hybrid wing body fuselage section in support of the Environmentally Responsible Aviation Project at NASA. This paper describes development of the multi-bay pressure box test using the COLTS facility. The multi-bay test article will be subjected to mechanical loads and internal pressure loads up to design ultimate load. Mechanical and pressure loads will be applied independently in some tests and simultaneously in others.

Impact Testing and Simulation of Composite Airframe Structures

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Littell, Justin D.; Horta, Lucas G.; Annett, Martin S.; Fasanella, Edwin L.; Seal, Michael D., II

2014-01-01

Dynamic tests were performed at NASA Langley Research Center on composite airframe structural components of increasing complexity to evaluate their energy absorption behavior when subjected to impact loading. A second objective was to assess the capabilities of predicting the dynamic response of composite airframe structures, including damage initiation and progression, using a state-of-the-art nonlinear, explicit transient dynamic finite element code, LS-DYNA. The test specimens were extracted from a previously tested composite prototype fuselage section developed and manufactured by Sikorsky Aircraft Corporation under the US Army's Survivable Affordable Repairable Airframe Program (SARAP). Laminate characterization testing was conducted in tension and compression. In addition, dynamic impact tests were performed on several components, including I-beams, T-sections, and cruciform sections. Finally, tests were conducted on two full-scale components including a subfloor section and a framed fuselage section. These tests included a modal vibration and longitudinal impact test of the subfloor section and a quasi-static, modal vibration, and vertical drop test of the framed fuselage section. Most of the test articles were manufactured of graphite unidirectional tape composite with a thermoplastic resin system. However, the framed fuselage section was constructed primarily of a plain weave graphite fabric material with a thermoset resin system. Test data were collected from instrumentation such as accelerometers and strain gages and from full-field photogrammetry.

Advanced composites structural concepts and materials technologies for primary aircraft structures: Design/manufacturing concept assessment

NASA Technical Reports Server (NTRS)

Chu, Robert L.; Bayha, Tom D.; Davis, HU; Ingram, J. ED; Shukla, Jay G.

1992-01-01

Composite Wing and Fuselage Structural Design/Manufacturing Concepts have been developed and evaluated. Trade studies were performed to determine how well the concepts satisfy the program goals of 25 percent cost savings, 40 percent weight savings with aircraft resizing, and 50 percent part count reduction as compared to the aluminum Lockheed L-1011 baseline. The concepts developed using emerging technologies such as large scale resin transfer molding (RTM), automatic tow placed (ATP), braiding, out-of-autoclave and automated manufacturing processes for both thermoset and thermoplastic materials were evaluated for possible application in the design concepts. Trade studies were used to determine which concepts carry into the detailed design development subtask.

Cost-efficient manufacturing of composite structures

NASA Technical Reports Server (NTRS)

Freeman, W. Tom; Davis, John G.; Johnston, Norman J.

1991-01-01

The Advanced Composites Technology (ACT) program is seeking research breakthroughs that will allow structures made of graphite epoxy materials to replace metals in the wings and fuselages of future aircrafts. NASA's goals are to reduce acquisition cost by 20 to 25 percent, structural weight for a resized aircraft by 40 to 50 percent, and the number of parts by half compared to current production aluminum aircraft. The innovative structural concepts, materials, and fabrication techniques emerging from the ACT program are described, and the relationship between aerospace developments and industrial, commercial, and sporting goods applications are discussed.

Automated composite ellipsoid modelling for high frequency GTD analysis

NASA Technical Reports Server (NTRS)

Sze, K. Y.; Rojas, R. G.; Klevenow, F. T.; Scheick, J. T.

1991-01-01

The preliminary results of a scheme currently being developed to fit a composite ellipsoid to the fuselage of a helicopter in the vicinity of the antenna location are discussed under the assumption that the antenna is mounted on the fuselage. The parameters of the close-fit composite ellipsoid would then be utilized as inputs into NEWAIR3, a code programmed in FORTRAN 77 for high frequency Geometrical Theory of Diffraction (GTD) Analysis of the radiation of airborne antennas.

Blended-Wing-Body (BWB) Fuselage Structural Design for Weight Reduction

NASA Technical Reports Server (NTRS)

Mukhopadhyay, V.

2005-01-01

Structural analysis and design of efficient pressurized fuselage configurations for the advanced Blended-Wing-Body (BWB) flight vehicle is a challenging problem. Unlike a conventional cylindrical pressurized fuselage, stress level in a box type BWB fuselage is an order of magnitude higher, because internal pressure primarily results in bending stress instead of skin-membrane stress. In addition, resulting deformation of aerodynamic surface could significantly affect performance advantages provided by lifting body. The pressurized composite conformal multi-lobe tanks of X-33 type space vehicle also suffered from similar problem. In the earlier BWB design studies, Vaulted Ribbed Shell (VLRS), Flat Ribbed Shell (FRS); Vaulted shell Honeycomb Core (VLHC) and Flat sandwich shell Honeycomb Core (FLHC) concepts were studied. The flat and vaulted ribbed shell concepts were found most efficient. In a recent study, a set of composite sandwich panel and cross-ribbed panel were analyzed. Optimal values of rib and skin thickness, rib spacing, and panel depth were obtained for minimal weight under stress and buckling constraints. In addition, a set of efficient multi-bubble fuselage (MBF) configuration concept was developed. The special geometric configuration of this concept allows for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls, while the outer-ribbed shell prevents buckling due to external resultant compressive loads. The initial results from these approximate finite element analyses indicate progressively lower maximum stresses and deflections compared to the earlier study. However, a relative comparison of the FEM weight per unit floor area of the segment unit indicates that the unit weights are still relatively higher that the conventional B777 type cylindrical or A380 type elliptic fuselage design. Due to the manufacturing concern associated with multi-bubble fuselage, a Y braced box-type fuselage alternative with special resin-film injected (RFI) stitched carbon composite with foam-core was designed by Boeing under a NASA research contract for the 480 passenger version. It is shown that this configuration can be improved to a modified multi-bubble fuselage which has better stress distribution, for same material and dimension.

Experimental investigation of the crashworthiness of scaled composite sailplane fuselages

NASA Technical Reports Server (NTRS)

Kampf, Karl-Peter; Crawley, Edward F.; Hansman, R. John, Jr.

1989-01-01

The crash dynamics and energy absorption of composite sailplane fuselage segments undergoing nose-down impact were investigated. More than 10 quarter-scale structurally similar test articles, typical of high-performance sailplane designs, were tested. Fuselages segments were fabricated of combinations of fiberglass, graphite, Kevlar, and Spectra fabric materials. Quasistatic and dynamic tests were conducted. The quasistatic tests were found to replicate the strain history and failure modes observed in the dynamic tests. Failure modes of the quarter-scale model were qualitatively compared with full-scale crash evidence and quantitatively compared with current design criteria. By combining material and structural improvements, substantial increases in crashworthiness were demonstrated.

Test and analysis results for composite transport fuselage and wing structures

NASA Technical Reports Server (NTRS)

Deaton, Jerry W.; Kullerd, Susan M.; Madan, Ram C.; Chen, Victor L.

1992-01-01

Automated tow placement (ATP) and stitching of dry textile composite preforms followed by resin transfer molding (RTM) are being studied as cost effective manufacturing processes for obtaining damage tolerant fuselage and wing structures for transport aircraft. Data are presented to assess the damage tolerance of ATP and RTM fuselage elements with stitched-on stiffeners from compression tests of impacted three J-stiffened panels and from stiffener pull-off tests. Data are also presented to assess the damage tolerance of RTM wing elements which had stitched skin and stiffeners from impacted single stiffener and three blade stiffened compression tests and stiffener pull-off tests.

Multi-Terrain Impact Testing and Simulation of a Composite Energy Absorbing Fuselage Section

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Lyle, Karen H.; Sparks, Chad E.; Sareen, Ashish K.

2007-01-01

Comparisons of the impact performance of a 5-ft diameter crashworthy composite fuselage section were investigated for hard surface, soft soil, and water impacts. The fuselage concept, which was originally designed for impacts onto a hard surface only, consisted of a stiff upper cabin, load bearing floor, and an energy absorbing subfloor. Vertical drop tests were performed at 25-ft/s onto concrete, soft-soil, and water at NASA Langley Research Center. Comparisons of the peak acceleration values, pulse durations, and onset rates were evaluated for each test at specific locations on the fuselage. In addition to comparisons of the experimental results, dynamic finite element models were developed to simulate each impact condition. Once validated, these models can be used to evaluate the dynamic behavior of subfloor components for improved crash protection for hard surface, soft soil, and water impacts.

Multi-Terrain Impact Testing and Simulation of a Composite Energy Absorbing Fuselage Section

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Lyle, Karen H.; Sparks, Chad E.; Sareen, Ashish K.

2004-01-01

Comparisons of the impact performance of a 5-ft diameter crashworthy composite fuselage section were investigated for hard surface, soft soil, and water impacts. The fuselage concept, which was originally designed for impacts onto a hard surface only, consisted of a stiff upper cabin, load bearing floor, and an energy absorbing subfloor. Vertical drop tests were performed at 25-ft/s onto concrete, soft-soil, and water at NASA Langley Research Center. Comparisons of the peak acceleration values, pulse durations, and onset rates were evaluated for each test at specific locations on the fuselage. In addition to comparisons of the experimental results, dynamic finite element models were developed to simulate each impact condition. Once validated, these models can be used to evaluate the dynamic behavior of subfloor components for improved crash protection for hard surface, soft soil, and water impacts.

Test and analysis results for composite transport fuselage and wing structures

NASA Technical Reports Server (NTRS)

Deaton, Jerry W.; Kullerd, Susan M.; Madan, Ram C.; Chen, Victor L.

1992-01-01

Automated tow placement (ATP) and stitching of dry textile composite preforms followed by resin transfer molding (RTM) are being investigated by researchers at NASA LaRC and Douglas Aircraft Company as cost-effective manufacturing processes for obtaining damage tolerant fuselage and wing structures for transport aircraft. The Douglas work is being performed under a NASA contract entitled 'Innovative Composites Aircraft Primary Structures (ICAPS)'. Data are presented in this paper to assess the damage tolerance of ATP and RTM fuselage elements with stitched-on stiffeners from compression tests of impacted three-J-stiffened panels and from stiffener pull-off tests. Data are also presented to assess the damage tolerance of RTM wing elements which had stitched skin and stiffeners from impacted single stiffener and three blade-stiffened compression tests and stiffener pull-off tests.

Novel Composites for Wing and Fuselage Applications

NASA Technical Reports Server (NTRS)

Suarez, J. A.; Buttitta, C.

1996-01-01

Design development was successfully completed for textile preforms with continuous cross-stiffened epoxy panels with cut-outs. The preforms developed included 3-D angle interlock weaving of graphite structural fibers impregnated by resin film infiltration (RFI) and shown to be structurally suitable under conditions requiring minimum acquisition costs. Design guidelines/analysis methodology for such textile structures are given. The development was expanded to a fuselage side-panel component of a subsonic commercial airframe and found to be readily scalable. The successfully manufactured panel was delivered to NASA Langley for biaxial testing. This report covers the work performed under Task 3 -- Cross-Stiffened Subcomponent; Task 4 -- Design Guidelines/Analysis of Textile-Reinforced Composites; and Task 5 -- Integrally Woven Fuselage Panel.

Advanced emergency openings for commercial aircraft

NASA Technical Reports Server (NTRS)

Bement, L. J.; Schimmel, M. L.

1985-01-01

Explosively actuated openings in composite panels are proposed to enhance passenger survivability within commercial aircraft by providing improvements in emergency openings, fuselage venting, and fuel dump. The concept is to embed a tiny, highly stable explosive cord in the periphery of a load-carrying composite panel; on initiation of the cord, the panel is fractured to create a well-defined opening. The panel would be installed in the sides of the fuselage for passenger egress, in the top of the fuselage for smoke venting, and in the bottoms of the fuel cells for fuel dump. Described are the concerns with the use of explosive systems, safety improvements, advantages, experimental results, and recommended approach to gain acceptance and develop this concept.

Simulating the Impact Response of Full-Scale Composite Airframe Structures

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Littell, Justin D.; Seal, Michael D.

2012-01-01

NASA Langley Research Center obtained a composite helicopter cabin structure in 2010 from the US Army's Survivable Affordable Repairable Airframe Program (SARAP) that was fabricated by Sikorsky Aircraft Corporation. The cabin had been subjected to a vertical drop test in 2008 to evaluate a tilting roof concept to limit the intrusion of overhead masses into the fuselage cabin. Damage to the cabin test article was limited primarily to the roof. Consequently, the roof area was removed and the remaining structure was cut into test specimens including a large subfloor section and a forward framed fuselage section. In 2011, NASA and Sikorsky entered into a cooperative research agreement to study the impact responses of composite airframe structures and to evaluate the capabilities of the explicit transient dynamic finite element code, LS-DYNA®, to simulate these responses including damage initiation and progressive failure. Most of the test articles were manufactured of graphite unidirectional tape composite with a thermoplastic resin system. However, the framed fuselage section was constructed primarily of a plain weave graphite fabric material with a thermoset resin system. Test data were collected from accelerometers and full-field photogrammetry. The focus of this paper will be to document impact testing and simulation results for the longitudinal impact of the subfloor section and the vertical drop test of the forward framed fuselage section.

Development of stitching reinforcement for transport wing panels

NASA Technical Reports Server (NTRS)

Palmer, Raymond J.; Dow, Marvin B.; Smith, Donald L.

1991-01-01

The NASA Advanced Composites Technology (ACT) program has the objective of providing the technology required to obtain the full benefit of weight savings and performance improvements offered by composite primary aircraft structures. Achieving the objective is dependent upon developing composite materials and structures which are damage tolerant and economical to manufacture. Researchers are investigating stitching reinforcement combined with resin transfer molding to produce materials meeting the ACT program objective. Research is aimed at materials, processes, and structural concepts for application in both transport wings and fuselages, but the emphasis to date has been on wing panels. Empirical guidelines are being established for stitching reinforcement in structures designed for heavy loads. Results are presented from evaluation tests investigating stitching types, threads, and density (penetrations per square inch). Tension strength, compression strength, and compression after impact data are reported.

Correlation of AH-1G airframe flight vibration data with a coupled rotor-fuselage analysis

NASA Technical Reports Server (NTRS)

Sangha, K.; Shamie, J.

1990-01-01

The formulation and features of the Rotor-Airframe Comprehensive Analysis Program (RACAP) is described. The analysis employs a frequency domain, transfer matrix approach for the blade structural model, a time domain wake or momentum theory aerodynamic model, and impedance matching for rotor-fuselage coupling. The analysis is applied to the AH-1G helicopter, and a correlation study is conducted on fuselage vibration predictions. The purpose of the study is to evaluate the state-of-the-art in helicopter fuselage vibration prediction technology. The fuselage vibration predicted using RACAP are fairly good in the vertical direction and somewhat deficient in the lateral/longitudinal directions. Some of these deficiencies are traced to the fuselage finite element model.

Finite Element Analysis of Composite Aircraft Fuselage Frame

NASA Astrophysics Data System (ADS)

Dandekar, Aditya Milind

Composites have been introduced in aircraft industries, for their stronger, stiffer, and lighter properties than their metal-alloys counterparts. The general purpose of an aircraft is to transport commercial or military payload. Aircraft frames primarily maintains the shape of fuselage and prevent instability of the structure. Fuselage is similar as wing in construction which consist of longitudinal elements (longerons and stringers), transverse elements (frames and bulkheads) and its external skin. The fuselage is subjected to forces such as the wing reactions, landing gear reaction, empennage reaction, inertia forces subjected due to size and weight, internal pressure forces due to high altitude. Frames also ensure fail-safe design against skin crack propagation due to hoops stress. Ideal fuselage frames cross section is often circular ring shape with a frame cap of Z section. They are mainly made up of light alloy commonly used is aluminium alloys such as Al-2024, Al-7010, Al-7050, Al-7175. Aluminium alloys have good strength to density ratios in compression and bending of thin plate. A high strength to weight ratio of composite materials can result in a lighter aircraft structure or better safety factor. This research focuses on analysis of fuselage frame under dynamic load condition with change in material. Composites like carbon fibre reinforced plastics [CFRP] and glass fibre reinforced plastics [GFRP] are compared with traditional aluminium alloy Al-7075. The frame is subjected to impact test by dropping it at a velocity of 30 ft. / secs from a height of 86 inch from its centre of gravity. These parameters are considered in event of failure of landing gear, and an aircraft is subject to belly landing or gear-up landing. The shear flow is calculated due to impact force which acts in radial direction. The frame is analysed under static structural and explicit dynamic load conditions. Geometry is created in ANSYS Design Modeler. Analysis setup is created using ANSYS Explicit Dynamic (AUTODYN) and ANSYS Composite PrepPost (ACP-Pre) modules. Shear flow and Stress Flow equations are solved by generating a MATLAB code.

Compression Strength of Composite Primary Structural Components

NASA Technical Reports Server (NTRS)

Johnson, Eric R.; Starnes, James H., Jr. (Technical Monitor)

2000-01-01

The focus of research activities under NASA Grant NAG-1-2035 was the response and failure of thin-walled structural components. The research is applicable to the primary load carrying structure of flight vehicles, with particular emphasis on fuselage and wing'structure. Analyses and tests were performed that are applicable to the following structural components an aft pressure bulkhead, or a composite pressure dome, pressure cabin damage containment, and fuselage frames subject to crash-type loads.

Structural analysis of Aircraft fuselage splice joint

NASA Astrophysics Data System (ADS)

Udaya Prakash, R.; Kumar, G. Raj; Vijayanandh, R.; Senthil Kumar, M.; Ramganesh, T.

2016-09-01

In Aviation sector, composite materials and its application to each component are one of the prime factors of consideration due to the high strength to weight ratio, design flexibility and non-corrosive so that the composite materials are widely used in the low weight constructions and also it can be treated as a suitable alternative to metals. The objective of this paper is to estimate and compare the suitability of a composite skin joint in an aircraft fuselage with different joints by simulating the displacement, normal stress, vonmises stress and shear stress with the help of numerical solution methods. The reference Z-stringer component of this paper is modeled by CATIA and numerical simulation is carried out by ANSYS has been used for splice joint presents in the aircraft fuselage with three combinations of joints such as riveted joint, bonded joint and hybrid joint. Nowadays the stringers are using to avoid buckling of fuselage skin, it has joined together by rivets and they are connected end to end by splice joint. Design and static analysis of three-dimensional models of joints such as bonded, riveted and hybrid are carried out and results are compared.

Impact damage resistance of composite fuselage structure, part 1

NASA Technical Reports Server (NTRS)

Dost, E. F.; Avery, W. B.; Ilcewicz, L. B.; Grande, D. H.; Coxon, B. R.

1992-01-01

The impact damage resistance of laminated composite transport aircraft fuselage structures was studied experimentally. A statistically based designed experiment was used to examine numerous material, laminate, structural, and extrinsic (e.g., impactor type) variables. The relative importance and quantitative measure of the effect of each variable and variable interactions on responses including impactor dynamic response, visibility, and internal damage state were determined. The study utilized 32 three-stiffener panels, each with a unique combination of material type, material forms, and structural geometry. Two manufacturing techniques, tow placement and tape lamination, were used to build panels representative of potential fuselage crown, keel, and lower side-panel designs. Various combinations of impactor variables representing various foreign-object-impact threats to the aircraft were examined. Impacts performed at different structural locations within each panel (e.g., skin midbay, stiffener attaching flange, etc.) were considered separate parallel experiments. The relationship between input variables, measured damage states, and structural response to this damage are presented including recommendations for materials and impact test methods for fuselage structure.

Standard test evaluation of graphite fiber/resin matrix composite materials for improved toughness

NASA Technical Reports Server (NTRS)

Chapman, Andrew J.

1984-01-01

Programs sponsored by NASA with the commercial transport manufacturers to develop a technology data base are required to design and build composite wing and fuselage structures. To realize the full potential of composite structures in these strength critical designs, material systems having improved ductility and interlaminar toughness are being sought. To promote systematic evaluation of new materials, NASA and the commercial transport manufacturers have selected and standardized a set of five common tests. These tests evaluate open hole tension and compression performance, compression performance after impact at an energy level of 20 ft-lb, and resistance to delamination. Ten toughened resin matrix/graphite fiber composites were evaluated using this series of tests, and their performance is compared with a widely used composite system.

Probabilistic evaluation of fuselage-type composite structures

NASA Technical Reports Server (NTRS)

Shiao, Michael C.; Chamis, Christos C.

1992-01-01

A methodology is developed to computationally simulate the uncertain behavior of composite structures. The uncertain behavior includes buckling loads, natural frequencies, displacements, stress/strain etc., which are the consequences of the random variation (scatter) of the primitive (independent random) variables in the constituent, ply, laminate and structural levels. This methodology is implemented in the IPACS (Integrated Probabilistic Assessment of Composite Structures) computer code. A fuselage-type composite structure is analyzed to demonstrate the code's capability. The probability distribution functions of the buckling loads, natural frequency, displacement, strain and stress are computed. The sensitivity of each primitive (independent random) variable to a given structural response is also identified from the analyses.

Development of stitched/RTM composite primary structures

NASA Technical Reports Server (NTRS)

Kullerd, Susan M.; Dow, Marvin B.

1992-01-01

The goal of the NASA Advanced Composites Technology (ACT) Program is to provide the technology required to gain the full benefit of weight savings and performance offered by composite primary structures. Achieving the goal is dependent on developing composite materials and structures which are damage tolerant and economical to manufacture. Researchers at NASA LaRC and Douglas Aircraft Company are investigating stitching reinforcement combined with resin transfer molding (RTM) to create structures meeting the ACT program goals. The Douglas work is being performed under a NASA contract entitled Innovative Composites Aircraft Primary Structures (ICAPS). The research is aimed at materials, processes and structural concepts for application in both transport wings and fuselages. Empirical guidelines are being established for stitching reinforcement in primary structures. New data are presented in this paper for evaluation tests of thick (90-ply) and thin (16-ply) stitched laminates, and from selection tests of RTM composite resins. Tension strength, compression strength and post-impact compression strength data are reported. Elements of a NASA LaRC program to expand the science base for stitched/RTM composites are discussed.

Aeroelastic tailoring and structural optimization of joined-wing configurations

NASA Astrophysics Data System (ADS)

Lee, Dong-Hwan

2002-08-01

Methodology for integrated aero-structural design was developed using formal optimization. ASTROS (Automated STRuctural Optimization System) was used as an analyzer and an optimizer for performing joined-wing weight optimization with stress, displacement, cantilever or body-freedom flutter constraints. As a pre/post processor, MATLAB was used for generating input file of ASTROS and for displaying the results of the ASTROS. The effects of the aeroelastic constraints on the isotropic and composite joined-wing weight were examined using this developed methodology. The aeroelastic features of a joined-wing aircraft were examined using both the Rayleigh-Ritz method and a finite element based aeroelastic stability and weight optimization procedure. Aircraft rigid-body modes are included to analyze of body-freedom flutter of the joined-wing aircraft. Several parametric studies were performed to determine the most important parameters that affect the aeroelastic behavior of a joined-wing aircraft. The special feature of a joined-wing aircraft is body-freedom flutter involving frequency interaction of the first elastic mode and the aircraft short period mode. In most parametric study cases, the body-freedom flutter speed was less than the cantilever flutter speed that is independent of fuselage inertia. As fuselage pitching moment of inertia was increased, the body-freedom flutter speed increased. When the pitching moment of inertia reaches a critical value, transition from body-freedom flutter to cantilever flutter occurred. The effects of composite laminate orientation on the front and rear wings of a joined-wing configuration were studied. An aircraft pitch divergence mode, which occurred because of forward movement of center of pressure due to wing deformation, was found. Body-freedom flutter and cantilever-like flutter were also found depending on combination of front and rear wing ply orientations. Optimized wing weight behaviors of the planar and non-planar configurations with isotropic and composite materials were investigated. Wing weight optimization of the composite joined-wing result in less weight compared to the metallic wing. Fuselage flexibility affects joined-wing flutter characteristics. Elastic mode shapes of the wing were affected by fuselage deformation and change the flutter speeds compared to the rigid fuselage. Body-freedom flutter speeds decrease as fuselage flexibility increases. Optimum wing weights increase as fuselage flexibility increases. Flutter analysis of a box wing configuration investigated the effects of center of gravity location and pitch moment of inertia on flutter speed.

Progress on Developing Sonic Infrared Imaging for Defect Detection in Composite Structures

NASA Astrophysics Data System (ADS)

Han, Xiaoyan; He, Qi; Li, Wei; Newaz, Golam; Favro, Lawrence D.; Thomas, Robert L.

2010-02-01

At last year's QNDE conference, we presented our development of Sonic IR imaging technology in metal structures, with results from both experimental studies and theoretical computing. In the latest aircraft designs, such as the B787 from Boeing, composites have become the major materials in structures such as the fuselage and wings. This is in contrast to composites' use only in auxiliary components such as flaps and spoilers in the past. With today's advanced technology of fabrication, it is expected the new materials can be put in use in even more aircraft structures due to its light weight and high strength (high strength-to-weight ratio), high specific stiffness, tailorability of properties, design flexibility etc. Especially, with increases in fuel cost, reducing the aircraft's body weight becomes more and more appealing. In this presentation, we describe the progress on our development of Sonic IR imaging for aircraft composite structures. In particular, we describe the some unexpected results discovered while modeling delaminations. These results were later experimentally verified with an engineered delamination.

Use of nondestructive inspection and fiber optic sensing for damage characterization in carbon fiber fuselage structure

NASA Astrophysics Data System (ADS)

Neidigk, Stephen; Le, Jacqui; Roach, Dennis; Duvall, Randy; Rice, Tom

2014-04-01

To investigate a variety of nondestructive inspection technologies and assess impact damage characteristics in carbon fiber aircraft structure, the FAA Airworthiness Assurance Center, operated by Sandia National Labs, fabricated and impact tested two full-scale composite fuselage sections. The panels are representative of structure seen on advanced composite transport category aircraft and measured approximately 56"x76". The structural components consisted of a 16 ply skin, co-cured hat-section stringers, fastened shear ties and frames. The material used to fabricate the panels was T800 unidirectional pre-preg (BMS 8-276) and was processed in an autoclave. Simulated hail impact testing was conducted on the panels using a high velocity gas gun with 2.4" diameter ice balls in collaboration with the University of California San Diego (UCSD). Damage was mapped onto the surface of the panels using conventional, hand deployed ultrasonic inspection techniques, as well as more advanced ultrasonic and resonance scanning techniques. In addition to the simulated hail impact testing performed on the panels, 2" diameter steel tip impacts were used to produce representative impact damage which can occur during ground maintenance operations. The extent of impact damage ranges from less than 1 in2 to 55 in2 of interply delamination in the 16 ply skin. Substructure damage on the panels includes shear tie cracking and stringer flange disbonding. It was demonstrated that the fiber optic distributed strain sensing system is capable of detecting impact damage when bonded to the backside of the fuselage.

Heavy Class Helicopter Fuselage Model Drag Reduction by Active Flow Control Systems

NASA Astrophysics Data System (ADS)

De Gregorio, F.

2017-08-01

A comprehensive experimental investigation of helicopter blunt fuselage drag reduction using active flow control is being carried out within the European Clean Sky program. The objective is to demonstrate the capability of several active flow technologies to decrease fuselage drag by alleviating the flow separation occurring in the rear area of some helicopters. The work is performed on a simplified blunt fuselage at model-scale. Two different flow control actuators are considered for evaluation: steady blowing, unsteady blowing (or pulsed jets). Laboratory tests of each individual actuator are first performed to assess their performance and properties. The fuselage model is then equipped with these actuators distributed in 3 slots located on the ramp bottom edge. This paper addresses the promising results obtained during the wind-tunnel campaign, since significant drag reductions are achieved for a wide range of fuselage angles of attack and yaw angles without detriment of the other aerodynamic characteristics.

Water Impact Test and Simulation of a Composite Energy Absorbing Fuselage Section

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Sparks, Chad; Sareen, Ashish

2003-01-01

In March 2002, a 25-ft/s vertical drop test of a composite fuselage section was conducted onto water. The purpose of the test was to obtain experimental data characterizing the structural response of the fuselage section during water impact for comparison with two previous drop tests that were performed onto a rigid surface and soft soil. For the drop test, the fuselage section was configured with ten 100-lb. lead masses, five per side, that were attached to seat rails mounted to the floor. The fuselage section was raised to a height of 10-ft. and dropped vertically into a 15-ft. diameter pool filled to a depth of 3.5-ft. with water. Approximately 70 channels of data were collected during the drop test at a 10-kHz sampling rate. The test data were used to validate crash simulations of the water impact that were developed using the nonlinear, explicit transient dynamic codes, MSC.Dytran and LS-DYNA. The fuselage structure was modeled using shell and solid elements with a Lagrangian mesh, and the water was modeled with both Eulerian and Lagrangian techniques. The fluid-structure interactions were executed using the fast general coupling in MSC.Dytran and the Arbitrary Lagrange-Euler (ALE) coupling in LS-DYNA. Additionally, the smooth particle hydrodynamics (SPH) meshless Lagrangian technique was used in LS-DYNA to represent the fluid. The simulation results were correlated with the test data to validate the modeling approach. Additional simulation studies were performed to determine how changes in mesh density, mesh uniformity, fluid viscosity, and failure strain influence the test-analysis correlation.

Effects of floor location on response of composite fuselage frames

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Jones, Lisa E.; Fasanella, Edwin L.

1992-01-01

Experimental and analytical results are presented which show the effect of floor placement on the structural response and strength of circular fuselage frames constructed of graphite-epoxy composite material. The research was conducted to study the behavior of conventionally designed advanced composite aircraft components. To achieve desired new designs which incorporate improved energy absorption capabilities requires an understanding of how these conventional designs behave under crash type loadings. Data are presented on the static behavior of the composite structure through photographs of the frame specimen, experimental strain distributions, and through analytical data from composite structural models. An understanding of this behavior can aid the dynamist in predicting the crash behavior of these structures and may assist the designer in achieving improved designs for energy absorption and crash behavior of future structures.

Full-Scale Test and Analysis of a PRSEUS Fuselage Panel to Assess Damage-Containment Features

NASA Technical Reports Server (NTRS)

Bergan, Andrew; Bakuckas, John G.; Lovejoy, Andrew E.; Jegley, Dawn C.; Linton, Kim A.; Korkosz, Gregory; Awerbuch, Jonathan; Tan, Tein-Min

2011-01-01

Stitched composite technology has the potential to substantially decrease structural weight through enhanced damage containment capabilities. The most recent generation of stitched composite technology, the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, has been shown to successfully arrest damage at the sub-component level through tension testing of a three stringer panel with damage in the form of a two-bay notch. In a joint effort undertaken by the National Aeronautics and Space Administration (NASA), the Federal Aviation Administration (FAA), and the Boeing Company, further studies are being conducted to characterize the damage containment features of the PRSEUS concept. A full-scale residual strength test will be performed on a fuselage panel to determine if the load capacity will meet strength, deformation, and damage tolerance requirements. A curved panel was designed, fabricated, and prepared for residual strength testing. A pre-test Finite Element Model (FEM) was developed using design allowables from previous test programs to predict test panel deformation characteristics and margins of safety. Three phases of testing with increasing damage severity include: (1) as manufactured; (2) barely visible impact damage (BVID) and visible impact damage (VID); and (3) discrete source damage (DSD) where the panel will be loaded to catastrophic failure. This paper presents the background information, test plan, and experimental procedure. This paper is the first of several future articles reporting the test preparations, results, and analysis conducted in the test program.

Composite Fuselage Technology

NASA Technical Reports Server (NTRS)

Lagace, Paul A.

1999-01-01

Work was conducted over a ten-year period to address the central issue of damage in primary load-bearing aircraft composite structure, specifically fuselage structure. This included the three facets of damage resistance, damage tolerance, and damage arrest. Experimental, analytical, and numerical work was conducted in order to identify and better understand the mechanisms that control the structural behavior of fuselage structures in their response to the three aspects of damage. Furthermore, work was done to develop straightforward design methodologies that can be employed by structural designers in preliminary design stages to make intelligent choices concerning the material, layup, and structural configurations so that a more efficient structure with structural integrity can be designed and built. Considerable progress was made towards achieving these goals via this work. In regard to damage tolerance considerations, the following were identified as important effects: composite layup and associated orthotropy/structural anisotropy, specifics of initial local damage mechanisms, role of longitudinal versus hoop stress, and large deformation and associated geometric nonlinearity. Means were established to account for effects of radius and for the nonlinear response. In particular, nondimensional parameters were identified to characterize the importance of nonlinearity in the response of pressurized cylinders. This led to the establishment of a iso-nonlinear-error plot for reference in structural design. Finally, in the case of damage tolerance, the general approach of the original methodology to predict the failure pressure involving extending basic plate failure data by accounting for the local stress intensification was accomplished for the general case by accounting for the mechanisms noted by utilizing the capability of the STAGS finite element code and numerically calculating the local stress intensification for the particular configuration to be considered. For the issue of damage arrest, placement of and configuration of stiffeners (including stiffener curvature), and magnitude and orientation of principal strains due to local bending were found to be key considerations. Means were established to account for stiffener effectiveness quantitatively based on radius, slit size, stiffener curvature' and relative bending stifffiesses involved. Geometric nonlinearity was also found to play an - 24 - important role here. Furthermore, it was determined that damage propagation is controlled by different mechanisms (hoop stress versus flapping stress and the associated factors involved in each) depending upon the direction of damage propagation. This latter item results in an inability to scale these phenomena in one test due to the different factors involved. Finally, the importance of shell curvature and associated instability in response to transverse loading including impact were found to be important considerations in damage resistance. A technique, involving asymmetric meshing of a finite element mesh, was developed to predict this behavior and showed excellent correlation with experimental results. Further details of these ten years of work are presented herein with references made to the fourteen documents produced during this work where full details can be found. Implications of this work are discussed and recommendations made. Although it is clear that there is more work to be done to fully understand composite fuselage technology and specifically the overall issue of damage in primary load-bearing composite structures, important understanding and capability has been extended via this work.

Vibro-Acoustics Modal Testing at NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Pappa, Richard S.; Pritchard, Jocelyn I.; Buehrle, Ralph D.

1999-01-01

This paper summarizes on-going modal testing activities at the NASA Langley Research Center for two aircraft fuselage structures: a generic "aluminum testbed cylinder" (ATC) and a Beechcraft Starship fuselage (BSF). Subsequent acoustic tests will measure the interior noise field created by exterior mechanical and acoustic sources. These test results will provide validation databases for interior noise prediction codes on realistic aircraft fuselage structures. The ATC is a 12-ft-long, all-aluminum, scale model assembly. The BSF is a 40-ft-long, all-composite, complete aircraft fuselage. To date, two of seven test configurations of the ATC and all three test configurations of the BSF have been completed. The paper briefly describes the various test configurations, testing procedure, and typical results for frequencies up to 250 Hz.

Crash Testing of Helicopter Airframe Fittings

NASA Technical Reports Server (NTRS)

Clarke, Charles W.; Townsend, William; Boitnott, Richard

2004-01-01

As part of the Rotary Wing Structures Technology Demonstration (RWSTD) program, a surrogate RAH-66 seat attachment fitting was dynamically tested to assess its response to transient, crash impact loads. The dynamic response of this composite material fitting was compared to the performance of an identical fitting subjected to quasi-static loads of similar magnitude. Static and dynamic tests were conducted of both smaller bench level and larger full-scale test articles. At the bench level, the seat fitting was supported in a steel fixture, and in the full-scale tests, the fitting was integrated into a surrogate RAH-66 forward fuselage. Based upon the lessons learned, an improved method to design, analyze, and test similar composite material fittings is proposed.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is moved into place on Atlantis. It is one of two OMS pods attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is moved into place on Atlantis. It is one of two OMS pods attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is suspended in air as it is moved toward Atlantis for installation. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is suspended in air as it is moved toward Atlantis for installation. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is moved closer to Atlantis for installation. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, an orbital maneuvering system (OMS) pod is moved closer to Atlantis for installation. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians make adjustments to the orbital maneuvering system (OMS) pod being installed on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians make adjustments to the orbital maneuvering system (OMS) pod being installed on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods is being moved for installation on Atlantis. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods is being moved for installation on Atlantis. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians move an orbital maneuvering system (OMS) pod into the correct position on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians move an orbital maneuvering system (OMS) pod into the correct position on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

Hybrid Wing-Body (HWB) Pressurized Fuselage Modeling, Analysis, and Design for Weight Reduction

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

2012-01-01

This paper describes the interim progress for an in-house study that is directed toward innovative structural analysis and design of next-generation advanced aircraft concepts, such as the Hybrid Wing-Body (HWB) and the Advanced Mobility Concept-X flight vehicles, for structural weight reduction and associated performance enhancement. Unlike the conventional, skin-stringer-frame construction for a cylindrical fuselage, the box-type pressurized fuselage panels in the HWB undergo significant deformation of the outer aerodynamic surfaces, which must be minimized without significant structural weight penalty. Simple beam and orthotropic plate theory is first considered for sizing, analytical verification, and possible equivalent-plate analysis with appropriate simplification. By designing advanced composite stiffened-shell configurations, significant weight reduction may be possible compared with the sandwich and ribbed-shell structural concepts that have been studied previously. The study involves independent analysis of the advanced composite structural concepts that are presently being developed by The Boeing Company for pressurized HWB flight vehicles. High-fidelity parametric finite-element models of test coupons, panels, and multibay fuselage sections, were developed for conducting design studies and identifying critical areas of potential failure. Interim results are discussed to assess the overall weight/strength advantages.

Detecting damage in full-scale honeycomb sandwich composite curved fuselage panels through frequency response

NASA Astrophysics Data System (ADS)

Leone, Frank A., Jr.; Ozevin, Didem; Mosinyi, Bao; Bakuckas, John G., Jr.; Awerbuch, Jonathan; Lau, Alan; Tan, Tein-Min

2008-03-01

Preliminary tests were conducted using frequency response (FR) characteristics to determine damage initiation and growth in a honeycomb sandwich graphite/epoxy curved panel. This investigation was part of a more general study investigating the damage tolerance characteristics of several such panels subjected to quasi-static internal pressurization combined with hoop and axial loading. The panels were tested at the Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility located at the Federal Aviation Administration William J. Hughes Technical Center in Atlantic City, NJ. The overall program objective was to investigate the damage tolerance characteristics of full-scale composite curved aircraft fuselage panels and the evolution of damage under quasi-static loading up to failure. This paper focuses on one aspect of this comprehensive investigation: the effect of state-of-damage on the characteristics of the frequency response of the subject material. The results presented herein show that recording the frequency response could be used for real-time monitoring of damage growth and in determining damage severity in full-scale composites fuselage aircraft structures.

Full-scale testing and progressive damage modeling of sandwich composite aircraft fuselage structure

NASA Astrophysics Data System (ADS)

Leone, Frank A., Jr.

A comprehensive experimental and computational investigation was conducted to characterize the fracture behavior and structural response of large sandwich composite aircraft fuselage panels containing artificial damage in the form of holes and notches. Full-scale tests were conducted where panels were subjected to quasi-static combined pressure, hoop, and axial loading up to failure. The panels were constructed using plain-weave carbon/epoxy prepreg face sheets and a Nomex honeycomb core. Panel deformation and notch tip damage development were monitored during the tests using several techniques, including optical observations, strain gages, digital image correlation (DIC), acoustic emission (AE), and frequency response (FR). Additional pretest and posttest inspections were performed via thermography, computer-aided tap tests, ultrasound, x-radiography, and scanning electron microscopy. The framework to simulate damage progression and to predict residual strength through use of the finite element (FE) method was developed. The DIC provided local and full-field strain fields corresponding to changes in the state-of-damage and identified the strain components driving damage progression. AE was monitored during loading of all panels and data analysis methodologies were developed to enable real-time determination of damage initiation, progression, and severity in large composite structures. The FR technique has been developed, evaluating its potential as a real-time nondestructive inspection technique applicable to large composite structures. Due to the large disparity in scale between the fuselage panels and the artificial damage, a global/local analysis was performed. The global FE models fully represented the specific geometries, composite lay-ups, and loading mechanisms of the full-scale tests. A progressive damage model was implemented in the local FE models, allowing the gradual failure of elements in the vicinity of the artificial damage. A set of modifications to the definitions of the local FE model boundary conditions is proposed and developed to address several issues related to the scalability of progressive damage modeling concepts, especially in regards to full-scale fuselage structures. Notable improvements were observed in the ability of the FE models to predict the strength of damaged composite fuselage structures. Excellent agreement has been established between the FE model predictions and the experimental results recorded by DIC, AE, FR, and visual observations.

Testing and Analysis of a Composite Non-Cylindrical Aircraft Fuselage Structure. Part 1; Ultimate Design Loads

NASA Technical Reports Server (NTRS)

Przekop, Adam; Jegley, Dawn C.; Lovejoy, Andrew E.; Rouse, Marshall; Wu, Hsi-Yung T.

2016-01-01

The Environmentally Responsible Aviation Project aimed to develop aircraft technologies enabling significant fuel burn and community noise reductions. Small incremental changes to the conventional metallic alloy-based 'tube and wing' configuration were not sufficient to achieve the desired metrics. One airframe concept identified by the project as having the potential to dramatically improve aircraft performance was a composite-based hybrid wing body configuration. Such a concept, however, presented inherent challenges stemming from, among other factors, the necessity to transfer wing loads through the entire center fuselage section which accommodates a pressurized cabin confined by flat or nearly flat panels. This paper discusses finite element analysis and testing of a large-scale hybrid wing body center section structure developed and constructed to demonstrate that the Pultruded Rod Stitched Efficient Unitized Structure concept can meet these challenging demands of the next generation airframes. Part I of the paper considers the five most critical load conditions, which are internal pressure only and positive and negative g-loads with and without internal pressure. Analysis results are compared with measurements acquired during testing. Performance of the test article is found to be closely aligned with predictions and, consequently, able to support the hybrid wing body design loads in pristine and barely visible impact damage conditions.

Testing and Analysis of a Composite Non-Cylindrical Aircraft Fuselage Structure . Part II; Severe Damage

NASA Technical Reports Server (NTRS)

Przekop, Adam; Jegley, Dawn C.; Lovejoy, Andrew E.; Rouse, Marshall; Wu, Hsi-Yung T.

2016-01-01

The Environmentally Responsible Aviation Project aimed to develop aircraft technologies enabling significant fuel burn and community noise reductions. Small incremental changes to the conventional metallic alloy-based 'tube and wing' configuration were not sufficient to achieve the desired metrics. One airframe concept identified by the project as having the potential to dramatically improve aircraft performance was a composite-based hybrid wing body configuration. Such a concept, however, presented inherent challenges stemming from, among other factors, the necessity to transfer wing loads through the entire center fuselage section which accommodates a pressurized cabin confined by flat or nearly flat panels. This paper discusses a finite element analysis and the testing of a large-scale hybrid wing body center section structure developed and constructed to demonstrate that the Pultruded Rod Stitched Efficient Unitized Structure concept can meet these challenging demands of the next generation airframes. Part II of the paper considers the final test to failure of the test article in the presence of an intentionally inflicted severe discrete source damage under the wing up-bending loading condition. Finite element analysis results are compared with measurements acquired during the test and demonstrate that the hybrid wing body test article was able to redistribute and support the required design loads in a severely damaged condition.

Advanced tow placement of composite fuselage structure

NASA Technical Reports Server (NTRS)

Anderson, Robert L.; Grant, Carroll G.

1992-01-01

The Hercules NASA ACT program was established to demonstrate and validate the low cost potential of the automated tow placement process for fabrication of aircraft primary structures. The program is currently being conducted as a cooperative program in collaboration with the Boeing ATCAS Program. The Hercules advanced tow placement process has been in development since 1982 and was developed specifically for composite aircraft structures. The second generation machine, now in operation at Hercules, is a production-ready machine that uses a low cost prepreg tow material form to produce structures with laminate properties equivalent to prepreg tape layup. Current program activities are focused on demonstration of the automated tow placement process for fabrication of subsonic transport aircraft fuselage crown quadrants. We are working with Boeing Commercial Aircraft and Douglas Aircraft during this phase of the program. The Douglas demonstration panels has co-cured skin/stringers, and the Boeing demonstration panel is an intricately bonded part with co-cured skin/stringers and co-bonded frames. Other aircraft structures that were evaluated for the automated tow placement process include engine nacelle components, fuselage pressure bulkheads, and fuselage tail cones. Because of the cylindrical shape of these structures, multiple parts can be fabricated on one two placement tool, thus reducing the cost per pound of the finished part.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers discuss the next step in moving the orbital maneuvering system (OMS) pod behind them. The OMS pod will be installed on Atlantis. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, workers discuss the next step in moving the orbital maneuvering system (OMS) pod behind them. The OMS pod will be installed on Atlantis. Two OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians make final adjustments to the orbital maneuvering system (OMS) pod being installed on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, technicians make final adjustments to the orbital maneuvering system (OMS) pod being installed on Atlantis. The OMS pod is one of two that are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods is lifted off its stand to move it toward Atlantis for installation. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

NASA Image and Video Library

2003-10-30

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods is lifted off its stand to move it toward Atlantis for installation. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts.

Comparisons of the Impact Responses of a 1/5-Scale Model and a Full-Scale Crashworthy Composite Fuselage Section

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.; Lyle, Karen H.

2003-01-01

A 25-fps vertical drop test of a 1/5-scale model composite fuselage section was conducted to replicate a previous test of a full-scale fuselage section. The purpose of the test was to obtain experimental data characterizing the impact response of the 1/5-scale model fuselage section for comparison with the corresponding full-scale data. This comparison is performed to assess the scaling procedures and to determine if scaling effects are present. For the drop test, the 1/5-scale model fuselage section was configured in a similar manner as the full-scale section, with lead masses attached to the floor through simulated seat rails. Scaled acceleration and velocity responses are compared and a general assessment of structural damage is made. To further quantify the data correlation, comparisons of the average acceleration data are made as a function of floor location and longitudinal position. Also, the percentage differences in the velocity change (area under the acceleration curve) and the velocity change squared (proportional to kinetic energy) are compared as a function of floor location. Finally, correlation coefficients are calculated for corresponding 1/5- and full-scale data channels and these values are plotted versus floor location. From a scaling perspective, the differences between the 1/5- and full-scale tests are relatively small, indicating that appropriate scaling procedures were used in fabricating the test specimens and in conducting the experiments. The small differences in the scaled test data are attributed to minor scaling anomalies in mass, potential energy, and impact attitude.

Preliminary Weight Savings Estimate for a Commercial Transport Wing Using Rod-Stiffened Stitched Composite Technology

NASA Technical Reports Server (NTRS)

Lovejoy, Andrew E.

2015-01-01

A structural concept called pultruded rod stitched efficient unitized structure (PRSEUS) was developed by the Boeing Company to address the complex structural design aspects associated with a pressurized hybrid wing body (HWB) aircraft configuration. While PRSEUS was an enabling technology for the pressurized HWB structure, limited investigation of PRSEUS for other aircraft structures, such as circular fuselages and wings, has been done. Therefore, a study was undertaken to investigate the potential weight savings afforded by using the PRSEUS concept for a commercial transport wing. The study applied PRSEUS to the Advanced Subsonic Technology (AST) Program composite semi-span test article, which was sized using three load cases. The initial PRSEUS design was developed by matching cross-sectional stiffnesses for each stringer/skin combination within the wing covers, then the design was modified to ensure that the PRSEUS design satisfied the design criteria. It was found that the PRSEUS wing design exhibited weight savings over the blade-stiffened composite AST Program wing of nearly 9%, and a weight savings of 49% and 29% for the lower and upper covers, respectively, compared to an equivalent metallic wing.

The Influences of Lamination Angles on the Interior Noise Levels of an Aircraft

NASA Technical Reports Server (NTRS)

Fernholz, Christian M.; Robinson, Jay H.

1996-01-01

The feasibility of reducing the interior noise levels of an aircraft passenger cabin through optimization of the composite lay up of the fuselage is investigated. MSC/NASTRAN, a commercially available finite element code, is used to perform the dynamic analysis and subsequent optimization of the fuselage. The numerical calculation of sensitivity of acoustic pressure to lamination angle is verified using a simple thin, cylindrical shell with point force excitations as noise sources. The thin shell used represents a geometry similar to the fuselage and analytic solutions are available for the cylindrical thin shell equations of motion. Optimization of lamination angle for the reduction of interior noise is performed using a finite element model of an actual aircraft fuselage. The aircraft modeled for this study is the Beech Starship. Point forces simulate the structure borne noise produced by the engines and are applied to the fuselage at the wing mounting locations. These forces are the noise source for the optimization problem. The acoustic pressure response is reduced at a number of points in the fuselage and over a number of frequencies. The objective function is minimized with the constraint that it be larger than the maximum sound pressure level at the response points in the passenger cabin for all excitation frequencies in the range of interest. Results from the study of the fuselage model indicate that a reduction in interior noise levels is possible over a finite frequency range through optimal configuration of the lamination angles in the fuselage. Noise reductions of roughly 4 dB were attained. For frequencies outside the optimization range, the acoustic pressure response may increase after optimization. The effects of changing lamination angle on the overall structural integrity of the airframe are not considered in this study.

Active Aerodynamic Load Reduction on a Rotorcraft Fuselage With Rotor Effects: A CFD Validation Effort

NASA Technical Reports Server (NTRS)

Allan, Brian G.; Schaeffler, Norman W.; Jenkins, Luther N.; Yao, Chung-Sheng; Wong, Oliver D.; Tanner, Philip E.

2015-01-01

A rotorcraft fuselage is typically designed with an emphasis on operational functionality with aerodynamic efficiency being of secondary importance. This results in a significant amount of drag during high-speed forward flight that can be a limiting factor for future high-speed rotorcraft designs. To enable higher speed flight, while maintaining a functional fuselage design (i.e., a large rear cargo ramp door), the NASA Rotary Wing Project has conducted both experimental and computational investigations to assess active flow control as an enabling technology for fuselage drag reduction. This paper will evaluate numerical simulations of a flow control system on a generic rotorcraft fuselage with a rotor in forward flight using OVERFLOW, a structured mesh Reynolds-averaged Navier-Stokes flow solver developed at NASA. The results are compared to fuselage forces, surface pressures, and PN flow field data obtained in a wind tunnel experiment conducted at the NASA Langley 14-by 22-Foot Subsonic Tunnel where significant drag and download reductions were demonstrated using flow control. This comparison showed that the Reynolds-averaged Navier-Stokes flow solver was unable to predict the fuselage forces and pressure measurements on the ramp for the baseline and flow control cases. While the CFD was able to capture the flow features, it was unable to accurately predict the performance of the flow control.

Damage Arresting Composites for Shaped Vehicles

NASA Technical Reports Server (NTRS)

Velicki, Alex

2009-01-01

This report describes the development of a novel structural solution that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body configurations that are described in NASA NRA subtopic A2A.3, "Materials and Structures for Wing Components and Non-Circular Fuselage." The phase I portion of this task includes a comprehensive finite element model-based structural sizing exercise performed using the BWB airplane configuration to generate internal loads and fuselage panel weights for an advanced Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) structural concept. An accompanying element-level test program is also described which substantiates the analytical results and calculation methods used in the trade study. The phase II plan for the continuation of this research is also included herein.

Hybrid Wing-Body Pressurized Fuselage and Bulkhead, Design and Optimization

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

2013-01-01

The structural weight reduction of a pressurized Hybrid Wing-Body (HWB) fuselage is a serious challenge. Hence, research and development are presently being continued at NASA under the Environmentally Responsible Aviation (ERA) and Subsonic Fixed Wing (SFW) projects in collaboration with the Boeing Company, Huntington Beach and Air Force Research Laboratory (AFRL). In this paper, a structural analysis of the HWB fuselage and bulkhead panels is presented, with the objectives of design improvement and structural weight reduction. First, orthotropic plate theories for sizing, and equivalent plate analysis with appropriate simplification are considered. Then parametric finite-element analysis of a fuselage section and bulkhead are conducted using advanced stitched composite structural concepts, which are presently being developed at Boeing for pressurized HWB flight vehicles. With this advanced stiffened-shell design, structural weights are computed and compared to the thick sandwich, vaulted-ribbed-shell, and multi-bubble stiffened-shell structural concepts that had been studied previously. The analytical and numerical results are discussed to assess the overall weight/strength advantages.

Damage-Tolerance Characteristics of Composite Fuselage Sandwich Structures with Thick Facesheets

NASA Technical Reports Server (NTRS)

McGowan, David M.; Ambur, Damodar R.

1997-01-01

Damage tolerance characteristics and results from experimental and analytical studies of a composite fuselage keel sandwich structure subjected to low-speed impact damage and discrete-source damage are presented. The test specimens are constructed from graphite-epoxy skins borided to a honeycomb core, and they are representative of a highly loaded fuselage keel structure. Results of compression-after-impact (CAI) and notch-length sensitivity studies of 5-in.-wide by 10-in.long specimens are presented. A correlation between low-speed-impact dent depth, the associated damage area, and residual strength for different impact-energy levels is described; and a comparison of the strength for undamaged and damaged specimens with different notch-length-to-specimen-width ratios is presented. Surface strains in the facesheets of the undamaged specimens as well as surface strains that illustrate the load redistribution around the notch sites in the notched specimens are presented and compared with results from finite element analyses. Reductions in strength of as much as 53.1 percent for the impacted specimens and 64.7 percent for the notched specimens are observed.

Nondestructive inspection of bonded composite doublers for aircraft

DOE Office of Scientific and Technical Information (OSTI.GOV)

Roach, D.; Moore, D.; Walkington, P.

1996-12-31

One major thrust in FAA`s National Aging Aircraft Research Program is to foster new technologies in civil aircraft maintenance. Recent DOD and other government developments in using bonded composite doublers on metal structures support the need for validation of such doubler applications on US certificated airplanes. In this study, a specific composite application was chosen on an L-1011 aircraft. Primary inspection requirements for these doublers include identifying disbonds between composite laminate and aluminum parent material, and delaminations in the composite laminate. Surveillance of cracks or corrosion in the parent aluminum material beneath the double is also a concern. No singlemore » NDI method can inspect for every flaw type, therefore we need to know NDI capabilities and limitations. This paper reports on a series of NDI tests conducted on laboratory test structures and on a fuselage section from a retired L-1011. Application of ultrasonics, x-ray, and eddy current to composite doublers and results from test specimens loaded to provide a changing flaw profile, are presented in this paper. Development of appropriate inspection calibration standards are also discussed.« less

Development of thermoplastic composite aircraft structures

NASA Technical Reports Server (NTRS)

Renieri, Michael P.; Burpo, Steven J.; Roundy, Lance M.; Todd, Stephanie A.; Kim, H. J.

1992-01-01

Efforts focused on the use of thermoplastic composite materials in the development of structural details associated with an advanced fighter fuselage section with applicability to transport design. In support of these designs, mechanics developments were conducted in two areas. First, a dissipative strain energy approach to material characterization and failure prediction, developed at the Naval Research Laboratory, was evaluated as a design/analysis tool. Second, a finite element formulation for thick composites was developed and incorporated into a lug analysis method which incorporates pin bending effects. Manufacturing concepts were developed for an upper fuel cell cover. A detailed trade study produced two promising concepts: fiber placement and single-step diaphragm forming. Based on the innovative design/manufacturing concepts for the fuselage section primary structure, elements were designed, fabricated, and structurally tested. These elements focused on key issues such as thick composite lugs and low cost forming of fastenerless, stiffener/moldine concepts. Manufacturing techniques included autoclave consolidation, single diaphragm consolidation (SDCC) and roll-forming.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is lowered toward a transporter. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is lowered toward a transporter. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

KENNEDY SPACE CENTER, FLA. - Technicians in the Orbiter Processing Facility oversee removal of one of two orbital maneuvering system (OMS) pods from Endeavour. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - Technicians in the Orbiter Processing Facility oversee removal of one of two orbital maneuvering system (OMS) pods from Endeavour. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility prepare to remove one of two orbital maneuvering system (OMS) pods from Endeavour. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility prepare to remove one of two orbital maneuvering system (OMS) pods from Endeavour. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is suspended overhead. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is suspended overhead. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is lowered onto a transporter. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

NASA Image and Video Library

2003-09-05

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, one of two orbital maneuvering system (OMS) pods removed from Endeavour is lowered onto a transporter. The OMS pods are attached to the upper aft fuselage left and right sides. Fabricated primarily of graphite epoxy composite and aluminum, each pod is 21.8 feet long and 11.37 feet wide at its aft end and 8.41 feet wide at its forward end, with a surface area of approximately 435 square feet. Each pod houses the Reaction Control System propulsion components used for inflight maneuvering and is attached to the aft fuselage with 11 bolts. OMS pods are removed during Orbiter Major Modifications. Once removed, the OMS pods undergo in-depth structural inspections, system checks and the thrusters are changed out.

Multi-Terrain Vertical Drop Tests of a Composite Fuselage Section

NASA Technical Reports Server (NTRS)

Kellas, Sotiris; Jackson, Karen E.

2008-01-01

A 5-ft-diameter composite fuselage section was retrofitted with four identical blocks of deployable honeycomb energy absorber and crash tested on two different surfaces: soft soil, and water. The drop tests were conducted at the 70-ft. drop tower at the Landing and Impact Research (LandIR) Facility of NASA Langley. Water drop tests were performed into a 15-ft-diameter pool of water that was approximately 42-in. deep. For the soft soil impact, a 15-ft-square container filled with fine-sifted, unpacked sand was located beneath the drop tower. All drop tests were vertical with a nominally flat attitude with respect to the impact surface. The measured impact velocities were 37.4, and 24.7-fps for soft soil and water, respectively. A fuselage section without energy absorbers was also drop tested onto water to provide a datum for comparison with the test, which included energy absorbers. In order to facilitate this type of comparison and to ensure fuselage survivability for the no-energy-absorber case, the velocity of the water impact tests was restricted to 25-fps nominal. While all tests described in this paper were limited to vertical impact velocities, the implications and design challenges of utilizing external energy absorbers during combined forward and vertical impact velocities are discussed. The design, testing and selection of a honeycomb cover, which was required in soft surface and water impacts to transmit the load into the honeycomb cell walls, is also presented.

Helicopter aeroelastic stability and response - Current topics and future trends

NASA Technical Reports Server (NTRS)

Friedmann, Peretz P.

1990-01-01

This paper presents several current topics in rotary wing aeroelasticity and concludes by attempting to anticipate future trends and developments. These topics are: (1) the role of geometric nonlinearities; (2) structural modeling, and aeroelastic analysis of composite rotor blades; (3) aeroelastic stability and response in forward flight; (4) modeling of coupled rotor/fuselage aeromechanical problems and their active control; and (5) the coupled rotor-fuselage vibration problem and its alleviation by higher harmonic control. Selected results illustrating the fundamental aspects of these topics are presented. Future developments are briefly discussed.

Simulating the Impact Response of Composite Airframe Components

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Littell, Justin D.; Fasanella, Edwin L.

2014-01-01

In 2010, NASA Langley Research Center obtained residual hardware from the US Army's Survivable Affordable Repairable Airframe Program (SARAP). The hardware consisted of a composite fuselage section that was representative of the center section of a Black Hawk helicopter. The section was fabricated by Sikorsky Aircraft Corporation and designated the Test Validation Article (TVA). The TVA was subjected to a vertical drop test in 2008 to evaluate a tilting roof concept to limit the intrusion of overhead mass items, such as the rotor transmission, into the fuselage cabin. As a result of the 2008 test, damage to the hardware was limited primarily to the roof. Consequently, when the post-test article was obtained in 2010, the roof area was removed and the remaining structure was cut into six different types of test specimens including: (1) tension and compression coupons for material property characterization, (2) I-beam sections, (3) T-sections, (4) cruciform sections, (5) a large subfloor section, and (6) a forward framed fuselage section. In 2011, NASA and Sikorsky entered into a cooperative research agreement to study the impact responses of composite airframe structures and to evaluate the capabilities of the explicit transient dynamic finite element code, LS-DYNA®, to simulate these responses including damage initiation and progressive failure. Finite element models of the composite specimens were developed and impact simulations were performed. The properties of the composite material were represented using both a progressive in-plane damage model (Mat 54) and a continuum damage mechanics model (Mat 58) in LS-DYNA. This paper provides test-analysis comparisons of time history responses and the location and type of damage for representative I-beam, T-section, and cruciform section components.

Studying impact damage on carbon-fiber reinforced aircraft composite panels with sonicir

DOE Office of Scientific and Technical Information (OSTI.GOV)

Han Xiaoyan; Zhang Ding; He Qi

2011-06-23

Composites are becoming more important materials in commercial aircraft structures such as the fuselage and wings with the new B787 Dreamliner from Boeing which has the target to utilize 50% by weight of composite materials. Carbon-fiber reinforced composites are the material of choice in aircraft structures. This is due to their light weight and high strength (high strength-to-weight ratio), high specific stiffness, tailorability of properties, design flexibility etc. Especially, by reducing the aircraft's body weight by using such lighter structures, the cost of fuel can be greatly reduced with the high jet fuel price for commercial airlines. However, these compositesmore » are prone to impact damage and the damage may occur without any observable sign on the surface, yet resulting in delaminations and disbonds that may occur well within the layers. We are studying the impact problem with carbon-fiber reinforced composite panels and developing SonicIR for this application as a fast and wide-area NDE technology. In this paper, we present our results in studying composite structures including carbon-fiber reinforced composite materials, and preliminary quantitative studies on delamination type defect depth identification in the panels.« less

Review of Aircraft Crash Structural Response Research.

DTIC Science & Technology

1982-08-01

structures consisting of conventional built-up metallic construction and those consisting of advanced composite materials were of interest. The latter...increasing importance. Some recent theoretical and experimental studies of the behavior of composite - material structures subjected to severe static...dynamic, and/or impact conditions are noted. Such topics as crashworthiness testing ot composite fuselage structures, the impact resistance of graphite and

Vibroacoustic Model Validation for a Curved Honeycomb Composite Panel

NASA Technical Reports Server (NTRS)

Buehrle, Ralph D.; Robinson, Jay H.; Grosveld, Ferdinand W.

2001-01-01

Finite element and boundary element models are developed to investigate the vibroacoustic response of a curved honeycomb composite sidewall panel. Results from vibroacoustic tests conducted in the NASA Langley Structural Acoustic Loads and Transmission facility are used to validate the numerical predictions. The sidewall panel is constructed from a flexible honeycomb core sandwiched between carbon fiber reinforced composite laminate face sheets. This type of construction is being used in the development of an all-composite aircraft fuselage. In contrast to conventional rib-stiffened aircraft fuselage structures, the composite panel has nominally uniform thickness resulting in a uniform distribution of mass and stiffness. Due to differences in the mass and stiffness distribution, the noise transmission mechanisms for the composite panel are expected to be substantially different from those of a conventional rib-stiffened structure. The development of accurate vibroacoustic models will aide in the understanding of the dominant noise transmission mechanisms and enable optimization studies to be performed that will determine the most beneficial noise control treatments. Finite element and boundary element models of the sidewall panel are described. Vibroacoustic response predictions are presented for forced vibration input and the results are compared with experimental data.

EC02-0188-3

NASA Image and Video Library

2002-07-12

Technician Shawn Warren carefully smoothes out the composite skin of an instrument fairingatop the upper fuselage of the Altair unmanned aerial vehicle (UAV) at General Atomics Aeronautical Systems, Inc., facility at Adelanto, Calif.

AST Composite Wing Program: Executive Summary

NASA Technical Reports Server (NTRS)

Karal, Michael

2001-01-01

The Boeing Company demonstrated the application of stitched/resin infused (S/RFI) composite materials on commercial transport aircraft primary wing structures under the Advanced Subsonic technology (AST) Composite Wing contract. This report describes a weight trade study utilizing a wing torque box design applicable to a 220-passenger commercial aircraft and was used to verify the weight savings a S/RFI structure would offer compared to an identical aluminum wing box design. This trade study was performed in the AST Composite Wing program, and the overall weight savings are reported. Previous program work involved the design of a S/RFI-base-line wing box structural test component and its associated testing hardware. This detail structural design effort which is known as the "semi-span" in this report, was completed under a previous NASA contract. The full-scale wing design was based on a configuration for a MD-90-40X airplane, and the objective of this structural test component was to demonstrate the maturity of the S/RFI technology through the evaluation of a full-scale wing box/fuselage section structural test. However, scope reductions of the AST Composite Wing Program pre-vented the fabrication and evaluation of this wing box structure. Results obtained from the weight trade study, the full-scale test component design effort, fabrication, design development testing, and full-scale testing of the semi-span wing box are reported.

Volumetric pattern analysis of fuselage-mounted airborne antennas. Ph.D. Thesis; [prediction analysis techniques for antenna radiation patterns of microwave antennas on commercial aircraft

NASA Technical Reports Server (NTRS)

Yu, C. L.

1976-01-01

A volumetric pattern analysis of fuselage-mounted airborne antennas at high frequencies was investigated. The primary goal of the investigation was to develop a numerical solution for predicting radiation patterns of airborne antennas in an accurate and efficient manner. An analytical study of airborne antenna pattern problems is presented in which the antenna is mounted on the fuselage near the top or bottom. Since this is a study of general-type commercial aircraft, the aircraft was modeled in its most basic form. The fuselage was assumed to be an infinitely long perfectly conducting elliptic cylinder in its cross-section and a composite elliptic cylinder in its elevation profile. The wing, cockpit, stabilizers (horizontal and vertical) and landing gear are modeled by "N" sided bent or flat plates which can be arbitrarily attached to the fuselage. The volumetric solution developed utilizes two elliptic cylinders, namely, the roll plane and elevation plane models to approximate the principal surface profile (longitudinal and transverse) at the antenna location. With the belt concept and the aid of appropriate coordinate system transformations the solution can be used to predict the volumetric patterns of airborne antennas in an accurate and efficient manner. Applications of this solution to various airborne antenna problems show good agreement with scale model measurements. Extensive data are presented for a microwave landing antenna system.

Small transport aircraft technology propeller study

NASA Technical Reports Server (NTRS)

Black, B. M.; Magliozzi, B.; Rohrbach, C.

1983-01-01

A study to define potential benefits of advanced technology propeller for 1985-1990 STAT commuter airplanes was completed. Two baselines, a Convair, 30 passenger, 0.47 Mach number airplane and a Lockheed, 50 passenger, 0.70 Mach number airplane, were selected from NASA-Ames sponsored airframe contracts. Parametric performance, noise level, weight and cost trends for propellers with varying number of blades, activity factor, camber and diameter incorporating blade sweep, tip proplets, advanced composite materials, advanced airfoils, advanced prevision synchrophasing and counter-rotation are presented. The resulting DOC, fuel burned, empty weight and acquisition cost benefits are presented for resizings of the two baseline airplanes. Six-bladed propeller having advanced composite blades, advanced airfoils, tip proplets and advanced prevision synchrophasers provided the maximum DOC improvements for both airplanes. DOC and fuel burned were reduced by 8.3% and 17.0% respectively for the Convair airplane and by 24.9% and 41.2% respectively for the Lockheed airplane. The larger reductions arose from a baseline definition with very heavy fuselage acoustic treatment. An alternate baseline, with a cabin noise 13dB in excess of the objective, was also studied.

Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

NASA Technical Reports Server (NTRS)

Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

2015-01-01

The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test predictions. This paper documents the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

Advanced manufacturing development of a composite empennage component for L-1011 aircraft. Phase 3: Production readiness verification testing

NASA Technical Reports Server (NTRS)

Jackson, A.; Sandifer, J.; Sandorff, P.; Vancleave, R.

1984-01-01

Twenty-two specimens of each of two key structural elements of the Advance Composite Vertical Fin (ACVF) were fabricated and tested. One element represented the front spar at the fuselage attachment area and the other element represented the cover at the fuselage joint area. Ten specimens of each element were selected for static testing. The coefficient of variation resulting from the tests was 3.28 percent for the ten cover specimens and 6.11 percent for the ten spar specimens, which compare well with metallic structures. The remaining twelve cover and twelve spar specimens were durability tested in environmental chambers which permitted the temperature and humidity environment to be cycled as well as the applied loads. Results of the durability tests indicated that such components will survive the service environment.

Analysis for the Progressive Failure Response of Textile Composite Fuselage Frames

NASA Technical Reports Server (NTRS)

Johnson, Eric R.; Boitnott, Richard L. (Technical Monitor)

2002-01-01

A part of aviation accident mitigation is a crashworthy airframe structure, and an important measure of merit for a crashworthy structure is the amount of kinetic energy that can be absorbed in the crush of the structure. Prediction of the energy absorbed from finite element analyses requires modeling the progressive failure sequence. Progressive failure modes may include material degradation, fracture and crack growth, and buckling and collapse. The design of crashworthy airframe components will benefit from progressive failure analyses that have been validated by tests. The subject of this research is the development of a progressive failure analysis for a textile composite, circumferential fuselage frame subjected to a quasi-static, crash-type load. The test data for the frame are reported, and these data are used to develop and to validate methods for the progressive failure response.

Open Circuit Resonant (SansEC) Sensor for Composite Damage Detection and Diagnosis in Aircraft Lightning Environments

NASA Technical Reports Server (NTRS)

Wang, Chuantong; Dudley, Kenneth L.; Szatkowski, George N.

2012-01-01

Composite materials are increasingly used in modern aircraft for reducing weight, improving fuel efficiency, and enhancing the overall design, performance, and manufacturability of airborne vehicles. Materials such as fiberglass reinforced composites (FRC) and carbon-fiber-reinforced polymers (CFRP) are being used to great advantage in airframes, wings, engine nacelles, turbine blades, fairings, fuselage and empennage structures, control surfaces and coverings. However, the potential damage from the direct and indirect effects of lightning strikes is of increased concern to aircraft designers and operators. When a lightning strike occurs, the points of attachment and detachment on the aircraft surface must be found by visual inspection, and then assessed for damage by maintenance personnel to ensure continued safe flight operations. In this paper, a new method and system for aircraft in-situ damage detection and diagnosis are presented. The method and system are based on open circuit (SansEC) sensor technology developed at NASA Langley Research Center. SansEC (Sans Electric Connection) sensor technology is a new technical framework for designing, powering, and interrogating sensors to detect damage in composite materials. Damage in composite material is generally associated with a localized change in material permittivity and/or conductivity. These changes are sensed using SansEC. Unique electrical signatures are used for damage detection and diagnosis. NASA LaRC has both experimentally and theoretically demonstrated that SansEC sensors can be effectively used for in-situ composite damage detection.

Evaluation of a Composite Sandwich Fuselage Side Panel with Damage and Subjected to Internal Pressure

NASA Technical Reports Server (NTRS)

Rouse, Marshall; Ambur, Damodar R.; Bodine, Jerry; Dopker, Bernhard

1997-01-01

The results from an experimental and analytical study of a composite sandwich fuselage side panel for a transport aircraft are presented. The panel has two window cutouts and three frames, and has been evaluated with internal pressure loads that generate biaxial tension loading conditions. Design limit load and design ultimate load tests have been performed on the graphite-epoxy sandwich panel with the middle frame removed to demonstrate the suitability of this two-frame design for supporting the prescribed biaxial loading conditions with twice the initial frame spacing of 20 inches. The two-frame panel was damaged by cutting a notch that originates at the edge of a cutout and extends in the panel hoop direction through the window-belt area. This panel with a notch was tested in a combined-load condition to demonstrate the structural damage tolerance at the design limit load condition. The two panel configurations successfully satisfied all design load requirements in the experimental part of the study, and the three-frame and two-frame panel responses are fully explained by the analysis results. The results of this study suggest that there is potential for using sandwich structural concepts with greater than the usual 20-in.-wide frame spacing to further reduce aircraft fuselage structural weight.

Tension fracture of laminates for transport fuselage. Part 1: Material screening

NASA Technical Reports Server (NTRS)

Walker, T. H.; Avery, W. B.; Ilcewicz, L. B.; Poe, C. C., Jr.; Harris, C. E.

1992-01-01

Transport fuselage structures are designed to contain pressure following a large penetrating damage event. Application of composites to fuselage structures requires a data base and supporting analysis on tension damage tolerance. Tests with 430 fracture specimens were used to accomplish the following: (1) identify critical material and laminate variables affecting notch sensitivity, (2) evaluate composite failure criteria, and (3) recommend a screening test method. Variables studied included fiber type, matrix toughness, lamination manufacturing process, and intraply hybridization. The laminates found to have the lowest notch sensitivity were manufactured using automated tow placement. This suggests a possible relationship between the stress distribution and repeatable levels of material inhomogeneity that are larger than found in traditional tape laminates. Laminates with the highest notch sensitivity consisted of toughened matrix materials that were resistant to a splitting phenomena that reduces stress concentrations in load bearing plies. Parameters for conventional fracture criteria were found to increase with the crack length of the smallest notch sizes studied. Most materials and laminate combinations followed less than a square root singularity for the largest crack sizes studied. Specimen geometry, notch type, and notch size were evaluated in developing a screening test procedure. Results indicate that a range of notch sizes must be tested to determine notch sensitivity.

Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

NASA Technical Reports Server (NTRS)

Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

2016-01-01

The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. LS-DYNA® was used to predict the thickness of the composite shield required to prevent blade penetration. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test LS-DYNA predictions. This paper documents the analysis conducted to predict the required thickness of a composite shield, the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

Thermal Analysis by Numerical Methods of Debonding Effects near the Crack Tip under Composite Repairs

NASA Astrophysics Data System (ADS)

Tsamasphyros, G. J.; Kanderakis, G. N.; Marioli-Riga, Z. P.

2003-05-01

Composite patch repair of metallic structures has become a rapidly grown technology in the aerospace field due to the demand for significant increases in the useful life of both military and civilian aircraft. This has led to significant advances overall in the repair technology of cracked metallic structures. Adhesively bonded composite reinforcements offer remarkable advantages such as mechanical efficiency, repair time, cost reduction, high structural integrity, repair inspectability, damage tolerance to further causes of future strains, anticorrosion and antifretting properties. However, because of the different nature and properties of the materials that form a repair (metals, composites, adhesives), side-effects may occur: debonding due to high stress concentration in the vicinity of the crack, thermal residual stresses because of different thermal expansion coefficients of the adherents, etc. In this paper a three-dimensional finite elements analysis of the area around a patch repaired crack of a typical aircraft fuselage is performed, taking into account both the properties and the geometry of the involved materials. Examined in this case are 2024-T3 aluminum alloy as base material, FM-73 as the adhesive system and F4/5521 boron/epoxy prepreg as the patch material. Through the thickness stresses near the crack tip and along the patch edges with and without temperature effects are calculated and debonding near the crack tip is examined. Finally, the calculated results are compared with existing theories.

Advances in crash dynamics for aircraft safety

NASA Astrophysics Data System (ADS)

Guida, M.; Marulo, F.; Abrate, S.

2018-04-01

This paper studies the ability of the fuselage's lower lobe to absorb the energy during a crash landing, where the introduction of the composite materials can improve the crash survivability thanks to the crushing capability of structural parts to limit the effects of deceleration on the occupants. Providing a protective shell around the occupants and minimizing the risks of injuries during and immediately after the crash in the post-crash regime is a safety requirement. This study consists of: (1) numerical and experimental investigations on small components to verify design concepts using high performance composite materials; (2) analyses of full scale crashes of fuselage lower lobes. This paper outlines an approach for demonstrating the crashworthiness characteristics of the airframe performing a drop test at low velocity impact to validate a numerical model obtained by assembling structural components and materials' properties previously obtained by testing coupons and sub-elements.

Analysis for the Progressive Failure Response of Textile Composite Fuselage Frames

NASA Technical Reports Server (NTRS)

Johnson, Eric R.; Boitnott, Richard L. (Technical Monitor)

2002-01-01

A part of aviation accident mitigation is a crash worthy airframe structure, and an important measure of merit for a crash worthy structure is the amount of kinetic energy that can be absorbed in the crush of the structure. Prediction of the energy absorbed from finite element analyses requires modeling the progressive failure sequence. Progressive failure modes may include material degradation, fracture and crack growth, and buckling and collapse. The design of crash worthy airframe components will benefit from progressive failure analyses that have been validated by tests. The subject of this research is the development of a progressive failure analysis for textile composite. circumferential fuselage frames subjected to a quasi-static, crash-type load. The test data for these frames are reported, and these data, along with stub column test data, are to be used to develop and to validate methods for the progressive failure response.

Test and Analyses of a Composite Multi-Bay Fuselage Panel Under Uni-Axial Compression

NASA Technical Reports Server (NTRS)

Li, Jian; Baker, Donald J.

2004-01-01

A composite panel containing three stringers and two frames cut from a vacuum-assisted resin transfer molded (VaRTM) stitched fuselage article was tested under uni-axial compression loading. The stringers and frames divided the panel into six bays with two columns of three bays each along the compressive loading direction. The two frames were supported at the ends with pins to restrict the out-of-plane translation. The free edges of the panel were constrained by knife-edges. The panel was modeled with shell finite elements and analyzed with ABAQUS nonlinear solver. The nonlinear predictions were compared with the test results in out-of-plane displacements, back-to-back surface strains on stringer flanges and back-to-back surface strains at the centers of the skin-bays. The analysis predictions were in good agreement with the test data up to post-buckling.

Monitoring of bolted joints using piezoelectric active-sensing for aerospace applications

DOE Office of Scientific and Technical Information (OSTI.GOV)

Park, Gyuhae; Farrar, Charles R; Park, Chan - Yik

2010-01-01

This paper is a report of an initial investigation into tracking and monitoring the integrity of bolted joints using piezoelectric active-sensors. The target application of this study is a fitting lug assembly of unmanned aerial vehicles (UAVs), where a composite wing is mounted to a UAV fuselage. The SHM methods deployed in this study are impedance-based SHM techniques, time-series analysis, and high-frequency response functions measured by piezoelectric active-sensors. Different types of simulated damage are introduced into the structure, and the capability of each technique is examined and compared. Additional considerations encountered in this initial investigation are made to guide furthermore » thorough research required for the successful field deployment of this technology.« less

Active infrared thermal imaging technology to detect the corrosion defects in aircraft cargo door

NASA Astrophysics Data System (ADS)

Chen, Dapeng; Zhang, Cunlin; Zeng, Zhi; Xing, Chunfei; Li, Yanhong

2009-11-01

Aircraft fuselage material corrosion problems have been major aviation security issues, which hinder the development of aviation industry. How can we use non-destructive testing methods to detect the internal corrosion defects from the outside of the fuselage, to find the hidden safety problems in advance and update the defective equipment and materials, has great significance for the prevention of accidents. Nowadays, the active infrared thermal imaging technology as a new nondestructive technology has been gradually used on a wide variety of materials, such as composite, metal and so on. This article makes use of this technology on an aircraft cargo door specimen to detect the corrosion defects. Firstly, use High-energy flash pulse to excite the specimen, and use the thermal image processing software to splice the thermal images, so the thermal images of the overall specimen can be showed. Then, heat the defects by ultrasonic excitation, this will cause vibration and friction or thermoelastic effects in the places of defects, so the ultrasonic energy will dissipate into heat and manifested in the uneven temperature of surface. An Infrared camera to capture the changes of temperature of material surface, send data to the computer and records the thermal information of the defects. Finally, extracting data and drawing infrared radiation-time curve of some selected points of interest to analyze the signal changes in heat of defects further more. The results of the experiments show that both of the two ways of heat excitation show a clear position and shape of defects, and the ultrasonic method has more obvious effect of excitation to the defects, and a higher signal to noise ratio than the flash pulse excitation, but flash pulse method do not contact the specimen in the process of excitation, and shows the location and shape of defects in the overall of the specimen has its advantages.

Damage Tolerance of Pre-Stressed Composite Panels Under Impact Loads

NASA Astrophysics Data System (ADS)

Johnson, Alastair F.; Toso-Pentecôte, Nathalie; Schueler, Dominik

2014-02-01

An experimental test campaign studied the structural integrity of carbon fibre/epoxy panels preloaded in tension or compression then subjected to gas gun impact tests causing significant damage. The test programme used representative composite aircraft fuselage panels composed of aerospace carbon fibre toughened epoxy prepreg laminates. Preload levels in tension were representative of design limit loads for fuselage panels of this size, and maximum compression preloads were in the post-buckle region. Two main impact scenarios were considered: notch damage from a 12 mm steel cube projectile, at velocities in the range 93-136 m/s; blunt impact damage from 25 mm diameter glass balls, at velocities 64-86 m/s. The combined influence of preload and impact damage on panel residual strengths was measured and results analysed in the context of damage tolerance requirements for composite aircraft panels. The tests showed structural integrity well above design limit loads for composite panels preloaded in tension and compression with visible notch impact damage from hard body impact tests. However, blunt impact tests on buckled compression loaded panels caused large delamination damage regions which lowered plate bending stiffness and reduced significantly compression strengths in buckling.

Technician Shawn Warren carefully smoothes out the composite skin of an instrument fairing atop the

NASA Technical Reports Server (NTRS)

2002-01-01

Technician Shawn Warren carefully smoothes out the composite skin of an instrument fairing atop the upper fuselage of the Altair unmanned aerial vehicle (UAV) at General Atomics Aeronautical Systems, Inc., facility at Adelanto, Calif. General Atomics Aeronautical Systems, Inc., is developing the Altair version of its Predator B unmanned reconnaissance aircraft under NASA's Environmental Research Aircraft and Sensor Technology (ERAST) project. NASA plans to use the Altair as a technology demonstrator to validate a variety of command and control technologies for UAVs, as well as demonstrate the capability to perform a variety of Earth science missions. The Altair is designed to carry an 700-lb. payload of scientific instruments and imaging equipment for as long as 32 hours at up to 52,000 feet altitude. Eleven-foot extensions have been added to each wing, giving the Altair an overall wingspan of 86 feet with an aspect ratio of 23. It is powered by a 700-hp. rear-mounted TPE-331-10 turboprop engine, driving a three-blade propeller. Altair is scheduled to begin flight tests in the fourth quarter of 2002, and be acquired by NASA following successful completion of basic airworthiness tests in early 2003 for evaluation of over-the-horizon control, detect, see and avoid and other technologies required to allow UAVs to operate safely with other aircraft in the national airspace.

Vibroacoustic Tailoring of a Rod-Stiffened Composite Fuselage Panel with Multidisciplinary Considerations

NASA Technical Reports Server (NTRS)

Allen, Albert R.; Przekop, Adam

2015-01-01

An efficient multi-objective design tailoring procedure seeking to improve the vibroacoustic performance of a fuselage panel while maintaining or reducing weight is presented. The structure considered is the pultruded rod stitched efficient unitized structure, a highly integrated composite structure concept designed for a noncylindrical, next-generation flight vehicle fuselage. Modifications to a baseline design are evaluated within a six-parameter design space including spacing, flange width, and web height for both frame and stringer substructure components. The change in sound power radiation attributed to a design change is predicted using finite-element models sized and meshed for analyses in the 500 Hz, 1 kHz, and 2 kHz octave bands. Three design studies are carried out in parallel while considering a diffuse acoustic field excitation and two types of turbulent boundary-layer excitation. Kriging surrogate models are used to reduce the computational costs of resolving the vibroacoustic and weight objective Pareto fronts. The resulting Pareto optimal designs are then evaluated under a static pressurization ultimate load to assess structural strength and stability. Results suggest that choosing alternative configurations within the considered design space can reduce weight and improve vibroacoustic performance without compromising strength and stability of the structure under the static load condition considered, but the tradeoffs are significantly influenced by the spatial characteristics of the assumed excitation field.

West Europe Report, Science and Technology, No. 152.

DTIC Science & Technology

1983-08-12

Technology (created 3 years ago by the CGE group), who recently won a contract to supply an automated ticketing system for Buffalo’s subway . But...plastic sandwich construction. The strut junction fairing between the struts and the lower side of the wing, the upper side of the fuselage and the

Developing the MD Explorer

NASA Astrophysics Data System (ADS)

Howie, Philip V.

1993-04-01

The MD Explorer is an eight-seat twin-turbine engine helicopter which is being developed using integrated product definition (IPD) team methodology. New techniques include NOTAR antitorque system for directional control, a composite fuselage, an all-composite bearingless main rotor, and digital cockpit displays. Three-dimensional CAD models are the basis of the entire Explorer design. Solid models provide vendor with design clarification, removing much of the normal drawing interpretation errors.

Finite Element Model Development For Aircraft Fuselage Structures

NASA Technical Reports Server (NTRS)

Buehrle, Ralph D.; Fleming, Gary A.; Pappa, Richard S.; Grosveld, Ferdinand W.

2000-01-01

The ability to extend the valid frequency range for finite element based structural dynamic predictions using detailed models of the structural components and attachment interfaces is examined for several stiffened aircraft fuselage structures. This extended dynamic prediction capability is needed for the integration of mid-frequency noise control technology. Beam, plate and solid element models of the stiffener components are evaluated. Attachment models between the stiffener and panel skin range from a line along the rivets of the physical structure to a constraint over the entire contact surface. The finite element models are validated using experimental modal analysis results.

Integrated Technology Rotor Methodology Assessment Workshop Held in Moffett Field, California on 21-22 Jun 1983

DTIC Science & Technology

1988-06-01

Fuseaje mode%, 6/un/im NA 10 10 10 6/10 6/9 6/10 6/10 4/0 NA rigid body / elastic Aerodynamics on NA NA NA NA NA CP CP CP CP NA NA fuselage Pylon CP PR GDOF... motions were cyclic and body modes using the moving-block allowed to decay freely. The body pitch and roll analysis. A complete discussion of the model...conditions. Quick-acting and slow- o acting (self-centering) snubbers were intalled to 0 0 arrest the fuselage motion divergences or co lock 0 0 out body

Damage Arresting Composites

NASA Technical Reports Server (NTRS)

Jegley, Dawn C.; Davis, Pamela A.

2015-01-01

Under NASA's Environmentally Responsible Aviation Project (ERA) the most promising vehicle concepts and technologies that can simultaneously reduce aircraft fuel use, community noise, and emissions are being evaluated. Two key factors to accomplishing these goals are reducing structural weight and moving away from the traditional tube and wing aircraft configuration to a shape that has improved lift and less drag. The hybrid wing body (HWB) configuration produces more lift and less drag by smoothly joining the wings to the center fuselage section so it provides aerodynamic advantages. This shape, however, presents structural challenges with its pressurized, non-circular cabin subjected to aerodynamic flight loads. In the HWB, the structure of the center section where the passenger cabin would be located must support large in-plane loads as well as internal pressure on nearly-flat panels and right-angle joints. This structural arrangement does not lend itself to simple, efficient designs. Traditional aluminum and even state-of-the-art composites do not provide a solution to this challenge.

Assessment of the Noise Reduction Potential of Advanced Subsonic Transport Concepts for NASA's Environmentally Responsible Aviation Project

NASA Technical Reports Server (NTRS)

Thomas, Russell H.; Burley, Casey L.; Nickol, Craig L.

2016-01-01

Aircraft system noise is predicted for a portfolio of NASA advanced concepts with 2025 entry-into-service technology assumptions. The subsonic transport concepts include tube-and-wing configurations with engines mounted under the wing, over the wing nacelle integration, and a double deck fuselage with engines at a mid-fuselage location. Also included are hybrid wing body aircraft with engines upstream of the fuselage trailing edge. Both advanced direct drive engines and geared turbofan engines are modeled. Recent acoustic experimental information was utilized in the prediction for several key technologies. The 301-passenger class hybrid wing body with geared ultra high bypass engines is assessed at 40.3 EPNLdB cumulative below the Stage 4 certification level. Other hybrid wing body and unconventional tube-and-wing configurations reach levels of 33 EPNLdB or more below the certification level. Many factors contribute to the system level result; however, the hybrid wing body in the 301-passenger class, as compared to a tubeand- wing with conventional engine under wing installation, has 11.9 EPNLdB of noise reduction due to replacing reflection with acoustic shielding of engine noise sources. Therefore, the propulsion airframe aeroacoustic interaction effects clearly differentiate the unconventional configurations that approach levels close to or exceed the 42 EPNLdB goal.

Technologies for Highly Manoeuvrable Aircraft (Les Technologies pour les Aeronefs a Haute Manoeuvrabilite)

DTIC Science & Technology

1994-03-01

Demonstrator did not address increased angles of atack, Analysis predicts rcduced directial stability, at high manoeuves at high anaje of snack were a... extruded aluminium fuselage, 2.6Smm thick include variations of both slot geoetry and free alumninium sheet flying surfaces and a fineness ratio 4

Weight estimation techniques for composite airplanes in general aviation industry

NASA Technical Reports Server (NTRS)

Paramasivam, T.; Horn, W. J.; Ritter, J.

1986-01-01

Currently available weight estimation methods for general aviation airplanes were investigated. New equations with explicit material properties were developed for the weight estimation of aircraft components such as wing, fuselage and empennage. Regression analysis was applied to the basic equations for a data base of twelve airplanes to determine the coefficients. The resulting equations can be used to predict the component weights of either metallic or composite airplanes.

Optimal Design and Damage Tolerance Verification of an Isogrid Structure for Helicopter Application

NASA Technical Reports Server (NTRS)

Baker, Donald J.; Fudge, Jack; Ambur, Damodar R.; Kassapoglou, Christos

2003-01-01

A composite isogrid panel design for application to a rotorcraft fuselage is presented. An optimum panel design for the lower fuselage of the rotorcraft that is subjected to combined in-plane compression and shear loads was generated using a design tool that utilizes a smeared-stiffener theory in conjunction with a genetic algorithm. A design feature was introduced along the edges of the panel that facilitates introduction of loads into the isogrid panel without producing undesirable local bending gradients. A low-cost manufacturing method for the isogrid panel that incorporates these design details is also presented. Axial compression tests were conducted on the undamaged and low-speed impact damaged panels to demonstrate the damage tolerance of this isogrid panel. A combined loading test fixture was designed and utilized that allowed simultaneous application of compression and shear loads to the test specimen. Results from finite element analyses are presented for the isogrid panel designs and these results are compared with experimental results. This study illustrates the isogrid concept to be a viable candidate for application to the helicopter lower fuselage structure.

Response of Composite Fuselage Sandwich Side Panels Subjected to Internal Pressure and Axial Tension

NASA Technical Reports Server (NTRS)

Rouse, Marshall; Ambur, Damodar R.; Dopker, Bernard; Shah, Bharat

1998-01-01

The results from an experimental and analytical study of two composite sandwich fuselage side panels for a transport aircraft are presented. Each panel has two window cutouts and three frames and utilizes a distinctly different structural concept. These panels have been evaluated with internal pressure loads that generate biaxial tension loading conditions. Design limit load and design ultimate load tests have been performed on both panels. One of the sandwich panels was tested with the middle frame removed to demonstrate the suitability of this two-frame design for supporting the prescribed biaxial loading conditions with twice the initial frame spacing of 20 inches. A damage tolerance study was conducted on the two-frame panel by cutting a notch in the panel that originates at the edge of a cutout and extends in the panel hoop direction through the window-belt area. This panel with a notch was tested in a combined-load condition to demonstrate the structural damage tolerance at the design limit load condition. Both the sandwich panel designs successfully satisfied all desired load requirements in the experimental part of the study, and experimental results from the two-frame panel with and without damage are fully explained by the analytical results. The results of this study suggest that there is potential for using sandwich structural concepts with greater than the usual 20-in. wide frame spacing to further reduce aircraft fuselage structural weight.

Tension fracture of laminates for transport fuselage. Part 1: Material screening

NASA Technical Reports Server (NTRS)

Walker, T. H.; Avery, W. B.; Ilcewicz, L. B.; Poe, C. C., Jr.; Harris, C. E.

1992-01-01

Transport fuselage structures are designed to contain pressure following a large penetrating damage event. Applications of composites to fuselage structures require a database and supporting analysis on tension damage tolerance. Tests with 430 fracture specimens were used to accomplish the following: (1) identify critical material and laminate variables affecting notch sensitivity; (2) evaluate composite failure criteria; and (3) recommend a screening test method. Variables studied included fiber type, matrix toughness, lamination manufacturing process, and intraply hybridization. The laminates found to have the lowest notch sensitivity were manufactured using automated tow placement. This suggests a possible relationship between the stress distribution and repeatable levels of material inhomogeneity that are larger than found in traditional tape laminates. Laminates with the highest notch sensitivity consisted of toughened matrix materials that were resistant to a splitting phenomena that reduces stress concentrations in major load bearing plies. Parameters for conventional fracture criteria were found to increase with crack length for the smallest notch sizes studied. Most material and laminate combinations followed less than a square root singularity for the largest crack sizes studied. Specimen geometry, notch type, and notch size were evaluated in developing a screening test procedure. Traitional methods of correcting for specimen finite width were found to be lacking. Results indicate that a range of notch sizes must be tested to determine notch sensitivity. Data for a single small notch size (0.25 in. diameter) was found to give no indication of the sensitivity of a particular material and laminate layup to larger notch sizes.

Variable Geometry Aircraft Wing Supported by Struts And/Or Trusses

NASA Technical Reports Server (NTRS)

Melton, John E. (Inventor); Dudley, Michael R. (Inventor)

2016-01-01

The present invention provides an aircraft having variable airframe geometry for accommodating efficient flight. The aircraft includes an elongated fuselage, an oblique wing pivotally connected with said fuselage, a wing pivoting mechanism connected with said oblique wing and said fuselage, and a brace operably connected between said oblique wing and said fuselage. The present invention also provides an aircraft having an elongated fuselage, an oblique wing pivotally connected with said fuselage, a wing pivoting mechanism connected with said oblique wing and said fuselage, a propulsion system pivotally connected with said oblique wing, and a brace operably connected between said propulsion system and said fuselage.

Nondestructive Evaluation of Large Scale Composite Components.

DTIC Science & Technology

1988-01-01

primary aircraft structure. Composites are now used in critical wing and fuselage...but are still very fuctional . 28 .",’ , ’. --’. .- % .’’-,,’ . ’-.’. .- ",-’,2" - ’-.% ’_’ .’- ’" -,’ ,.,’ . ,- .’ %’%’%’% % -"%"% " "% % % ", ’ r...geometry does not permit the MAUS to scan. For those flaws which were difficult to detect or which we were not able to detect, the primary

Computing induced velocity perturbations due to a helicopter fuselage in a free stream

NASA Technical Reports Server (NTRS)

Berry, John D.; Althoff, Susan L.

1989-01-01

The velocity field of a representative helicopter fuselage in a free stream is computed. Perturbation velocities due to the fuselage are computed in a plan above the location of the helicopter rotor (rotor removed). The velocity perturbations computed by a source-panel model of the fuselage are compared with experimental measurements taken with a laser velocimeter. Three paneled fuselage models are studied: fuselage shape, fuselage shape with hub shape, and a body of revolution. The velocity perturbations computed for both fuselage shape models agree well with the measured velocity field except in the close vicinity of the rotor hub. In the hub region, without knowing the extent of separation, modeling of the effective source shape is difficult. The effects of the fuselage perturbations are not well-predicted with a simplified ellipsoid fuselage. The velocity perturbations due to the fuselage at the plane of the measurements have magnitudes of less than 8 percent of free-stream velocity. The velocity perturbations computed by the panel method are tabulated for the same locations at which previously reported rotor-inflow velocity measurements were made.

Composite structural materials

NASA Technical Reports Server (NTRS)

Ansell, G. S.; Loewy, R. G.; Wiberley, S. E.

1983-01-01

Progress and plans are reported for investigations of: (1) the mechanical properties of high performance carbon fibers; (2) fatigue in composite materials; (3) moisture and temperature effects on the mechanical properties of graphite-epoxy laminates; (4) the theory of inhomogeneous swelling in epoxy resin; (5) numerical studies of the micromechanics of composite fracture; (6) free edge failures of composite laminates; (7) analysis of unbalanced laminates; (8) compact lug design; (9) quantification of Saint-Venant's principles for a general prismatic member; (10) variation of resin properties through the thickness of cured samples; and (11) the wing fuselage ensemble of the RP-1 and RP-2 sailplanes.

Unsteady transonic potential flow over a flexible fuselage

NASA Technical Reports Server (NTRS)

Gibbons, Michael D.

1993-01-01

A flexible fuselage capability has been developed and implemented within version 1.2 of the CAP-TSD code. The capability required adding time dependent terms to the fuselage surface boundary conditions and the fuselage surface pressure coefficient. The new capability will allow modeling the effect of a flexible fuselage on the aeroelastic stability of complex configurations. To assess the flexible fuselage capability several steady and unsteady calculations have been performed for slender fuselages with circular cross-sections. Steady surface pressures are compared with experiment at transonic flight conditions. Unsteady cross-sectional lift is compared with other analytical results at a low subsonic speed and a transonic case has been computed. The comparisons demonstrate the accuracy of the flexible fuselage modifications.

Wind Tunnel Investigation of Fuselage Stability in Yaw with Various Arrangements of Fins

NASA Technical Reports Server (NTRS)

Hoggard, H Page, Jr

1940-01-01

An investigation was made in the 7-by-10 foot wind tunnel to determine the effects of dorsal-type fins and various arrangements of fins on the aerodynamic characteristics of a streamline circular fuselage. Comparative plots of the aerodynamic characteristics of the fuselage alone and the fuselage with various fin arrangements are given to show their effects on coefficients of yawing moment, drag, and lateral force. Results are also given for one case in which a rear fin on a circular fuselage was faired with modeling clay to obtain a fuselage shape with the same side elevation as the fuselage with the unfaired fin but with an elliptical cross section over the rearward portion of the fuselage. The results indicated that fin area to the rear of the center of gravity of the fuselage was beneficial in reducing the magnitude of the unstable yawing moments at large angles of yaw; whereas, fin area forward of the center of gravity was harmful. The dorsal-type fin was more effective for increasing the yawing stability of the fuselage than was a smoothly faired rearward portion with the same side elevation as the fuselage with the unfaired dorsal type fin. The minimum drag coefficient and the slope of the curve of yawing-moment coefficient of the fuselage at zero yaw were unaffected by the addition of the fins, within the experimental accuracy of the tests.

Fiber Optic Rosette Strain Gauge Development and Application on a Large-Scale Composite Structure

NASA Technical Reports Server (NTRS)

Moore, Jason P.; Przekop, Adam; Juarez, Peter D.; Roth, Mark C.

2015-01-01

A detailed description of the construction, application, and measurement of 196 FO rosette strain gauges that measured multi-axis strain across the outside upper surface of the forward bulkhead component of a multibay composite fuselage test article is presented. A background of the FO strain gauge and the FO measurement system as utilized in this application is given and results for the higher load cases of the testing sequence are shown.

An unsteady helicopter rotor: Fuselage interaction analysis

NASA Technical Reports Server (NTRS)

Lorber, Peter F.; Egolf, T. Alan

1988-01-01

A computational method was developed to treat unsteady aerodynamic interactions between a helicopter rotor, wake, and fuselage and between the main and tail rotors. An existing lifting line prescribed wake rotor analysis and a source panel fuselage analysis were coupled and modified to predict unsteady fuselage surface pressures and airloads. A prescribed displacement technique is used to position the rotor wake about the fuselage. Either a rigid blade or an aeroelastic blade analysis may be used to establish rotor operating conditions. Sensitivity studies were performed to determine the influence of the wake fuselage geometry on the computation. Results are presented that describe the induced velocities, pressures, and airloads on the fuselage and on the rotor. The ability to treat arbitrary geometries is demonstrated using a simulated helicopter fuselage. The computational results are compared with fuselage surface pressure measurements at several locations. No experimental data was available to validate the primary product of the analysis: the vibratory airloads on the entire fuselage. A main rotor-tail rotor interaction analysis is also described, along with some hover and forward flight.

Main Propulsion Test Article (MPTA)

NASA Technical Reports Server (NTRS)

Snoddy, Cynthia

2010-01-01

Scope: The Main Propulsion Test Article integrated the main propulsion subsystem with the clustered Space Shuttle Main Engines, the External Tank and associated GSE. The test program consisted of cryogenic tanking tests and short- and long duration static firings including gimbaling and throttling. The test program was conducted on the S1-C test stand (Position B-2) at the National Space Technology Laboratories (NSTL)/Stennis Space Center. 3 tanking tests and 20 hot fire tests conducted between December 21 1 1977 and December 17, 1980 Configuration: The main propulsion test article consisted of the three space shuttle main engines, flightweight external tank, flightweight aft fuselage, interface section and a boilerplate mid/fwd fuselage truss structure.

Wind tunnel investigations of glider fuselages with different waistings and wing arrangements

NASA Technical Reports Server (NTRS)

Radespiel, R.

1983-01-01

The parameters fuselage pinch glider wing arrangement and fuselage leading edge radius of nine glider configurations were investigated in wind tunnel tests. Laminar separation bubbles were found on strongly recessed fuselages. These separations in the juncture between fuselage and wing are essential in the prevention of harmful aerodynamic drag. Drag reduction was measured with increasing pinch and the wing arrangement in the rear. These results are only valid for laminar flow on the fuselage leading edge.

Impact damage resistance of composite fuselage structure, part 2

NASA Technical Reports Server (NTRS)

Dost, Ernest F.; Finn, Scott R.; Murphy, Daniel P.; Huisken, Amy B.

1993-01-01

The strength of laminated composite materials may be significantly reduced by foreign object impact induced damage. An understanding of the damage state is required in order to predict the behavior of structure under operational loads or to optimize the structural configuration. Types of damage typically induced in laminated materials during an impact event include transverse matrix cracking, delamination, and/or fiber breakage. The details of the damage state and its influence on structural behavior depend on the location of the impact. Damage in the skin may act as a soft inclusion or affect panel stability, while damage occurring over a stiffener may include debonding of the stiffener flange from the skin. An experiment to characterize impact damage resistance of fuselage structure as a function of structural configuration and impact threat was performed. A wide range of variables associated with aircraft fuselage structure such as material type and stiffener geometry (termed, intrinsic variables) and variables related to the operating environment such as impactor mass and diameter (termed, extrinsic variables) were studied using a statistically based design-of-experiments technique. The experimental design resulted in thirty-two different 3-stiffener panels. These configured panels were impacted in various locations with a number of impactor configurations, weights, and energies. The results obtained from an examination of impacts in the skin midbay and hail simulation impacts are documented. The current discussion is a continuation of that work with a focus on nondiscrete characterization of the midbay hail simulation impacts and discrete characterization of impact damage for impacts over the stiffener.

Modeling Lightning Impact Thermo-Mechanical Damage on Composite Materials

NASA Astrophysics Data System (ADS)

Muñoz, Raúl; Delgado, Sofía; González, Carlos; López-Romano, Bernardo; Wang, De-Yi; LLorca, Javier

2014-02-01

Carbon fiber-reinforced polymers, used in primary structures for aircraft due to an excellent strength-to-weight ratio when compared with conventional aluminium alloy counterparts, may nowadays be considered as mature structural materials. Their use has been extended in recent decades, with several aircraft manufacturers delivering fuselages entirely manufactured with carbon composites and using advanced processing technologies. However, one of the main drawbacks of using such composites entails their poor electrical conductivity when compared with aluminium alloy competitors that leads to lightning strikes being considered a significant threat during the service life of the aircraft. Traditionally, this problem was overcome with the use of a protective copper/bronze mesh that added additional weight and reduced the effectiveness of use of the material. Moreover, this traditional sizing method is based on vast experimental campaigns carried out by subjecting composite panels to simulated lightning strike events. While this method has proven its validity, and is necessary for certification of the structure, it may be optimized with the aid provided by physically based numerical models. This paper presents a model based on the finite element method that includes the sources of damage observed in a lightning strike, such as thermal damage caused by Joule overheating and electromagnetic/acoustic pressures induced by the arc around the attachment points. The results of the model are compared with lightning strike experiments carried out in a carbon woven composite.

Reliability, Risk and Cost Trade-Offs for Composite Designs

NASA Technical Reports Server (NTRS)

Shiao, Michael C.; Singhal, Surendra N.; Chamis, Christos C.

1996-01-01

Risk and cost trade-offs have been simulated using a probabilistic method. The probabilistic method accounts for all naturally-occurring uncertainties including those in constituent material properties, fabrication variables, structure geometry and loading conditions. The probability density function of first buckling load for a set of uncertain variables is computed. The probabilistic sensitivity factors of uncertain variables to the first buckling load is calculated. The reliability-based cost for a composite fuselage panel is defined and minimized with respect to requisite design parameters. The optimization is achieved by solving a system of nonlinear algebraic equations whose coefficients are functions of probabilistic sensitivity factors. With optimum design parameters such as the mean and coefficient of variation (representing range of scatter) of uncertain variables, the most efficient and economical manufacturing procedure can be selected. In this paper, optimum values of the requisite design parameters for a predetermined cost due to failure occurrence are computationally determined. The results for the fuselage panel analysis show that the higher the cost due to failure occurrence, the smaller the optimum coefficient of variation of fiber modulus (design parameter) in longitudinal direction.

An unsteady rotor/fuselage interaction method

NASA Technical Reports Server (NTRS)

Egolf, T. Alan; Lorber, Peter F.

1987-01-01

An analytical method has been developed to treat unsteady helicopter rotor, wake, and fuselage interaction aerodynamics. An existing lifting line/prescribed wake rotor analysis and a source panel fuselage analysis were modified to predict vibratory fuselage airloads. The analyses were coupled through the induced flow velocities of the rotor and wake on the fuselage and the fuselage on the rotor. A prescribed displacement technique was used to distort the rotor wake about the fuselage. Sensitivity studies were performed to determine the influence of wake and body geometry on the computed airloads. Predicted and measured mean and unsteady pressures on a cylindrical body in the wake of a two-bladed rotor were compared. Initial results show good qualitative agreement.

75 FR 47182 - Airworthiness Directives; GROB-WERKE GMBH & CO KG Models G102 ASTIR CS and G102 STANDARD ASTIR...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-08-05

... fuselage tail which could cause a displacement of the sailplane centre of gravity and consequently may lead... tanks could run down into the fuselage and fuselage tail which could cause a displacement of the... tanks could run down into the fuselage and fuselage tail which could cause a displacement of the...

DOE Office of Scientific and Technical Information (OSTI.GOV)

Degenhardt, R.; PFH, Private University of Applied Sciences Goettingen, Composite Engineering Campus Stade; Araujo, F. C. de

European aircraft industry demands for reduced development and operating costs. Structural weight reduction by exploitation of structural reserves in composite aerospace structures contributes to this aim, however, it requires accurate and experimentally validated stability analysis of real structures under realistic loading conditions. This paper presents different advances from the area of computational stability analysis of composite aerospace structures which contribute to that field. For stringer stiffened panels main results of the finished EU project COCOMAT are given. It investigated the exploitation of reserves in primary fibre composite fuselage structures through an accurate and reliable simulation of postbuckling and collapse. Formore » unstiffened cylindrical composite shells a proposal for a new design method is presented.« less

RLV-TD Flight Measured Aeroacoustic Levels and its Comparison with Predictions

NASA Astrophysics Data System (ADS)

Manokaran, K.; Prasath, M.; Venkata Subrahmanyam, B.; Ganesan, V. R.; Ravindran, Archana; Babu, C.

2017-12-01

The Reusable Launch Vehicle-Technology Demonstrator (RLV-TD) is a wing body configuration successfully flight tested. One of the important flight measurements is the acoustic levels. There were five external microphones, mounted on the fuselage-forebody, wing, vertical tail, inter-stage (ITS) and core base shroud to measure the acoustic levels from lift-off to splash down. In the ascent phase, core base shroud recorded the overall maximum at both lift-off and transonic conditions. In-flight noise levels measured on the wing is second highest, followed by fuselage and vertical tail. Predictions for flight trajectory compare well at all locations except for vertical tail (4.5 dB). In the descent phase, maximum measured OASPL occurs at transonic condition for the wing, followed by vertical tail and fuselage. Predictions for flight trajectory compare well at all locations except for wing (- 6.0 dB). Spectrum comparison is good in the ascent phase compared to descent phase. Roll Reaction control system (RCS) thruster firing signature is seen in the acoustic measurements on the wing and vertical tail during lift-off.

Design of Composite Structures for Reliability and Damage Tolerance

NASA Technical Reports Server (NTRS)

Rais-Rohani, Masoud

1999-01-01

A summary of research conducted during the first year is presented. The research objectives were sought by conducting two tasks: (1) investigation of probabilistic design techniques for reliability-based design of composite sandwich panels, and (2) examination of strain energy density failure criterion in conjunction with response surface methodology for global-local design of damage tolerant helicopter fuselage structures. This report primarily discusses the efforts surrounding the first task and provides a discussion of some preliminary work involving the second task.

ARC-2008-ACD08-0108-001

NASA Image and Video Library

2008-05-23

Next-Generation Aircraft, Pratt and Whitney Ultra-High Bypass Integration test at NASA Ames 11ft. wind tunnel (test 11-0182) assess the interaction effects of a scaled Pratt & Whitney geared turbofan on a Boeing 737-800 fuselage in an effort to use emerging technologies to make next-generation airliners quieter, more fuel efficient and lower on emissions. (printed in Aviation Week & Space Technology April 8, 2011 issue)

78 FR 68775 - Special Conditions: Airbus, Model A350-900 Series Airplane; Composite Fuselage In-Flight Fire...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-11-15

... Model A350-900 series airplanes. This airplane will have a novel or unusual design features associated... Model A350-900 series because of a novel or unusual design feature, special conditions are prescribed... Airbus Model A350-900 series airplane will incorporate the following novel or unusual design features...

Nonlinear Finite Element Analysis of a Composite Non-Cylindrical Pressurized Aircraft Fuselage Structure

NASA Technical Reports Server (NTRS)

Przekop, Adam; Wu, Hsi-Yung T.; Shaw, Peter

2014-01-01

The Environmentally Responsible Aviation Project aims to develop aircraft technologies enabling significant fuel burn and community noise reductions. Small incremental changes to the conventional metallic alloy-based 'tube and wing' configuration are not sufficient to achieve the desired metrics. One of the airframe concepts that might dramatically improve aircraft performance is a composite-based hybrid wing body configuration. Such a concept, however, presents inherent challenges stemming from, among other factors, the necessity to transfer wing loads through the entire center fuselage section which accommodates a pressurized cabin confined by flat or nearly flat panels. This paper discusses a nonlinear finite element analysis of a large-scale test article being developed to demonstrate that the Pultruded Rod Stitched Efficient Unitized Structure concept can meet these challenging demands of the next generation airframes. There are specific reasons why geometrically nonlinear analysis may be warranted for the hybrid wing body flat panel structure. In general, for sufficiently high internal pressure and/or mechanical loading, energy related to the in-plane strain may become significant relative to the bending strain energy, particularly in thin-walled areas such as the minimum gage skin extensively used in the structure under analysis. To account for this effect, a geometrically nonlinear strain-displacement relationship is needed to properly couple large out-of-plane and in-plane deformations. Depending on the loading, this nonlinear coupling mechanism manifests itself in a distinct manner in compression- and tension-dominated sections of the structure. Under significant compression, nonlinear analysis is needed to accurately predict loss of stability and postbuckled deformation. Under significant tension, the nonlinear effects account for suppression of the out-of-plane deformation due to in-plane stretching. By comparing the present results with the previously published preliminary linear analysis, it is demonstrated in the present paper that neglecting nonlinear effects for the structure and loads of interest can lead to appreciable loss in analysis fidelity.

Structural-acoustic coupling in aircraft fuselage structures

NASA Technical Reports Server (NTRS)

Mathur, Gopal P.; Simpson, Myles A.

1992-01-01

Results of analytical and experimental investigations of structural-acoustic coupling phenomenon in an aircraft fuselage are described. The structural and acoustic cavity modes of DC-9 fuselage were determined using a finite element approach to vibration analysis. Predicted structural and acoustic dispersion curves were used to determine possible occurrences of structural-acoustic coupling for the fuselage. An aft section of DC-9 aircraft fuselage, housed in an anechoic chamber, was used for experimental investigations. The test fuselage was excited by a shaker and vibration response and interior sound field were measured using accelerometer and microphone arrays. The wavenumber-frequency structural and cavity response maps were generated from the measured data. Analysis and interpretation of the spatial plots and wavenumber maps provided the required information on modal characteristics, fuselage response and structural-acoustic coupling.

Application of Carbon Fibre Truss Technology to the Fuselage Structure of the SKYLON Spaceplane

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

A reusable SSTO spaceplane employing dual mode airbreathing/rocket engines, such as SKYLON, has a voluminous fuselage in order to accommodate the considerable quantities of hydrogen fuel needed for the ascent. The loading intensity which this fuselage has to withstand is relatively low due to the modest in-flight inertial accelerations coupled with the very low density of liquid hydrogen. Also the requirement to accommo- date considerable temperature differentials between the internal cryogenic tankage and the aerodynamically heated outer skin of the vehicle imposes an additional design constraint that results in an optimum fuselage structural concept very different to conventional aircraft or rocket practice. Several different structural con- cepts exist for the primary loadbearing structure. This paper explores the design possibilities of the various types and explains why an independent near ambient temperature CFRP truss structure was selected for the SKYLON vehicle. The construction of such a truss structure, at a scale not witnessed since the days of the airship, poses a number of manufacturing and design difficulties. In particular the construction of the nodes and their attachment to the struts is considered to be a key issue. This paper describes the current design status of the overall truss geometry, strut construction and manufacturing route, and the final method of assembly. The results of a preliminary strut and node test programme are presented which give confidence that the design targets will eventually be met.

Characteristics of Flow Past Fuselages and Wing-Fuselage Systems of Gliders

NASA Technical Reports Server (NTRS)

Ostrowski, Jerzy; Litwinczyk, Mieczyslaw; Turkowski, Lukasz

1980-01-01

The results are presented for visualization tests and measurements of the velocity field in diffusion regions (with a positive pressure gradient) for fuselages and transition regions between the wing and the fuselage. Wind tunnel and flight tests were performed. Specific emphasis was placed on examining the secondary flow influencing separation acceleration and the influence of the geometrical form of the wing fuselage system manifested by the occurrence of secondary flows of various types.

Full-Scale Test and Analysis Results of a PRSEUS Fuselage Panel to Assess Damage Containment Features

NASA Technical Reports Server (NTRS)

Bergan, Andrew; Bakuckas, John G., Jr.; Lovejoy, Andrew; Jegley, Dawn; Linton, Kim; Neal, Bert; Korkosz, Gregory; Awerbuch, Jonathan; Tan, Tein-Min

2012-01-01

Integrally stitched composite technology is an area that shows promise in enhancing the structural integrity of aircraft and aerospace structures. The most recent generation of this technology is the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept. The goal of the PRSEUS concept relevant to this test is to provide damage containment capability for composite structures while reducing overall structural weight. The National Aeronautics and Space Administration (NASA), the Federal Aviation Administration (FAA), and The Boeing Company have partnered in an effort to assess the damage containment features of a full-scale curved PRSEUS panel using the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. A single PRSEUS test panel was subjected to axial tension, internal pressure, and combined axial tension and internal pressure loads. The test results showed excellent performance of the PRSEUS concept. No growth of Barely Visible Impact Damage (BVID) was observed after ultimate loads were applied. With a two-bay notch severing the central stringer, damage was contained within the two-bay region well above the required limit load conditions. Catastrophic failure was well above the ultimate load level. Information describing the test panel and procedure has been previously presented, so this paper focuses on the experimental procedure, test results, nondestructive inspection results, and preliminary test and analysis correlation.

Program user's manual for an unsteady helicopter rotor-fuselage aerodynamic analysis

NASA Technical Reports Server (NTRS)

Lorber, Peter F.

1988-01-01

The Rotor-Fuselage Analysis is a method of calculating the aerodynamic reaction between a helicopter rotor and fuselage. This manual describes the structure and operation of the computer programs that make up the Rotor-Fuselage Analysis, programs which prepare the input and programs which display the output.

Overview Of Structural Behavior and Occupant Responses from Crash Test of a Composite Airplane

NASA Technical Reports Server (NTRS)

Jones, Lisa E.; Carden, Huey D.

1995-01-01

As part of NASA's composite structures crash dynamics research, a general aviation aircraft with composite wing, fuselage and empennage (but with metal subfloor structure) was crash tested at the NASA Langley Research Center Impact Research Facility. The test was conducted to determine composite aircraft structural behavior for crash loading conditions and to provide a baseline for a similar aircraft test with a modified subfloor. Structural integrity and cabin volume were maintained. Lumbar loads for dummy occupants in energy absorbing seats wer substantially lower than those in standard aircraft seats; however, loads in the standard seats were much higher that those recorded under similar conditions for an all-metallic aircraft.

Relationship between mechanical-property and energy-absorption trends for composite tubes

NASA Technical Reports Server (NTRS)

Farley, Gary L.

1992-01-01

U.S. Army helicopters are designed to dissipate prescribed levels of crash impact kinetic energy without compromising the integrity of the fuselage. Because of the complexity of the energy-absorption process it is imperative for designers of energy-absorbing structures to develop an in-depth understanding of how and why composite structures absorb energy. A description of the crushing modes and mechanisms of energy absorption for composite tubes and beams is presented. Three primary crushing modes of composite structures including transverse shearing, lamina bending, and local buckling are described. The experimental data presented show that fiber and matrix mechanical properties and laminate stiffness and strength mechanical properties cannot reliably predict the energy-absorption response of composite tubes.

Wind-tunnel investigation of effect of interference on lateral-stability characteristics of four NACA 23012 wings, an elliptical and a circular fuselage and vertical fins

NASA Technical Reports Server (NTRS)

House, Rufus O; Wallace, Arthur R

1941-01-01

Report presents the results of a wind-tunnel investigation of the effect of wing-fuselage interference on lateral-stability characteristics made in the NACA 7 by 10-foot wind tunnel on four fuselages and two fins, representing high-wing, low-wing, and midwing monoplanes. The fuselages are of circular and elliptical cross section. The wings have rounded tips and, in plan form, one is rectangular and the three are tapered 3:1 with various amounts of sweep. The rate of change in the coefficients of rolling moment, yawing moment, and lateral force with angle of yaw is given in a form to show the increment caused by wing-fuselage interference for the model with no fin and the effect of wing-fuselage interference on fin effectiveness. Results for the fuselage-fin combination and the wing tested alone are also given.

Fuselage shell and cavity response measurements on a DC-9 test section

NASA Technical Reports Server (NTRS)

Simpson, M. A.; Mathur, G. P.; Cannon, M. R.; Tran, B. N.; Burge, P. L.

1991-01-01

A series of fuselage shell and cavity response measurements conducted on a DC-9 aircraft test section are described. The objectives of these measurements were to define the shell and cavity model characteristics of the fuselage, understand the structural-acoustic coupling characteristics of the fuselage, and measure the response of the fuselage to different types of acoustic and vibration excitation. The fuselage was excited with several combinations of acoustic and mechanical sources using interior and exterior loudspeakers and shakers, and the response to these inputs was measured with arrays of microphones and accelerometers. The data were analyzed to generate spatial plots of the shell acceleration and cabin acoustic pressure field, and corresponding acceleration and pressure wavenumber maps. Analysis and interpretation of the spatial plots and wavenumber maps provided the required information on modal characteristics, structural-acoustic coupling, and fuselage response.

Study for prediction of rotor/wake/fuselage interference, part 1

NASA Technical Reports Server (NTRS)

Clark, D. R.; Maskew, B.

1985-01-01

A method was developed which allows the fully coupled calculation of fuselage and rotor airloads for typical helicopter configurations in forward flight. To do this, an iterative solution is carried out based on a conventional panel representation of the fuselage and a blade element representation of the rotor where fuselage and rotor singularity strengths are determined simultaneously at each step and the rotor wake is allowed to relax (deform) in response to changes in rotor wake loading and fuselage presence. On completion of the iteration, rotor loading and inflow, fuselage singularity strength (and, hence, pressure and velocity distributions) and rotor wake are all consistent. The results of a fully coupled calculation of the flow around representative helicopter configurations are presented. The effect of fuselage components on the rotor flow field and the overall wake structure is detailed and the aerodynamic interference between the different parts of the aircraft is discussed.

Progress Towards Fuselage Drag Reduction via Active Flow Control: A Combined CFD and Experimental Effort

NASA Technical Reports Server (NTRS)

Schaeffler, Norman W.; Allan, Brian G.; Lienard, Caroline; LePape, Arnaud

2010-01-01

A combined computational and experimental effort has been undertaken to study fuselage drag reduction on a generic, non-proprietary rotorcraft fuselage by the application of active ow control. Fuselage drag reduction is an area of research interest to both the United States and France and this area is being worked collaboratively as a task under the United States/France Memorandum of Agreement on Helicopter Aeromechanics. In the first half of this task, emphasis is placed on the US generic fuselage, the ROBIN-mod7, with the experimental work being conducted on the US side and complementary US and French CFD analysis of the baseline and controlled cases. Fuselage simulations were made using Reynolds-averaged Navier-Stokes ow solvers and with multiple turbulence models. Comparisons were made to experimental data for numerical simulations of the isolated fuselage and for the fuselage as installed in the tunnel, which includes modeling of the tunnel contraction, walls, and support fairing. The numerical simulations show that comparisons to the experimental data are in good agreement when the tunnel and model support are included. The isolated fuselage simulations compare well to each other, however, there is a positive shift in the centerline pressure when compared to the experiment. The computed flow separation locations on the rear ramp region had only slight differences with and without the tunnel walls and model support. For the simulations, the flow control slots were placed at several locations around the flow separation lines as a series of eight slots that formed a nearly continuous U-shape. Results from the numerical simulations resulted in an estimated 35% fuselage drag reduction from a steady blowing flow control configuration and a 26% drag reduction for unsteady zero-net-mass flow control configuration. Simulations with steady blowing show a delayed flow separation at the rear ramp of the fuselage that increases the surface pressure acting on the ramp, thus decreasing the overall fuselage pressure drag.

Field-incidence transmission of treated orthotropic and laminated composite panels

NASA Technical Reports Server (NTRS)

Koval, L. R.

1983-01-01

In an effort to improve understanding of the phenomenon of noise transmission through the sidewalls of an aircraft fuselage, an analytical model was developed for the field incidence transmission loss of an orthotropic or laminated composite infinite panel with layers of various noise insulation treatments. The model allows for four types of treatments, impervious limp septa, orthotropic trim panels, porous blankets, and air spaces, while it also takes into account the effects of forward speed. Agreement between the model and transmission loss data for treated panels is seen to be fairly good overall. In comparison with transmission loss data for untreated composite panels, excellent agreement occurred.

Nonlinear Deformation and Stability of a Noncircular Cylindrical Shell Under Combined Loading with Bending and Twisting Moments

NASA Astrophysics Data System (ADS)

Belov, V. K.; Zheleznov, L. P.; Ognyanova, T. S.

2018-03-01

A previously developed technique is used to solve problems of strength and stability of discretely reinforced noncircular cylindrical shells made of a composite material with allowance for the moments and nonlinearity of their subcritical stress-strain state. Stability of a reinforced bay of the aircraft fuselage made of a composite material under combined loading with bending and twisting moments is studied. The effects of straining nonlinearity, stiffness of longitudinal ribs, and shell thickness on the critical loads that induce shell buckling are analyzed.

Safe-life and damage-tolerant design approaches for helicopter structures

NASA Technical Reports Server (NTRS)

Reddick, H. K., Jr.

1983-01-01

The safe-life and damage-tolerant design approaches discussed apply to both metallic and fibrous composite helicopter structures. The application of these design approaches to fibrous composite structures is emphasized. Safe-life and damage-tolerant criteria are applied to all helicopter flight critical components, which are generally categorized as: dynamic components with a main and tail rotor system, which includes blades, hub and rotating controls, and drive train which includes transmission, and main and interconnecting rotor shafts; and the airframe, composed of the fuselage, aerodynamic surfaces, and landing gear.

Finite Element Model Development and Validation for Aircraft Fuselage Structures

NASA Technical Reports Server (NTRS)

Buehrle, Ralph D.; Fleming, Gary A.; Pappa, Richard S.; Grosveld, Ferdinand W.

2000-01-01

The ability to extend the valid frequency range for finite element based structural dynamic predictions using detailed models of the structural components and attachment interfaces is examined for several stiffened aircraft fuselage structures. This extended dynamic prediction capability is needed for the integration of mid-frequency noise control technology. Beam, plate and solid element models of the stiffener components are evaluated. Attachment models between the stiffener and panel skin range from a line along the rivets of the physical structure to a constraint over the entire contact surface. The finite element models are validated using experimental modal analysis results. The increased frequency range results in a corresponding increase in the number of modes, modal density and spatial resolution requirements. In this study, conventional modal tests using accelerometers are complemented with Scanning Laser Doppler Velocimetry and Electro-Optic Holography measurements to further resolve the spatial response characteristics. Whenever possible, component and subassembly modal tests are used to validate the finite element models at lower levels of assembly. Normal mode predictions for different finite element representations of components and assemblies are compared with experimental results to assess the most accurate techniques for modeling aircraft fuselage type structures.

Amphibious Airplane

NASA Technical Reports Server (NTRS)

1986-01-01

The airplane pictured is the new Air Shark I, a four-place amphibian that makes extensive use of composite materials and cruises at close to 200 miles per hour under power from a 200-horsepower engine. Air Shark I is a "homebuilt" airplane, assembled from a kit of parts and components furnished by Freedom Master Corporation, Satellite Beach, Florida. The airplane incorporates considerable NASA technology and its construction benefited from research assistance provided by Kennedy Space Center (KSC) In designing the Shark, company president Arthur M. Lueck was able to draw on NASA's aeronautical technology bank through KSC's computerized "recon" library. As a result of his work at KSC, the wing of the Air Shark I is a new airfoil developed by Langley Research Center for light aircraft. In addition, Lueck opted for NASA-developed "winglets," vertical extensions of the wing that reduce drag by smoothing air turbulence at the wingtips. The NASA technology bank also contributed to the hull design. Lueck is considering application of NASA laminar flow technology-means of smoothing the airflow over wing and fuselage-to later models for further improvement of the Shark's aerodynamic efficiency. A materials engineer, Lueck employed his own expertise in designing and selecting the materials for the composite segments, which include all structural members, exposed surfaces and many control components. The materials are fiber reinforced plastics, or FRP They offer a high strength-to-weight ratio, with a nominal strength rating about one and a half times that of structural steel. They provide other advantages: the materials can be easily molded into finished shapes without expensive tooling or machining, and they are highly corrosion resistant. The first homebuilt to be offered by Freedom Master, Air Shark I completed air and water testing in mid-1985 and the company launched production of kits.

Caracterisation experimentale de la transmission acoustique de structures aeronautiques

NASA Astrophysics Data System (ADS)

Pointel, Vincent

Le confort des passagers à l'intérieur des avions pendant le vol est un axe en voie d'amélioration constante. L'augmentation de la proportion des matériaux composites dans la fabrication des structures aéronautiques amène de nouvelles problématiques à résoudre. Le faible amortissement de ces structures, en contre partie de leur poids/raideur faible, est non favorable sur le plan acoustique, ce qui oblige les concepteurs à devoir trouver des moyens d'amélioration. De plus, les mécanismes de transmission du son au travers d'un système double paroi de type aéronautique ne sont pas complètement compris, c'est la raison qui motive cette étude. L'objectif principal de ce projet est de constituer une base de données pour le partenaire industriel de ce projet : Bombardier Aéronautique. En effet, les données expérimentales de performance d'isolation acoustique, de systèmes complets représentatifs d'un fuselage d'avion sont très rares dans la littérature scientifique. C'est pourquoi une méthodologie expérimentale est utilisée dans ce projet. Deux conceptions différentes de fuselage sont comparées. La première possède une peau (partie extérieure du fuselage) métallique raidie, alors que la deuxième est constituée d'un panneau sandwich composite. Dans les deux cas, un panneau de finition de fabrication sandwich est utilisé. Un traitement acoustique en laine de verre est placé à l'intérieur de chacun des fuselages. Des isolateurs vibratoires sont utilisés pour connecter les deux panneaux du fuselage. La simulation en laboratoire de la couche limite turbulente, qui est la source d'excitation prépondérante pendant la phase de vol, n'est pas encore possible hormis en soufflerie. C'est pourquoi deux cas d'excitation sont considérés pour essayer d'approcher cette sollicitation : une excitation mécanique (pot vibrant) et une acoustique (champ diffus). La validation et l'analyse des résultats sont effectuées par le biais des logiciels NOVA et VAONE, utilisés par le partenaire industriel de ce projet. Un des objectifs secondaires est de valider le modèle double paroi implémenté dans NOVA. L'investigation de l'effet de compression local du traitement acoustique, sur la perte par transmission d'une simple paroi, montre que cette action n'a aucun effet bénéfique notable. D'autre part, il apparaît que la raideur des isolateurs vibratoires a un lien direct avec les performances d'isolation du système double paroi. Le système double paroi avec peau composite semble moins sensible à ce paramètre. Le modèle double paroi de NOVA donne de bons résultats concernant le système double paroi avec une peau métallique. Des écarts plus importants sont observés en moyennes et hautes fréquences dans le cas du système avec une peau composite. Cependant, la bonne tendance de la prédiction au vu de la complexité de la structure est plutôt prometteuse. Mots-clés : Expérimental, avion, double paroi, panneaux composites, perte par transmission, transmission aérienne, transmission solidienne, isolateur vibratoire.

Analysis and Design of Fuselage Structures Including Residual Strength Prediction Methodology

NASA Technical Reports Server (NTRS)

Knight, Norman F.

1998-01-01

The goal of this research project is to develop and assess methodologies for the design and analysis of fuselage structures accounting for residual strength. Two primary objectives are included in this research activity: development of structural analysis methodology for predicting residual strength of fuselage shell-type structures; and the development of accurate, efficient analysis, design and optimization tool for fuselage shell structures. Assessment of these tools for robustness, efficient, and usage in a fuselage shell design environment will be integrated with these two primary research objectives.

Static and fatigue testing of full-scale fuselage panels fabricated using a Therm-X(R) process

NASA Technical Reports Server (NTRS)

Dinicola, Albert J.; Kassapoglou, Christos; Chou, Jack C.

1992-01-01

Large, curved, integrally stiffened composite panels representative of an aircraft fuselage structure were fabricated using a Therm-X process, an alternative concept to conventional two-sided hard tooling and contour vacuum bagging. Panels subsequently were tested under pure shear loading in both static and fatigue regimes to assess the adequacy of the manufacturing process, the effectiveness of damage tolerant design features co-cured with the structure, and the accuracy of finite element and closed-form predictions of postbuckling capability and failure load. Test results indicated the process yielded panels of high quality and increased damage tolerance through suppression of common failure modes such as skin-stiffener separation and frame-stiffener corner failure. Finite element analyses generally produced good predictions of postbuckled shape, and a global-local modelling technique yielded failure load predictions that were within 7% of the experimental mean.

Reusable Space Vehicle Ground Operations Baseline Conceptual Model

DTIC Science & Technology

2004-03-01

Micro .........................................................................129 Figure B-91. Lower Forward Fuselage Macro...129 Figure B-92. Lower Forward Fuselage Micro ...Fuselage Micro ................................................................130 Figure B-95. Lower Wing Macro

Effects of boundary layer refraction and fuselage scattering on fuselage surface noise from advanced turboprop propellers

NASA Technical Reports Server (NTRS)

Mcaninch, G. L.; Rawls, J. W., Jr.

1984-01-01

An acoustic disturbance's propagation through a boundary layer is discussed with a view to the analysis of the acoustic field generated by a propfan rotor incident to the fuselage of an aircraft. Applying the parallel flow assumption, the resulting partial differential equations are reduced to an ordinary acoustic pressure differential equation by means of the Fourier transform. The methods used for the solution of this equation include those of Frobenius and of analytic continuation; both yield exact solutions in series form. Two models of the aircraft fuselage-boundary layer system are considered, in the first of which the fuselage is replaced by a flat plate and the acoustic field is assumed to be two-dimensional, while in the second the fuselage is a cylinder in a fully three-dimensional acoustic field. It is shown that the boundary layer correction improves theory-data comparisons over simple application of a pressure-doubling rule at the fuselage.

Study for prediction of rotor/wake/fuselage interference. Part 2: Program users guide

NASA Technical Reports Server (NTRS)

Clark, D. R.; Maskew, B.

1985-01-01

A method was developed which permits the fully coupled calculation of fuselage and rotor airloads for typical helicopter configurations in forward flight. To do this, an iterative solution is carried out based on a conventional panel representation of the fuselage and a blade element representation of the rotor where fuselage and rotor singularity strengths are determined simultaneously at each step and the rotor wake is allowed to relax (deform) in response to changes in rotor wake loading and fuselage presence. On completion of the iteration, rotor loading and inflow, fuselage singularity strength (and, hence, pressure and velocity distributions) and rotor wake are all consistent. The results of a fully coupled calculation of the flow around representative helicopter configurations are presented. The effect of fuselage components on the rotor flow field and the overall wake structure is discussed as well as the aerodynamic interference between the different parts of the aircraft. Details of the computer program are given.

77 FR 46185 - United States v. United Technologies Corporation and Goodrich Corporation; Proposed Final...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-02

... for generating power for all the in-flight systems that run on electricity, including pumping breathable air into the fuselage, operating the lights, and running the navigation and communication... turning a propeller blade on a turboprop engine, a rotor shaft on a turboshaft engine, or a fan in front...

Fuselage mounted anti-collision lights utilizing high power LEDs

NASA Astrophysics Data System (ADS)

Lundberg, John; Machi, Nicolo; Mangum, Scott; Singer, Jeffrey

2005-09-01

As LEDs continue to improve in efficacy and total light output, they are increasingly finding their way in to new applications in the aviation industry as well as adjacent markets. One function that is particularly challenging and may reap substantial benefits from this new technology is the fuselage mounted anti-collision light. Anti-collision lights provide conspicuity for the aircraft by periodically emitting bright flashes of light. The color, light distribution and intensity levels for these lights are all closely regulated through Federal Aviation Regulation (FAR) documents. These lighting requirements, along with thermal, environmental and aerodynamic requirements, drive the overall design. In this paper, we will discuss the existing technologies used in anti-collision lights and the advantages and challenges associated with an LED solution. Particular attention will be given to the optical, thermal, electrical and aerodynamic aspects associated with an LED approach. A specific case study will be presented along with some of the challenges that have arisen during the design process. These challenges include the addition of an integrated covert anti-collision lighting.

Damage Arresting Composites for Shaped Vehicles - Phase II Final Report

NASA Technical Reports Server (NTRS)

Velicki, Alex; Yovanof, Nicolette; Baraja, Jaime; Linton, Kim; Li, Victor; Hawley, Arthur; Thrash, Patrick; DeCoux, Steve; Pickell, Robert

2011-01-01

This report describes the development of a novel structural concept, Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS), that addresses the demanding fuselage loading requirements for the Hybrid Wing or Blended Wing Body (BWB) airplane configuration. In addition to the analytical studies, a three specimen test program was also completed to assess the concept under axial tension loading, axial compression loading, and internal pressure loading.

Interior and exterior fuselage noise measured on NASA's C-8a augmentor wing jet-STOL research aircraft

NASA Technical Reports Server (NTRS)

Shovlin, M. D.

1977-01-01

Interior and exterior fuselage noise levels were measured on NASA's C-8A Augmentor Wing Jet-STOL Research Aircraft in order to provide design information for the Quiet Short-Haul Research Aircraft (QSRA), which will use a modified C-8A fuselage. The noise field was mapped by 11 microphones located internally and externally in three areas: mid-fuselage, aft fuselage, and on the flight deck. Noise levels were recorded at four power settings varying from takeoff to flight idle and were plotted in one-third octave band spectra. The overall sound pressure levels of the external noise field were compared to previous tests and found to correlate well with engine primary thrust levels. Fuselage values were 145 + or - 3 dB over the aircraft's normal STOL operating range.

Al-Li Alloy 1441 for Fuselage Applications

NASA Technical Reports Server (NTRS)

Bird, R. K.; Dicus, D. L.; Fridlyander, J. N.; Sandler, V. S.

2000-01-01

A cooperative investigation was conducted to evaluate Al-Cu-Mg-Li alloy 1441 for long service life fuselage applications. Alloy 1441 is currently being used for fuselage applications on the Russian Be-103 amphibious aircraft, and is expected to be used for fuselage skin on a new Tupolev business class aircraft. Alloy 1441 is cold-rollable and has several attributes that make it attractive for fuselage skin applications. These attributes include lower density and higher specific modulus with similar strength as compared to conventional Al-Cu-Mg alloys. Cold-rolled 1441 Al-Li sheet specimens were tested at NASA Langley Research Center (LaRC) and at the All-Russia Institute of Aviation Materials (VIAM) in Russia to evaluate tensile properties, fracture toughness, impact resistance, fatigue life and fatigue crack growth rate. In addition, fuselage panels were fabricated by Tupolev Design Bureau (TDB) using 1441 skins and Al-Zn-Mg-Cu alloy stiffeners. The panels were subjected to cyclic pressurization fatigue tests at TDB and at LaRC to simulate fuselage pressurization/depressurization during aircraft service. This paper discusses the results from this investigation.

Rotor-Fuselage Interaction: Analysis and Validation with Experiment

NASA Technical Reports Server (NTRS)

Berry, John D.; Bettschart, Nicolas

1997-01-01

The problem of rotor-fuselage aerodynamic interaction has to be considered in industry applications from various aspects. First, in order to increase helicopter speed and reduce operational costs, rotorcraft tend to be more and more compact, with a main rotor closer to the fuselage surface. This creates significant perturbations both on the main rotor and on the fuselage, including steady and unsteady effects due to blade and wake passage and perturbed inflow at the rotor disk. Furthermore,the main rotor wake affects the tail boom, empennage and anti-torque system. This has important consequences for helicopter control and vibrations at low speeds and also on tail rotor acoustics (main rotor wake-tail rotor interactions). This report describes the US Army-France MOD cooperative work on this problem from both the theoretical and experimental aspects. Using experimental 3D velocity field and fuselage surface pressure measurements, three codes that model the interactions of a helicopter rotor with a fuselage are compared. These comparisons demonstrate some of the strengths and weaknesses of current models for the combined rotor-fuselage analysis.

Pressure Distribution Over the Fuselage of a PW-9 Pursuit Airplane in Flight

NASA Technical Reports Server (NTRS)

Rhode, Richard V; Lundquist, Eugene E

1932-01-01

This report presents the results obtained from pressure distribution tests on the fuselage of a PW-9 pursuit airplane in a number of conditions of flight. The investigation was made to determine the contribution of the fuselage to the total lift in conditions considered critical for the wing structure, and also to determine whether the fuselage loads acting simultaneously with the maximum tail loads were of such a character as to be of concern with respect to the structural design of other parts of the airplane. The results show that the contribution of the fuselage toward the total lift is small on this airplane. Aerodynamic loads on the fuselage are, in general, unimportant from the structural viewpoint, and in most cases they are of such character that, if neglected, a conservative design results. In spins, aerodynamic forces on the fuselage produce diving moments of appreciable magnitude and yawing moments of small magnitude, but opposing the rotation of the airplane. A table of cowling pressures for various maneuvers is included in the report.

Analytical and experimental investigation of aircraft metal structures reinforced with filamentary composites. Phase 3: Major component development

NASA Technical Reports Server (NTRS)

Bryson, L. L.; Mccarty, J. E.

1973-01-01

Analytical and experimental investigations, performed to establish the feasibility of reinforcing metal aircraft structures with advanced filamentary composites, are reported. Aluminum-boron-epoxy and titanium-boron-epoxy were used in the design and manufacture of three major structural components. The components were representative of subsonic aircraft fuselage and window belt panels and supersonic aircraft compression panels. Both unidirectional and multidirectional reinforcement concepts were employed. Blade penetration, axial compression, and inplane shear tests were conducted. Composite reinforced structural components designed to realistic airframe structural criteria demonstrated the potential for significant weight savings while maintaining strength, stability, and damage containment properties of all metal components designed to meet the same criteria.

Taxiing, Take-Off, and Landing Simulation of the High Speed Civil Transport Aircraft

NASA Technical Reports Server (NTRS)

Reaves, Mercedes C.; Horta, Lucas G.

1999-01-01

The aircraft industry jointly with NASA is studying enabling technologies for higher speed, longer range aircraft configurations. Higher speeds, higher temperatures, and aerodynamics are driving these newer aircraft configurations towards long, slender, flexible fuselages. Aircraft response during ground operations, although often overlooked, is a concern due to the increased fuselage flexibility. This paper discusses modeling and simulation of the High Speed Civil Transport aircraft during taxiing, take-off, and landing. Finite element models of the airframe for various configurations are used and combined with nonlinear landing gear models to provide a simulation tool to study responses to different ground input conditions. A commercial computer simulation program is used to numerically integrate the equations of motion and to compute estimates of the responses using an existing runway profile. Results show aircraft responses exceeding safe acceptable human response levels.

MM-122: High speed civil transport

NASA Technical Reports Server (NTRS)

Demarest, Bill; Anders, Kurt; Manchec, John; Yang, Eric; Overgaard, Dan; Kalkwarf, Mike

1992-01-01

The rapidly expanding Pacific Rim market along with other growing markets indicates that the future market potential for a high speed civil transport is great indeed. The MM-122 is the answer to the international market desire for a state of the art, long range, high speed civil transport. It will carry 250 passengers a distance of 5200 nm at over twice the speed of sound. The MM-122 is designed to incorporate the latest technologies in the areas of control systems, propulsions, aerodynamics, and materials. The MM-122 will accomplish these goals using the following design parameters. First, a double delta wing planform with highly swept canards and an appropriately area ruled fuselage will be incorporated to accomplish desired aerodynamic characteristics. Propulsion will be provided by four low bypass variable cycle turbofan engines. A quad-redundant fly-by-wire flight control system will be incorporated to provide appropriate static stability and level 1 handling qualities. Finally, the latest in conventional metallic and modern composite materials will be used to provide desired weight and performance characteristics. The MM-122 incorporates the latest in technology and cost minimization techniques to provide a viable solution to this future market potential.

Coupled rotor and fuselage equations of motion

NASA Technical Reports Server (NTRS)

Warmbrodt, W.

1979-01-01

The governing equations of motion of a helicopter rotor coupled to a rigid body fuselage are derived. A consistent formulation is used to derive nonlinear periodic coefficient equations of motion which are used to study coupled rotor/fuselage dynamics in forward flight. Rotor/fuselage coupling is documented and the importance of an ordering scheme in deriving nonlinear equations of motion is reviewed. The nature of the final equations and the use of multiblade coordinates are discussed.

The vibration characteristics of a coupled helicopter rotor-fuselage by a finite element analysis

NASA Technical Reports Server (NTRS)

Rutkowski, M. J.

1983-01-01

The dynamic coupling between the rotor system and the fuselage of a simplified helicopter model in hover was analytically investigated. Mass, aerodynamic damping, and elastic and centrifugal stiffness matrices are presented for the analytical model; the model is based on a beam finite element, with polynomial mass and stiffness distributions for both the rotor and fuselage representations. For this analytical model, only symmetric fuselage and collective blade degrees of freedom are treated. Real and complex eigen-analyses are carried out to obtain coupled rotor-fuselage natural modes and frequencies as a function of rotor speed. Vibration response results are obtained for the coupled system subjected to a radially uniform, harmonic blade loading. The coupled response results are compared with response results from an uncoupled analysis in which hub loads for an isolated rotor system subjected to the same sinusoidal blade loading as the coupled system are applied to a free-free fuselage.

Evaluation of a metal fuselage frame selectively reinforced with filamentary composites for space shuttle application

NASA Technical Reports Server (NTRS)

Oken, S.; Skoumal, D. E.; Straayer, J. W.

1974-01-01

The development of metal structures reinforced with filamentary composites as a weight saving feature of the space shuttle components is discussed. A frame was selected for study that was representative of the type of construction used in the bulk frames of the orbiter vehicle. Theoretical and experimental investigations were conducted. Component tests were performed to evaluate the critical details used in the designs and to provide credibility to the weight saving results. A model frame was constructed of the reinforced metal material to provide a final evaluation of the construction under realistic load conditions.

Investigation of Fuselage Structure Subject to Widespread Fatigue Damage

DOT National Transportation Integrated Search

1996-01-01

This report documents the results of the "Investigation of Fuselage Structure Subject to Widespread Fatigue Damage" contract. The primary program objective was to obtain data on airplane fuselage structures subject to multiple site damage (MSD) in an...

The SKYLON Spaceplane

NASA Astrophysics Data System (ADS)

Varvill, R.; Bond, A.

SKYLON is a single stage to orbit (SSTO) winged spaceplane designed to give routine low cost access to space. At a gross takeoff weight of 275 tonnes of which 220 tonnes is propellant the vehicle is capable of placing 12 tonnes into an equatorial low Earth orbit. The vehicle configuration consists of a slender fuselage containing the propellant tankage and payload bay with delta wings located midway along the fuselage carrying the SABRE engines in axisymmetric nacelles on the wingtips. The vehicle takes off and lands horizontally on it's own undercarriage. The fuselage is constructed as a multilayer structure consisting of aeroshell, insulation, structure and tankage. SKYLON employs extant or near term materials technology in order to minimise development cost and risk. The SABRE engines have a dual mode capability. In rocket mode the engine operates as a closed cycle liquid oxygen/liquid hydrogen high specific impulse rocket engine. In airbreathing mode (from takeoff to Mach 5) the liquid oxygen flow is replaced by atmospheric air, increasing the installed specific impulse 3-6 fold. The airflow is drawn into the engine via a 2 shock axisymmetric intake and cooled to cryogenic temperatures prior to compression. The hydrogen fuel flow acts as a heat sink for the closed cycle helium loop before entering the main combustion chamber.

Wind tunnel investigation of helicopter-rotor wake effects on three helicopter fuselage models

NASA Technical Reports Server (NTRS)

Wilson, J. C.; Mineck, R. E.

1975-01-01

The effects of rotor wake on helicopter fuselage aerodynamic characteristics were investigated in the Langley V/STOL tunnel. Force, moment, and pressure data were obtained on three fuselage models at various combinations of windspeed, sideslip angle, and pitch angle. The data show that the influence of rotor wake on the helicopter fuselage yawing moment imposes a significant additional thrust requirement on the tail rotor of a single-rotor helicopter at high sideslip angles.

Structural Concepts Study of Non-circular Fuselage Configurations

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivel

1996-01-01

A preliminary study of structural concepts for noncircular fuselage configurations is presented. For an unconventional flying-wing type aircraft, in which the fuselage is inside the wing, multiple fuselage bays with non-circular sections need to be considered. In a conventional circular fuselage section, internal pressure is carried efficiently by a thin skin via hoop tension. If the section is non-circular, internal pressure loads also induce large bending stresses. The structure must also withstand additional bending and compression loads from aerodynamic and gravitational forces. Flat and vaulted shell structural configurations for such an unconventional, non-circular pressurized fuselage of a large flying-wing were studied. A deep honeycomb sandwich-shell and a ribbed double-wall shell construction were considered. Combinations of these structural concepts were analyzed using both analytical and simple finite element models of isolated sections for a comparative conceptual study. Weight, stress, and deflection results were compared to identify a suitable configuration for detailed analyses. The flat sandwich-shell concept was found preferable to the vaulted shell concept due to its superior buckling stiffness. Vaulted double-skin ribbed shell configurations were found to be superior due to their weight savings, load diffusion, and fail-safe features. The vaulted double-skin ribbed shell structure concept was also analyzed for an integrated wing-fuselage finite element model. Additional problem areas such as wing-fuselage junction and pressure-bearing spar were identified.

Design, Fabrication and Test of Composite Curved Frames for Helicopter Fuselage Structure

NASA Technical Reports Server (NTRS)

Lowry, D. W.; Krebs, N. E.; Dobyns, A. L.

1984-01-01

Aspects of curved beam effects and their importance in designing composite frame structures are discussed. The curved beam effect induces radial flange loadings which in turn causes flange curling. This curling increases the axial flange stresses and induces transverse bending. These effects are more important in composite structures due to their general inability to redistribute stresses by general yielding, such as in metal structures. A detailed finite element analysis was conducted and used in the design of composite curved frame specimens. Five specimens were statically tested and compared with predicted and test strains. The curved frame effects must be accurately accounted for to avoid premature fracture; finite element methods can accurately predict most of the stresses and no elastic relief from curved beam effects occurred in the composite frames tested. Finite element studies are presented for comparative curved beam effects on composite and metal frames.

Technology development status at McDonnell Douglas

NASA Technical Reports Server (NTRS)

Rowe, W. T.

1981-01-01

The significant technology items of the Concorde and the conceptual MCD baseline advanced supersonic transport are compared. The four major improvements are in the areas of range performance, structures (materials), aerodynamics, and in community noise. Presentation charts show aerodynamic efficiency; the reoptimized wing; low scale lift/drag ratio; control systems; structural modeling and analysis; weight and cost comparisons for superplasticity diffusion bonded titanium sandwich structures and for aluminum brazed titanium honeycomb structures; operating cost reduction; suppressor nozzles; noise reduction and range; the bicone inlet; a market summary; environmental issues; high priority items; the titanium wing and fuselage test components; and technology validation.

Dual-band infrared (DBIR) imaging inspections of Boeing 737 and KC-135 aircraft panels

DOE Office of Scientific and Technical Information (OSTI.GOV)

Del Grande, N.K.; Dolan, K.W.; Durbin, P.F.

1993-08-27

We apply dual-band infrared (DBIR) imaging as a dynamic thermal tomography tool for wide area inspection of a Boeing 737 aircraft, and several Boeing KC-135 aircraft panels. Our analyses are discussed in this report. After flash-heating the aircraft skin, we record synchronized DBIR images every 40 ms, from onset to 8 seconds after the heat flash. We analyze selective DBIR image ratios which enhance surface temperature contrast and remove surface-emissivity clutter (from dirt, dents, tape, markings, ink, sealants, uneven paint, paint stripper, exposed metal and roughness variations). The Boeing 737 and KC-135 aircraft fuselage panels have varying percent thickness lossesmore » from corrosion. We established the correlation of percent thickness loss with surface temperature rise (above ambient) for a partially corroded F-18 wing box structure and several aluminum reference panels. Based on this correlation, lap splice temperatures rise 1{degrees}C per 24 {plus_minus} 5 % material loss at 0.4 s after the heat flash. We show tables, charts and temperature maps of typical lap splice material losses for the riveted (and bonded) Boeing 737, and the riveted (but unbonded) Boeing KC-135. We map the fuselage composite thermal inertia, based on the (inverse) slope of the surface temperature versus inverse square root of time. Composite thermal inertia maps characterize shallow skin defects within the lap splice at early times (<0.3 s) and deeper skin defects within the lap splice at late times (>0.4 s). Late time composite thermal inertia maps depict where corrosion-related thickness losses occur. Lap splice sites on a typical Boeing KC-135 panel with low composite thermal inertia values had high skin-thickness losses from corrosion.« less

Dynamic thermal tomography for nondestructive inspection of aging aircraft

DOE Office of Scientific and Technical Information (OSTI.GOV)

Del Grande, N.K.; Dolan, K.W.; Durbin, P.F.

1993-11-01

The authors apply dual-band infrared (DBIR) imaging as a dynamic thermal tomography tool for wide area inspection of a Boeing 737 aircraft and several Boeing KC-135 aircraft panels. The analyses are discussed in this report. After flash-heating the aircraft skin, they record synchronized DBIR images every 40 ms, from onset to 8 seconds after the heat flash. They analyze selective DBIR image ratios which enhance surface temperature contrast and remove surface-emissivity clutter. The Boeing 737 and KC-135 aircraft fuselage panels have varying percent thickness losses from corrosion. They established the correlation of percent thickness loss with surface temperature rise (abovemore » ambient) for a partially corroded F-18 wing box structure and several aluminum plates which had 6 to 60% thickness losses at milled flat-bottom hole sites. Based on this correlation, lap splice temperatures rise 1C per 24 {plus_minus} 5% material loss at 0.4 s after the heat flash. They tabulate and map corrosion-related percent thickness loss effects for the riveted Boeing 737, and the riveted Boeing KKC-135. They map the fuselage composite thermal inertia, based on the (inverse) slope of the surface temperature versus inverse square root of time. Composite thermal inertia maps characterized shallow skin defects within the lap splice at early times (< 0.3 s) and deeper skin defects within the lap splice at late times (> 0.4 s). Late time composite thermal inertia maps depict where corrosion-related thickness losses occur (e.g., on the inside of the Boeing 737 lap splice, beneath the galley and the latrine). Lap splice sites on a typical Boeing KC-135 panel with low composite thermal inertia values had high skin-thickness losses from corrosion.« less

Drop test analysis of fuselage section of R80 commuter aircraft by using finite element method

NASA Astrophysics Data System (ADS)

Anggono, Agus Dwi; Ardianto, Adik Nofa Rochma Wahyu

2017-04-01

In commercial aerospace development, feasibility accidents design or crashworthiness is a major concern in aviation safety. Fuselage structure plays an important role in absorbing energy during an accident. The research aims are to determine drop test phenomenon on the fuselage, to investigate deformation occurred in the structure of the fuselage, and to know the influence of the airframe falls position to the stress strain which occurred in the structure of the fuselage. This research was conducted by varying the fall angle of the fuselage in a vertical position or 0° and 15°. Fuselage design was modeled by using SolidWorks. Then the model is imported to the Abaqus for drop test simulation. From the simulation results, it can be obtained the phenomenon of deformation on the structure of the fuselage when it comes in contact with the rigid ground. The high deformation occurs shows the structure capabilities in order to absorb the impact. It could be happened because the deformation is influenced by internal energy and strain energy. The various positions shows the structure capability in order to withstand impact loads during periods of 4-8 seconds and the maximum deformation was reached in 12 seconds. The experiment on the vertical position and the position falls of 15° angle was delivered the highest stress strain. The stress was 483 MPa in struts section, 400.78 MPa in skin section, 358.28 MPa in the floor and 483 MPa in the cargo frame section.

Evaluation of structural design concepts for an arrow-wing supersonic cruise aircraft

NASA Technical Reports Server (NTRS)

Sakata, I. F.; Davis, G. W.

1977-01-01

An analytical study was performed to determine the best structural approach for design of primary wing and fuselage structure of a Mach 2.7 arrow wing supersonic cruise aircraft. Concepts were evaluated considering near term start of design. Emphasis was placed on the complex interactions between thermal stress, static aeroelasticity, flutter, fatigue and fail safe design, static and dynamic loads, and the effects of variations in structural arrangements, concepts and materials on these interactions. Results indicate that a hybrid wing structure incorporating low profile convex beaded and honeycomb sandwich surface panels of titanium alloy 6Al-4V were the most efficient. The substructure includes titanium alloy spar caps reinforced with boron polyimide composites. The fuselage shell consists of hat stiffened skin and frame construction of titanium alloy 6Al-4V. A summary of the study effort is presented, and a discussion of the overall logic, design philosophy and interaction between the analytical methods for supersonic cruise aircraft design are included.

Lessons in the Design and Characterization Testing of the Semi-Span Super-Sonic Transport (S4T) Wind-Tunnel Model

NASA Technical Reports Server (NTRS)

2012-01-01

This paper focuses on some of the more challenging design processes and characterization tests of the Semi-Span Super-Sonic Transport (S4T)-Active Controls Testbed (ACT). The model was successfully tested in four entries in the National Aeronautics and Space Administration Langley Transonic Dynamics Tunnel to satisfy the goals and objectives of the Fundamental Aeronautics Program Supersonic Project Aero-Propulso-Servo-Elastic effort. Due to the complexity of the S4T-ACT, only a small sample of the technical challenges for designing and characterizing the model will be presented. Specifically, the challenges encountered in designing the model include scaling the Technology Concept Airplane to model scale, designing the model fuselage, aileron actuator, and engine pylons. Characterization tests included full model ground vibration tests, wing stiffness measurements, geometry measurements, proof load testing, and measurement of fuselage static and dynamic properties.

Conceptual Design of a Single-Aisle Turboelectric Commercial Transport With Fuselage Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Welstead, Jason R.; Felder, James L.

2016-01-01

A single-aisle commercial transport concept with a turboelectric propulsion system architecture was developed assuming entry into service in 2035 and compared to a similar technology conventional configuration. The turboelectric architecture consisted of two underwing turbofans with generators extracting power from the fan shaft and sending it to a rear fuselage, axisymmetric, boundary layer ingesting fan. Results indicate that the turbo- electric concept has an economic mission fuel burn reduction of 7%, and a design mission fuel burn reduction of 12% compared to the conventional configuration. An exploration of the design space was performed to better understand how the turboelectric architecture changes the design space, and system sensitivities were run to determine the sensitivity of thrust specific fuel consumption at top of climb and propulsion system weight to the motor power, fan pressure ratio, and electrical transmission efficiency of the aft boundary layer ingesting fan.

General Aerodynamic Characteristics of a Research Model with High Disk Loading Direct Lifting Fan Mounted in Fuselage

NASA Image and Video Library

1960-10-26

3/4 Low front view of fuselage and fan. Showing jet engine hanging below. Lift fan powered by jet exhaust. General Aerodynamic Characteristics of a Research Model with High Disk Loading Direct Lifting Fan Mounted in Fuselage

Behavior Of Aircraft Components Under Crash-Type Loads

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Boitnott, Richard L.; Fasanella, Edwin L.

1993-01-01

Report presents overview of research involving use of concepts of aircraft elements and substructures not necessarily designed or optimized with respect to energy-absorption or crash-loading considerations. Experimental and analytical data presented in report indicate some general trends in failure behaviors of class of composite-material structures including individual fuselage frames, skeleton subfloors with stringers and floor beams but without skin covering, and subfloors with skin added to frame/stringer arrangement.

Effects of Hydrogen Peroxide on Common Aviation Textiles

DTIC Science & Technology

2009-08-01

efficacious (complete kill of 106 CFU of the spore forming Geobacillus stearothermophilus ) in a narrow-body aircraft fuselage (3), as well as wide-body...disinfectant/ sterilant for transportation vehicles like aircraft, buses, subway trains, ambulances, etc. Although the biological efficacy of STERIS...hydrogen peroxide (VHP)1 technology is of particular interest due to rapid sterilization , easy usage, intrinsic environmental friendliness (i .e

Interference of Wing and Fuselage from Tests of 209 Combinations in the NACA Variable-Density Tunnel

NASA Technical Reports Server (NTRS)

Jacobs, Eastman N; Ward, Kenneth E

1936-01-01

This report presents the results of tests of 209 simple wing-fuselage combinations made in the NACA variable-density wind tunnel to provide information regarding the effects of aerodynamic interference between wings and fuselages at a large value of Reynolds number.

VIEW OF BOEING 737200 FUSELAGE FROM TOP LEVEL OF TAIL ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

VIEW OF BOEING 737-200 FUSELAGE FROM TOP LEVEL OF TAIL DOCK AND. A NEW SAFETY CABLE FROM THE TAIL DOCK WILL ALLOW INSPECTORS TO WALK UP AND DOWN THE FUSELAGE TO CHECK FOR CRACKS OR MISSING FASTENERS. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

The "Turkey Buzzard" glider

NASA Technical Reports Server (NTRS)

Miller, Roy G; Brown, D T

1923-01-01

The "Turkey Buzzard" is a semi-internally braced monoplane (Fig. 1). The wing is placed above the fuselage for two important aerodynamical reasons: first, because this position minimizes the mutual interference between the wing and the fuselage, and, second, useful lifting surface is utilized with the wing passing over the fuselage instead of through it.

Aircraft Configured for Flight in an Atmosphere Having Low Density

NASA Technical Reports Server (NTRS)

Teter, Jr., John E. (Inventor); Croom, Mark A. (Inventor); Smith, Stephen C. (Inventor); Gelhausen, Paul A. (Inventor); Hunter, Craig A. (Inventor); Riddick, Steven E. (Inventor); Guynn, Mark D. (Inventor); Paddock, David A. (Inventor)

2012-01-01

An aircraft is configured for flight in an atmosphere having a low density. The aircraft includes a fuselage, a pair of wings, and a rear stabilizer. The pair of wings extends from the fuselage in opposition to one another. The rear stabilizer extends from the fuselage in spaced relationship to the pair of wings. The fuselage, the wings, and the rear stabilizer each present an upper surface opposing a lower surface. The upper and lower surfaces have X, Y, and Z coordinates that are configured for flight in an atmosphere having low density.

Crack Growth Simulation and Residual Strength Prediction in Airplane Fuselages

NASA Technical Reports Server (NTRS)

Chen, Chuin-Shan; Wawrzynek, Paul A.; Ingraffea, Anthony R.

1999-01-01

The objectives were to create a capability to simulate curvilinear crack growth and ductile tearing in aircraft fuselages subjected to widespread fatigue damage and to validate with tests. Analysis methodology and software program (FRANC3D/STAGS) developed herein allows engineers to maintain aging aircraft economically, while insuring continuous airworthiness, and to design more damage-tolerant aircraft for the next generation. Simulations of crack growth in fuselages were described. The crack tip opening angle (CTOA) fracture criterion, obtained from laboratory tests, was used to predict fracture behavior of fuselage panel tests. Geometrically nonlinear, elastic-plastic, thin shell finite element crack growth analyses were conducted. Comparisons of stress distributions, multiple stable crack growth history, and residual strength between measured and predicted results were made to assess the validity of the methodology. Incorporation of residual plastic deformations and tear strap failure was essential for accurate residual strength predictions. Issue related to predicting crack trajectory in fuselages were also discussed. A directional criterion, including T-stress and fracture toughness orthotropy, was developed. Curvilinear crack growth was simulated in coupon and fuselage panel tests. Both T-stress and fracture toughness orthotropy were essential to predict the observed crack paths. Flapping of fuselages were predicted. Measured and predicted results agreed reasonable well.

Study to investigate design, fabrication and test of low cost concepts for large hybrid composite helicopter fuselage, phase 2

NASA Technical Reports Server (NTRS)

Adams, K. M.; Lucas, J. J.

1977-01-01

The development of a frame/stringer/skin fabrication technique for composite airframe construction was studied as a low cost approach to the manufacturer of larger helicopter airframe components. A center cabin aluminum airframe section of the Sikorsky CH-53D, was selected for evaluation as a composite structure. The design, as developed, is composed of a woven KEVLAR R-49/epoxy skin and graphite/epoxy frames and stringers. The single cure concept is made possible by the utilization of pre-molded foam cores, over which the graphite/epoxy pre-impregnated frame and stringer reinforcements are positioned. Bolted composite channel sections were selected as the optimum joint construction. The applicability of the single cure concept to larger realistic curved airframe sections, and the durability of the composite structure in a realistic spectrum fatigue environment, was described.

Flight service evaluation of an advanced composite empennage component on commercial transport aircraft. Phase 1: Engineering development

NASA Technical Reports Server (NTRS)

Ary, A.; Axtell, C.; Fogg, L.; Jackson, A.; James, A. M.; Mosesian, B.; Vanderwier, J.; Vanhamersveld, J.

1976-01-01

The empennage component selected for this program is the vertical fin box of the L-1011 aircraft. The box structure extends from the fuselage production joint to the tip rib and includes the front and rear spars. Various design options were evaluated to arrive at a configuration which would offer the highest potential for satisfying program objectives. The preferred configuration selected consists of a hat-stiffened cover with molded integrally stiffened spars, aluminum trussed composite ribs, and composite miniwich web ribs with integrally molded caps. Material screening tests were performed to select an advanced composite material system for the Advanced Composite Vertical Fin (ACFV) that would meet the program requirements from the standpoint of quality, reproducibility, and cost. Preliminary weight and cost analysis were made, targets established, and tracking plans developed. These include FAA certification, ancillary test program, quality control, and structural integrity control plans.

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 5: Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Magliozzi, B.; Hanson, D. B.

1991-01-01

An analysis of tone noise propagation through a boundary layer and fuselage scattering effects was derived. This analysis is a three dimensional and the complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The outer wave field is constructed analytically from an incident wave appropriate to the source and a scattered wave in the standard Hankel function form. For the incident wave, an existing function - domain propeller noise radiation theory is used. In the boundary layer region, the wave equation is solved by numerical methods. The theoretical analysis is embodied in a computer program which allows the calculation of correction factors for the fuselage scattering and boundary layer refraction effects. The effects are dependent on boundary layer profile, flight speed, and frequency. Corrections can be derived for any point on the fuselage, including those on the opposite side from the source. The theory was verified using limited cases and by comparing calculations with available measurements from JetStar tests of model prop-fans. For the JetStar model scale, the boundary layer refraction effects produce moderate fuselage pressure reinforcements aft of and near the plane of rotation and significant attenuation forward of the plane of rotation at high flight speeds. At lower flight speeds, the calculated boundary layer effects result in moderate amplification over the fuselage area of interest. Apparent amplification forward of the plane of rotation is a result of effective changes in the source directivity due to boundary layer refraction effects. Full scale effects are calculated to be moderate, providing fuselage pressure amplification of about 5 dB at the peak noise location. Evaluation using available noise measurements was made under high-speed, high-altitude flight conditions. Comparisons of calculations made of free field noise, using a current frequency-domain propeller noise prediction method, and fuselage effects using this new procedure show good agreement with fuselage measurements over a wide range of flight speeds and frequencies. Correction factors for the JetStar measurements made on the fuselage are provided in an Appendix.

Lightning protection technology for small general aviation composite material aircraft

NASA Technical Reports Server (NTRS)

Plumer, J. A.; Setzer, T. E.; Siddiqi, S.

1993-01-01

An on going NASA (Small Business Innovative Research) SBIR Phase II design and development program will produce the first lightning protected, fiberglass, General Aviation aircraft that is available as a kit. The results obtained so far in development testing of typical components of the aircraft kit, such as the wing and fuselage panels indicate that the lightning protection design methodology and materials chosen are capable of protecting such small composite airframes from lightning puncture and structural damage associated with severe threat lightning strikes. The primary objective of the program has been to develop a lightening protection design for full scale test airframe and verify its adequacy with full scale laboratory testing, thus enabling production and sale of owner-built, lightning-protected, Stoddard-Hamilton Aircraft, Inc. Glasair II airplanes. A second objective has been to provide lightning protection design guidelines for the General Aviation industry, and to enable these airplanes to meet lightening protection requirements for certification of small airplanes. This paper describes the protection design approaches and development testing results obtained thus far in the program, together with design methodology which can achieve the design goals listed above. The presentation of this paper will also include results of some of the full scale verification tests, which will have been completed by the time of this conference.

77 FR 48423 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-14

... Company Model 737-500 series airplanes. This AD was prompted by reports of chem-mill step cracking on the aft lower lobe fuselage skins. This AD requires inspections of the fuselage skin at the chem- mill... 22686). That NPRM proposed to require inspections of the fuselage skin at the chem-mill steps, and...

Surface generation and editing operations applied to structural support of aerospace vehicle fuselages. M.S. Thesis

NASA Technical Reports Server (NTRS)

Schwartz, Susan K.

1992-01-01

The Solid Modeling Aerospace Research Tool (SMART) is a computer aided design tool used in aerospace vehicle design. Modeling of structural components using SMART includes the representation of the transverse or cross-wise elements of a vehicle's fuselage, ringframes, and bulkheads. Ringframes are placed along a vehicle's fuselage to provide structural support and maintain the shape of the fuselage. Bulkheads are also used to maintain shape, but are placed at locations where substantial structural support is required. Given a Bezier curve representation of a cross sectional cut through a vehicle's fuselage and/or an internal tank, this project produces a first-guess Bezier patch representation of a ringframe or bulkhead at the cross-sectional position. The grid produced is later used in the structural analysis of the vehicle. The graphical display of the generated patches allows the user to edit patch control points in real time. Constraints considered in the patch generation include maintaining 'square-like' patches and placement of longitudinal, or lengthwise along the fuselage, structural elements called longerons.

On the prediction of auto-rotational characteristics of light airplane fuselages

NASA Technical Reports Server (NTRS)

Pamadi, B. N.; Taylor, L. W., Jr.

1984-01-01

A semi-empirical theory is presented for the estimation of aerodynamic forces and moments acting on a steadily rotating (spinning) airplane fuselage, with a particular emphasis on the prediction of its auto-rotational behavior. This approach is based on an extension of the available analytical methods for high angle of attack and side-slip and then coupling this procedure with strip theory for application to a rotating airplane fuselage. The analysis is applied to the fuselage of a light general aviation airplane and the results are shown to be in fair agreement with experimental data.

Aerodynamic Interaction Effects of a Helicopter Rotor and Fuselage

NASA Technical Reports Server (NTRS)

Boyd, David D., Jr.

1999-01-01

A three year Cooperative Research Agreements made in each of the three years between the Subsonic Aerodynamics Branch of the NASA Langley Research Center and the Virginia Polytechnic Institute and State University (Va. Tech) has been completed. This document presents results from this three year endeavor. The goal of creating an efficient method to compute unsteady interactional effects between a helicopter rotor and fuselage has been accomplished. This paper also includes appendices to support these findings. The topics are: 1) Rotor-Fuselage Interactions Aerodynamics: An Unsteady Rotor Model; and 2) Rotor/Fuselage Unsteady Interactional Aerodynamics: A New Computational Model.

Noise-reduction measurements of stiffened and unstiffened cylindrical models of an airplane fuselage

NASA Technical Reports Server (NTRS)

Willis, C. M.; Mayes, W. H.

1984-01-01

Noise-reduction measurements are presented for a stiffened and an unstiffened model of an airplane fuselage. The cylindrical models were tested in a reverberant-field noise environment over a frequency range from 20 Hz to 6 kHz. An unstiffened metal fuselage provided more noise reduction than a fuselage having the same sidewall weight divided between skin and stiffening stringers and ring frames. The addition of acoustic insulation to the models tended to smooth out the interior-noise spectrum by reducing or masking the noise associated with the structural response at some of the resonant frequencies.

Coupled rotor/fuselage dynamic analysis of the AH-1G helicopter and correlation with flight vibrations data

NASA Technical Reports Server (NTRS)

Corrigan, J. C.; Cronkhite, J. D.; Dompka, R. V.; Perry, K. S.; Rogers, J. P.; Sadler, S. G.

1989-01-01

Under a research program designated Design Analysis Methods for VIBrationS (DAMVIBS), existing analytical methods are used for calculating coupled rotor-fuselage vibrations of the AH-1G helicopter for correlation with flight test data from an AH-1G Operational Load Survey (OLS) test program. The analytical representation of the fuselage structure is based on a NASTRAN finite element model (FEM), which has been developed, extensively documented, and correlated with ground vibration test. One procedure that was used for predicting coupled rotor-fuselage vibrations using the advanced Rotorcraft Flight Simulation Program C81 and NASTRAN is summarized. Detailed descriptions of the analytical formulation of rotor dynamics equations, fuselage dynamic equations, coupling between the rotor and fuselage, and solutions to the total system of equations in C81 are included. Analytical predictions of hub shears for main rotor harmonics 2p, 4p, and 6p generated by C81 are used in conjunction with 2p OLS measured control loads and a 2p lateral tail rotor gearbox force, representing downwash impingement on the vertical fin, to excite the NASTRAN model. NASTRAN is then used to correlate with measured OLS flight test vibrations. Blade load comparisons predicted by C81 showed good agreement. In general, the fuselage vibration correlations show good agreement between anslysis and test in vibration response through 15 to 20 Hz.

Crash Simulation of a Vertical Drop Test of a B737 Fuselage Section with Overhead Bins and Luggage

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.

2004-01-01

The focus of this paper is to describe a crash simulation of a 30-ft/s vertical drop test of a Boeing 737 (B737) fuselage section. The drop test of the 10-ft. long fuselage section of a B737 aircraft was conducted in November of 2000 at the FAA Technical Center in Atlantic City, NJ. The fuselage section was outfitted with two different commercial overhead stowage bins. In addition, 3,229-lbs. of luggage were packed in the cargo hold to represent a maximum take-off weight condition. The main objective of the test was to evaluate the response and failure modes of the overhead stowage bins in a narrow-body transport fuselage section when subjected to a severe, but survivable, impact. A secondary objective of the test was to generate experimental data for correlation with the crash simulation. A full-scale 3-dimensional finite element model of the fuselage section was developed and a crash simulation was conducted using the explicit, nonlinear transient dynamic code, MSC.Dytran. Pre-test predictions of the fuselage and overhead bin responses were generated for correlation with the drop test data. A description of the finite element model and an assessment of the analytical/experimental correlation are presented. In addition, suggestions for modifications to the model to improve correlation are proposed.

Assessing inspection sensitivity as it relates to damage tolerance in composite rotor hubs

NASA Astrophysics Data System (ADS)

Roach, Dennis P.; Rackow, Kirk

2001-08-01

Increasing niche applications, growing international markets, and the emergence of advanced rotorcraft technology are expected to greatly increase the population of helicopters over the next decade. In terms of fuselage fatigue, helicopters show similar trends as fixed-wing aircraft. The highly unsteady loads experienced by rotating wings not only directly affect components in the dynamic systems but are also transferred to the fixed airframe structure. Expanded use of rotorcraft has focused attention on the use of new materials and the optimization of maintenance practices. The FAA's Airworthiness Assurance Center (AANC) at Sandia National Labs has joined with Bell Helicopter andother agencies in the rotorcraft industry to evaluate nondestructive inspection (NDI) capabilities in light of the damage tolerance of assorted rotorcraft structure components. Currently, the program's emphasis is on composite rotor hubs. The rotorcraft industry is constantly evaluating new types of lightweight composite materials that not only enhance the safety and reliability of rotor components but also improve performance and extended operating life as well. Composite rotor hubs have led to the use of bearingless rotor systems that are less complex and require less maintenance than their predecessors. The test facility described in this paper allows the structural stability and damage tolerance of composite hubs to be evaluated using realistic flight load spectrums of centrifugal force and bending loads. NDI was integrated into the life-cycle fatigue tests in order to evaluate flaw detection sensitivity simultaneously wiht residual strength and general rotor hub peformance. This paper will describe the evolving use of damage tolerance analysis (DTA) to direct and improve rotorcraft maintenance along with the related use of nondestructive inspections to manage helicopter safety. OVeralll, the data from this project will provide information to improve the producibility, inspectability, serviceability, and cost effectively of rotorcraft components.

Prediction and Measurement of the Vibration and Acoustic Radiation of Panels Subjected to Acoustic Loading

NASA Technical Reports Server (NTRS)

Turner, Travis L.; Rizzi, Stephen A.

1995-01-01

Interior noise and sonic fatigue are important issues in the development and design of advanced subsonic and supersonic aircraft. Conventional aircraft typically employ passive treatments, such as constrained layer damping and acoustic absorption materials, to reduce the structural response and resulting acoustic levels in the aircraft interior. These techniques require significant addition of mass and only attenuate relatively high frequency noise transmitted through the fuselage. Although structural acoustic coupling is in general very important in the study of aircraft fuselage interior noise, analysis of noise transmission through a panel supported in an infinite rigid baffle (separating two semi-infinite acoustic domains) can be useful in evaluating the effects of active/adaptive materials, complex loading, etc. Recent work has been aimed at developing adaptive and/or active methods of controlling the structural acoustic response of panels to reduce the transmitted noise1. A finite element formulation was recently developed to study the dynamic response of shape memory alloy (SMA) hybrid composite panels (conventional composite panel with embedded SMA fibers) subject to combined acoustic and thermal loads2. Further analysis has been performed to predict the far-field acoustic radiation using the finite element dynamic panel response prediction3. The purpose of the present work is to validate the panel vibration and acoustic radiation prediction methods with baseline experimental results obtained from an isotropic panel, without the effect of SMA.

75 FR 4480 - Airworthiness Directives; BAE Systems (Operations) Limited Model BAe 146 and Avro 146-RJ Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-28

... condition as: Reports have been received of finding corrosion at the Frame 29 wing-to-fuselage attachment... structural integrity of Frame 29 and the wing-to-fuselage attachment. * * * * * The unsafe condition is degradation of the structural integrity of Frame 29 and the wing-to-fuselage attachment, which could result in...

Some studies on the aerodynamic effect of the gap between airplane wings and fuselages

NASA Technical Reports Server (NTRS)

Ober, Shatswell

1929-01-01

The general result indicated by this study is that if desirable from any viewpoint the gap between wing and fuselage may be closed without detrimental aerodynamic effects, and with a given monoplane there is less drag if the wing is directly on top of the fuselage than if it is parasol.

Sound Modeling Simplifies Vehicle Noise Management

NASA Technical Reports Server (NTRS)

2015-01-01

Under two SBIR contracts with Langley Research Center, Ann Arbor, Michigan-based Comet Technology Corporation developed Comet EnFlow, a software program capable of predicting both high- and low-frequency noise and vibration behavior in plane fuselages and other structures. The company now markets the software to airplane, automobile, and ship manufacturers, and Langley has found an unexpected use for it in leak detection on the International Space Station.

Simulation of Flow around Isolated Helicopter Fuselage

NASA Astrophysics Data System (ADS)

Kusyumov, A. N.; Mikhailov, S. A.; Romanova, E. V.; Garipov, A. O.; Nikolaev, E. I.; Barakos, G.

2013-04-01

Low fuselage drag has always been a key target of helicopter manufacturers. Therefore, this paper focuses on CFD predictions of the drag of several components of a typical helicopter fuselage. In the first section of the paper, validation of the obtained CFD predictions is carried out using wind tunnel measurements. The measurements were carried out at the Kazan National Research Technical University n.a. A. Tupolev. The second section of the paper is devoted to the analysis of drag contributions of several components of the ANSAT helicopter prototype fuselage using the RANS approach. For this purpose, several configurations of fuselages are considered with different levels of complexity including exhausts and skids. Depending on the complexity of the considered configuration and CFD mesh both the multi-block structured HMB solver and the unstructured commercial tool Fluent are used. Finally, the effect of an actuator disk on the predicted drag is addressed.

Evaluation of Pressurization Fatigue Life of 1441 Al-li Fuselage Panel

NASA Technical Reports Server (NTRS)

Bird, R. Keith; Dicus, Dennis I.; Fridlyander, Joseph; Davydov, Valentin

1999-01-01

A study was conducted to evaluate the pressurization fatigue life of fuselage panels with skins fabricated from 1441 Al-Li, an attractive new Russian alloy. The study indicated that 1441 Al-Li has several advantages over conventional aluminum fuselage skin alloy with respect to fatigue behavior. Smooth 1441 Al-Li sheet specimens exhibited a fatigue endurance limit similar to that for 1163 Al (Russian version of 2024 Al) sheet. Notched 1441 Al-Li sheet specimens exhibited greater fatigue strength and longer fatigue life than 1163 Al. In addition, Tu-204 fuselage panels fabricated by Tupolev Design Bureau using Al-Li skin and ring frames with riveted 7000-series aluminum stiffeners had longer pressurization fatigue lives than did panels constructed from conventional aluminum alloys. Taking into account the lower density of this alloy, the results suggest that 1441 Al-Li has the potential to improve fuselage performance while decreasing structural weight.

Failure behavior of generic metallic and composite aircraft structural components under crash loads

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Robinson, Martha P.

1990-01-01

Failure behavior results are presented from crash dynamics research using concepts of aircraft elements and substructure not necessarily designed or optimized for energy absorption or crash loading considerations. To achieve desired new designs incorporating improved energy absorption capabilities often requires an understanding of how more conventional designs behave under crash loadings. Experimental and analytical data are presented which indicate some general trends in the failure behavior of a class of composite structures including individual fuselage frames, skeleton subfloors with stringers and floor beams without skin covering, and subfloors with skin added to the frame-stringer arrangement. Although the behavior is complex, a strong similarity in the static/dynamic failure behavior among these structures is illustrated through photographs of the experimental results and through analytical data of generic composite structural models.

76 FR 54399 - Airworthiness Directives; The Boeing Company Model 737-200, -200C, -300, -400, and -500 Series...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-09-01

... result in sudden fracture and failure of the fuselage skin panels, and consequent rapid decompression of..., which could result in sudden fracture and failure of the fuselage skin panels, and consequent rapid... fuselage skin panels at the chem-milled steps, which could result in sudden fracture and failure of the...

78 FR 40640 - Airworthiness Directives; Agusta S.p.A. (Type Certificate Currently Held by AgustaWestland S.p.A...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-08

... AW139 helicopters. The existing AD currently requires inspecting the fuselage frame to detect fatigue... interval for inspecting the fuselage frame for a fatigue crack. This proposed AD would require inspecting... detect a fatigue crack that could result in failure of the fuselage frame and subsequent loss of control...

An analysis of available data on effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes

NASA Technical Reports Server (NTRS)

Wollner, Bertram C

1949-01-01

Available information on the effects of wing-fuselage-tail and wing-nacelle interference on the distribution of the air load among components of airplanes is analyzed. The effects of wing and nacelle incidence, horizontal andvertical position of wing and nacelle, fuselage shape, wing section and filleting are considered. Where sufficient data were unavailable to determine the distribution of the air load, the change in lift caused by interference between wing and fuselage was found. This increment is affected to the greatest extent by vertical wing position.

Fuselage ventilation due to wind flow about a postcrash aircraft

NASA Technical Reports Server (NTRS)

Stuart, J. W.

1980-01-01

Postcrash aircraft fuselage fire development, dependent on the internal and external fluid dynamics is discussed. The natural ventilation rate, a major factor in the internal flow patterns and fire development is reviewed. The flow about the fuselage as affected by the wind and external fire is studied. An analysis was performend which estimated the rates of ventilation produced by the wind for a limited idealized environmental configuration. The simulation utilizes the empirical pressure coefficient distribution of an infinite circular cylinder near a wall with its boundary later flow to represent the atmospheric boundary layer. The resulting maximum ventilation rate for two door size openings, with varying circumferential location in a common 10 mph wind was an order of magnitude greater than the forced ventilation specified in full scale fire testing. The parameter discussed are: (1) fuselage size and shape, (2) fuselage orientation and proximity to the ground, (3) fuselage-openings size and location, (4) wind speed and direction, and (5) induced flow of the external fire plume is recommended. The fire testing should be conducted to a maximum ventilation rate at least an order of magnitude greater than the inflight air conditioning rates.

Secondary Wing System for Use on an Aircraft

NASA Technical Reports Server (NTRS)

Smith, Brian E. (Inventor)

1999-01-01

A secondary wing system for use on an aircraft augments the lift, stability, and control of the aircraft at subsonic speeds. The secondary wing system includes a mechanism that allows the canard to be retracted within the contour of the aircraft fuselage from an operational position to a stowed position. The top surface of the canard is exposed to air flow in the stowed position, and is contoured to integrate aerodynamically and smoothly within the contour of the fuselage when the canard is retracted for high speed flight. The bottom portion of the canard is substantially flat for rotation into a storage recess within the fuselage. The single canard rotates about a vertical axis at its spanwise midpoint. The canard can be positioned between a range of sweep angles during flight and a stowed position in which its span is substantially parallel to the aircraft fuselage. The canard can be deployed and retracted during flight. The deployment mechanism includes a circular mounting ring and drive mechanism that connects the canard with the fuselage and permits it to rotate and to change incidence. The deployment mechanism further includes retractable fairings which serve to streamline the wing when it is retracted into the top of the fuselage.

Crack Growth Simulation and Residual Strength Prediction in Airplane Fuselages

NASA Technical Reports Server (NTRS)

Chen, Chuin-Shan; Wawrzynek, Paul A.; Ingraffea, Anthony R.

1999-01-01

This is the final report for the NASA funded project entitled "Crack Growth Prediction Methodology for Multi-Site Damage." The primary objective of the project was to create a capability to simulate curvilinear fatigue crack growth and ductile tearing in aircraft fuselages subjected to widespread fatigue damage. The second objective was to validate the capability by way of comparisons to experimental results. Both objectives have been achieved and the results are detailed herein. In the first part of the report, the crack tip opening angle (CTOA) fracture criterion, obtained and correlated from coupon tests to predict fracture behavior and residual strength of built-up aircraft fuselages, is discussed. Geometrically nonlinear, elastic-plastic, thin shell finite element analyses are used to simulate stable crack growth and to predict residual strength. Both measured and predicted results of laboratory flat panel tests and full-scale fuselage panel tests show substantial reduction of residual strength due to the occurrence of multi-site damage (MSD). Detailed comparisons of n stable crack growth history, and residual strength between the predicted and experimental results are used to assess the validity of the analysis methodology. In the second part of the report, issues related to crack trajectory prediction in thin shells; an evolving methodology uses the crack turning phenomenon to improve the structural integrity of aircraft structures are discussed, A directional criterion is developed based on the maximum tangential stress theory, but taking into account the effect of T-stress and fracture toughness orthotropy. Possible extensions of the current crack growth directional criterion to handle geometrically and materially nonlinear problems are discussed. The path independent contour integral method for T-stress evaluation is derived and its accuracy is assessed using a p- and hp-version adaptive finite element method. Curvilinear crack growth is simulated in coupon tests and in full-scale fuselage panel tests. Both T-stress and fracture toughness orthotropy are found to be essential to predict the observed crack paths. The analysis methodology and software program (FRANC3D/STAGS) developed herein allows engineers to maintain aging aircraft economically while insuring continuous airworthiness. Consequently, it will improve the technology to support the safe operation of the current aircraft fleet as well as the design of more damage-tolerant aircraft for the next generation fleet.

Wind Tunnel Test of a Risk-Reduction Wing/Fuselage Model to Examine Juncture-Flow Phenomena

NASA Technical Reports Server (NTRS)

Kegerise, Michael A.; Neuhart, Dan H.

2016-01-01

A wing/fuselage wind-tunnel model was tested in the Langley 14- by 22-foot Subsonic Wind Tunnel in preparation for a highly-instrumented Juncture Flow Experiment to be conducted in the same facility. This test, which was sponsored by the NASA Transformational Tool and Technologies Project, is part of a comprehensive set of experimental and computational research activities to develop revolutionary, physics-based aeronautics analysis and design capability. The objectives of this particular test were to examine the surface and off-body flow on a generic wing/body combination to: 1) choose a final wing for a future, highly instrumented model, 2) use the results to facilitate unsteady pressure sensor placement on the model, 3) determine the area to be surveyed with an embedded laser-doppler velocimetry (LDV) system, 4) investigate the primary juncture corner- flow separation region using particle image velocimetry (PIV) to see if the particle seeding is adequately entrained and to examine the structure in the separated region, and 5) to determine the similarity of observed flow features with those predicted by computational fluid dynamics (CFD). This report documents the results of the above experiment that specifically address the first three goals. Multiple wing configurations were tested at a chord Reynolds number of 2.4 million. Flow patterns on the surface of the wings and in the region of the wing/fuselage juncture were examined using oil- flow visualization and infrared thermography. A limited number of unsteady pressure sensors on the fuselage around the wing leading and trailing edges were used to identify any dynamic effects of the horseshoe vortex on the flow field. The area of separated flow in the wing/fuselage juncture near the wing trailing edge was observed for all wing configurations at various angles of attack. All of the test objectives were met. The staff of the 14- by 22-foot Subsonic Wind Tunnel provided outstanding support and delivered exceptional value to the experiment, which exceeded expectations. The results of this test will directly inform the planning for the first of a series of instrumented-model tests at the same Reynolds number. These tests will be performed on a slightly larger-scale model with the selected wing, and will include off-body measurements with LDV and PIV, steady and unsteady pressure measurements, and the flow-visualization techniques that are discussed in this report.

75 FR 35356 - Airworthiness Directives; The Boeing Company Model 747-100, 747-200B, and 747-200F Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-06-22

...-related skin cracks and corrosion of the skin panel lap joints in the fuselage upper lobe, and repair if... inspections of the fuselage skin at the upper lobe skin lap joints for cracks and evidence of corrosion, and... correct fatigue cracking and corrosion in the fuselage upper lobe skin lap joints, which could lead to...

Interactive Inverse Design Optimization of Fuselage Shape for Low-Boom Supersonic Concepts

NASA Technical Reports Server (NTRS)

Li, Wu; Shields, Elwood; Le, Daniel

2008-01-01

This paper introduces a tool called BOSS (Boom Optimization using Smoothest Shape modifications). BOSS utilizes interactive inverse design optimization to develop a fuselage shape that yields a low-boom aircraft configuration. A fundamental reason for developing BOSS is the need to generate feasible low-boom conceptual designs that are appropriate for further refinement using computational fluid dynamics (CFD) based preliminary design methods. BOSS was not developed to provide a numerical solution to the inverse design problem. Instead, BOSS was intended to help designers find the right configuration among an infinite number of possible configurations that are equally good using any numerical figure of merit. BOSS uses the smoothest shape modification strategy for modifying the fuselage radius distribution at 100 or more longitudinal locations to find a smooth fuselage shape that reduces the discrepancies between the design and target equivalent area distributions over any specified range of effective distance. For any given supersonic concept (with wing, fuselage, nacelles, tails, and/or canards), a designer can examine the differences between the design and target equivalent areas, decide which part of the design equivalent area curve needs to be modified, choose a desirable rate for the reduction of the discrepancies over the specified range, and select a parameter for smoothness control of the fuselage shape. BOSS will then generate a fuselage shape based on the designer's inputs in a matter of seconds. Using BOSS, within a few hours, a designer can either generate a realistic fuselage shape that yields a supersonic configuration with a low-boom ground signature or quickly eliminate any configuration that cannot achieve low-boom characteristics with fuselage shaping alone. A conceptual design case study is documented to demonstrate how BOSS can be used to develop a low-boom supersonic concept from a low-drag supersonic concept. The paper also contains a study on how perturbations in the equivalent area distribution affect the ground signature shape and how new target area distributions for low-boom signatures can be constructed using superposition of equivalent area distributions derived from the Seebass-George-Darden (SGD) theory.

Geometrically nonlinear analysis of laminated elastic structures

NASA Technical Reports Server (NTRS)

Reddy, J. N.

1984-01-01

Laminated composite plates and shells that can be used to model automobile bodies, aircraft wings and fuselages, and pressure vessels among many other were analyzed. The finite element method, a numerical technique for engineering analysis of structures, is used to model the geometry and approximate the solution. Various alternative formulations for analyzing laminated plates and shells are developed and their finite element models are tested for accuracy and economy in computation. These include the shear deformation laminate theory and degenerated 3-D elasticity theory for laminates.

Validated Feasibility Study of Integrally Stiffened Metallic Fuselage Panels for Reducing Manufacturing Costs

NASA Technical Reports Server (NTRS)

Pettit, R. G.; Wang, J. J.; Toh, C.

2000-01-01

The continual need to reduce airframe cost and the emergence of high speed machining and other manufacturing technologies has brought about a renewed interest in large-scale integral structures for aircraft applications. Applications have been inhibited, however, because of the need to demonstrate damage tolerance, and by cost and manufacturing risks associated with the size and complexity of the parts. The Integral Airframe Structures (IAS) Program identified a feasible integrally stiffened fuselage concept and evaluated performance and manufacturing cost compared to conventional designs. An integral skin/stiffener concept was produced both by plate hog-out and near-net extrusion. Alloys evaluated included 7050-T7451 plate, 7050-T74511 extrusion, 6013-T6511 extrusion, and 7475-T7351 plate. Mechanical properties, structural details, and joint performance were evaluated as well as repair, static compression, and two-bay crack residual strength panels. Crack turning behavior was characterized through panel tests and improved methods for predicting crack turning were developed. Manufacturing cost was evaluated using COSTRAN. A hybrid design, made from high-speed machined extruded frames that are mechanically fastened to high-speed machined plate skin/stringer panels, was identified as the most cost-effective manufacturing solution. Recurring labor and material costs of the hybrid design are up to 61 percent less than the current technology baseline.

The FC-1D: The profitable alternative Flying Circus Commercial Aviation Group

NASA Technical Reports Server (NTRS)

Meza, Victor J.; Alvarez, Jaime; Harrington, Brook; Lujan, Michael A.; Mitlyng, David; Saroughian, Andy; Silva, Alex; Teale, Tim

1994-01-01

The FC-1D was designed as an advanced solution for a low cost commercial transport meeting or exceeding all of the 1993/1994 AIAA/Lockheed request for proposal requirements. The driving philosophy behind the design of the FC-1D was the reduction of airline direct operating costs. Every effort was made during the design process to have the customer in mind. The Flying Circus Commercial Aviation Group targeted reductions in drag, fuel consumption, manufacturing costs, and maintenance costs. Flying Circus emphasized cost reduction throughout the entire design program. Drag reduction was achieved by implementation of the aft nacelle wing configuration to reduce cruise drag and increase cruise speeds. To reduce induced drag, rather than increasing the wing span of the FC-1D, spiroids were included in the efficient wing design. Profile and friction drag are reduced by using riblets in place of paint around the fuselage and empennage of the FC-1D. Choosing a single aisle configuration enabled the Flying Circus to optimize the fuselage diameter. Thus, reducing fuselage drag while gaining high structural efficiency. To further reduce fuel consumption a weight reduction program was conducted through the use of composite materials. An additional quality of the FC-1D is its design for low cost manufacturing and assembly. As a result of this design attribute, the FC-1D will have fewer parts which reduces weight as well as maintenance and assembly costs. The FC-1D is affordable and effective, the apex of commercial transport design.

Advanced manufacturing development of a composite empennage component for L-1011 aircraft

NASA Technical Reports Server (NTRS)

1979-01-01

Work on process verification and tooling development continued. The cover process development was completed with the decision to proceed with low resin content prepreg material (34 + or - 3% by weight) in the fabrication of production readiness verification test (PRVT) specimens and the full-scale covers. The structural integrity of the cover/joint design was verified with the successful test of the cover attachment to fuselage ancillary test specimen (H25). Failure occurred, as predicted, in the skin panel away from the fuselage joint at 141 percent of the design ultimate load. With the successful completion of the H25 test, the PRVT cover specimens, which are identical to the H25 ancillary test specimen, were cleared for production. Eight of the twenty cover specimens were fabricated and are in preparation for test. All twenty of the PRVT spar specimens were fabricated and also were prepared for test. The environmental chambers used in the durability test of ten cover and ten spar PRVT specimens were completed and installed in the load reaction frames.

Prediction of the interior noise levels of high-speed propeller-driven aircraft

NASA Technical Reports Server (NTRS)

Rennison, D. C.; Wilby, J. F.; Wilby, E. G.

1980-01-01

The theoretical basis for an analytical model developed to predict the interior noise levels of high-speed propeller-driven airplanes is presented. Particular emphasis is given to modeling the transmission of discrete tones through a fuselage element into a cavity, estimates for the mean and standard deviation of the acoustic power flow, the coupling between a non-homogeneous excitation and the fuselage vibration response, and the prediction of maximum interior noise levels. The model allows for convenient examination of the various roles of the excitation and fuselage structural characteristics on the fuselage vibration response and the interior noise levels, as is required for the design of model or prototype noise control validation tests.

Evaluation of viscous drag reduction schemes for subsonic transports

NASA Technical Reports Server (NTRS)

Marino, A.; Economos, C.; Howard, F. G.

1975-01-01

The results are described of a theoretical study of viscous drag reduction schemes for potential application to the fuselage of a long-haul subsonic transport aircraft. The schemes which were examined included tangential slot injection on the fuselage and various synergetic combinations of tangential slot injection and distributed suction applied to wing and fuselage surfaces. Both passive and mechanical (utilizing turbo-machinery) systems were examined. Overall performance of the selected systems was determined at a fixed subsonic cruise condition corresponding to a flight Mach number of free stream M = 0.8 and an altitude of 11,000 m. The nominal aircraft to which most of the performance data was referenced was a wide-body transport of the Boeing 747 category. Some of the performance results obtained with wing suction are referenced to a Lockheed C-141 Star Lifter wing section. Alternate designs investigated involved combinations of boundary layer suction on the wing surfaces and injection on the fuselage, and suction and injection combinations applied to the fuselage only.

Theoretical prediction of thick wing and pylon-fuselage-fanpod-nacelle aerodynamic characteristics at subcritical speeds. Part 1: Theory and results

NASA Technical Reports Server (NTRS)

Tulinius, J. R.

1974-01-01

The theoretical development and the comparison of results with data of a thick wing and pylon-fuselage-fanpod-nacelle analysis are presented. The analysis utilizes potential flow theory to compute the surface velocities and pressures, section lift and center of pressure, and the total configuration lift, moment, and vortex drag. The skin friction drag is also estimated in the analysis. The perturbation velocities induced by the wing and pylon, fuselage and fanpod, and nacelle are represented by source and vortex lattices, quadrilateral vortices, and source frustums, respectively. The strengths of these singularities are solved for simultaneously including all interference effects. The wing and pylon planforms, twists, cambers, and thickness distributions, and the fuselage and fanpod geometries can be arbitrary in shape, provided the surface gradients are smooth. The flow through nacelle is assumed to be axisymmetric. An axisymmetric center engine hub can also be included. The pylon and nacelle can be attached to the wing, fuselage, or fanpod.

The characterization of widespread fatigue damage in fuselage structure

NASA Technical Reports Server (NTRS)

Piascik, Robert S.; Willard, Scott A.; Miller, Matthew

1994-01-01

The characteristics of widespread fatigue damage (WSFD) in fuselage riveted structure were established by detailed nondestructive and destructive examinations of fatigue damage contained in a full size fuselage test article. The objectives of this were to establish an experimental data base for validating emerging WSFD analytical prediction methodology and to identify first order effects that contribute to fatigue crack initiation and growth. Detailed examinations were performed on a test panel containing four bays of a riveted lap splice joint. The panel was removed from a full scale fuselage test article after receiving 60,000 full pressurization cycles. The results of in situ examinations document the progression of fuselage skin fatigue crack growth through crack linkup. Detailed tear down examinations and fractography of the lap splice joint region revealed fatigue crack initiation sites, crack morphology, and crack linkup geometry. From this large data base, distributions of crack size and locations are presented and discussions of operative damage mechanisms are offered.

The characterization of widespread fatigue damage in fuselage structure

NASA Technical Reports Server (NTRS)

Piascik, Robert S.; Willard, Scott A.; Miller, Matthew

1994-01-01

The characteristics of widespread fatigue damage (WSFD) in fuselage riveted structure were established by detailed nondestructive and destructive examinations of fatigue damage contained in a full size fuselage test article. The objectives of this work were to establish an experimental data base for validating emerging WSFD analytical prediction methodology and to identify first order effects that contribute to fatigue crack initiation and growth. Detailed examinations were performed on a test panel containing four bays of a riveted lap splice joint. The panel was removed from a full scale fuselage test article after receiving 60,000 full pressurization cycles. The results of in situ examinations document the progression of fuselage skin fatigue crack growth through crack linkup. Detailed tear down examinations and fractography of the lap splice joint region revealed fatigue crack initiation sites, crack morphology and crack linkup geometry. From this large data base, distributions of crack size and locations are presented and discussions of operative damage mechanisms are offered.

A 0.15-scale study of configuration effects on the aerodynamic interaction between main rotor and fuselage

NASA Technical Reports Server (NTRS)

Trept, Ted

1984-01-01

Hover and forward flight tests were conducted to investigate the mutual aerodynamic interaction between the main motor and fuselage of a conventional helicopter configuration. A 0.15-scale Model 222 two-bladed teetering rotor was combined with a 0.15-scale model of the NASA Ames 40x80-foot wind tunnel 1500 horsepower test stand fairing. Configuration effects were studied by modifying the fairing to simulate a typical helicopter forebody. Separation distance between rotor and body were also investigated. Rotor and fuselage force and moment as well as pressure data are presented in graphical and tabular format. Data was taken over a range of thrust coefficients from 0.002 to 0.007. In forward flight speed ratio was varied from 0.1 to 0.3 with shaft angle varying from +4 to -12 deg. The data show that the rotors effect on the fuselage may be considerably more important to total aircraft performance than the effect of the fuselage on the rotor.

Theoretical study of nonadiabatic boundary-layer stabilization times in a cryogenic wind tunnel for typical stainless steel wing and fuselage models

NASA Technical Reports Server (NTRS)

Johnson, C. B.

1980-01-01

The time varying effect of nonadiabatic wall conditions on boundary layer properties was studied for a two dimensional wing section and an axisymmetric fuselage. The wing and fuselage sections are representative of the wing root chord and fuselage of a typical transport model for the National Transonic Facility. The analysis was made with a solid wing and three fuselage configurations (one solid and two hollow with varying skin thicknesses) all made from AISI type 310S stainless steel. The displacement thickness and local skin friction were investigated at a station on the model in terms of the time required for these two boundary layer properties to reach an adiabatic wall condition after a 50 K step change in total temperature. The analysis was made for a free stream Mach number of 0.85, a total temperature of 117 K, and stagnation pressures of 2, 6, and 9 atm.

Crash Simulation of a Boeing 737 Fuselage Section Vertical Drop Test

NASA Technical Reports Server (NTRS)

Fasanella, Edwin L.; Jackson, Karen E.; Jones, Yvonne T.; Frings, Gary; Vu, Tong

2004-01-01

A 30-ft/s vertical drop test of a fuselage section of a Boeing 737 aircraft was conducted in October of 1999 at the FAA Technical Center in Atlantic City, NJ. This test was performed to evaluate the structural integrity of a conformable auxiliary fuel tank mounted beneath the floor and to determine its effect on the impact response of the airframe structure and the occupants. The test data were used to compare with a finite element simulation of the fuselage structure and to gain a better understanding of the impact physics through analytical/experimental correlation. To perform this simulation, a full-scale 3-dimensional finite element model of the fuselage section was developed using the explicit, nonlinear transient-dynamic finite element code, MSC.Dytran. The emphasis of the simulation was to predict the structural deformation and floor-level acceleration responses obtained from the drop test of the B737 fuselage section with the auxiliary fuel tank.

Ground-based and in-flight simulator studies of flight characteristics of a twin-fuselage passenger transport airplane during approach and landing

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Smith, P. M.; Neely, W. R., Jr.; Deal, P. L.; Yenni, K. R.

1985-01-01

Six-degree-of-freedom ground-based and in-flight simulator studies were conducted to evaluate the low-speed flight characteristics of a twin-fuselage passenger transport airplane and to compare these characteristics with those of a large, single-fuselage (reference) transport configuration similar to the Lockheed C-5A airplane. The primary piloting task was the approach and landing task. The results of this study indicated that the twin-fuselage transport concept had acceptable but unsatisfactory longitudinal and lateral-directional low-speed flight characteristics, and that stability and control augmentation would be required in order to improve the handling qualities. Through the use of rate-command/attitude-hold augmentation in the pitch and roll axes, and the use of several turn coordination features, the handling qualities of the simulated transport were improved appreciably. The in-flight test results showed excellent agreement with those of the six-degree-of-freedom ground-based simulator handling qualities tests. As a result of the in-flight simulation study, a roll-control-induced normal-acceleration criterion was developed. The handling qualities of the augmented twin-fuselage passenger transport airplane exhibited an improvement over the handling characteristics of the reference (single-fuselage) transport.

Electrically conductive carbon fibre-reinforced composite for aircraft lightning strike protection

NASA Astrophysics Data System (ADS)

Katunin, Andrzej; Krukiewicz, Katarzyna; Turczyn, Roman; Sul, Przemysław; Bilewicz, Marcin

2017-05-01

Aircraft elements, especially elements of exterior fuselage, are subjected to damage caused by lightning strikes. Due to the fact that these elements are manufactured from polymeric composites in modern aircraft, and thus, they cannot conduct electrical charges, the lightning strikes cause burnouts in composite structures. Therefore, the effective lightning strike protection for such structures is highly desired. The solution presented in this paper is based on application of organic conductive fillers in the form of intrinsically conducting polymers and carbon fabric in order to ensure electrical conductivity of whole composite and simultaneously retain superior mechanical properties. The presented studies cover synthesis and manufacturing of the electrically conductive composite as well as its characterization with respect to mechanical and electrical properties. The performed studies indicate that the proposed material can be potentially considered as a constructional material for aircraft industry, which characterizes by good operational properties and low cost of manufacturing with respect to current lightning strike protection materials solutions.

Impact evaluation of composite floor sections

NASA Technical Reports Server (NTRS)

Boitnott, Richard L.; Fasanella, Edwin L.

1989-01-01

Graphite-epoxy floor sections representative of aircraft fuselage construction were statically and dynamically tested to evaluate their response to crash loadings. These floor sections were fabricated using a frame-stringer design typical of present aluminum aircraft without features to enhance crashworthiness. The floor sections were tested as part of a systematic research program developed to study the impact response of composite components of increasing complexity. The ultimate goal of the research program is to develop crashworthy design features for future composite aircraft. Initially, individual frames of six-foot diameter were tested both statically and dynamically. The frames were then used to construct built-up floor sections for dynamic tests at impact velocities of approximately 20 feet/sec to simulate survivable crash velocities. In addition, static tests were conducted to gain a better understanding of the failure mechanisms seen in the dynamic tests.

Effects of wing leading-edge flap deflections on subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing

NASA Technical Reports Server (NTRS)

Henderson, W. P.

1978-01-01

An investigation was conducted to determine the effects of wing leading-edge flap deflections on the subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing. The tests were conducted at Mach numbers from 0.40 to 0.85, corresponding to Reynolds numbers (based on wing mean geometric chord) of 2.37 x 1,000,000 to 4.59 x 1,000,000 and at angles of attack from -3 deg to 22 deg. The configurations under study included a wing-fuselage configuration and a wing-fuselage-strake configuration. Each configuration had multisegmented, constant-chord leading-edge flaps which could be deflected independently or in various combinations.

The History of Structural Fatigue Testing at Fishermans Bend Australia

DTIC Science & Technology

2005-10-01

Aviation DH de Havilland Aircraft (UK) DOA Department of Aviation DSTO Defence Science and Technology Organisation ESH Equivalent Service Hours...gave early evidence of the truth of Mr. Wills’s predictions and greatly influenced the fatigue research programme. In October 1951 a de Havilland (DH...crack growth [22]. At about this time (in 1953) fatigue failures occurred in the fuselage of the DH Comet which led to the loss of two aircraft with

Advanced composites structural concepts and materials technologies for primary aircraft structures: Structural response and failure analysis

NASA Technical Reports Server (NTRS)

Dorris, William J.; Hairr, John W.; Huang, Jui-Tien; Ingram, J. Edward; Shah, Bharat M.

1992-01-01

Non-linear analysis methods were adapted and incorporated in a finite element based DIAL code. These methods are necessary to evaluate the global response of a stiffened structure under combined in-plane and out-of-plane loading. These methods include the Arc Length method and target point analysis procedure. A new interface material model was implemented that can model elastic-plastic behavior of the bond adhesive. Direct application of this method is in skin/stiffener interface failure assessment. Addition of the AML (angle minus longitudinal or load) failure procedure and Hasin's failure criteria provides added capability in the failure predictions. Interactive Stiffened Panel Analysis modules were developed as interactive pre-and post-processors. Each module provides the means of performing self-initiated finite elements based analysis of primary structures such as a flat or curved stiffened panel; a corrugated flat sandwich panel; and a curved geodesic fuselage panel. This module brings finite element analysis into the design of composite structures without the requirement for the user to know much about the techniques and procedures needed to actually perform a finite element analysis from scratch. An interactive finite element code was developed to predict bolted joint strength considering material and geometrical non-linearity. The developed method conducts an ultimate strength failure analysis using a set of material degradation models.

An overview of the NASA textile composites program

NASA Technical Reports Server (NTRS)

Dexter, H. Benson

1993-01-01

The NASA Langley Research Center is conducting and sponsoring research to explore the benefits of textile reinforced composites for civil transport aircraft primary structures. The objective of this program is to develop and demonstrate the potential of affordable textile reinforced composite materials to meet design properties and damage tolerance requirements of advanced aircraft structures. In addition to in-house research, the program includes major participation by the aircraft industry and aerospace textile companies. The major program elements include development of textile preforms, processing science, mechanics of materials, experimental characterization of materials, and development and evaluation of textile reinforced composite structural elements and subcomponents. The NASA Langley in-house research is focused on science-based understanding of resin transfer molding (RTM), development of powder-coated towpreg processes, analysis methodology, and development of a performance database on textile reinforced composites. The focus of the textile industry participation is on development of multidirectional, damage-tolerant preforms, and the aircraft industry participation is in the areas of innovative design concepts, cost-effective fabrication, and testing of textile reinforced composite structural elements and subcomponents. Textile processes such as 3-D weaving, 2-D and 3-D braiding, and knitting/stitching are being compared with conventional laminated tape processes for improved damage tolerance. Through-the-thickness reinforcements offer significant damage tolerance improvements. However, these gains must be weighed against potential loss in in-plane properties such as strength and stiffness. Analytical trade studies are underway to establish design guidelines for the application of textile material forms to meet specific loading requirements. Fabrication and testing of large structural components are required to establish the full potential of textile reinforced composite materials. The goals of the NASA Langley-sponsored research program are to demonstrate technology readiness with subscale composite components by 1995 and to verify the performance of full-scale composite primary aircraft structural components by 1997. The status of textile reinforced composite structural elements under development by Boeing, Douglas, Lockheed, and Grumman are presented. Included are braided frames and woven/stitched wing and fuselage panels.

Energy Finite Element Analysis Developments for Vibration Analysis of Composite Aircraft Structures

NASA Technical Reports Server (NTRS)

Vlahopoulos, Nickolas; Schiller, Noah H.

2011-01-01

The Energy Finite Element Analysis (EFEA) has been utilized successfully for modeling complex structural-acoustic systems with isotropic structural material properties. In this paper, a formulation for modeling structures made out of composite materials is presented. An approach based on spectral finite element analysis is utilized first for developing the equivalent material properties for the composite material. These equivalent properties are employed in the EFEA governing differential equations for representing the composite materials and deriving the element level matrices. The power transmission characteristics at connections between members made out of non-isotropic composite material are considered for deriving suitable power transmission coefficients at junctions of interconnected members. These coefficients are utilized for computing the joint matrix that is needed to assemble the global system of EFEA equations. The global system of EFEA equations is solved numerically and the vibration levels within the entire system can be computed. The new EFEA formulation for modeling composite laminate structures is validated through comparison to test data collected from a representative composite aircraft fuselage that is made out of a composite outer shell and composite frames and stiffeners. NASA Langley constructed the composite cylinder and conducted the test measurements utilized in this work.

Vertical drop test of a transport fuselage section located forward of the wing

NASA Technical Reports Server (NTRS)

Williams, M. S.; Hayduk, R. J.

1983-01-01

A Boeing 707 fuselage section was drop tested at the NASA Langley Research Center to measure structural, seat, and occupant response to vertical crack loads. Post-test inspection showed that the section bottom collapsed inward approximately 2 ft. Preliminary data traces indicated maximum normal accelerations of 20 g on the fuselage bottom, 10 to 12 g on the cabin floor, and 6.5 to 8 g in the pelvises of the anthropomorphic dummies.

NASTRAN data generation of helicopter fuselages using interactive graphics. [preprocessor system for finite element analysis using IBM computer

NASA Technical Reports Server (NTRS)

Sainsbury-Carter, J. B.; Conaway, J. H.

1973-01-01

The development and implementation of a preprocessor system for the finite element analysis of helicopter fuselages is described. The system utilizes interactive graphics for the generation, display, and editing of NASTRAN data for fuselage models. It is operated from an IBM 2250 cathode ray tube (CRT) console driven by an IBM 370/145 computer. Real time interaction plus automatic data generation reduces the nominal 6 to 10 week time for manual generation and checking of data to a few days. The interactive graphics system consists of a series of satellite programs operated from a central NASTRAN Systems Monitor. Fuselage structural models including the outer shell and internal structure may be rapidly generated. All numbering systems are automatically assigned. Hard copy plots of the model labeled with GRID or elements ID's are also available. General purpose programs for displaying and editing NASTRAN data are included in the system. Utilization of the NASTRAN interactive graphics system has made possible the multiple finite element analysis of complex helicopter fuselage structures within design schedules.

Residual Strength Pressure Tests and Nonlinear Analyses of Stringer- and Frame-Stiffened Aluminum Fuselage Panels with Longitudinal Cracks

NASA Technical Reports Server (NTRS)

Young, Richard D.; Rouse, Marshall; Ambur, Damodar R.; Starnes, James H., Jr.

1999-01-01

The results of residual strength pressure tests and nonlinear analyses of stringer- and frame-stiffened aluminum fuselage panels with longitudinal cracks are presented. Two types of damage are considered: a longitudinal crack located midway between stringers, and a longitudinal crack adjacent to a stringer and along a row of fasteners in a lap joint that has multiple-site damage (MSD). In both cases, the longitudinal crack is centered on a severed frame. The panels are subjected to internal pressure plus axial tension loads. The axial tension loads are equivalent to a bulkhead pressure load. Nonlinear elastic-plastic residual strength analyses of the fuselage panels are conducted using a finite element program and the crack-tip-opening-angle (CTOA) fracture criterion. Predicted crack growth and residual strength results from nonlinear analyses of the stiffened fuselage panels are compared with experimental measurements and observations. Both the test and analysis results indicate that the presence of MSD affects crack growth stability and reduces the residual strength of stiffened fuselage shells with long cracks.

Measurement of Insertion Loss of an Acoustic Treatment in the Presence of Additional Uncorrelated Sound Sources

NASA Technical Reports Server (NTRS)

Klos, Jacob; Palumbo, Daniel L.

2003-01-01

A method to intended for measurement of the insertion loss of an acoustic treatment applied to an aircraft fuselage in-situ is documented in this paper. Using this method, the performance of a treatment applied to a limited portion of an aircraft fuselage can be assessed even though the untreated fuselage also radiates into the cabin, corrupting the intensity measurement. This corrupting noise in the intensity measurement incoherent with the panel vibration of interest is removed by correlating the intensity to reference transducers such as accelerometers. Insertion loss of the acoustic treatments is estimated from the ratio of correlated intensity measurements with and without a treatment applied. In the case of turbulent boundary layer excitation of the fuselage, this technique can be used to assess the performance of noise control methods without requiring treatment of the entire fuselage. Several experimental studies and numerical simulations have been conducted, and results from three case studies are documented in this paper. Conclusions are drawn about the use of this method to study aircraft sidewall treatments.

Residual Strength Pressure Tests and Nonlinear Analyses of Stringer-and Frame-Stiffened Aluminum Fuselage Panels with Longitudinal Cracks

NASA Technical Reports Server (NTRS)

Young, Richard D.; Rouse, Marshall; Ambur, Damodar R.; Starnes, James H., Jr.

1998-01-01

The results of residual strength pressure tests and nonlinear analyses of stringer- and frame-stiffened aluminum fuselage panels with longitudinal cracks are presented. Two types of damage are considered: a longitudinal crack located midway between stringers, and a longitudinal crack adjacent to a stringer and along a row of fasteners in a lap joint that has multiple-site damage (MSD). In both cases, the longitudinal crack is centered on a severed frame. The panels are subjected to internal pressure plus axial tension loads. The axial tension loads are equivalent to a bulkhead pressure load. Nonlinear elastic-plastic residual strength analyses of the fuselage panels are conducted using a finite element program and the crack-tip-opening-angle (CTOA) fracture criterion. Predicted crack growth and residual strength results from nonlinear analyses of the stiffened fuselage panels are compared with experimental measurements and observations. Both the test and analysis results indicate that the presence of MSD affects crack growth stability and reduces the residual strength of stiffened fuselage shells with long cracks.

Simulator study of flight characteristics of a large twin-fuselage cargo transport airplane during approach and landing

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Deal, P. L.; Keyser, G. L., Jr.; Smith, P. M.

1983-01-01

A six degree-of-freedom, ground-based simulator study was conducted to evaluate the low speed flight characteristics of a twin fuselage cargo transport airplane and to compare these characteristics with those of a large, single fuselage (reference) transport configuration which was similar to the Lockheed C-5C airplane. The primary piloting task was the approach and landing. The results indicated that in order to achieve "acceptable' low speed handling qualities on the twin fuselage concept, considerable stability and control augmentation was required, and although the augmented airplane could be landed safely under adverse conditions, the roll performance of the aircraft had to be improved appreciably before the handling qualities were rated as being "satisfactory.' These ground-based simulation results indicated that a value of t sub phi = 30 (time required to bank 30 deg) less than 6 sec should result in "acceptable' roll response characteristics, and when t sub phi = 30 is less than 3.8 sec, "satisfactory' roll response should be attainable on such large and unusually configured aircraft as the subject twin fuselage cargo transport concept.

Structural health monitoring and impact detection for primary aircraft structures

NASA Astrophysics Data System (ADS)

Kosters, Eric; van Els, Thomas J.

2010-04-01

The increasing use of thermoplastic carbon fiber-reinforced plastic (CFRP) materials in the aerospace industry for primary aircraft structures, such as wing leading-edge surfaces and fuselage sections, has led to rapid growth in the field of structural health monitoring (SHM). Impact, vibration, and load can all cause failure, such as delamination and matrix cracking, in composite materials. Moreover, the internal material damage can occur without being visible to the human eye, making inspection of and clear insight into structural integrity difficult using currently available evaluation methods. Here, we describe the detection of impact and its localization in materials and structures by high-speed interrogation of multiple-fiber Bragg grating (FBG) sensors mounted on a composite aircraft component.

Computational Aerodynamic Modeling of Small Quadcopter Vehicles

NASA Technical Reports Server (NTRS)

Yoon, Seokkwan; Ventura Diaz, Patricia; Boyd, D. Douglas; Chan, William M.; Theodore, Colin R.

2017-01-01

High-fidelity computational simulations have been performed which focus on rotor-fuselage and rotor-rotor aerodynamic interactions of small quad-rotor vehicle systems. The three-dimensional unsteady Navier-Stokes equations are solved on overset grids using high-order accurate schemes, dual-time stepping, low Mach number preconditioning, and hybrid turbulence modeling. Computational results for isolated rotors are shown to compare well with available experimental data. Computational results in hover reveal the differences between a conventional configuration where the rotors are mounted above the fuselage and an unconventional configuration where the rotors are mounted below the fuselage. Complex flow physics in forward flight is investigated. The goal of this work is to demonstrate that understanding of interactional aerodynamics can be an important factor in design decisions regarding rotor and fuselage placement for next-generation multi-rotor drones.

Subsonic and supersonic longitudinal stability and control characteristics of an aft-tail fighter configuration with cambered and uncambered wings and cambered fuselage

NASA Technical Reports Server (NTRS)

Dollyhigh, S. M.

1977-01-01

The longitudinal aerodynamic characteristics of a fighter airplane concept has been determined through an investigation over a Mach number range from 0.50 to 2.16. The configuration incorporates a cambered fuselage with a single external compression horizontal ramp inlet, a clipped arrow wing, twin horizontal tails, and a single vertical tail. The wing camber surface was optimized in drag due to lift and was designed to be self trimming at Mach 1.40 and at a lift coefficient of 0.20. The fuselage was cambered to preserve the design wing loadings on the part of the theoretical wing enclosed by the fuselage. An uncambered of flat wing of the same planform and thickness ratio distribution was also tested.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Roach, Dennis Patrick; Rackow, Kirk A.

A typical aircraft can experience over 2,000 fatigue cycles (cabin pressurizations) and even greater flight hours in a single year. An unavoidable by-product of aircraft use is that crack, impact, and corrosion flaws develop throughout the aircraft's skin and substructure elements. Economic barriers to the purchase of new aircraft have placed even greater demands on efficient and safe repair methods. The use of bonded composite doublers offers the airframe manufacturers and aircraft maintenance facilities a cost effective method to safely extend the lives of their aircraft. Instead of riveting multiple steel or aluminum plates to facilitate an aircraft repair, itmore » is now possible to bond a single Boron-Epoxy composite doubler to the damaged structure. The FAA's Airworthiness Assurance Center at Sandia National Labs (AANC), Boeing, and Federal Express completed a pilot program to validate and introduce composite doubler repair technology to the U.S. commercial aircraft industry. This project focused on repair of DC-10 fuselage structure and its primary goal was to demonstrate routine use of this repair technology using niche applications that streamline the design-to-installation process. As composite doubler repairs gradually appear in the commercial aircraft arena, successful flight operation data is being accumulated. These commercial aircraft repairs are not only demonstrating the engineering and economic advantages of composite doubler technology but they are also establishing the ability of commercial maintenance depots to safely adopt this repair technique. This report presents the array of engineering activities that were completed in order to make this technology available for widespread commercial aircraft use. Focused laboratory testing was conducted to compliment the field data and to address specific issues regarding damage tolerance and flaw growth in composite doubler repairs. Fatigue and strength tests were performed on a simulated wing repair using a substandard design and a flawed installation. In addition, the new Sol-Gel surface preparation technique was evaluated. Fatigue coupon tests produced Sol-Gel results that could be compared with a large performance database from conventional, riveted repairs. It was demonstrated that not only can composite doublers perform well in severe off-design conditions (low doubler stiffness and presence of defects in doubler installation) but that the Sol-Gel surface preparation technique is easier and quicker to carry out while still producing optimum bonding properties. Nondestructive inspection (NDI) methods were developed so that the potential for disbond and delamination growth could be monitored and crack growth mitigation could be quantified. The NDI methods were validated using full-scale test articles and the FedEx aircraft installations. It was demonstrated that specialized NDI techniques can detect flaws in composite doubler installations before they reach critical size. Probability of Detection studies were integrated into the FedEx training in order to quantify the ability of aircraft maintenance depots to properly monitor these repairs. In addition, Boeing Structural Repair and Nondestructive Testing Manuals were modified to include composite doubler repair and inspection procedures. This report presents the results from the FedEx Pilot Program that involved installation and surveillance of numerous repairs on operating aircraft. Results from critical NDI evaluations are reported in light of damage tolerance assessments for bonded composite doublers. This work has produced significant interest from airlines and aircraft manufacturers. The successful Pilot Program produced flight performance history to establish the durability of bonded composite patches as a permanent repair on commercial aircraft structures. This report discusses both the laboratory data and Pilot Program results from repair installations on operating aircraft to introduce composite doubler repairs into mainstream commercial aircraft use.« less

The effect of air flow, panel curvature, and internal pressurization on field-incidence transmission loss. [acoustic propagation through aircraft fuselage

NASA Technical Reports Server (NTRS)

Koval, L. R.

1975-01-01

In the context of sound transmission through aircraft fuselage panels, equations for the field-incidence transmission loss (TL) of a single-walled panel are derived that include the effects of external air flow, panel curvature, and internal fuselage pressurization. These effects are incorporated into the classical equations for the TL of single panels, and the resulting double integral for field-incidence TL is numerically evaluated for a specific set of parameters.

Development of stitched/RTM primary structures for transport aircraft

NASA Technical Reports Server (NTRS)

Hawley, Arthur V.

1993-01-01

This report covers work accomplished in the Innovative Composite Aircraft Primary Structure (ICAPS) program. An account is given of the design criteria and philosophy that guides the development. Wing and fuselage components used as a baseline for development are described. The major thrust of the program is to achieve a major cost breakthrough through development of stitched dry preforms and resin transfer molding (RTM), and progress on these processes is reported. A full description is provided on the fabrication of the stitched RTM wing panels. Test data are presented.

Damage Detection Response Characteristics of Open Circuit Resonant (SansEC) Sensors

NASA Technical Reports Server (NTRS)

Dudley, Kenneth L.; Szatkowski, George N.; Smith, Laura J.; Koppen, Sandra V.; Ely, Jay J.; Nguyen, Truong X.; Wang, Chuantong; Ticatch, Larry A.; Mielnik, John J.

2013-01-01

The capability to assess the current or future state of the health of an aircraft to improve safety, availability, and reliability while reducing maintenance costs has been a continuous goal for decades. Many companies, commercial entities, and academic institutions have become interested in Integrated Vehicle Health Management (IVHM) and a growing effort of research into "smart" vehicle sensing systems has emerged. Methods to detect damage to aircraft materials and structures have historically relied on visual inspection during pre-flight or post-flight operations by flight and ground crews. More quantitative non-destructive investigations with various instruments and sensors have traditionally been performed when the aircraft is out of operational service during major scheduled maintenance. Through the use of reliable sensors coupled with data monitoring, data mining, and data analysis techniques, the health state of a vehicle can be detected in-situ. NASA Langley Research Center (LaRC) is developing a composite aircraft skin damage detection method and system based on open circuit SansEC (Sans Electric Connection) sensor technology. Composite materials are increasingly used in modern aircraft for reducing weight, improving fuel efficiency, and enhancing the overall design, performance, and manufacturability of airborne vehicles. Materials such as fiberglass reinforced composites (FRC) and carbon-fiber-reinforced polymers (CFRP) are being used to great advantage in airframes, wings, engine nacelles, turbine blades, fairings, fuselage structures, empennage structures, control surfaces and aircraft skins. SansEC sensor technology is a new technical framework for designing, powering, and interrogating sensors to detect various types of damage in composite materials. The source cause of the in-service damage (lightning strike, impact damage, material fatigue, etc.) to the aircraft composite is not relevant. The sensor will detect damage independent of the cause. Damage in composite material is generally associated with a localized change in material permittivity and/or conductivity. These changes are sensed using SansEC. The unique electrical signatures (amplitude, frequency, bandwidth, and phase) are used for damage detection and diagnosis. An operational system and method would incorporate a SansEC sensor array on select areas of the aircraft exterior surfaces to form a "Smart skin" sensing surface. In this paper a new method and system for aircraft in-situ damage detection and diagnosis is presented. Experimental test results on seeded fault damage coupons and computational modeling simulation results are presented. NASA LaRC has demonstrated with individual sensors that SansEC sensors can be effectively used for in-situ composite damage detection of delamination, voids, fractures, and rips. Keywords: Damage Detection, Composites, Integrated Vehicle Health Monitoring (IVHM), Aviation Safety, SansEC Sensors

Recommendations for numerical solution of reinforced-panel and fuselage-ring problems

NASA Technical Reports Server (NTRS)

Hoff, N J; Libby, Paul A

1949-01-01

Procedures are recommended for solving the equations of equilibrium of reinforced panels and isolated fuselage rings as represented by the external loads and the operations table established according to Southwell's method. From the solution of these equations the stress distribution can be easily determined. The method of systematic relaxations, the matrix-calculus method, and several other methods applicable in special cases are discussed. Definite recommendations are made for obtaining the solution of reinforced-panel problems which are generally designated as shear lag problems. The procedures recommended are demonstrated in the analysis of a number of panels. In the case of fuselage rings it is not possible to make definite recommendations for the solution of the equilibrium equations for all rings and loadings. However, suggestions based on the latest experience are made and demonstrated on several rings.

Structural Acoustic Characteristics of Aircraft and Active Control of Interior Noise

NASA Technical Reports Server (NTRS)

Fuller, C. R.

1998-01-01

The reduction of aircraft cabin sound levels to acceptable values still remains a topic of much research. The use of conventional passive approaches has been extensively studied and implemented. However performance limits of these techniques have been reached. In this project, new techniques for understanding the structural acoustic behavior of aircraft fuselages and the use of this knowledge in developing advanced new control approaches are investigated. A central feature of the project is the Aircraft Fuselage Test Facility at Va Tech which is based around a full scale Cessna Citation III fuselage. The work is divided into two main parts; the first part investigates the use of an inverse technique for identifying dominant fuselage vibrations. The second part studies the development and implementation of active and active-passive techniques for controlling aircraft interior noise.

An analysis of the impact of cabin floor angle restrictions on L/D for a typical supersonic transport

NASA Technical Reports Server (NTRS)

Radkey, R. L.

1974-01-01

High floor angles at cruise have been identified as a significant problem facing airline and public acceptance of a supersonic transport. In order to explore the relationship between cruise performances and floor angle, four related wing-fuselage design and integration studies have been conducted. The studies were: (1) a fuselage camber study in which perturbations in the fuselage camber distribution were examined with a baseline wing, (2) a wing optimization study in which wings were optimized for minimum drag at C sub L's less than the design C sub L. These wings were optimized as wing planform camber surfaces alone and evaluated with a baseline fuselage, (3) a second wing optimization study in which wings were optimized for minimum drag at C sub L's less than the design C sub L but for this study the wings were optimized in the presence of the baseline fuselage, and (4) a third wing optimization study in which wings were optimized for minmum drag subject to C sub M constraints designed to produce more positive C sub MO's, thereby reducing trim drag. The studies indicated that it was not possible to both improve the aircraft cruise L/D and substantially reduce the cruise floor angle. The studies did indicate that the cruise floor angle could be reduced by reducing the fuselage incidence relative to the wing, but the reduction in floor angle was accompanied by a substantial reduction in L/D.

Simulating the Response of a Composite Honeycomb Energy Absorber. Part 1; Dynamic Crushing of Components and Multi-Terrain Impacts

NASA Technical Reports Server (NTRS)

Jackson, Karen E.; Fasanella, Edwin L.; Polanco, Michael A.

2012-01-01

This paper describes the experimental and analytical evaluation of an externally deployable composite honeycomb structure that is designed to attenuate impact energy during helicopter crashes. The concept, designated the Deployable Energy Absorber (DEA), utilizes an expandable Kevlar (Registered Trademark) honeycomb to dissipate kinetic energy through crushing. The DEA incorporates a unique flexible hinge design that allows the honeycomb to be packaged and stowed until needed for deployment. Experimental evaluation of the DEA included dynamic crush tests of multi-cell components and vertical drop tests of a composite fuselage section, retrofitted with DEA blocks, onto multi-terrain. Finite element models of the test articles were developed and simulations were performed using the transient dynamic code, LSDYNA (Registered Trademark). In each simulation, the DEA was represented using shell elements assigned two different material properties: Mat 24, an isotropic piecewise linear plasticity model, and Mat 58, a continuum damage mechanics model used to represent laminated composite fabrics. DEA model development and test-analysis comparisons are presented.

Validation of an interior noise prediction model for a composite cylinder

NASA Technical Reports Server (NTRS)

Beyer, Todd B.; Grosveld, Ferdinand W.

1987-01-01

An acoustic modal analysis has been performed in the cavity of a composite cylinder model of an aircraft fuselage. The filament wound, composite shell is 12 feet long and 5.5 feet in diameter. A one-half-in. thick plywood floor is attached to the shell 69 deg from the vertical centerline through the bottom of the shell. The acoustic modal frequencies were obtained from a sound pressure level and phase survey conducted throughout the interior volume bounded by the floor, endcaps and stiffened shell, while being excited by white noise from a loudspeaker source. The measured acoustic resonance frequencies and mode shapes compare well with analytical predictions from the Propeller Aircraft Interior Noise (PAIN) model. Details of the theory and derivation of the acoustic characteristics have been included. Reverberation time measurements, using the integrated impulse technique, have been performed to determine acoustic loss factors. These measured loss factors have been input to the PAIN program in order to more accurately predict the space-averaged interior noise of the composite cylinder.

A Conceptual Aerospace Vehicle Structural System Modeling, Analysis and Design Process

NASA Technical Reports Server (NTRS)

Mukhopadhyay, Vivek

2007-01-01

A process for aerospace structural concept analysis and design is presented, with examples of a blended-wing-body fuselage, a multi-bubble fuselage concept, a notional crew exploration vehicle, and a high altitude long endurance aircraft. Aerospace vehicle structures must withstand all anticipated mission loads, yet must be designed to have optimal structural weight with the required safety margins. For a viable systems study of advanced concepts, these conflicting requirements must be imposed and analyzed early in the conceptual design cycle, preferably with a high degree of fidelity. In this design process, integrated multidisciplinary analysis tools are used in a collaborative engineering environment. First, parametric solid and surface models including the internal structural layout are developed for detailed finite element analyses. Multiple design scenarios are generated for analyzing several structural configurations and material alternatives. The structural stress, deflection, strain, and margins of safety distributions are visualized and the design is improved. Over several design cycles, the refined vehicle parts and assembly models are generated. The accumulated design data is used for the structural mass comparison and concept ranking. The present application focus on the blended-wing-body vehicle structure and advanced composite material are also discussed.

Cost model relationships between textile manufacturing processes and design details for transport fuselage elements

NASA Technical Reports Server (NTRS)

Metschan, Stephen L.; Wilden, Kurtis S.; Sharpless, Garrett C.; Andelman, Rich M.

1993-01-01

Textile manufacturing processes offer potential cost and weight advantages over traditional composite materials and processes for transport fuselage elements. In the current study, design cost modeling relationships between textile processes and element design details were developed. Such relationships are expected to help future aircraft designers to make timely decisions on the effect of design details and overall configurations on textile fabrication costs. The fundamental advantage of a design cost model is to insure that the element design is cost effective for the intended process. Trade studies on the effects of processing parameters also help to optimize the manufacturing steps for a particular structural element. Two methods of analyzing design detail/process cost relationships developed for the design cost model were pursued in the current study. The first makes use of existing databases and alternative cost modeling methods (e.g. detailed estimating). The second compares design cost model predictions with data collected during the fabrication of seven foot circumferential frames for ATCAS crown test panels. The process used in this case involves 2D dry braiding and resin transfer molding of curved 'J' cross section frame members having design details characteristic of the baseline ATCAS crown design.

77 FR 36134 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-06-18

... chem- mill steps, which could result in sudden fracture and failure of the fuselage skin panels, and...-mill steps, which could result in sudden fracture and failure of the fuselage skin panels, and...

Aircraft drag prediction and reduction. Addendum 1: Computational drag analyses and minimization; mission impossible?

NASA Technical Reports Server (NTRS)

Slooff, J. W.

1986-01-01

The Special Course on Aircraft Drag Prediction was sponsored by the AGARD Fluid Dynamics Panel and the von Karman Institute and presented at the von Karman Institute, Rhode-Saint-Genese, Belgium, on 20 to 23 May 1985 and at the NASA Langley Research Center, Hampton, Virginia, USA, 5 to 6 August 1985. The course began with a general review of drag reduction technology. Then the possibility of reduction of skin friction through control of laminar flow and through modification of the structure of the turbulence in the boundary layer were discussed. Methods for predicting and reducing the drag of external stores, of nacelles, of fuselage protuberances, and of fuselage afterbodies were then presented followed by discussion of transonic drag rise. The prediction of viscous and wave drag by a method matching inviscid flow calculations and boundary layer integral calculations, and the reduction of transonic drag through boundary layer control are also discussed. This volume comprises Paper No. 9 Computational Drag Analyses and Minimization: Mission Impossible, which was not included in AGARD Report 723 (main volume).

Helicopter Pitch-Control Mechanism Reduces Vibration

NASA Technical Reports Server (NTRS)

Lemont, H.

1986-01-01

Large forces accommodated without increasing weight of helicopter structure. New mechanism yields stiffer control and improves accuracy of pitch changes under load. As result, heavy casting not for gearbox, nor extra reinforcing members needed for fuselage bulkheads, stringers, skin, and other parts. In new mechanism, reaction forces developed in rotor hub. Long load paths to gearbox and fuselage elminated. Reaction member rigidly attached to hub and rotates with it. At lower end of reaction member, bearing forms bridge to fuselage through stationary beam and antirotation link. Beam connected to reaction plate through rods.

Closeup oblique view of the aft fuselage of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up oblique view of the aft fuselage of the Orbiter Discovery looking forward and port with the Space Shuttle Main Engines (SSME) and Orbiter Maneuvering System/Reaction Control System pods still in place. However. the heat shields have been removed from the SSMEs providing a good view toward the interior of the aft fuselage. This image was taken inside the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Closeup oblique view of the aft fuselage of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up oblique view of the aft fuselage of the Orbiter Discovery looking forward and starboard with the Space Shuttle Main Engines (SSME) and Orbiter Maneuvering System/Reaction Control System pods removed. The openings for the SSMEs have been covered with a flexible barrier to create a positive pressure envelope inside of the aft fuselage. This image was taken inside the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Closeup oblique view of the aft fuselage of the Orbiter ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up oblique view of the aft fuselage of the Orbiter Discovery looking forward and starboard with the Space Shuttle Main Engines (SSME) and Orbiter Maneuvering System/Reaction Control System pods still in place. However. the heat shields have been removed from the SSMEs providing a good view toward the interior of the aft fuselage. This image was taken inside the Orbiter Processing Facility at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Experimental Investigation of the Effect of Vertical-tail Size and Length and of Fuselage Shape and Length on the Static Lateral Stability Characteristics of a Model with 45 Degree Sweptback Wing and Tail Surfaces

NASA Technical Reports Server (NTRS)

Queijo, M J; Wolhart, Walter D

1951-01-01

An investigation was made to determine the effects of vertical-tail size and length and of fuselage shape and length on the static lateral stability characteristics of a model with wing and vertical tails having the quarter-chord lines swept back 45 degrees. The results indicate that the directional instability of the various isolated fuselages was about two-thirds as large as that predicted by classical theory.

Evaluation of the fuselage lap joint fatigue and terminating action repair

NASA Technical Reports Server (NTRS)

Samavedam, Gopal; Thomson, Douglas; Jeong, David Y.

1994-01-01

Terminating action is a remedial repair which entails the replacement of shear head countersunk rivets with universal head rivets which have a larger shank diameter. The procedure was developed to eliminate the risk of widespread fatigue damage (WFD) in the upper rivet row of a fuselage lap joint. A test and evaluation program has been conducted by Foster-Miller, Inc. (FMI) to evaluate the terminating action repair of the upper rivet row of a commercial aircraft fuselage lap splice. Two full scale fatigue tests were conducted on fuselage panels using the growth of fatigue cracks in the lap joint. The second test was performed to evaluate the effectiveness of the terminating action repair. In both tests, cyclic pressurization loading was applied to the panels while crack propagation was recorded at all rivet locations at regular intervals to generate detailed data on conditions of fatigue crack initiation, ligament link-up, and fuselage fracture. This program demonstrated that the terminating action repair substantially increases the fatigue life of a fuselage panel structure and effectively eliminates the occurrence of cracking in the upper rivet row of the lap joint. While high cycle crack growth was recorded in the middle rivet row during the second test, failure was not imminent when the test was terminated after cycling to well beyond the service life. The program also demonstrated that the initiation, propagation, and linkup of WFD in full-scale fuselage structures can be simulated and quantitatively studied in the laboratory. This paper presents an overview of the testing program and provides a detailed discussion of the data analysis and results. Crack distribution and propagation rates and directions as well as frequency of cracking are presented for both tests. The progression of damage to linkup of adjacent cracks and to eventual overall panel failure is discussed. In addition, an assessment of the effectiveness of the terminating action repair and the occurrence of cracking in the middle rivet row is provided, and conclusions of practical interest are drawn.

Etude numerique et experimentale de la reponse vibro-acoustique des structures raidies a des excitations aeriennes et solidiennes

NASA Astrophysics Data System (ADS)

Mejdi, Abderrazak

Les fuselages des avions sont generalement en aluminium ou en composite renforces par des raidisseurs longitudinaux (lisses) et transversaux (cadres). Les raidisseurs peuvent etre metalliques ou en composite. Durant leurs differentes phases de vol, les structures d'avions sont soumises a des excitations aeriennes (couche limite turbulente : TBL, champs diffus : DAF) sur la peau exterieure dont l'energie acoustique produite se transmet a l'interieur de la cabine. Les moteurs, montes sur la structure, produisent une excitation solidienne significative. Ce projet a pour objectifs de developper et de mettre en place des strategies de modelisations des fuselages d'avions soumises a des excitations aeriennes et solidiennes. Tous d'abord, une mise a jour des modeles existants de la TBL apparait dans le deuxieme chapitre afin de mieux les classer. Les proprietes de la reponse vibro-acoustique des structures planes finies et infinies sont analysees. Dans le troisieme chapitre, les hypotheses sur lesquelles sont bases les modeles existants concernant les structures metalliques orthogonalement raidies soumises a des excitations mecaniques, DAF et TBL sont reexamines en premier lieu. Ensuite, une modelisation fine et fiable de ces structures est developpee. Le modele est valide numeriquement a l'aide des methodes des elements finis (FEM) et de frontiere (BEM). Des tests de validations experimentales sont realises sur des panneaux d'avions fournis par des societes aeronautiques. Au quatrieme chapitre, une extension vers les structures composites renforcees par des raidisseurs aussi en composites et de formes complexes est etablie. Un modele analytique simple est egalement implemente et valide numeriquement. Au cinquieme chapitre, la modelisation des structures raidies periodiques en composites est beaucoup plus raffinee par la prise en compte des effets de couplage des deplacements planes et transversaux. L'effet de taille des structures finies periodiques est egalement pris en compte. Les modeles developpes ont permis de conduire plusieurs etudes parametriques sur les proprietes vibro-acoustiques des structures d'avions facilitant ainsi la tache des concepteurs. Dans le cadre de cette these, un article a ete publie dans le Journal of Sound and Vibration et trois autres soumis, respectivement aux Journal of Acoustical Society of America, International Journal of Solid Mechanics et au Journal of Sound and Vibration Mots cles : structures raidies, composites, vibro-acoustique, perte par transmission.

14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

Code of Federal Regulations, 2010 CFR

2010-01-01

... engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (Secs. 604, 605, 72 Stat. 778, 49 U.S.C...

14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

Code of Federal Regulations, 2013 CFR

2013-01-01

... engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (Secs. 604, 605, 72 Stat. 778, 49 U.S.C...

Nondestructive Inspection of Piper PA-25 Forward Spar Fuselage Attachment Fitting

DOT National Transportation Integrated Search

1995-09-01

The Federal Aviation Administration's (FAA's) Aging Aircraft NDI Validation : Center (AANC) at Sandia National Laboratoriess applied two nondestructive : inspection (NDI) techniques for the inspection of a Piper PA-25 forward spar : fuselage attachme...

Airbag system and method for facilitating emergency egress from an aircraft

NASA Technical Reports Server (NTRS)

Rawdon, Blaine K. (Inventor); Hawley, Arthur V. (Inventor)

2002-01-01

An airbag system for elevating the fuselage of an aircraft off a landing surface a sufficient degree to allow for emergency egress of passengers and crew through ventral emergency exit doors. An airbag assembly made up of a plurality of independent airbags is disposed within the aircraft. When activated, the airbag system deploys the airbags external of the aircraft that elevate the fuselage of the aircraft a sufficient degree to allow for utilizing the ventral emergency exit doors on the fuselage to enable evacuating the passengers and crew. An activation mechanism is connected to the inflation.devices associated with each of the airbags. The activation mechanism generates an electrical signal which activates the inflation devices, which in turn fill the airbags with a compressed fluid, thus expanding the airbags and lifting the fuselage. A crew member initiates the activation of the airbag system through one or more switches.

XSECT: A computer code for generating fuselage cross sections - user's manual

NASA Technical Reports Server (NTRS)

Ames, K. R.

1982-01-01

A computer code, XSECT, has been developed to generate fuselage cross sections from a given area distribution and wing definition. The cross sections are generated to match the wing definition while conforming to the area requirement. An iterative procedure is used to generate each cross section. Fuselage area balancing may be included in this procedure if desired. The code is intended as an aid for engineers who must first design a wing under certain aerodynamic constraints and then design a fuselage for the wing such that the contraints remain satisfied. This report contains the information necessary for accessing and executing the code, which is written in FORTRAN to execute on the Cyber 170 series computers (NOS operating system) and produces graphical output for a Tektronix 4014 CRT. The LRC graphics software is used in combination with the interface between this software and the PLOT 10 software.

Noise transmission and reduction in turboprop aircraft

NASA Astrophysics Data System (ADS)

MacMartin, Douglas G.; Basso, Gordon L.; Leigh, Barry

1994-09-01

There is considerable interest in reducing the cabin noise environment in turboprop aircraft. Various approaches have been considered at deHaviland Inc., including passive tuned-vibration absorbers, speaker-based noise cancellation, and structural vibration control of the fuselage. These approaches will be discussed briefly. In addition to controlling the noise, a method of predicting the internal noise is required both to evaluate potential noise reduction approaches, and to validate analytical design models. Instead of costly flight tests, or carrying out a ground simulation of the propeller pressure field, a much simpler reciprocal technique can be used. A capacitive scanner is used to measure the fuselage vibration response on a deHaviland Dash-8 fuselage, due to an internal noise source. The approach is validated by comparing this reciprocal noise transmission measurement with the direct measurement. The fuselage noise transmission information is then combined with computer predictions of the propeller pressure field data to predict the internal noise at two points.

A theoretical investigation of noise reduction through the cylindrical fuselage of a twin-engine, propeller-driven aircraft

NASA Technical Reports Server (NTRS)

Bhat, R. B.; Mixson, J. S.

1978-01-01

Interior noise in the fuselage of a twin-engine, propeller-driven aircraft with two propellers rotating in opposite directions is studied analytically. The fuselage was modeled as a stiffened cylindrical shell with simply supported ends, and the effects of stringers and frames were averaged over the shell surface. An approximate mathematical model of the propeller noise excitation was formulated which includes some of the propeller noise characteristics such as sweeping pressure waves around the sidewalls due to propeller rotation and the localized nature of the excitation with the highest levels near the propeller plane. Results are presented in the form of noise reduction, which is the difference between the levels of external and interior noise. The influence of propeller noise characteristics on the noise reduction was studied. The results indicate that the sweep velocity of the excitation around the fuselage sidewalls is critical to noise reduction.

Subsonic and supersonic longitudinal stability and control characteristics of an aft tail fighter configuration with cambered and uncambered wings and uncambered fuselage

NASA Technical Reports Server (NTRS)

Dollyhigh, S. M.

1974-01-01

An investigation has been made in the Mach number range from 0.20 to 2.16 to determine the longitudinal aerodynamic characteristics of a fighter airplane concept. The configuration concept employs a single fixed geometry inlet, a 50 deg leading-edge-angle clipped-arrow wing, a single large vertical tail, and low horizontal tails. The wing camber surface was optimized in drag due to lift and was designed to be self-trimming at Mach 1.40 and at a lift coefficient of 0.20. An uncambered or flat wing of the same planform and thickness ratio was also tested. However, for the present investigation, the fuselage was not cambered. Further tests should be made on a cambered fuselage version, which attempts to preserve the optimum wing loading on that part of the theoretical wing enclosed by the fuselage.

NDE scanning and imaging of aircraft structure

NASA Astrophysics Data System (ADS)

Bailey, Donald; Kepler, Carl; Le, Cuong

1995-07-01

The Science and Engineering Lab at McClellan Air Force Base, Sacramento, Calif. has been involved in the development and use of computer-based scanning systems for NDE (nondestructive evaluation) since 1985. This paper describes the history leading up to our current applications which employ eddy current and ultrasonic scanning of aircraft structures that contain both metallics and advanced composites. The scanning is performed using industrialized computers interfaced to proprietary acquisition equipment and software. Examples are shown that image several types of damage such as exfoliation and fuselage lap joint corrosion in aluminum, impact damage, embedded foreign material, and porosity in Kevlar and graphite epoxy composites. Image analysis techniques are reported that are performed using consumer oriented computer hardware and software that are not NDE specific and not expensive

Repair of Composites: Design Choices Leading to Lower Life-Cycle Cost

NASA Astrophysics Data System (ADS)

Kassapoglou, Christos; Rangelov, Konstantin; Rangelov, Svilen

2017-08-01

The fabrication cost of composite aircraft structures is revisited and the effect of part size on cost is examined with emphasis on design decisions which affect the ease of (bonded) repair and the total cost of the part and subsequent repairs. The case of moderately loaded stiffened fuselage or wing panels under compression is analysed in detail and the fabrication cost of the panel made as a single piece or as an assembly of smaller identical components or modules is determined. The cost of special purpose repairs for two different damage sizes is compared to removing and replacing damaged modules. Hand layup and automated processing are compared. It is found that for certain repair sizes removing and replacing modules leads to lower overall cost as compared to applying a special purpose repair.

Summary of AH-1G flight vibration data for validation of coupled rotor-fuselage analyses

NASA Technical Reports Server (NTRS)

Dompka, R. V.; Cronkhite, J. D.

1986-01-01

Under a NASA research program designated DAMVIBS (Design Analysis Methods for VIBrationS), four U. S. helicopter industry participants (Bell Helicopter, Boeing Vertol, McDonnell Douglas Helicopter, and Sikorsky Aircraft) are to apply existing analytical methods for calculating coupled rotor-fuselage vibrations of the AH-1G helicopter for correlation with flight test data from an AH-1G Operational Load Survey (OLS) test program. Bell Helicopter, as the manufacturer of the AH-1G, was asked to provide pertinent rotor data and to collect the OLS flight vibration data needed to perform the correlations. The analytical representation of the fuselage structure is based on a NASTRAN finite element model (FEM) developed by Bell which has been extensively documented and correlated with ground vibration tests.The AH-1G FEM was provided to each of the participants for use in their coupled rotor-fuselage analyses. This report describes the AH-1G OLS flight test program and provides the flight conditions and measured vibration data to be used by each participant in their correlation effort. In addition, the mechanical, structural, inertial and aerodynamic data for the AH-1G two-bladed teetering main rotor system are presented. Furthermore, modifications to the NASTRAN FEM of the fuselage structure that are necessary to make it compatible with the OLS test article are described. The AH-1G OLS flight test data was found to be well documented and provide a sound basis for evaluating currently existing analysis methods used for calculation of coupled rotor-fuselage vibrations.

Analytical Fuselage and Wing Weight Estimation of Transport Aircraft

DOT National Transportation Integrated Search

1996-05-01

A method of estimating the load-bearing fuselage weight and wing weight of transport aircraft based on fundamental structural principles has been developed. This method of weight estimation represents a compromise between the rapid assessment of comp...

Lateral-directional stability characteristics of a wing-fuselage configuration at angles of attack up to 44 deg

NASA Technical Reports Server (NTRS)

Henderson, W. P.; Huffman, J. K.

1974-01-01

An investigation has been conducted to determine the effects of configuration variables on the lateral-directional stability characteristics of a wing-fuselage configuration. The variables under study included variations in the location of a single center-line vertical tail and twin vertical tails, wing height, fuselage strakes, and horizontal tails. The study was conducted in the Langley high-speed 7-by 10-foot tunnel at a Mach number of 0.30, at angles of attack up to 44 deg and at sideslip angles of 0 deg and plus or minus 5 deg.

A mixed optimization method for automated design of fuselage structures.

NASA Technical Reports Server (NTRS)

Sobieszczanski, J.; Loendorf, D.

1972-01-01

A procedure for automating the design of transport aircraft fuselage structures has been developed and implemented in the form of an operational program. The structure is designed in two stages. First, an overall distribution of structural material is obtained by means of optimality criteria to meet strength and displacement constraints. Subsequently, the detailed design of selected rings and panels consisting of skin and stringers is performed by mathematical optimization accounting for a set of realistic design constraints. The practicality and computer efficiency of the procedure is demonstrated on cylindrical and area-ruled large transport fuselages.

A three-dimensional, compressible, laminar boundary-layer method for general fuselages. Volume 1: Numerical method

NASA Technical Reports Server (NTRS)

Wie, Yong-Sun

1990-01-01

A procedure for calculating 3-D, compressible laminar boundary layer flow on general fuselage shapes is described. The boundary layer solutions can be obtained in either nonorthogonal 'body oriented' coordinates or orthogonal streamline coordinates. The numerical procedure is 'second order' accurate, efficient and independent of the cross flow velocity direction. Numerical results are presented for several test cases, including a sharp cone, an ellipsoid of revolution, and a general aircraft fuselage at angle of attack. Comparisons are made between numerical results obtained using nonorthogonal curvilinear 'body oriented' coordinates and streamline coordinates.

Wing-Fuselage Interference, Tail Buffeting, and Air Flow About the Tail of a Low-Wing Monoplane

NASA Technical Reports Server (NTRS)

White, James A; Hood, Manley J

1935-01-01

This report presents the results of wind tunnel tests on a Mcdonnell Douglas airplane to determine the wing-fuselage interference of a low-wing monoplane. The tests included a study of tail buffeting and the air flow in the region of the tail. The airplane was tested with and without the propeller slipstream, both in the original condition and with several devices designed to reduce or eliminate tail buffeting. The devices used were wing-fuselage fillets, a NACA cowling, reflexed trailing edge of the wing, and stub auxiliary airfoils.

Detail view of the port side of the aft fuselage ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the port side of the aft fuselage of the Orbiter Discovery in the transfer aisle of the Vehicle Assembly Building at Kennedy Space Center with a lifting frame attached to the aft attach points of the orbiter. In this view, the Orbiter Maneuvering/Reaction Control Systems pod is in place. Also note the darker-colored trapezoidal aft fuselage access door and the T-0 umbilical panel to its right in the view. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Experimental modal analysis of the fuselage panels of an Aero Commander aircraft

NASA Technical Reports Server (NTRS)

Geisler, D.

1981-01-01

The reduction of interior noise in light aircraft was investigated with emphasis the thin fuselage sidewall. The approach used is theoretical and involves modeling of the sidewall panels and stiffeners. Experimental data obtained from tests investigating the effects of mass and stiffness treatments to the sidewalls are presented. The dynamic characteristics of treated panels are contrasted with the untreated sidewall panels using experimental modal analysis techniques. The results include the natural frequencies, modal dampling, and mode shapes of selected panels. Frequency response functions, data relating to the global fuselage response, and acoustic response are also presented.

Evaluation of the concept of pressure proof testing fuselage structures

NASA Technical Reports Server (NTRS)

Harris, Charles E.; Orringer, Oscar

1991-01-01

The FAA and NASA have recently completed independent technical evaluations of the concept of pressure proof testing the fuselage of commercial transport airplanes. The results of these evaluations are summarized. The objectives of the evaluations were to establish the potential benefit of the pressure proof test, to quantify the most desirable proof test pressure, and to quantify the required proof test interval. The focus of the evaluations was on multiple-site cracks extending from adjacent rivet holes of a typical fuselage longitudinal lap splice joint. The FAA and NASA do not support pressure proof testing the fuselage of aging commercial transport aircraft. The argument against proof testing is as follows: (1) a single proof test does not insure an indefinite life; therefore, the proof test must be repeated at regular intervals; (2) for a proof factor of 1.33, the required proof test interval must be below 300 flights to account for uncertainties in the evaluation; (3) conducting the proof test at a proof factor of 1.5 would considerably exceed the fuselage design limit load; therefore, it is not consistent with accepted safe practices; and (4) better safety can be assured by implementing enhanced nondestructive inspection requirements, and adequate reliability can be achieved by an inspection interval several times longer than the proof test interval.

Numerical simulation of helicopter engine plume in forward flight

NASA Technical Reports Server (NTRS)

Dimanlig, Arsenio C. B.; Vandam, Cornelis P.; Duque, Earl P. N.

1994-01-01

Flowfields around helicopters contain complex flow features such as large separated flow regions, vortices, shear layers, blown and suction surfaces and an inherently unsteady flow imposed by the rotor system. Another complicated feature of helicopters is their infrared signature. Typically, the aircraft's exhaust plume interacts with the rotor downwash, the fuselage's complicated flowfield, and the fuselage itself giving each aircraft a unique IR signature at given flight conditions. The goal of this project was to compute the flow about a realistic helicopter fuselage including the interaction of the engine air intakes and exhaust plume. The computations solve the Think-Layer Navier Stokes equations using overset type grids and in particular use the OVERFLOW code by Buning of NASA Ames. During this three month effort, an existing grid system of the Comanche Helicopter was to be modified to include the engine inlet and the hot engine exhaust. The engine exhaust was to be modeled as hot air exhaust. However, considerable changes in the fuselage geometry required a complete regriding of the surface and volume grids. The engine plume computations have been delayed to future efforts. The results of the current work consists of a complete regeneration of the surface and volume grids of the most recent Comanche fuselage along with a flowfield computation.

Winding trajectories of noncircular composite shells

NASA Astrophysics Data System (ADS)

Nikityuk, V. A.; Fedorov, V. V.

1995-07-01

An approach has been proposed for determination of the trajectory parameters of a layer formed by winding of continuous ribbons on a complicated surface. An algorithm has been developed for determining the geodesic trajectories of the reinforcement fiber arrangement, reinforcement angles, and geodesic deviation angles. Conditions have been formulated for positional stability of the ribbons on the surface and avoidance of gaps and overlapping between the ribbons along with restrictions to the surface form. Results are given for a calculation of the geodesic turn parameters on a fuselage surface, which is not a surface of revolution, of a light airplane.

Structural evaluation of curved stiffened composite panels fabricated using a THERM-Xsm process

NASA Technical Reports Server (NTRS)

Kassapoglou, Christos; Dinicola, Albert J.; Chou, Jack C.; Deaton, Jerry W.

1991-01-01

The use of composites in aircraft structures is often limited by material and manufacturing costs which, for some designs and applications, are prohibitively high. To increase the frequency of application of composites in primary airframe components alternative manufacturing processes are sought that reduce cost and/or enhance structural efficiency. One alternative process involves the use of THERM-Xsm as the pressure transfer medium during autoclave curing. THERM-Xsm, a silicon-based flow able polymer which behaves like a liquid under autoclave presssure, transmits quasi-hydrostatic pressure to all contacting surfaces of the part to be cured. Once the autoclave pressure is relieved, THERM-Xsm reverts back to the powdery solid state and can be reused many times. The THERM-Xsm process to be evaluated is depicted and consists of (1) enclosing the tool and part to be cured by a set of frames that create a box, (2) pouring THERM-Xsm powder onto the part and filling the box, and (3) placing a vacuum bag over the box assembly. In this program, a separating non-porous film (Teflon) was placed between the part to be cured and THERM-Xsm powder to avoid any contamination. The use of THERM-Xsm has two significant advantages over conventional manufacturing procedures. First, it eliminates complicated hard tooling since it guarantees uniform pressure transfer and thus, good compaction at complex structural details (such as frame-stiffener intersections and corners). Second, it greatly simplifies vacuum bagging, since once the part to be cured is covered by THERM-Xsm powder, the vacuum bag need only conform to a relatively flat shape reducing significantly the number of pleats required. A program is on-going at Sikorsky Aircraft to evaluate the structural performance of complex composite fuselage structures made with this THERM-Xsm process and to quantify the impact of THERM-Xsm on manufacturing labor hours and cost. The program involves fuselage panel optimization analysis, a building block test program where structural details representative of the full-scale article are analyzed and tested, and static and fatigue test/analysis of the full-scale test articles. The main results of this program are reported.

VIEW OF LEFT WING AND FUSELAGE FROM TOP LEVEL OF ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

VIEW OF LEFT WING AND FUSELAGE FROM TOP LEVEL OF TAIL DOCK STAND. LEADING AND TRAILING EDGE FLAPS ARE DOWN; AIELERONS ARE IN NEUTRAL. ENGINE COWLING OFF FOR HEAVY INSPECTION. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

A laboratory study of multiple site damage in fuselage lap splices

DOT National Transportation Integrated Search

1993-12-01

This report details an experimental study that was conducted to explore the causes of : fuselage lap splice multiple site damage (MSD), which has been observed in several : aging aircraft. MSD was partially responsible for the 1988 Aloha Airlines acc...

The Design and Operation of Suborbital Low Cost and Low Risk Vehicle to the Edge of Space (SOLVES)

NASA Astrophysics Data System (ADS)

Ridzuan Zakaria, Norul; Nasrun, Nasri; Rashidy Zulkifi, Mohd; Izmir Yamin, Mohd; Othman, Jamaludin; Rafidi Zakaria, Norul

2013-09-01

Inclusive in the planning of Spaceport Malaysia are 2 local suborbital vehicles development. One of the vehicles is called SOLVES or Suborbital Low Cost and Low Risk Vehicle to the Edge of Space. The emphasis on the design and operation of SOLVES is green and robotic technology, where both green technology and robotic technology are used to protect the environment and enhance safety. As SOLVES climbs, its center of gravity stabilizes and remains at the bottom as its propellant being used until it depletes, due to the position of the vehicle's passenger cabin and its engines at its lower end. It will reach 80km from sea level generally known as "the edge of space" due to its momentum although its propellant will be depleted at a lower altitude. As the suborbital vehicle descends tail first, its wings automatically extend and rotate at horizontal axes perpendicular to the fuselage. These naturally and passively rotating wings ensure controlled low velocity and stable descend of the vehicle. The passenger cabin also rotates automatically at a steady low speed at the centerline of its fuselage as it descends, caused naturally by the lift force, enabling its passengers a surrounding 360 degrees view. SOLVES is steered automatically to its landing point by an electrical propulsion system with a vectoring nozzle. The electrical propulsion minimizes space and weight and is free of pollution and noise. Its electrical power comes from a battery aided by power generated by the naturally rotating wings. When the vehicle lands, it is in the safest mode as its propellant is depleted and its center of gravity remains at the bottom of its cabin. The cabin, being located at the bottom of the fuselage, enables very convenient, rapid and safe entry and exit of its passengers. SOLVES will be a robotic suborbital vehicle with green technology. The vehicle will carry 4 passengers and each passenger will be trained to land the vehicle manually if the fully automated landing system fails and therefore it will be engineered for simple operation by trained passengers. However, for certification by aviation authorities the vehicle may be operational with 3 passengers and a pilot. A specific operation considered for SOLVES is navaloperation where the suborbital vehicle will be operating from a seaborne spaceport, probably a superyacht with spacepad for the vertical launching and landing of the vehicle. Such naval operation enables the vehicle to fly above exotic locations reachable by sea. SOLVES is also planned for further development into reusable rocket booster to carry small suborbiter to 160km from sea level, enables the passengers aboard the suborbiter to experience longer zero gravity time and more effective suborbital flight.

14 CFR 25.783 - Fuselage doors.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Fuselage doors. 25.783 Section 25.783 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... not latch after closing; and (ii) With jamming as a result of mechanical failure or blocking debris...

14 CFR 25.783 - Fuselage doors.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Fuselage doors. 25.783 Section 25.783 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... not latch after closing; and (ii) With jamming as a result of mechanical failure or blocking debris...

14 CFR 25.783 - Fuselage doors.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Fuselage doors. 25.783 Section 25.783 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... not latch after closing; and (ii) With jamming as a result of mechanical failure or blocking debris...

6. Detail of forward fuselage showing open cockpit hatch and ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

6. Detail of forward fuselage showing open cockpit hatch and ladder. View to southeast. - Offutt Air Force Base, Looking Glass Airborne Command Post, Looking Glass Aircraft, On Operational Apron covering northeast half of Project Looking Glass Historic District, Bellevue, Sarpy County, NE

76 FR 74667 - Airworthiness Directives; The Boeing Company Model 737-200, -200C, -300, -400, and -500 Series...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-12-01

... at the chem-milled steps, which could result in sudden fracture and failure of the fuselage skin... chem-milled steps, which could result in sudden fracture and failure of the fuselage skin panels, and...

14 CFR 25.783 - Fuselage doors.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Fuselage doors. 25.783 Section 25.783 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... not latch after closing; and (ii) With jamming as a result of mechanical failure or blocking debris...

14 CFR 25.783 - Fuselage doors.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Fuselage doors. 25.783 Section 25.783 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION AIRCRAFT AIRWORTHINESS... not latch after closing; and (ii) With jamming as a result of mechanical failure or blocking debris...

Mechanical characterization of 2D, 2D stitched, and 3D braided/RTM materials

NASA Technical Reports Server (NTRS)

Deaton, Jerry W.; Kullerd, Susan M.; Portanova, Marc A.

1993-01-01

Braided composite materials have potential for application in aircraft structures. Fuselage frames, floor beams, wing spars, and stiffeners are examples where braided composites could find application if cost effective processing and damage tolerance requirements are met. Another important consideration for braided composites relates to their mechanical properties and how they compare to the properties of composites produced by other textile composite processes being proposed for these applications. Unfortunately, mechanical property data for braided composites do not appear extensively in the literature. Data are presented in this paper on the mechanical characterization of 2D triaxial braid, 2D triaxial braid plus stitching, and 3D (through-the-thickness) braid composite materials. The braided preforms all had the same graphite tow size and the same nominal braid architectures, (+/- 30 deg/0 deg), and were resin transfer molded (RTM) using the same mold for each of two different resin systems. Static data are presented for notched and unnotched tension, notched and unnotched compression, and compression after impact strengths at room temperature. In addition, some static results, after environmental conditioning, are included. Baseline tension and compression fatigue results are also presented, but only for the 3D braided composite material with one of the resin systems.

Analytical Fuselage and Wing Weight Estimation of Transport Aircraft

NASA Technical Reports Server (NTRS)

Chambers, Mark C.; Ardema, Mark D.; Patron, Anthony P.; Hahn, Andrew S.; Miura, Hirokazu; Moore, Mark D.

1996-01-01

A method of estimating the load-bearing fuselage weight and wing weight of transport aircraft based on fundamental structural principles has been developed. This method of weight estimation represents a compromise between the rapid assessment of component weight using empirical methods based on actual weights of existing aircraft, and detailed, but time-consuming, analysis using the finite element method. The method was applied to eight existing subsonic transports for validation and correlation. Integration of the resulting computer program, PDCYL, has been made into the weights-calculating module of the AirCraft SYNThesis (ACSYNT) computer program. ACSYNT has traditionally used only empirical weight estimation methods; PDCYL adds to ACSYNT a rapid, accurate means of assessing the fuselage and wing weights of unconventional aircraft. PDCYL also allows flexibility in the choice of structural concept, as well as a direct means of determining the impact of advanced materials on structural weight. Using statistical analysis techniques, relations between the load-bearing fuselage and wing weights calculated by PDCYL and corresponding actual weights were determined.

An Airplane Design having a Wing with Fuselage Attached to Each Tip

NASA Technical Reports Server (NTRS)

Spearman, Leroy M.

2001-01-01

This paper describes the conceptual design of an airplane having a low aspect ratio wing with fuselages that are attached to each wing tip. The concept is proposed for a high-capacity transport as an alternate to progressively increasing the size of a conventional transport design having a single fuselage with cantilevered wing panels attached to the sides and tail surfaces attached at the rear. Progressively increasing the size of conventional single body designs may lead to problems in some area's such as manufacturing, ground-handling and aerodynamic behavior. A limited review will be presented of some past work related to means of relieving some size constraints through the use of multiple bodies. Recent low-speed wind-tunnel tests have been made of models representative of the inboard-wing concept. These models have a low aspect ratio wing with a fuselage attached to each tip. Results from these tests, which included force measurements, surface pressure measurements, and wake surveys, will be presented herein.

Coupled CFD/CSD Analysis of an Active-Twist Rotor in a Wind Tunnel with Experimental Validation

NASA Technical Reports Server (NTRS)

Massey, Steven J.; Kreshock, Andrew R.; Sekula, Martin K.

2015-01-01

An unsteady Reynolds averaged Navier-Stokes analysis loosely coupled with a comprehensive rotorcraft code is presented for a second-generation active-twist rotor. High fidelity Navier-Stokes results for three configurations: an isolated rotor, a rotor with fuselage, and a rotor with fuselage mounted in a wind tunnel, are compared to lifting-line theory based comprehensive rotorcraft code calculations and wind tunnel data. Results indicate that CFD/CSD predictions of flapwise bending moments are in good agreement with wind tunnel measurements for configurations with a fuselage, and that modeling the wind tunnel environment does not significantly enhance computed results. Actuated rotor results for the rotor with fuselage configuration are also validated for predictions of vibratory blade loads and fixed-system vibratory loads. Varying levels of agreement with wind tunnel measurements are observed for blade vibratory loads, depending on the load component (flap, lag, or torsion) and the harmonic being examined. Predicted trends in fixed-system vibratory loads are in good agreement with wind tunnel measurements.

A comparison of the structureborne and airborne paths for propfan interior noise

NASA Technical Reports Server (NTRS)

Eversman, W.; Koval, L. R.; Ramakrishnan, J. V.

1986-01-01

A comparison is made between the relative levels of aircraft interior noise related to structureborne and airborne paths for the same propeller source. A simple, but physically meaningful, model of the structure treats the fuselage interior as a rectangular cavity with five rigid walls. The sixth wall, the fuselage sidewall, is a stiffened panel. The wing is modeled as a simple beam carried into the fuselage by a large discrete stiffener representing the carry-through structure. The fuselage interior is represented by analytically-derived acoustic cavity modes and the entire structure is represented by structural modes derived from a finite element model. The noise source for structureborne noise is the unsteady lift generation on the wing due to the rotating trailing vortex system of the propeller. The airborne noise source is the acoustic field created by a propeller model consistent with the vortex representation. Comparisons are made on the basis of interior noise over a range of propeller rotational frequencies at a fixed thrust.

Comparison of Vee-Type and Conventional Tail Surfaces in Combination with Fuselage and Wing in the Variable-Density Tunnel

NASA Technical Reports Server (NTRS)

Greenberg, Harry

1941-01-01

The pitching and the yawing moments of a vee-type and a conventional type of tail surface were measured. The tests were made in the presence of a fuselage and a wing-fuselage combination in such a way as to determine the moments contributed by the tail surfaces. The results showed that the vee-type tail tested, with a dihedral angle of 35.3 deg, was about 71 percent as effective in pitch as the conventional tail and had a yawing-moment to pitching-moment ratio of 0.3. The conventional tail, the panels of which were all congruent to those of the vee-type tail, had a yawing-moment to pitching-moment ratio of 0.48. These ratios are in fair agreement with values calculated by methods shown in this and previous reports. The values of the measured moments were reduced from 15 to 25 percent of the calculated value by fuselage interference.

Wind tunnel investigation of helicopter rotor wake effects on three helicopter fuselage models

NASA Technical Reports Server (NTRS)

Wilson, J. C.; Mineck, R. E.

1974-01-01

The effects of rotor downwash on helicopter fuselage aerodynamic characteristics were investigated. A rotor model for generating the downwash was mounted close to each of three fuselage models. The main report presents the force and moment data in both graphical and tabular form and the pressure data in graphical form. This supplement presents the pressure data in tabular form. Each run or parameter sweep is identified by a unique run number. The data points in each run are identified by a point number. The pressure data can be matched to the force data by matching the run and point number.

Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [conducted in Ames 40- by 80-foot wind tunnel

NASA Technical Reports Server (NTRS)

Garland, D. B.

1980-01-01

Modifications were made to the model to improve longitudinal acceleration capability during transition from hovering to wing borne flight. A rearward deflection of the fuselage augmentor thrust vector is shown to be beneficial in this regard. Other agmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also shows negligible influence on the performance of the wing and of the fuselage augmentor.

The Growth of Multi-Site Fatigue Damage in Fuselage Lap Joints

NASA Technical Reports Server (NTRS)

Piascik, Robert S.; Willard, Scott A.

1999-01-01

Destructive examinations were performed to document the progression of multi-site damage (MSD) in three lap joint panels that were removed from a full scale fuselage test article that was tested to 60,000 full pressurization cycles. Similar fatigue crack growth characteristics were observed for small cracks (50 microns to 10 mm) emanating from counter bore rivets, straight shank rivets, and 100 deg counter sink rivets. Good correlation of the fatigue crack growth data base obtained in this study and FASTRAN Code predictions show that the growth of MSD in the fuselage lap joint structure can be predicted by fracture mechanics based methods.

Axial crack propagation and arrest in pressurized fuselage

NASA Technical Reports Server (NTRS)

Kosai, M.; Shimamoto, A.; Yu, C.-T.; Walker, S. I.; Kobayashi, A. S.; Tan, P.

1994-01-01

The crack arrest capability of a tear strap in a pressurized precracked fuselage was studied through instrumented axial rupture tests of small scale models of an idealized fuselage. Upon pressurization, rapid crack propagation initiated at an axial through crack along the stringer and immediately kinked due to the mixed modes 1 and 2 state caused by the one-sided opening of the crack flap. The diagonally running crack further turned at the tear straps. Dynamic finite element analysis of the rupturing cylinder showed that the crack kinked and also ran straight in the presence of a mixed mode state according to a modified two-parameter crack kinking criterion.

Detail view of the starboard side of the aft fuselage ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detail view of the starboard side of the aft fuselage of the Orbiter Discovery in the Orbiter Processing Facility at Kennedy Space Center with the Orbiter Maneuvering/Reaction Control Systems Pod removed and exposing the insulating foil used to protect the orbiter structure from the heat generated by the maneuvering and reaction control engines. Also note in the view that the aft fuselage access door has bee removed and also note the ground support equipment attached to the T-0 umbilical plate in the lower left of the view. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Continued development and correlation of analytically based weight estimation codes for wings and fuselages

NASA Technical Reports Server (NTRS)

Mullen, J., Jr.

1978-01-01

The implementation of the changes to the program for Wing Aeroelastic Design and the development of a program to estimate aircraft fuselage weights are described. The equations to implement the modified planform description, the stiffened panel skin representation, the trim loads calculation, and the flutter constraint approximation are presented. A comparison of the wing model with the actual F-5A weight material distributions and loads is given. The equations and program techniques used for the estimation of aircraft fuselage weights are described. These equations were incorporated as a computer code. The weight predictions of this program are compared with data from the C-141.

Interaction Metrics for Feedback Control of Sound Radiation from Stiffened Panels

NASA Technical Reports Server (NTRS)

Cabell, Randolph H.; Cox, David E.; Gibbs, Gary P.

2003-01-01

Interaction metrics developed for the process control industry are used to evaluate decentralized control of sound radiation from bays on an aircraft fuselage. The metrics are applied to experimentally measured frequency response data from a model of an aircraft fuselage. The purpose is to understand how coupling between multiple bays of the fuselage can destabilize or limit the performance of a decentralized active noise control system. The metrics quantitatively verify observations from a previous experiment, in which decentralized controllers performed worse than centralized controllers. The metrics do not appear to be useful for explaining control spillover which was observed in a previous experiment.

Novel Composites for Wing and Fuselage Applications. Task 1; Novel Wing Design Concepts

NASA Technical Reports Server (NTRS)

Suarez, J. A.; Buttitta, C.; Flanagan, G.; DeSilva, T.; Egensteiner, W.; Bruno, J.; Mahon, J.; Rutkowski, C.; Collins, R.; Fidnarick, R.;

Numerical simulation of unmanned aerial vehicle under centrifugal load and optimization of milling and planing

NASA Astrophysics Data System (ADS)

Chen, Yunsheng; Lu, Xinghua

2018-05-01

The mechanical parts of the fuselage surface of the UAV are easily fractured by the action of the centrifugal load. In order to improve the compressive strength of UAV and guide the milling and planing of mechanical parts, a numerical simulation method of UAV fuselage compression under centrifugal load based on discrete element analysis method is proposed. The three-dimensional discrete element method is used to establish the splitting tensile force analysis model of the UAV fuselage under centrifugal loading. The micro-contact connection parameters of the UAV fuselage are calculated, and the yield tensile model of the mechanical components is established. The dynamic and static mechanical model of the aircraft fuselage milling is analyzed by the axial amplitude vibration frequency combined method. The correlation parameters of the cutting depth on the tool wear are obtained. The centrifugal load stress spectrum of the surface of the UAV is calculated. The meshing and finite element simulation of the rotor blade of the unmanned aerial vehicle is carried out to optimize the milling process. The test results show that the accuracy of the anti - compression numerical test of the UAV is higher by adopting the method, and the anti - fatigue damage capability of the unmanned aerial vehicle body is improved through the milling and processing optimization, and the mechanical strength of the unmanned aerial vehicle can be effectively improved.

Mechanisms of Active Aerodynamic Load Reduction on a Rotorcraft Fuselage With Rotor Effects

NASA Technical Reports Server (NTRS)

Schaeffler, Norman W.; Allan, Brian G.; Jenkins, Luther N.; Yao, Chung-Sheng; Bartram, Scott M.; Mace, W. Derry; Wong, Oliver D.; Tanner, Philip E.

2016-01-01

The reduction of the aerodynamic load that acts on a generic rotorcraft fuselage by the application of active flow control was investigated in a wind tunnel test conducted on an approximately 1/3-scale powered rotorcraft model simulating forward flight. The aerodynamic mechanisms that make these reductions, in both the drag and the download, possible were examined in detail through the use of the measured surface pressure distribution on the fuselage, velocity field measurements made in the wake directly behind the ramp of the fuselage and computational simulations. The fuselage tested was the ROBIN-mod7, which was equipped with a series of eight slots located on the ramp section through which flow control excitation was introduced. These slots were arranged in a U-shaped pattern located slightly downstream of the baseline separation line and parallel to it. The flow control excitation took the form of either synthetic jets, also known as zero-net-mass-flux blowing, and steady blowing. The same set of slots were used for both types of excitation. The differences between the two excitation types and between flow control excitation from different combinations of slots were examined. The flow control is shown to alter the size of the wake and its trajectory relative to the ramp and the tailboom and it is these changes to the wake that result in a reduction in the aerodynamic load.

Performance of Optimized Actuator and Sensor Arrays in an Active Noise Control System

NASA Technical Reports Server (NTRS)

Palumbo, D. L.; Padula, S. L.; Lyle, K. H.; Cline, J. H.; Cabell, R. H.

1996-01-01

Experiments have been conducted in NASA Langley's Acoustics and Dynamics Laboratory to determine the effectiveness of optimized actuator/sensor architectures and controller algorithms for active control of harmonic interior noise. Tests were conducted in a large scale fuselage model - a composite cylinder which simulates a commuter class aircraft fuselage with three sections of trim panel and a floor. Using an optimization technique based on the component transfer functions, combinations of 4 out of 8 piezoceramic actuators and 8 out of 462 microphone locations were evaluated against predicted performance. A combinatorial optimization technique called tabu search was employed to select the optimum transducer arrays. Three test frequencies represent the cases of a strong acoustic and strong structural response, a weak acoustic and strong structural response and a strong acoustic and weak structural response. Noise reduction was obtained using a Time Averaged/Gradient Descent (TAGD) controller. Results indicate that the optimization technique successfully predicted best and worst case performance. An enhancement of the TAGD control algorithm was also evaluated. The principal components of the actuator/sensor transfer functions were used in the PC-TAGD controller. The principal components are shown to be independent of each other while providing control as effective as the standard TAGD.

Caul and method for bonding and curing intricate composite structures

NASA Technical Reports Server (NTRS)

Willden, Kurtis S. (Inventor); Goodno, Kenneth N. (Inventor)

1993-01-01

The invention disclosed here is a method for forming and curing an intricate structure of criss-crossing composite stringers and frames that are bonded to a skin panel. A structure constructed in accordance with the invention would be well-suited for use as a portion of an aircraft fuselage, a boat hull, or the like. The method is preferably practiced by applying uncured composite stringers to an uncured composite sheet panel. This is followed by placing cured frames crosswise over the stringers. The frames have openings at the locations where they intersect with the stringers which enables the frames to come into direct contact with the skin along most of their length. During the forming and curing process, the stringers are covered with a plurality of cauls, and the entire assembly of skin panel, stringers, frames and cauls is subjected to a vacuum bagging and curing process. The cauls serve to maintain both part shape and to control the flow of resin within the stringers as they are cured. Further, they probably eliminate the need for intermediate protective materials between the vacuum bag and the stringers.

Load transfer in the stiffener-to-skin joints of a pressurized fuselage

NASA Technical Reports Server (NTRS)

Johnson, Eric R.; Rastogi, Naveen

1995-01-01

Structural analyses are developed to determine the linear elastic and the geometrically nonlinear elastic response of an internally pressurized, orthogonally stiffened, composite material cylindrical shell. The configuration is a long circular cylindrical shell stiffened on the inside by a regular arrangement of identical stringers and identical rings. Periodicity permits the analysis of a unit cell model consisting of a portion of the shell wall centered over one stringer-ring joint. The stringer-ring-shell joint is modeled in an idealized manner; the stiffeners are mathematically permitted to pass through one another without contact, but do interact indirectly through their mutual contact with the shell at the joint. Discrete beams models of the stiffeners include a stringer with a symmetrical cross section and a ring with either a symmetrical or an asymmetrical open section. Mathematical formulations presented for the linear response include the effect of transverse shear deformations and the effect of warping of the ring's cross section due to torsion. These effects are important when the ring has an asymmetrical cross section because the loss of symmetry in the problem results in torsion and out-of-plane bending of the ring, and a concomitant rotation of the joint at the stiffener intersection about the circumferential axis. Data from a composite material crown panel typical of a large transport fuselage structure are used for two numerical examples. Although the inclusion of geometric nonlinearity reduces the 'pillowing' of the shell, it is found that bending is localized to a narrow region near the stiffener. Including warping deformation of the ring into the analysis changes the sense of the joint rotation. Transverse shear deformation models result in increased joint flexibility.

Novel Composites for Wing and Fuselage Applications: Speedy Nonlinear Analysis of Postbuckled Panels in Shear (SNAPPS)

NASA Technical Reports Server (NTRS)

Sharp, Dave; Sobel, Larry

1997-01-01

A simple and rapid analysis method, consisting of a number of modular, 'strength-of-materials-type' models, is presented for predicting the nonlinear response and stiffener separation of postbuckled, flat, composite, shear panels. The analysis determines the maximum principal tensile stress in the skin surface layer under to toe. Failure is said to occur when this stress reaches the mean transverse tensile strength of the layer. The analysis methodology consists of a number of closed-form equations that can easily be used in a 'hand analysis. For expediency, they have been programmed into a preliminary design code called SNAPPS (Speedy Nonlinear Analysis of Postbuckled Panels in Shear), which rapidly predicts postbuckling response of the panel for each value of the applied shear load. SNAPPS response and failure predictions were found to agree well with test results for three panels with widely different geometries, laminates and stiffnesses. Design guidelines are given for increasing the load-carrying capacity of stiffened, composite shear panels.

78 FR 72561 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-12-03

... and a skin crack in section 41 of the fuselage, and multiple reports of cracked or missing fastener heads. This AD requires repetitive inspections for cracking of the fuselage skin, discrepant fasteners... disbonding and subsequent cracking of the skin panels. DATES: This AD is effective January 7, 2014. The...

78 FR 28767 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-05-16

... a skin crack in section 41 of the fuselage, and multiple reports of cracked or missing fastener heads. This proposed AD would require repetitive inspections for cracking of the fuselage skin... airplane due to such disbonding and subsequent cracking of the skin panels. DATES: We must receive comments...

14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Fuselage, landing gear, and rotor pylon structures. 27.549 Section 27.549 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... rotor pylon structure must be designed as prescribed in this section. Resultant rotor forces may be...

14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Fuselage, landing gear, and rotor pylon structures. 27.549 Section 27.549 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... rotor pylon structure must be designed as prescribed in this section. Resultant rotor forces may be...

14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Fuselage, landing gear, and rotor pylon structures. 27.549 Section 27.549 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... rotor pylon structure must be designed as prescribed in this section. Resultant rotor forces may be...

14 CFR 29.549 - Fuselage and rotor pylon structures.

Code of Federal Regulations, 2011 CFR

2011-01-01

... flight conditions, must be considered. (c) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (d) [Reserved] (e) If approval for the use of 21/2-minute OEI power is requested, each engine...

14 CFR 29.549 - Fuselage and rotor pylon structures.

Code of Federal Regulations, 2013 CFR

2013-01-01

... flight conditions, must be considered. (c) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (d) [Reserved] (e) If approval for the use of 21/2-minute OEI power is requested, each engine...

14 CFR 29.549 - Fuselage and rotor pylon structures.

Code of Federal Regulations, 2010 CFR

2010-01-01

... flight conditions, must be considered. (c) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (d) [Reserved] (e) If approval for the use of 21/2-minute OEI power is requested, each engine...

14 CFR 29.549 - Fuselage and rotor pylon structures.

Code of Federal Regulations, 2012 CFR

2012-01-01

... flight conditions, must be considered. (c) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (d) [Reserved] (e) If approval for the use of 21/2-minute OEI power is requested, each engine...

14 CFR 29.549 - Fuselage and rotor pylon structures.

Code of Federal Regulations, 2014 CFR

2014-01-01

... flight conditions, must be considered. (c) Each engine mount and adjacent fuselage structure must be designed to withstand the loads occurring under accelerated flight and landing conditions, including engine torque. (d) [Reserved] (e) If approval for the use of 21/2-minute OEI power is requested, each engine...

Fracture and Fatigue Strength Evaluation of Multiple Site Damaged Aircraft Fuselages - Curved Panel Testing and Analysis.

DOT National Transportation Integrated Search

1994-01-01

As the fleet of commercial passenger aircraft is utilized beyond their design lives, flaws have developed at the rivet holes of fuselage lap joints on some aircraft, which can grow and link up into cracks of significant length. Understanding, predict...

77 FR 29857 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-21

... of cabin pressurization and the inability of the fuselage to withstand fail-safe loads. DATES: This... cabin pressurization and the inability of the fuselage to withstand fail-safe loads. (f) Compliance You... reference of certain publications listed in the AD as of June 25, 2012. ADDRESSES: For service information...

77 FR 63270 - Airworthiness Directives; Airbus Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-10-16

... missing fastener between certain stringers of the fuselage frame which connects the frame to a tee. This proposed AD would require a rototest inspection and modification or repair of the fuselage frame at the... Portal: Go to http://www.regulations.gov . Follow the instructions for submitting comments. Fax: (202...

77 FR 6522 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-02-08

...Rulemaking Portal: Go to http://www.regulations.gov . Follow the instructions for submitting comments. Fax... hold the door closed for the design loads. Fractured and missing latch pin retention bolts, if not... cargo door, fuselage frames, internal and external skin of the fuselage, cargo door frames, mid-span...

78 FR 65166 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-10-31

... the upper deck tension ties of the fuselage are subject to widespread fatigue damage. We are issuing this AD to prevent widespread fatigue damage of certain fuselage upper deck tension ties, which could result in reduced structural integrity of the airplane. DATES: This AD is effective December 5, 2013. The...

78 FR 43839 - Airworthiness Directives; the Boeing Company

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-22

... fuselage are subject to widespread fatigue damage. This proposed AD would retain the repetitive inspections... inspections. We are proposing this AD to prevent widespread fatigue damage of certain fuselage upper deck... comments on this proposed AD by September 5, 2013. ADDRESSES: You may send comments, using the procedures...

14 CFR 23.1201 - Fire extinguishing systems materials.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Fire extinguishing systems materials. 23... Powerplant Powerplant Fire Protection § 23.1201 Fire extinguishing systems materials. For all airplanes with engine(s) embedded in the fuselage or in pylons on the aft fuselage the following applies: (a) No...

14 CFR 23.1201 - Fire extinguishing systems materials.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Fire extinguishing systems materials. 23... Powerplant Powerplant Fire Protection § 23.1201 Fire extinguishing systems materials. For all airplanes with engine(s) embedded in the fuselage or in pylons on the aft fuselage the following applies: (a) No...

77 FR 5724 - Airworthiness Directives; Fokker Services B.V. Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-02-06

... proposed AD would require repetitive low frequency eddy current inspections of the forward fuselage butt... repetitive [low frequency eddy current] inspections of the forward fuselage butt joints for cracks and, when... effective date of this AD, whichever occurs later, do a low frequency eddy current inspection of the forward...

Recent advances in active noise and vibration control at NASA Langley Research Center

NASA Astrophysics Data System (ADS)

Gibbs, Gary P.; Cabell, Randolph H.; Palumbo, Daniel L.; Silcox, Richard J.; Turner, Travis L.

2002-11-01

Over the past 15 years NASA has investigated the use of active control technology for aircraft interior noise. More recently this work has been supported through the Advanced Subsonic Technology Noise Reduction Program (1994-2001), High Speed Research Program (1994-1999), and through the Quiet Aircraft Technology Program (2000-present). The interior environment is recognized as an important element in flight safety, crew communications and fatigue, as well as passenger comfort. This presentation will overview research in active noise and vibration control relating to interior noise being investigated by NASA. The research to be presented includes: active control of aircraft fuselage sidewall transmission due to turbulent boundary layer or jet noise excitation, active control of interior tones due to propeller excitation of aircraft structures, and adaptive stiffening of structures for noise, vibration, and fatigue control. Work on actuator technology ranging from piezoelectrics, shape memory actuators, and fluidic actuators will be described including applications. Control system technology will be included that is experimentally based, real-time, and adaptive.

Crashworthy airframe design concepts: Fabrication and testing

NASA Technical Reports Server (NTRS)

Cronkhite, J. D.; Berry, V. L.

1982-01-01

Crashworthy floor concepts applicable to general aviation aircraft metal airframe structures were investigated. Initially several energy absorbing lower fuselage structure concepts were evaluated. Full scale floor sections representative of a twin engine, general aviation airplane lower fuselage structure were designed and fabricated. The floors featured an upper high strength platform with an energy absorbing, crushable structure underneath. Eighteen floors were fabricated that incorporated five different crushable subfloor concepts. The floors were then evaluated through static and dynamic testing. Computer programs NASTRAN and KRASH were used for the static and dynamic analysis of the floor section designs. Two twin engine airplane fuselages were modified to incorporate the most promising crashworthy floor sections for test evaluation.

Turbulent flow around a wing-fuselage type juncture

NASA Technical Reports Server (NTRS)

Kubendran, L. R.; Mcmahon, H. M.; Hubbartt, J. E.

1985-01-01

The flow over a 58-mm-thick uniform-thickness winglike body having a 1.5:1 elliptical leading edge and joined to a large flat plate (representing an aircraft fuselage) is characterized experimentally at freestream velocity 15 m/s, corresponding to Reynolds number 940,000/m, using hot-wire anemometry. The results are presented graphically, and it is found that the horseshoe vortex formed by the separation of the fuselage boundary layer ahead of the wing leading edge is effective in transporting turbulence and modifying the mean-flow characteristics and the turbulent-stress distribution. It is suggested that the slenderness ratio of the leading edge is the dominant factor affecting the strength and location of the vortex.

A NASTRAN model of a large flexible swing-wing bomber. Volume 4: NASTRAN model development-fuselage structure

NASA Technical Reports Server (NTRS)

Mock, W. D.; Latham, R. A.

1982-01-01

The NASTRAN model plan for the fuselage structure was expanded in detail to generate the NASTRAN model for this substructure. The grid point coordinates were coded for each element. The material properties and sizing data for each element were specified. The fuselage substructure model was thoroughly checked out for continuity, connectivity, and constraints. This substructure was processed for structural influence coefficients (SIC) point loadings and the deflections were compared to those computed for the aircraft detail model. Finally, a demonstration and validation processing of this substructure was accomplished using the NASTRAN finite element program. The bulk data deck, stiffness matrices, and SIC output data were delivered.

Skid Landings of Airplanes on Rocker-Type Fuselages

NASA Technical Reports Server (NTRS)

Mayo, Wilbur L.

1961-01-01

A study is made of the landing of an airplane on a fuselage with "planned" curvature of its lower surface. Initial contact is considered to stop the vertical motion of a point remote from the center of gravity, thus causing rocking on the curved lower surface which converts sinking-speed energy into angular energy in pitch for dissipation by damping forces. Analysis is made of loads and motions for a given fuselage shape, and the contours required to give desired load histories are determined. Most of the calculations involve initial contact at the tail, but there are two cases of unflared landings with initial contact at the nose. The calculations are checked experimentally for the tail - low case.

STAGS Developments for Residual Strength Analysis Methods for Metallic Fuselage Structures

NASA Technical Reports Server (NTRS)

Young, Richard D.; Rose, Cheryl A.

2014-01-01

A summary of advances in the Structural Analysis of General Shells (STAGS) finite element code for the residual strength analysis of metallic fuselage structures, that were realized through collaboration between the structures group at NASA Langley, and Dr. Charles Rankin is presented. The majority of the advancements described were made in the 1990's under the NASA Airframe Structural Integrity Program (NASIP). Example results from studies that were conducted using the STAGS code to develop improved understanding of the nonlinear response of cracked fuselage structures subjected to combined loads are presented. An integrated residual strength analysis methodology for metallic structure that models crack growth to predict the effect of cracks on structural integrity is demonstrated

Propeller aircraft interior noise model

NASA Technical Reports Server (NTRS)

Pope, L. D.; Wilby, E. G.; Wilby, J. F.

1984-01-01

An analytical model was developed to predict the interior noise of propeller-driven aircraft. The fuselage model is that of a cylinder with a structurally-integral floor. The cabin sidewall is stiffened by stringers and ring frames, and the floor by longitudinal beams. The cabin interior is covered with a sidewall treatments consisting of layers of porous material and an impervious trim septum. Representation of the propeller pressure field is utilized as input data in the form of the propeller noise signature at a series of locations on a grid over the fuselage structure. Results obtained from the analytical model are compared with test data measured by NASA in a scale model cylindrical fuselage excited by a model propeller.

LEFT WING AND FUSELAGE FROM THIRD LEVEL OF TAIL DOCK ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

LEFT WING AND FUSELAGE FROM THIRD LEVEL OF TAIL DOCK STAND. THE WING IS PREPARED FOR BASIC LUBRICATION WITH E SPOILER BOARDS UP AND ALL SAFETY LOCKS IN PLACE TO PROTECT MECHANICS FROM INJURY. ON THE WING AN INSPECTOR CHECKS THE ACTUATORS. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

75 FR 32251 - Airworthiness Directives; PILATUS Aircraft Ltd. Model PC-7 Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-06-08

... contains regulatory documents #0;having general applicability and legal effect, most of which are keyed #0... right lower longerons between fuselage frames 1 and 1A. The possibility of corrosion is increased... corrosion at the bonding strap connections on the left and right lower longerons between fuselage frames 1...

75 FR 12150 - Airworthiness Directives; PILATUS Aircraft Ltd. Model PC-7 Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-15

... the left and right lower longerons between fuselage frames 1 and 1A. The possibility of corrosion is... methods: Federal eRulemaking Portal: Go to http://www.regulations.gov . Follow the instructions for... at the bonding strap connections on the left and right lower longerons between fuselage frames 1 and...

78 FR 14719 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-03-07

... that the fuselage skin just above certain lap splice locations is subject to widespread fatigue damage... fatigue cracking of the fuselage skin, which could result in reduced structural integrity of the airplane.... Hand Delivery: Deliver to Mail address above between 9 a.m. and 5 p.m., Monday through Friday, except...

76 FR 72853 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2011-11-28

... detect and correct fatigue cracking of the fuselage skin panels at the chem-mill steps, which could... correct fatigue cracking of the fuselage skin panels at the chem-mill steps, which could result in sudden... STA 400 and STA 460 on certain airplanes. The cause of the cracking was fatigue due to high- tension...

Propeller aircraft interior noise model: User's manual for computer program

NASA Technical Reports Server (NTRS)

Wilby, E. G.; Pope, L. D.

1985-01-01

A computer program entitled PAIN (Propeller Aircraft Interior Noise) has been developed to permit calculation of the sound levels in the cabin of a propeller-driven airplane. The fuselage is modeled as a cylinder with a structurally integral floor, the cabin sidewall and floor being stiffened by ring frames, stringers and floor beams of arbitrary configurations. The cabin interior is covered with acoustic treatment and trim. The propeller noise consists of a series of tones at harmonics of the blade passage frequency. Input data required by the program include the mechanical and acoustical properties of the fuselage structure and sidewall trim. Also, the precise propeller noise signature must be defined on a grid that lies in the fuselage skin. The propeller data are generated with a propeller noise prediction program such as the NASA Langley ANOPP program. The program PAIN permits the calculation of the space-average interior sound levels for the first ten harmonics of a propeller rotating alongside the fuselage. User instructions for PAIN are given in the report. Development of the analytical model is presented in NASA CR 3813.

Two-dimensional aerodynamic characteristics of several polygon-shaped cross-sectional models applicable to helicopter fuselages

NASA Technical Reports Server (NTRS)

Kelley, Henry L.; Crowell, Cynthia A.; Wilson, John C.

1992-01-01

A wind-tunnel investigation was conducted to determine 2-D aerodynamic characteristics of nine polygon-shaped models applicable to helicopter fuselages. The models varied from 1/2 to 1/5 scale and were nominally triangular, diamond, and rectangular in shape. Side force and normal force were obtained at increments of angle of flow incidence from -45 to 90 degrees. The data were compared with results from a baseline UH-60 tail-boom cross-section model. The results indicate that the overall shapes of the plots of normal force and side force were similar to the characteristic shape of the baseline data; however, there were important differences in magnitude. At a flow incidence of 0 degrees, larger values of normal force for the polygon models indicate an increase in fuselage down load of 1 to 2.5 percent of main-rotor thrust compared with the baseline value. Also, potential was indicated among some of the configurations to produce high fuselage side forces and yawing moments compared with the baseline model.

Wing planform effects at supersonic speeds for an advanced fighter configuration

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1984-01-01

Four advanced fighter configurations, which differed in wing planform and airfoil shape, were investigated in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Supersonic data were obtained on the four uncambered wings, which were each attached to a single fighter fuselage. The fuselage geometry varied in cross-sectional shape and had two side-mounted, flow-through, half-axisymmetric inlets. Twin vertical tails were attached to the fuselage. The four planforms tested were a 65 deg delta wing, a combination of a 20 deg trapezoidal wing and a 45 deg horizontal tail, a 70 deg/30 deg cranked wing, and a 70 deg/66 deg crank wing, where the angle values refer to the leading-edge sweep angle of the lifting-surface planform. Planform effects on a single fuselage representative of an advanced fighter aircraft were studied. Results show that the highly swept cranked wings exceeded the aerodynamic performance levels, at low lift coefficients, of the 65 deg delta wing and the 20 deg trapezoidal wing at trimmed and untrimmed conditions.

Test results from large wing and fuselage panels

NASA Technical Reports Server (NTRS)

Madan, Ram C.; Voldman, Mike

1993-01-01

This paper presents the first results in an assessment of the strength, stiffness, and damage tolerance of stiffened wing and fuselage subcomponents. Under this NASA funded program, 10 large wing and fuselage panels, variously fabricated by automated tow placement and dry-stitched preform/resin transfer molding, are to be tested. The first test of an automated tow placement six-longeron fuselage panel under shear load was completed successfully. Using NASTRAN finite-element analysis the stiffness of the panel in the linear range prior to buckling was predicted within 3.5 percent. A nonlinear analysis predicted the buckling load within 10 percent and final failure load within 6 percent. The first test of a resin transfer molding six-stringer wing panel under compression was also completed. The panel failed unexpectedly in buckling because of inadequate supporting structure. The average strain was 0.43 percent with a line load of 20.3 kips per inch of width. This strain still exceeds the design allowable strains. Also, the stringers did not debond before failure, which is in contrast to the general behavior of unstitched panels.

Experimental measurement of structural power flow on an aircraft fuselage

NASA Technical Reports Server (NTRS)

Cuschieri, J. M.

1989-01-01

An experimental technique is used to measure the structural power flow through an aircraft fuselage with the excitation near the wing attachment location. Because of the large number of measurements required to analyze the whole of an aircraft fuselage, it is necessary that a balance be achieved between the number of measurement transducers, the mounting of these transducers, and the accuracy of the measurements. Using four transducers mounted on a bakelite platform, the structural intensity vectors at locations distributed throughout the fuselage are measured. To minimize the errors associated with using a four transducers technique the measurement positions are selected away from bulkheads and stiffeners. Because four separate transducers are used, with each transducer having its own drive and conditioning amplifiers, phase errors are introduced in the measurements that can be much greater than the phase differences associated with the measurements. To minimize these phase errors two sets of measurements are taken for each position with the orientation of the transducers rotated by 180 deg and an average taken between the two sets of measurements. Results are presented and discussed.

AFTI/F-16 in banked flight

NASA Technical Reports Server (NTRS)

1989-01-01

This photo depicts the AFTI F-16 in the configuration used midway through the program. The sensor pods were added to the fuselage, but the chin canards remained in place. Painted in non-standard gray tones, it carried Sidewinder air-to-air missles on its wingtips. During the 1980s and 1990s, NASA and the U.S. Air Force participated in a joint program to integrate and demonstrate new avionics technologies to improve close air support capabilities in next-generation aircraft. The testbed aircraft, seen here in flight over the desert at NASA's Dryden Flight Research Center, Edwards, California, was called the Advanced Fighter Technology Integration (AFTI) F-16. The tests demonstrated technologies to improve navigation and the pilot's ability to find and destroy enemy ground targets day or night, including adverse weather. The aircraft--an F-16A Fighting Falcon (Serial #75-0750)--underwent numerous modifications. A relatively low-cost testbed, it evaluated the feasability of advanced, intergrated-sensor, avionics, and flight control technologies. During the first phase of the AFTI/F-16 program, which began in 1983, the aircraft demonstrated voice-actuated commands, helmet-mounted sights, flat turns, and selective fuselage pointing using forward-mounted canards and a triplex digital flight control computer system. The second phase of research, which began in the summer of 1991, demonstrated advanced technologies and capabilities to find and destroy ground targets day or night, and in adverse weather while using maneuverability and speed at low altitude. This phase was known as the close air support and battlefield air interdiction (CAS/BAI) phase. Finally, the aircraft was used to assess the Automatic Ground Collision Avoidance System (Auto - GCAS), a joint project with the Swedish Government. For these tests, the pilot flew the aircraft directly toward the ground, simulating a total loss of control. The GCAS was designed to take command in such emergencies and bring the aircraft back to level flight. The AFTI F-16 program ended at Dryden on November 4, 1997 after 15 years and over 700 research flights. The USAF continued to fly the aircraft until retiring it to the Air Force Museum on January 9, 2001.

Fuselage boundary-layer refraction of fan tones radiated from an installed turbofan aero-engine.

PubMed

Gaffney, James; McAlpine, Alan; Kingan, Michael J

2017-03-01

A distributed source model to predict fan tone noise levels of an installed turbofan aero-engine is extended to include the refraction effects caused by the fuselage boundary layer. The model is a simple representation of an installed turbofan, where fan tones are represented in terms of spinning modes radiated from a semi-infinite circular duct, and the aircraft's fuselage is represented by an infinitely long, rigid cylinder. The distributed source is a disk, formed by integrating infinitesimal volume sources located on the intake duct termination. The cylinder is located adjacent to the disk. There is uniform axial flow, aligned with the axis of the cylinder, everywhere except close to the cylinder where there is a constant thickness boundary layer. The aim is to predict the near-field acoustic pressure, and in particular, to predict the pressure on the cylindrical fuselage which is relevant to assess cabin noise. Thus no far-field approximations are included in the modelling. The effect of the boundary layer is quantified by calculating the area-averaged mean square pressure over the cylinder's surface with and without the boundary layer included in the prediction model. The sound propagation through the boundary layer is calculated by solving the Pridmore-Brown equation. Results from the theoretical method show that the boundary layer has a significant effect on the predicted sound pressure levels on the cylindrical fuselage, owing to sound radiation of fan tones from an installed turbofan aero-engine.

Investigation of an underslung half-cone inlet with compression surface mounted outboard from fuselage at Mach numbers of 1.5, 1.8, and 2.0

NASA Technical Reports Server (NTRS)

Yeager, Richard A; Gertsma, Laurence W

1958-01-01

An investigation was conducted to determine the performance of an underslung half-cone inlet mounted on a missile forebody model with the compression surface outboard from the fuselage. The inlet was designed for shock-on-lip operation at Mach number 2.0 with 25 degree half-angle spike. The cowling was attached to the fuselage through the boundary-layer plow and served as part of the fuselage boundary-layer diverter system. The performance of the half-cone inlet was compared with that of a scoop-type inlet and a normal-wedge inlet on a maximum-thrust-minus-drag basis. The increase in pressure recovery obtained with the half-cone inlet over that obtained with the reference inlets offset the slightly higher drags observed over the Mach number range for the half-cone so that the performance of this configuration was equal to that of the other inlets at Mach number 2.0 and was slightly superior at the lower Mach numbers. For a particular configuration, a peak pressure recovery of 0.879 was obtained at Mach number 2.0, zero angle of attack, and 4-percent throat bleed; the subcritical stability was 16 percent. Use of a fuselage-mounted boundary-layer splitter plate ahead of the inlet did not improve the stability. Subcritical distortion values were below 10 percent for all configurations. (author)

Filament winding - Waking the sleeping giant

NASA Technical Reports Server (NTRS)

Freeman, W. T., Jr.; Stein, B. A.

1985-01-01

The use of filament winding (FW) in the production of aerospace composite structures is examined. The FW process applies spools of fiber and prepreg tow or prepreg tape to a male mandrel; the process is more efficient and cost effective than metallic construction. The fibers used in FW and the curing process are explained. The reduced storage and fabrication costs that result from FW are discussed. The use of FW to produce a filament-wound case for a solid rocket motor and the substructure and skin of an aircraft fuselage are described. Areas which require further development in order to expand the use of FW are listed and discussed.

Helicopter rotor dynamics and aeroelasticity - Some key ideas and insights

NASA Technical Reports Server (NTRS)

Friedmann, Peretz P.

1990-01-01

Four important current topics in helicopter rotor dynamics and aeroelasticity are discussed: (1) the role of geometric nonlinearities in rotary-wing aeroelasticity; (2) structural modeling, free vibration, and aeroelastic analysis of composite rotor blades; (3) modeling of coupled rotor/fuselage areomechanical problems and their active control; and (4) use of higher-harmonic control for vibration reduction in helicopter rotors in forward flight. The discussion attempts to provide an improved fundamental understanding of the current state of the art. In this way, future research can be focused on problems which remain to be solved instead of producing marginal improvements on problems which are already understood.

Cabin-fuselage-wing structural design concept with engine installation

NASA Technical Reports Server (NTRS)

Ariotti, Scott; Garner, M.; Cepeda, A.; Vieira, J.; Bolton, D.

1993-01-01

The purpose of this project is to provide a fuselage structural assembly and wing structural design that will be able to withstand the given operational parameters and loads provided by Federal Aviation Regulation Part 23 (FAR 23) and the Statement of Work (SOW). The goal is to provide a durable lightweight structure that will transfer the applied loads through the most efficient load path. Areas of producibility and maintainability of the structure will also be addressed. All of the structural members will also meet or exceed the desired loading criteria, along with providing adequate stiffness, reliability, and fatigue life as stated in the SOW. Considerations need to be made for control system routing and cabin heating/ventilation. The goal of the wing structure and carry through structure is also to provide a simple, lightweight structure that will transfer the aerodynamic forces produced by the wing, tailboom, and landing gear. These forces will be channeled through various internal structures sized for the pre-determined loading criteria. Other considerations were to include space for flaps, ailerons, fuel tanks, and electrical and control system routing. The difficulties encountered in the fuselage design include expanding the fuselage cabin to accept a third occupant in a staggered configuration and providing ample volume for their safety. By adding a third person the CG of aircraft will move forward so the engine needs to be moved aft to compensate for the difference in the moment. This required the provisions of a ring frame structure for the new position of the engine mount. The difficulties encountered in the wing structural design include resizing the wing for the increased capacity and weight, and compensating for a large torsion produced by the tail boom by placing a great number of stiffeners inside the boom, which will result in the relocation of the fuel tank. Finally, an adequate carry through structure for the wing and fuselage interface will be designed to effectively transmit loads through the fuselage.

Semi-active control of helicopter vibration using controllable stiffness and damping devices

NASA Astrophysics Data System (ADS)

Anusonti-Inthra, Phuriwat

Semi-active concepts for helicopter vibration reduction are developed and evaluated in this dissertation. Semi-active devices, controllable stiffness devices or controllable orifice dampers, are introduced; (i) in the blade root region (rotor-based concept) and (ii) between the rotor and the fuselage as semi-active isolators (in the non-rotating frame). Corresponding semi-active controllers for helicopter vibration reduction are also developed. The effectiveness of the rotor-based semi-active vibration reduction concept (using stiffness and damping variation) is demonstrated for a 4-bladed hingeless rotor helicopter in moderate- to high-speed forward flight. A sensitivity study shows that the stiffness variation of root element can reduce hub vibrations when proper amplitude and phase are used. Furthermore, the optimal semi-active control scheme can determine the combination of stiffness variations that produce significant vibration reduction in all components of vibratory hub loads simultaneously. It is demonstrated that desired cyclic variations in properties of the blade root region can be practically achieved using discrete controllable stiffness devices and controllable dampers, especially in the flap and lag directions. These discrete controllable devices can produce 35--50% reduction in a composite vibration index representing all components of vibratory hub loads. No detrimental increases are observed in the lower harmonics of blade loads and blade response (which contribute to the dynamic stresses) and controllable device internal loads, when the optimal stiffness and damping variations are introduced. The effectiveness of optimal stiffness and damping variations in reducing hub vibration is retained over a range of cruise speeds and for variations in fundamental rotor properties. The effectiveness of the semi-active isolator is demonstrated for a simplified single degree of freedom system representing the semi-active isolation system. The rotor, represented by a lumped mass under harmonic force excitation, is supported by a spring and a parallel damper on the fuselage (assumed to have infinite mass). Properties of the spring or damper can then be controlled to reduce transmission of the force into the fuselage or the support structure. This semi-active isolation concept can produce additional 30% vibration reduction beyond the level achieved by a passive isolator. Different control schemes (i.e. open-loop, closed-loop, and closed-loop adaptive schemes) are developed and evaluated to control transmission of vibratory loads to the support structure (fuselage), and it is seen that a closed-loop adaptive controller is required to retain vibration reduction effectiveness when there is a change in operating condition. (Abstract shortened by UMI.)

POSTOP: Postbuckled open-stiffener optimum panels, user's manual

NASA Technical Reports Server (NTRS)

Biggers, S. B.; Dickson, J. N.

1984-01-01

The computer program POSTOP developed to serve as an aid in the analysis and sizing of stiffened composite panels that may be loaded in the postbuckling regime, is intended for the preliminary design of metal or composite panels with open-section stiffeners, subjected to multiple combined biaxial compression (or tension), shear and normal pressure load cases. Longitudinal compression, however, is assumed to be the dominant loading. Temperature, initial bow eccentricity and load eccentricity effects are included. The panel geometry is assumed to be repetitive over several bays in the longitudinal (stiffener) direction as well as in the transverse direction. Analytical routines are included to compute panel stiffnesses, strains, local and panel buckling loads, and skin/stiffener interface stresses. The resulting program is applicable to stiffened panels as commonly used in fuselage, wing, or empennage structures. The capabilities and limitations of the code are described. Instructions required to use the program and several example problems are included.

Soft hub for bearingless rotors

NASA Technical Reports Server (NTRS)

Dixon, Peter G. C.

1991-01-01

Soft hub concepts which allow the direct replacement of articulated rotor systems by bearingless types without any change in controllability or need for reinforcement to the drive shaft and/or transmission/fuselage attachments of the helicopter were studied. Two concepts were analyzed and confirmed for functional and structural feasibility against a design criteria and specifications established for this effort. Both systems are gimballed about a thrust carrying universal elastomeric bearing. One concept includes a set of composite flexures for drive torque transmittal from the shaft to the rotor, and another set (which is changeable) to impart hub tilting stiffness to the rotor system as required to meet the helicopter application. The second concept uses a composite bellows flexure to drive the rotor and to augment the hub stiffness provided by the elastomeric bearing. Each concept was assessed for weight, drag, ROM cost, and number of parts and compared with the production BO-105 hub.

78 FR 70214 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-11-25

... cracks at the shear tie end fastener locations of the fuselage frames, and repair of cracks if necessary... fastener locations of the fuselage frames, and repair of cracks if necessary. The NPRM also proposed to... the frame shear ties, as stated in the NPRM (78 FR 33012, June 3, 2013). We agree with the request and...

Effects of propeller rotation direction on airplane interior noise levels

NASA Technical Reports Server (NTRS)

Willis, C. M.; Mayes, W. H.; Daniels, E. F.

1985-01-01

Interior noise measurements for upsweeping and downsweeping movement of the propeller blade tips past the fuselage were made on a twin-engine airplane and on two simplified fuselage models. Changes in interior noise levels of as much as 8 dB reversal of propeller rotation direction were measured for some configurations and test conditions.

77 FR 20572 - Airworthiness Directives; BAE Systems (Operations) Limited Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-04-05

... at the water trap/air drier unit of the forward discharge valve due to corrosion. This proposed AD... failure of the fuselage skin, leading to a possible sudden loss of cabin pressure. DATES: We must receive... surface anomalies (bulges and/or dents) of the fuselage skin at the water trap/air drier unit of the...

75 FR 48623 - Airworthiness Directives; The Boeing Company Model 767-300 Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-08-11

... fuselage skin and backup structure at the lower very high frequency (VHF) antenna cutout at station 1197... section 46 fuselage lower skin around the periphery of the VHF antenna baseplate at station 1197 + 99. We... antenna baseplate at station 1197 + 99. One operator reported 5 cracks, with a maximum length of 11 inches...

NASTRAN DMAP Fuzzy Structures Analysis: Summary of Research

NASA Technical Reports Server (NTRS)

Sparrow, Victor W.

2001-01-01

The main proposed tasks of Cooperative Agreement NCC1-382 were: (1) developing MSC/NASTRAN DMAP language scripts to implement the Soize fuzzy structures approach for modeling the dynamics of complex structures; (2) benchmarking the results of the new code to those for a cantilevered beam in the literature; and (3) testing and validating the new code by comparing the fuzzy structures results to NASA Langley experimental and conventional finite element results for two model test structures representative of aircraft fuselage sidewall construction: (A) a small aluminum test panel (SLP, single longeron panel) with a single longitudinal stringer attached with bolts; and (B) a 47 by 72 inch flat aluminum fuselage panel (AFP, aluminum fuselage panel) including six longitudinal stringers and four frame stiffeners attached with rivets.

Design of a 4-seat, general aviation, electric aircraft

NASA Astrophysics Data System (ADS)

Rajagopalan, Arvindhakshan

Range and payload of current electric aircraft is limited primarily due to low energy density of batteries. However, recent advances in battery technology promise storage of more than 1 kWh of energy per kilogram of weight in the near future. This kind of energy storage makes possible the design of an electric aircraft comparable to, if not better than existing state-of-the art general aviation aircraft powered by internal combustion engines. This thesis explores through parametric studies the effect of lift-to-drag ratio, flight speed, and cruise altitude on required thrust power and battery energy and presents the conceptual and preliminary design of a four-seat, general aviation electric aircraft with a takeoff weight of 1750 kg, a range of 800 km, and a cruise speed of 200 km/h. An innovative configuration design will take full advantage of the electric propulsion system, while a Lithium-Polymer battery and a DC brush less motor will provide the power. Advanced aerodynamics will explore the greatest possible extend of laminar flow on the fuselage, the wing, and the empennage surfaces to minimize drag, while advanced composite structures will provide the greatest possible savings on empty weight. The proposed design is intended to be certifiable under current FAR 23 requirements.

Evaluation of Environmental Profiles for Reliability Demonstration

DTIC Science & Technology

1974-11-01

Center, Attn: RBRS, Griffiss AFB, NY 13441. RADC ltr, 3 Mar 1977 ä>> 2s **« ** BEEN 0EUmTE0 ’T8 USE «■ DISCLOSURE. WSTRIBüTION SWTE)( tNT ^ *^VE0 F0R...source for this aircraft is two turbo-jet engines with afterburner , mounted in the aft fuselage, having no fuselage exhaust impingement. The vibration

75 FR 61977 - Airworthiness Directives; The Boeing Company Model 747-100, 747-200B, and 747-200F Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2010-10-07

... fatigue-related skin cracks and corrosion of the skin panel lap joints in the fuselage upper lobe, and... of corrosion, and related investigative and corrective actions. This AD reduces the maximum interval... and correct fatigue cracking and corrosion in the fuselage upper lobe skin lap joints, which could...

An oblique view of the forward fuselage and port side ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

An oblique view of the forward fuselage and port side of the Orbiter Discovery while mounted atop the 76-wheeled orbiter transfer system as it is being rolled from the Orbiter Processing Facility to the Vehicle Assembly Building at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

An oblique view of the forward fuselage and starboard side ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

An oblique view of the forward fuselage and starboard side of the Orbiter Discovery while mounted atop the 76-wheeled orbiter transfer system as it is being rolled from the Orbiter Processing Facility to the Vehicle Assembly Building at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Nonstandard and Higher-Order Finite-Difference Methods for Electromagnetics

DTIC Science & Technology

2009-10-26

Simplified Fuselage filled with 90 passengers. . . . . . . . . 135 4.4. A top view photograph of the expanded polystyrene passenger support, and the... expanded polystyrene supports. . . . . . . . . . . . . . . . . . . . . . . 140 4.10. Measured S11 (the exterior antenna) of the simplified fuselage...escape. To keep the passengers in their designated locations and upright, an expanded polystyrene support system was made. In a sheet of 1” thick

76 FR 2281 - Airworthiness Directives; BAE SYSTEMS (OPERATIONS) LIMITED Model BAe 146 Airplanes, and Model...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-01-13

... this AD, perform an external eddy current inspection of the forward fuselage skin to detect cracking... paragraphs (j)(1), (j)(2), and (j)(3) of this AD, do an external eddy current inspection of the forward... this AD, do an external eddy current inspection of the forward fuselage skin to detect cracking, in...

Propagation of propeller tone noise through a fuselage boundary layer

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Magliozzi, B.

1984-01-01

In earlier experimental and analytical studies, it was found that the boundary layer on an aircraft could provide significant shielding from propeller noise at typical transport airplane cruise Mach numbers. In this paper a new three-dimensional theory is described that treats the combined effects of refraction and scattering by the fuselage and boundary layer. The complete wave field is solved by matching analytical expressions for the incident and scattered waves in the outer flow to a numerical solution in the boundary layer flow. The model for the incident waves is a near-field frequency-domain propeller source theory developed previously for free field studies. Calculations for an advanced turboprop (Prop-Fan) model flight test at 0.8 Mach number show a much smaller than expected pressure amplification at the noise directivity peak, strong boundary layer shielding in the forward quadrant, and shadowing around the fuselage. Results are presented showing the difference between fuselage surface and free-space noise predictions as a function of frequency and Mach number. Comparison of calculated and measured effects obtained in a Prop-Fan model flight test show good agreement, particularly near and aft of the plane of rotation at high cruise Mach number.

Skylon Aerodynamics and SABRE Plumes

NASA Technical Reports Server (NTRS)

Mehta, Unmeel; Afosmis, Michael; Bowles, Jeffrey; Pandya, Shishir

2015-01-01

An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown-a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

Experimental measurement of structural power flow on an aircraft fuselage

NASA Technical Reports Server (NTRS)

Cuschieri, J. M.

1991-01-01

An experimental technique is used to measure structural intensity through an aircraft fuselage with an excitation load applied near one of the wing attachment locations. The fuselage is a relatively large structure, requiring a large number of measurement locations to analyze the whole of the structure. For the measurement of structural intensity, multiple point measurements are necessary at every location of interest. A tradeoff is therefore required between the number of measurement transducers, the mounting of these transducers, and the accuracy of the measurements. Using four transducers mounted on a bakelite platform, structural intensity vectors are measured at locations distributed throughout the fuselage. To minimize the errors associated with using the four transducer technique, the measurement locations are selected to be away from bulkheads and stiffeners. Furthermore, to eliminate phase errors between the four transducer measurements, two sets of data are collected for each position, with the orientation of the platform with the four transducers rotated by 180 degrees and an average taken between the two sets of data. The results of these measurements together with a discussion of the suitability of the approach for measuring structural intensity on a real structure are presented.

Full-Scale Wind-Tunnel Investigation of the Drag Characteristics of an HU2K Helicopter Fuselage

NASA Technical Reports Server (NTRS)

Scallion, William I.

1963-01-01

An investigation was conducted in the Langley full-scale tunnel to determine the drag characteristics of the HU2K helicopter fuselage. The effects of body shape, engine operation, appendages, and leakage on the model drag were determined. The results of the tests showed that the largest single contribution to the parasite drag was that of the rotor hub installation which produced about 80 percent of the drag of the sealed and faired production body. Fairings on the rotor hub and blade retentions, or a cleaned-up hub and retentions, appeared to be the most effective single modifications tested. The total drag of all protuberances and air leakage also contributed a major part of the drag - an 83-percent increase over the drag of the sealed and faired production body. An additional increment of drag was caused by the basic shape of the fuselage - 19 percent more than the drag obtained when the fuselage shape was extensively refaired. Another sizable increment of drag was caused by the engine oil-cooler exit which gave a drag of 8 percent of that of the sealed and faired production body.

Arrow-wing supersonic cruise aircraft structural design concepts evaluation. Volume 3: Sections 12 through 14

NASA Technical Reports Server (NTRS)

Sakata, I. F.; Davis, G. W.

1975-01-01

The design of an economically viable supersonic cruise aircraft requires the lowest attainable structural-mass fraction commensurate with the selected near-term structural material technology. To achieve this goal of minimum structural-mass fraction, various combinations of promising wing and fuselage primary structure were analyzed for the load-temperature environment applicable to the arrow wing configuration. This analysis was conducted in accordance with the design criteria specified and included extensive use of computer-aided analytical methods to screen the candidate concepts and select the most promising concepts for the in-depth structural analysis.

X-Wing 25 Foot Diameter Lockheed Model Whirl Test Report.

DTIC Science & Technology

1983-01-17

REPORT ( U) UNITED TECHNOLOGIES CORP STRATFORD CT SIKORSKY AIRCRAFT DIV J P PERSCHBACHER 17 JAN 83 UNCLASSIFIED SER-5i872 MDA93-8i-C-28i F/G 13 N E ...4Approved kw, ptahhc wml~ e I efti bution Unlimited C.E~ ET Eq ’ - ’~JUNZ. 1 9 8 3 k ’ ~3 06 01 100 SA3RE ,.i. ’ ,’ ’- ". -’ " .. ,. -- . -. -- s...close proximity fuselage This report will comment on the first four items; item ( e )-ll be addressed by Boeing Vertol. 1.2 Background C This rotor design

The NASA supercritical-wing technology

NASA Technical Reports Server (NTRS)

Bartlett, D. W.; Patterson, J. C., Jr.

1978-01-01

A number of high aspect ratio supercritical wings in combination with a representative wide body type fuselage were tested in the Langley 8 foot transonic pressure tunnel. The wing parameters investigated include aspect ratio, sweep, thickness to chord ratio, and camber. Subsequent to these initial series of tests, a particular wing configuration was selected for further study and development. Tests on the selected wing involved the incorporation of a larger inboard trailing edge extension, an inboard leading edge extension, and flow through nacelles. Range factors for the various supercritical wing configurations are compared with those for a reference wide body transport configuration.

The cetaceopteryx: A global range military transport aircraft

NASA Technical Reports Server (NTRS)

Brivkalns, Chad; English, Nicole; Kazemi, Tahmineh; Kopel, Kim; Kroger, Seth; Ortega, ED

1993-01-01

This paper presents a design of a military transport aircraft capable of carrying 800,000 lbs of payload from any point in the United States to any other point in the world. Such massive airlift requires aggressive use of advanced technology and a unique configuration. The Cetaceopteyx features a joined wing, canard and six turbofan engines. The aircraft has a cost 1.07 billion (1993) dollars each. This paper presents in detail the mission description, preliminary sizing, aircraft configuration, wing design, fuselage design, empennage design, propulsion system, landing gear design, structures, drag, stability and control, systems layout, and cost analysis of the aircraft.

Conformal Lightweight Antenna Structures for Aeronautical Communication Technologies

NASA Technical Reports Server (NTRS)

Meador, Mary Ann

2017-01-01

This project is to develop antennas which enable beyond line of sight (BLOS) command and control for UAVs. We will take advantage of newly assigned provisional Ku-band spectrum for UAVs and use unique antenna designs to avoid interference with ground systems. This will involve designing antennas with high isotropic effective radiated power (EIRP) and ultra-low sidelobes. The antennas will be made with polymer aerogel as a substrate to both reduce weight and improve performance, as demonstrated in an Aero Seedling. In addition, designing the antennas to be conformal to the aircraft fuselage will reduce drag.

Sound transmission through lined, composite panel structures: Transversely isotropic poro-elastic model

NASA Astrophysics Data System (ADS)

Kim, Jeong-Woo

A joint experimental and analytical investigation of the sound transmission loss (STL) and two-dimensional free wave propagation in composite sandwich panels is presented here. An existing panel, a Nomex honeycomb sandwich panel, was studied in detail. For the purpose of understanding the typical behavior of sandwich panels, a composite structure comprising two aluminum sheets with a relatively soft, poro-elastic foam core was also constructed and studied. The cores of both panels were modeled using an anisotropic (transversely isotropic) poro-elastic material theory. Several estimation methods were used to obtain the material properties of the honeycomb core and the skin plates to be used in the numerical calculations. Appropriate values selected from among the estimates were used in the STL and free wave propagation models. The prediction model was then verified in two ways: first, the calculated wave speeds and STL of a single poro-elastic layer were numerically verified by comparison with the predictions of a previously developed isotropic model. Secondly, to physically validate the transversely isotropic model, the measured STL and the phase speeds of the sandwich panels were compared with their predicted values. To analyze the actual treatment of a fuselage structure, multi-layered configurations, including a honeycomb panel and several layers such as air gaps, acoustic blankets and membrane partitions, were formulated. Then, to find the optimal solution for improving the sound barrier performance of an actual fuselage system, air layer depth and glass fiber lining effects were investigated by using these multi-layer models. By using the free wave propagation model, the first anti-symmetric and symmetric modes of the sandwich panels were characterized to allow the identification of the coincidence frequencies of the sandwich panel. The behavior of the STL could then be clearly explained by comparison with the free wave propagation solutions. By performing a parameter study based both on the STL and free wave propagation speeds, the mass, stiffness and damping-controlled regions of the STL were identified. The structural factors that can be adjusted to improve STL performance were also identified.

Finite Element Analysis of an Energy Absorbing Sub-floor Structure

NASA Technical Reports Server (NTRS)

Moore, Scott C.

1995-01-01

As part of the Advanced General Aviation Transportation Experiments program, the National Aeronautics and Space Administration's Langley Research Center is conducting tests to design energy absorbing structures to improve occupant survivability in aircraft crashes. An effort is currently underway to design an Energy Absorbing (EA) sub-floor structure which will reduce occupant loads in an aircraft crash. However, a recent drop test of a fuselage specimen with a proposed EA sub-floor structure demonstrated that the effects of sectioning the fuselage on both the fuselage section's stiffness and the performance of the EA structure were not fully understood. Therefore, attempts are underway to model the proposed sub-floor structure on computers using the DYCAST finite element code to provide a better understanding of the structure's behavior in testing, and in an actual crash.

Phase 2 and 3 wind tunnel tests of the J-97 powered, external augmentor V/STOL model. [at Ames 40 by 80 wind tunnel

NASA Technical Reports Server (NTRS)

Garland, D. B.; Harris, J. L.

1980-01-01

Static and forward speed tests were made in a 40 multiplied by 80 foot wind tunnel of a large-scale, ejector-powered V/STOL aircraft model. Modifications were made to the model following earlier tests primarily to improve longitudinal acceleration capability during transition from hovering to wingborne flight. A rearward deflection of the fuselage augmentor thrust vector was shown to be beneficial in this regard. Other augmentor modifications were tested, notably the removal of both endplates, which improved acceleration performance at the higher transition speeds. The model tests again demonstrated minimal interference of the fuselage augmentor on aerodynamic lift. A flapped canard surface also showed negligible influence on the performance of the wing and of the fuselage augmentor.

Energy-Absorbing Beam Member

NASA Technical Reports Server (NTRS)

Littell, Justin D. (Inventor)

2017-01-01

An energy-absorbing (EA) beam member and having a cell core structure is positioned in an aircraft fuselage proximate to the floor of the aircraft. The cell core structure has a length oriented along a width of the fuselage, a width oriented along a length of the fuselage, and a depth extending away from the floor. The cell core structure also includes cell walls that collectively define a repeating conusoidal pattern of alternating respective larger and smaller first and second radii along the length of the cell core structure. The cell walls slope away from a direction of flight of the aircraft at a calibrated lean angle. An EA beam member may include the cell core structure and first and second plates along the length of the cell core structure on opposite edges of the cell material.

Measurement of static pressure on aircraft

NASA Technical Reports Server (NTRS)

Gracey, William

1958-01-01

Existing data on the errors involved in the measurement of static pressure by means of static-pressure tubes and fuselage vents are presented. The errors associated with the various design features of static-pressure tubes are discussed for the condition of zero angle of attack and for the case where the tube is inclined to flow. Errors which result from variations in the configuration of static-pressure vents are also presented. Errors due to the position of a static-pressure tube in the flow field of the airplane are given for locations ahead of the fuselage nose, ahead of the wing tip, and ahead of the vertical tail fin. The errors of static-pressure vents on the fuselage of an airplane are also presented. Various methods of calibrating static-pressure installations in flight are briefly discussed.

The Low-Noise Potential of Distributed Propulsion on a Catamaran Aircraft

NASA Technical Reports Server (NTRS)

Posey, Joe W.; Tinetti, A. F.; Dunn, M. H.

2006-01-01

The noise shielding potential of an inboard-wing catamaran aircraft when coupled with distributed propulsion is examined. Here, only low-frequency jet noise from mid-wing-mounted engines is considered. Because low frequencies are the most difficult to shield, these calculations put a lower bound on the potential shielding benefit. In this proof-of-concept study, simple physical models are used to describe the 3-D scattering of jet noise by conceptualized catamaran aircraft. The Fast Scattering Code is used to predict noise levels on and about the aircraft. Shielding results are presented for several catamaran type geometries and simple noise source configurations representative of distributed propulsion radiation. Computational analyses are presented that demonstrate the shielding benefits of distributed propulsion and of increasing the width of the inboard wing. Also, sample calculations using the FSC are presented that demonstrate additional noise reduction on the aircraft fuselage by the use of acoustic liners on the inboard wing trailing edge. A full conceptual aircraft design would have to be analyzed over a complete mission to more accurately quantify community noise levels and aircraft performance, but the present shielding calculations show that a large acoustic benefit could be achieved by combining distributed propulsion and liner technology with a twin-fuselage planform.

Further comparison of wind tunnel and airplane acoustic data for advanced design high speed propeller models

NASA Technical Reports Server (NTRS)

Dittmar, J. H.

1985-01-01

Comparisons were made between the SR-2 and SR-3 model propeller noise data taken in the NASA 8-by-6 wind tunnel, in the United Technologies Research Center (UTRC) anechoic tunnel, and with boom and fuselage microphones on the NASA Jetstar airplane. Plots of peak blade passage tone noise versus helical tip Mach number generally showed good agreement. The levels of the airplane fuselage data were somewhat lower than the boom data by an approximately uniform value. The curve shapes were similar except for the UTRC data which was flatter than the other sets. This was attributed to the UTRC data being taken at constant power while the other data were taken at constant advance ratio. General curves of the peak blade passage tone versus helical tip Mach number fit through all the data are also presented. Directivity shape comparisons at the cruise condition were similar for the airplane and 8-by-6 tunnel data. The UTRC data peaked farther forward but, when an angle correction was made for the different axial Mach number used in the UTRC tests, the shape was similar to the others. The general agreement of the data from the four configurations enables the formation of a good consensus of the noise from these propellers.

Technical and Economic Assessment of Span-Distributed Loading Cargo Aircraft Concepts

NASA Technical Reports Server (NTRS)

Johnston, W. M.; Muehlbauer, J. C.; Eudaily, R. R.; Farmer, B. T.; Monrath, J. F.; Thompson, S. G.

1976-01-01

A 700,000 kg (1,540,000-lb) aircraft with a cruise Mach number of 0.75 was found to be optimum for the specified mission parameters of a 272 155-kg (600,000-lb) payload, a 5560-km (3000-n.mi.) range, and an annual productivity of 113 billion revenue-ton km (67 billion revenue-ton n. mi.). The optimum 1990 technology level spanloader aircraft exhibited the minimum 15-year life-cycle costs, direct operating costs, and fuel consumption of all candidate versions. Parametric variations of wing sweep angle, thickness ratio, rows of cargo, and cargo density were investigated. The optimum aircraft had two parallel rows of 2.44 x 2.44-m (8 x 8-ft) containerized cargo with a density of 160 kg/cu m (10 lb/ft 3) carried throughout the entire 101-m (331-ft) span of the constant chord, 22-percent thick, supercritical wing. Additional containers or outsized equipment were carried in the 24.4-m (80-ft) long fuselage compartment preceding the wing. Six 284,000-N (64,000-lb) thrust engines were mounted beneath the 0.7-rad (40-deg) swept wing. Flight control was provided by a 36.6-m (120-ft) span canard surface mounted atop the forward fuselage, by rudders on the wingtip verticals and by outboard wing flaperons.

Special Aspects of Flight Dynamics of a Reusable Cryogenic Booster Stage

NASA Astrophysics Data System (ADS)

Klevanski, J.; Sippel, M.

2005-02-01

The semi-reusable Space Transportation System (STS) investigated in the German ASTRA research program consists of two reusable winged liquid fly-back booster stages called LFBB attached to the expendable Ariane 5 core at an upgraded technology level. The focus of the presented study is the in-depth research of the special aspects of flight dynamics of a LFBB in all flight phases and especially the investigation of the trimmablity, stability and controllability. The LFBB-layout is developed under consideration of tight structural and geometrical constraints of using European cryogenic rocket engine technology [1]. These design conditions result in the special LFBB layout features, such as its large diameter fuselage and a forward position of the air-breathing engines. The great variation in the location of the center of pressure for different Mach numbers caused by large diameter fuselage and a wide range of center of gravity position demand canards for trim and control purposes. The aerodynamic investigation loop performed by the DLR Institute of Aerodynamics and Flow Technology in Braunschweig [2],[3] and in the DLR Wind Tunnel in Cologne [6] allowed optimizing the key features of the aerodynamical layout, such as the canards shape and the wing profile. The main task of this study is the investigation of the longitudinal flight dynamics of the LFBB using the obtained aerodynamic data within a closed loop simulation taking into account rigid body equations of motion, control law and actuators by realistic assumption. This paper discusses the controllability of the whole STS during the ascent phase until separation including wind gust influence, guaranteeing of the structural constraints during re-entry phase and the trimmablity and controllability of the LFBB during the cruise return flight. The flight dynamics behavior of the LFBB is investigated for different center of gravity positions with reasonable margins.

Behavior of composite/metal aircraft structural elements and components under crash type loads: What are they telling us

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Boitnott, Richard L.; Fasanella, Edwin L.

1990-01-01

Failure behavior results are presented from crash dynamics research using concepts of aircraft elements and substructure not necessarily designed or optimized for energy absorption or crash loading considerations. To achieve desired new designs which incorporate improved energy absorption capabilities often requires an understanding of how more conventional designs behave under crash loadings. Experimental and analytical data are presented which indicate some general trends in the failure behavior of a class of composite structures which include individual fuselage frames, skeleton subfloors with stringers and floor beams but without skin covering, and subfloors with skin added to the frame-stringer arrangement. Although the behavior is complex, a strong similarity in the static and dynamic failure behavior among these structures is illustrated through photographs of the experimental results and through analytical data of generic composite structural models. It is believed that the similarity in behavior is giving the designer and dynamists much information about what to expect in the crash behavior of these structures and can guide designs for improving the energy absorption and crash behavior of such structures.

Behavior of composite/metal aircraft structural elements and components under crash type loads - What are they telling us?

NASA Technical Reports Server (NTRS)

Carden, Huey D.; Boitnott, Richard L.; Fasanella, Edwin L.

1990-01-01

Failure behavior results are presented from crash dynamics research using concepts of aircraft elements and substructure not necessarily designed or optimized for energy absorption or crash loading considerations. To achieve desired new designs which incorporate improved energy absorption capabilities often requires an understanding of how more conventional designs behave under crash loadings. Experimental and analytical data are presented which indicate some general trends in the failure behavior of a class of composite structures which include individual fuselage frames, skeleton subfloors with stringers and floor beams but without skin covering, and subfloors with skin added to the frame-stringer arrangement. Although the behavior is complex, a strong similarity in the static and dynamic failure behavior among these structures is illustrated through photographs of the experimental results and through analytical data of generic composite structural models. It is believed that the similarity in behavior is giving the designer and dynamists much information about what to expect in the crash behavior of these structures and can guide designs for improving the energy absorption and crash behavior of such structures.

Unique failure behavior of metal/composite aircraft structural components under crash type loads

NASA Technical Reports Server (NTRS)

Carden, Huey D.

1990-01-01

Failure behavior results are presented on some of the crash dynamics research conducted with concepts of aircraft elements and substructure which have not necessarily been designed or optimized for energy absorption or crash loading considerations. To achieve desired new designs which incorporate improved energy absorption capabilities often requires an understanding of how more conventional designs behave under crash type loadings. Experimental and analytical data are presented which indicate some general trends in the failure behavior of a class of composite structures which include individual fuselage frames, skeleton subfloors with stringers and floor beams but without skin covering, and subfloors with skin added to the frame-stringer arrangement. Although the behavior is complex, a strong similarity in the static/dynamic failure behavior among these structures is illustrated through photographs of the experimental results and through analytical data of generic composite structural models. It is believed that the thread of similarity in behavior is telling the designer and dynamists a great deal about what to expect in the crash behavior of these structures and can guide designs for improving the energy absorption and crash behavior of such structures.

Evaluation of Braided Stiffener Concepts for Transport Aircraft Wing Structure Applications

NASA Technical Reports Server (NTRS)

Deaton, Jerry W.; Dexter, H. Benson (Editor); Markus, Alan; Rohwer, Kim

1995-01-01

Braided composite materials have potential for application in aircraft structures. Stiffeners, wing spars, floor beams, and fuselage frames are examples where braided composites could find application if cost effective processing and damage requirements are met. Braiding is an automated process for obtaining near-net shape preforms for fabrication of components for structural applications. Previous test results on braided composite materials obtained at NASA Langley indicate that damage tolerance requirements can be met for some applications. In addition, the braiding industry is taking steps to increase the material through-put to be more competitive with other preform fabrication processes. Data are presented on the compressive behavior of three braided stiffener preform fabric constructions as determined from individual stiffener crippling test and three stiffener wide panel tests. Stiffener and panel fabrication are described and compression data presented for specimens tested with and without impact damage. In addition, data are also presented on the compressive behavior of the stitched stiffener preform construction currently being used by McDonnell Douglas Aerospace in the NASA ACT wing development program.

Characteristic evaluation of process parameters of friction stir welding of aluminium 2024 hybrid composites

NASA Astrophysics Data System (ADS)

Sadashiva, M.; Shivanand, H. K.; Vidyasagar, H. N.

2018-04-01

The Current work is aimed to investigate the effect of process parameters in friction stir welding of Aluminium 2024 base alloy and Aluminium 2024 matrix alloy reinforced with E Glass and Silicon Carbide reinforcements. The process involved a set of synthesis techniques incorporating stir casting methodology resulting in fabrication of the composite material. This composite material that is synthesized is then machined to obtain a plate of dimensions 100 mm * 50 mm * 6 mm. The plate is then friction stir welded at different set of parameters viz. the spindle speed of 600 rpm, 900 rpm and 1200 rpm and feed rate of 40 mm/min, 80 mm/min and 120 mm/min for analyzing the process capability. The study of the given set of parameters is predominantly important to understand the physics of the process that may lead to better properties of the joint, which is very much important in perspective to its use in advanced engineering applications, especially in aerospace domain that uses Aluminium 2024 alloy for wing and fuselage structures under tension.

Study to investigate design, fabrication and test of low cost concepts for large hybrid composite helicopter fuselage, phase 1

NASA Technical Reports Server (NTRS)

Adams, K. M.; Lucas, J. J.

1975-01-01

The development of a frame/stringer/skin fabrication technique for composite airframe construction was studied as a low cost approach to the manufacture of large helicopter airframe components. A center cabin aluminum airframe section of the Sikorsky CH-53D helicopter was selected for evaluation as a composite structure. The design, as developed, is composed of a woven KEVLAR-49/epoxy skin and graphite/epoxy frames and stringers. To support the selection of this specific design concept a materials study was conducted to develop and select a cure compatible graphite and KEVLAR-49/epoxy resin system, and a foam system capable of maintaining shape and integrity under the processing conditions established. The materials selected were, Narmco 5209/Thornel T-300 graphite, Narmco 5209/KEVLAR-49 woven fabric, and Stathane 8747 polyurethane foam. Eight specimens were fabricated, representative of the frame, stringer, and splice joint attachments. Evaluation of the results of analysis and test indicate that design predictions are good to excellent except for some conservatism of the complex frame splice.

Multistage aerospace craft. [perspective drawings of conceptual design

NASA Technical Reports Server (NTRS)

Kelly, D. L. (Inventor)

1973-01-01

A conceptual design of a multi-stage aerospace craft is presented. Two perspective views of the vehicle are developed to show the two component configuration with delta wing, four vertical tail surfaces, tricycle landing gear, and two rocket exhaust nozzles at the rear of the fuselage. Engines for propulsion in the atmosphere are mounted on the fuselage in front of the wing root attachment.

View of the forward fuselage and the reinforced carboncarbon nose ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

View of the forward fuselage and the reinforced carbon-carbon nose of the Orbiter Discovery looking aft while mounted atop the 76-wheeled orbiter transfer system as it is being rolled from the Orbiter Processing Facility to the Vehicle Assembly Building at Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Closeup view of the upper exterior of the forward fuselage ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Close-up view of the upper exterior of the forward fuselage of the Orbiter Discovery in the Orbiter Processing Facility at NASA's Kennedy Space Center. The view show a detail of the flight deck windows with protective covers installed to protect the window surfaces during processing. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Noise of a model counterrotation propeller with simulated fuselage and support pylon at takeoff/approach conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hughes, Christopher E.

1989-01-01

Two modern high-speed advanced counterrotation propellers, F7/A7 and F7/A3 were tested in the NASA Lewis Research Centers's 9- by 15-foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Both rotors were of similar diameter on the F7/A7 propeller, while the aft rotor diameter of the F7/A3 propeller was 85 percent of the forward propeller to reduce tip vortex-aft rotor interaction. The two propellers were designed for similar performance. The propellers were tested in both the clean configuration, and installed configuration consisting of a simulated upstream nacelle support pylon and fuselage section. Acoustic measurements were made with an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could make both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish propeller operating conditions. The propellers were run at blade setting angles (fron angle/rear angle) of 41.1/39.4 deg for the F7/A7 propeller, and 41.1/46.4 deg for the F7/A3 propeller. The forward rotors were tested over a range of tip speeds from 165 to 259 m/sec (540 to 850 ft/sec), and both propellers were tested at the maximum rotor-rotor spacing, based on pitch change axis separation, of 14.99 cm (5.90 in.). The data presented in this paper are for 0 deg propeller axis angle of attack. Results are presented for the baseline, pylon-alone, and strut + fuselage configurations. The presence of the simulated fuselage resulted in higher rotor-alone tone levels in a direction normal to the advancing propeller blade near the fuselage. A corresponding rotor-alone tone reduction was often observed 180 deg circumferentially from this region of increased noise. A significant rotor-alone increase for both rotors was observed diametrically opposite the fuselage. In some cases, interaction tone levels were likewise affected by the simulated installation.

Natural laminar flow hits smoother air

NASA Technical Reports Server (NTRS)

Holmes, B. J.

1985-01-01

Natural laminar flow (NLF) may be attained in aircraft with lower cost, weight, and maintenance penalties than active flow laminarization by means of a slot suction system. A high performance general aviation jet aircraft possessing a moderate degree of NLF over wing, fuselage, empennage and engine nacelles will accrue a 24 percent reduction in total aircraft drag in the cruise regime. NASA-Langley has conducted NLF research centered on the use of novel airfoil profiles as well as composite and milled aluminum alloy construction methods which minimize three-dimensional aerodynamic surface roughness and waviness. It is noted that higher flight altitudes intrinsically reduce unit Reynolds numbers, thereby minimizing turbulence for a given cruise speed.

Wall interaction effects for a full-scale helicopter rotor in the NASA Ames 80- by 120-foot wind tunnel

NASA Technical Reports Server (NTRS)

Shinoda, Patrick M.

1994-01-01

A full-scale helicopter rotor test was conducted in the NASA Ames 80- by 120-Foot Wind Tunnel with a four-bladed S-76 rotor system. This wind tunnel test generated a unique and extensive data base covering a wide range of rotor shaft angles-of-attack and rotor thrust conditions from 0 to 100 knots. Three configurations were tested: (1) empty tunnel; (2) test stand body (fuselage) and support system; and (3) fuselage and support system with rotor installed. Empty tunnel wall pressure data are evaluated as a function of tunnel speed to understand the baseline characteristics. Aerodynamic interaction effects between the fuselage and the walls of the tunnel are investigated by comparing wall, ceiling, and floor pressures for various tunnel velocities and fuselage angles-of-attack. Aerodynamic interaction effects between the rotor and the walls of the tunnel are also investigated by comparing wall, ceiling, and floor pressures for various rotor shaft angles, rotor thrust conditions, and tunnel velocities. Empty tunnel wall pressure data show good repeatability and are not affected by tunnel speed. In addition, the tunnel wall pressure profiles are not affected by the presence of the fuselage apart from a pressure shift. Results do not indicate that the tunnel wall pressure profiles are affected by the presence of the rotor. Significant changes in the wall, ceiling, and floor pressure profiles occur with changing tunnel speeds for constant rotor thrust and shaft angle conditions. Significant changes were also observed when varying rotor thrust or rotor shaft angle-of-attack. Other results indicate that dynamic rotor loads and blade motion are influenced by the presence of the tunnel walls at very low tunnel velocity and, together with the wall pressure data, provide a good indication of flow breakdown.

Pretest Report for the Full Span Propulsive Wing/Canard Model Test in the NASA Langley 4 x 7 Meter Low Speed Wind Tunnel Second Series Test

NASA Technical Reports Server (NTRS)

Stewart, V. R.

1986-01-01

A full span propulsive wing/canard model is to be tested in the NASA Langley Research Center (LaRC) 4 x 7 meter low speed wind tunnel. These tests are a continuation of the tests conducted in Feb. 1984, NASA test No.290, and are being conducted under NASA Contract NAS1-17171. The purpose of these tests is to obtain extensive lateral-directional data with a revised fuselage concept. The wings, canards, and vertical tail of this second test series model are the same as tested in the previous test period. The fuselage and internal flow path have been modified to better reflect an external configuration suitable for a fighter airplane. Internal ducting and structure were changed as required to provide test efficiency and blowing control. The model fuselage tested during the 1984 tests was fabricated with flat sides to provide multiple wing and canard placement variations. The locations of the wing and canard are important variables in configuration development. With the establishment of the desired relative placement of the lifting surfaces, a typically shaped fuselage has been fabricated for these tests. This report provides the information necessary for the second series tests of the propulsive wing/canard model. The discussion in this report is limited to that affected by the model changes and to the second series test program. The pretest report information for test 290 which is valid for the second series test was published in Rockwell report NR 83H-79. This report is presented as Appendix 1 and the modified fuselage stress report is presented as Appendix 2 to this pretest report.

Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Aerodynamic Performance Analysis of the HSR Technology Concept Airplane at Supersonic Cruise Speed

NASA Technical Reports Server (NTRS)

Ghaffari, Farhad

1999-01-01

Unstructured grid Euler computations, performed at supersonic cruise speed, are presented for a High Speed Civil Transport (HSCT) configuration, designated as the Technology Concept Airplane (TCA) within the High Speed Research (HSR) Program. The numerical results are obtained for the complete TCA cruise configuration which includes the wing, fuselage, empennage, diverters, and flow through nacelles at M (sub infinity) = 2.4 for a range of angles-of-attack and sideslip. Although all the present computations are performed for the complete TCA configuration, appropriate assumptions derived from the fundamental supersonic aerodynamic principles have been made to extract aerodynamic predictions to complement the experimental data obtained from a 1.675%-scaled truncated (aft fuselage/empennage components removed) TCA model. The validity of the computational results, derived from the latter assumptions, are thoroughly addressed and discussed in detail. The computed surface and off-surface flow characteristics are analyzed and the pressure coefficient contours on the wing lower surface are shown to correlate reasonably well with the available pressure sensitive paint results, particularly, for the complex flow structures around the nacelles. The predicted longitudinal and lateral/directional performance characteristics for the truncated TCA configuration are shown to correlate very well with the corresponding wind-tunnel data across the examined range of angles-of-attack and sideslip. The complementary computational results for the longitudinal and lateral/directional performance characteristics for the complete TCA configuration are also presented along with the aerodynamic effects due to empennage components. Results are also presented to assess the computational method performance, solution sensitivity to grid refinement, and solution convergence characteristics.

SMART Structures User's Guide - Version 3.0

NASA Technical Reports Server (NTRS)

Spangler, Jan L.

1996-01-01

Version 3.0 of the Solid Modeling Aerospace Research Tool (SMART Structures) is used to generate structural models for conceptual and preliminary-level aerospace designs. Features include the generation of structural elements for wings and fuselages, the integration of wing and fuselage structural assemblies, and the integration of fuselage and tail structural assemblies. The highly interactive nature of this software allows the structural engineer to move quickly from a geometry that defines a vehicle's external shape to one that has both external components and internal components which may include ribs, spars, longerons, variable depth ringframes, a floor, a keel, and fuel tanks. The geometry that is output is consistent with FEA requirements and includes integrated wing and empennage carry-through and frame attachments. This report provides a comprehensive description of SMART Structures and how to use it.

Transonic Flow Field Analysis for Wing-Fuselage Configurations

NASA Technical Reports Server (NTRS)

Boppe, C. W.

1980-01-01

A computational method for simulating the aerodynamics of wing-fuselage configurations at transonic speeds is developed. The finite difference scheme is characterized by a multiple embedded mesh system coupled with a modified or extended small disturbance flow equation. This approach permits a high degree of computational resolution in addition to coordinate system flexibility for treating complex realistic aircraft shapes. To augment the analysis method and permit applications to a wide range of practical engineering design problems, an arbitrary fuselage geometry modeling system is incorporated as well as methodology for computing wing viscous effects. Configuration drag is broken down into its friction, wave, and lift induced components. Typical computed results for isolated bodies, isolated wings, and wing-body combinations are presented. The results are correlated with experimental data. A computer code which employs this methodology is described.

An estimate of the noise shielding on the fuselage resulting from installing a short duct around an advanced propeller

NASA Technical Reports Server (NTRS)

Dittmar, James H.

1988-01-01

A simple barrier shielding model was used to estimate the amount of noise shielding on the fuselage that could result from installing a short duct around a wing-mounted advanced propeller. With the propeller located one-third of the duct length from the inlet, estimates for the maximum blade passing tone attenuation varied from 7 dB for a duct 0.25 propeller diameter long to 16.75 dB for a duct 1 diameter long. Attenuations for the higher harmonics would be even larger because of their shorter wavelengths relative to the duct length. These estimates show that the fuselage noise reduction potential of a ducted compared with an unducted propeller is significant. Even more reduction might occur if acoustic attenuation material were installed in the duct.

Structural response of transport airplanes in crash situations

NASA Technical Reports Server (NTRS)

Thomson, R. G.; Caiafa, C.

1983-01-01

This report highlights the results of contractural studies of transport accident data undertaken in a joint research program sponsored by the FAA and NASA. From these accident data studies it was concluded that the greatest potential for improved transport crashworthiness is in the reduction of fire related fatalities. Accident data pertaining to fuselage integrity, main landing gear collapse, fuel tank rupture, wing breaks, tearing of tank lower surfaces, and engine pod scrubbing are discussed. In those accidents where the energy absorbing protective capability of the fuselage structure is expended and the airplane experiences major structural damage, trauma caused fatalities are also discussed. The dynamic performance of current seat/restraint systems are examined but it is concluded that the accident data does not adequately define the relationship between occupant response and the dynamic interaction with the seat, floor and fuselage structure.

A model for rotorcraft flying qualities studies

NASA Technical Reports Server (NTRS)

Mittal, Manoj; Costello, Mark F.

1993-01-01

This paper outlines the development of a mathematical model that is expected to be useful for rotorcraft flying qualities research. A computer model is presented that can be applied to a range of different rotorcraft configurations. The algorithm computes vehicle trim and a linear state-space model of the aircraft. The trim algorithm uses non linear optimization theory to solve the nonlinear algebraic trim equations. The linear aircraft equations consist of an airframe model and a flight control system dynamic model. The airframe model includes coupled rotor and fuselage rigid body dynamics and aerodynamics. The aerodynamic model for the rotors utilizes blade element theory and a three state dynamic inflow model. Aerodynamics of the fuselage and fuselage empennages are included. The linear state-space description for the flight control system is developed using standard block diagram data.

Asymmetric Thrust Reversers

NASA Technical Reports Server (NTRS)

Chandler, Jesse M. (Inventor); Suciu, Gabriel L. (Inventor)

2018-01-01

An aircraft includes a propulsion supported within an aft portion of a fuselage A thrust reverser is mounted in the aft portion of the fuselage proximate the propulsion system for directing thrust in a direction to slow the aircraft. The thrust reverser includes an upper blocker door movable about a first pivot axis to a deployed position and a lower blocker door movable about a second pivot axis not parallel to the first pivot axis.

Fire tests for airplane interior materials

NASA Technical Reports Server (NTRS)

Tustin, E. A.

1980-01-01

Large scale, simulated fire tests of aircraft interior materials were carried out in salvaged airliner fuselage. Two "design" fire sources were selected: Jet A fuel ignited in fuselage midsection and trash bag fire. Comparison with six established laboratory fire tests show that some laboratory tests can rank materials according to heat and smoke production, but existing tests do not characterize toxic gas emissions accurately. Report includes test parameters and test details.

Wireless Local Area Network Performance Inside Aircraft Passenger Cabins

NASA Technical Reports Server (NTRS)

Whetten, Frank L.; Soroker, Andrew; Whetten, Dennis A.; Whetten, Frank L.; Beggs, John H.

2005-01-01

An examination of IEEE 802.11 wireless network performance within an aircraft fuselage is performed. This examination measured the propagated RF power along the length of the fuselage, and the associated network performance: the link speed, total throughput, and packet losses and errors. A total of four airplanes: one single-aisle and three twin-aisle airplanes were tested with 802.11a, 802.11b, and 802.11g networks.

Detailed view inside the aft fuselage of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

Detailed view inside the aft fuselage of the Orbiter Discovery showing the network of supply, distribution and feed lines to deliver fuel, oxidizer and other vital gasses and fluids to the Space Shuttle Main Engines (SSMEs). This photograph was taken in the Orbiter Processing Facility at the Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

General view of the aft fuselage of the Orbiter Discovery ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

General view of the aft fuselage of the Orbiter Discovery looking forward showing Space Shuttle Main Engines (SSMEs) installed in positions one and three and an SSME on the process of being installed in position two. This photograph was taken in the Orbiter Processing Facility at the Kennedy Space Center. - Space Transportation System, Orbiter Discovery (OV-103), Lyndon B. Johnson Space Center, 2101 NASA Parkway, Houston, Harris County, TX

Nonlifting wing-body combinations with certain geometric restraints having minimum wave drag at low supersonic speeds

NASA Technical Reports Server (NTRS)

Lomax, Harvard

1957-01-01

Several variational problems involving optimum wing and body combinations having minimum wave drag for different kinds of geometrical restraints are analyzed. Particular attention is paid to the effect on the wave drag of shortening the fuselage and, for slender axially symmetric bodies, the effect of fixing the fuselage diameter at several points or even of fixing whole portions of its shape.

Drag and Cooling with Various Forms of Cowling for a "Whirlwind" Radial Air-cooled Engine II

NASA Technical Reports Server (NTRS)

Weick, Fred E

1930-01-01

This report gives the results of the second portion of an investigation in the twenty-foot Propeller Research Tunnel of the National Advisory Committee for Aeronautics, on the cowling and cooling of a "Whirlwind" J-5 radial air-cooled engine. The first portion pertains to tests with a cabin fuselage. This report covers tests with several forms of cowling, including conventional types, individual fairings behind the cylinders, individual hoods over the over the cylinders, and the new N. A. C. A. complete cowling, all on an open cockpit fuselage. Drag tests were also made with a conventional engine nacelle, and with a nacelle having the new complete cowling. In the second part of the investigation the results found in the first part were substantiated. It was also found that the reduction in drag with the complete cowling over that with conventional cowling is greater with the smaller bodies than with the cabin fuselage; in fact, the gain in the case of the completely cowled nacelle is over twice that with the cabin fuselage. The individual fairings and hoods did not prove effective in reducing the drag. The results of flight tests on AT-5A airplane has been analyzed and found to agree very well with the results of the wind tunnel tests. (author)

Study of Semi-Span Model Testing Techniques

NASA Technical Reports Server (NTRS)

Gatlin, Gregory M.; McGhee, Robert J.

1996-01-01

An investigation has been conducted in the NASA Langley 14- by 22-Foot Subsonic Tunnel in order to further the development of semi-span testing capabilities. A twin engine, energy efficient transport (EET) model with a four-element wing in a takeoff configuration was used for this investigation. Initially a full span configuration was tested and force and moment data, wing and fuselage surface pressure data, and fuselage boundary layer measurements were obtained as a baseline data set. The semi-span configurations were then mounted on the wind tunnel floor, and the effects of fuselage standoff height and shape as well as the effects of the tunnel floor boundary layer height were investigated. The effectiveness of tangential blowing at the standoff/floor juncture as an active boundary-layer control technique was also studied. Results indicate that the semi-span configuration was more sensitive to variations in standoff height than to variations in floor boundary layer height. A standoff height equivalent to 30 percent of the fuselage radius resulted in better correlation with full span data than no standoff or the larger standoff configurations investigated. Undercut standoff leading edges or the use of tangential blowing in the standoff/ floor juncture improved correlation of semi-span data with full span data in the region of maximum lift coefficient.

Wind-Tunnel Investigation of the Effects on the Aerodynamic Characteristics of Modifications to a Model of a Bomb Mounted on a Wing-Fuselage Model and to a Model of the Bomb Alone

NASA Technical Reports Server (NTRS)

King, Thomas J., Jr.

1954-01-01

An investigation was conducted in the Langley high-speed 7- by 10-foot tunnel to determine effects of modifications to a bomb model (particularly with regard to drag) when mounted on a wing-fuselage model and tested at Mach numbers from 0.70 to 1.10. In addition, the static longitudinal stability characteristics of several configurations of a larger scale model of the bomb alone were obtained over a Mach number range from 0.50 to 0.95. The results obtained for the wing-fuselage-bomb model indicate that large reductions in installation drag were obtained for the wing-fuselage-bomb model when the flat nose of the basic bomb was replaced by rounded or pointed noses of various calibers. Shortening the mounting pylon gave further decreases in the installation drag. The tests of the bomb alone indicated that only the flat-nose configurations were stable over the greater part of the Mach number range. Nose-shape modifications which improved the drag also caused the bombs to become unstable at low angles of attack. The stability of the low-drag bomb configurations could be improved by lengthening the cylindrical portion of the body behind the center of gravity.

The vibro-acoustic response and analysis of a full-scale aircraft fuselage section for interior noise reduction.

PubMed

Herdic, Peter C; Houston, Brian H; Marcus, Martin H; Williams, Earl G; Baz, Amr M

2005-06-01

The surface and interior response of a Cessna Citation fuselage section under three different forcing functions (10-1000 Hz) is evaluated through spatially dense scanning measurements. Spatial Fourier analysis reveals that a point force applied to the stiffener grid provides a rich wavenumber response over a broad frequency range. The surface motion data show global structural modes (approximately < 150 Hz), superposition of global and local intrapanel responses (approximately 150-450 Hz), and intrapanel motion alone (approximately > 450 Hz). Some evidence of Bloch wave motion is observed, revealing classical stop/pass bands associated with stiffener periodicity. The interior response (approximately < 150 Hz) is dominated by global structural modes that force the interior cavity. Local intrapanel responses (approximately > 150 Hz) of the fuselage provide a broadband volume velocity source that strongly excites a high density of interior modes. Mode coupling between the structural response and the interior modes appears to be negligible due to a lack of frequency proximity and mismatches in the spatial distribution. A high degree-of-freedom finite element model of the fuselage section was developed as a predictive tool. The calculated response is in good agreement with the experimental result, yielding a general model development methodology for accurate prediction of structures with moderate to high complexity.

Wind-tunnel investigation of several high aspect-ratio supercritical wing configurations on a wide-body-type fuselage

NASA Technical Reports Server (NTRS)

Bartlett, D. W.

1977-01-01

An investigation was conducted in the Langley 8-foot transonic pressure tunnel on two aspect-ratio 11.95 supercritical wings that were tested in combination with a representative wide-body-type fuselage. The two supercritical wings have identical planforms for equal sweep angles and differ only in thickness. Each wing was tested at quarter-chord sweep angles of 27 deg and 30 deg. At the higher sweep angle, the aspect ratio is reduced to 11.36. At 27 deg of quarter-chord sweep, the thicker supercritical wing (SCW-1) has maximum streamwise thickness-to-chord ratios of 0.16 at the wing-fuselage juncture, 0.14 at the planform break station, and 0.12 at the tip. The thinner wing (SCW-2) has maximum streamwise thickness-to-chord ratios of 0.144, 0.12, and 0.10 at the same stations respectively. Tests were also conducted on the thinner supercritical wing at the 27 deg sweep angle with a 15.24 cm (6.0 in.) shorter span which results in an aspect ratio of 10.25. For comparison, data were obtained on a current wide-body transport wing (AR=7) that was tested on the same fuselage used with the supercritical wings.

High transonic speed transport aircraft study. [aerodynamic characteristics of single-fuselage, yawed-wing configuration

NASA Technical Reports Server (NTRS)

Kulfan, R. M.; Neumann, F. D.; Nisbet, J. W.; Mulally, A. R.; Murakami, J. K.; Noble, E. C.; Mcbarron, J. P.; Stalter, J. L.; Gimmestad, D. W.; Sussman, M. B.

1973-01-01

An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range=5560 Km (3000 nmi), payload-18 143 kg (40 000lb), Mach=1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211 828 Kg (467 000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226 796 Kg (500 000 lb). Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-phase development plan is recommended to establish the full potential of the yawed-wing concept.

AFTI/F-16 in flight

NASA Technical Reports Server (NTRS)

1989-01-01

Overhead photograph of the AFTI F-16 painted in a non-standard gray finish, taken during a research flight in 1989. The two sensor pods are visible on the fuselage just forward of the wings and one of the two chin canards can be seen as a light-colored triangle ahead of one of the pods. A Sidewinder air-to-air missile is mounted on each wing tip. During the 1980s and 1990s, NASA and the U.S. Air Force participated in a joint program to integrate and demonstrate new avionics technologies to improve close air support capabilities in next-generation aircraft. The testbed aircraft, seen here in flight over the desert at NASA's Dryden Flight Research Center, Edwards, California, was called the Advanced Fighter Technology Integration (AFTI) F-16. The tests demonstrated technologies to improve navigation and the pilot's ability to find and destroy enemy ground targets day or night, including adverse weather. The aircraft--an F-16A Fighting Falcon (Serial #75-0750)--underwent numerous modifications. A relatively low-cost testbed, it evaluated the feasability of advanced, intergrated-sensor, avionics, and flight control technologies. During the first phase of the AFTI/F-16 program, which began in 1983, the aircraft demonstrated voice-actuated commands, helmet-mounted sights, flat turns, and selective fuselage pointing using forward-mounted canards and a triplex digital flight control computer system. The second phase of research, which began in the summer of 1991, demonstrated advanced technologies and capabilities to find and destroy ground targets day or night, and in adverse weather while using maneuverability and speed at low altitude. This phase was known as the close air support and battlefield air interdiction (CAS/BAI) phase. Finally, the aircraft was used to assess the Automatic Ground Collision Avoidance System (Auto - GCAS), a joint project with the Swedish Government. For these tests, the pilot flew the aircraft directly toward the ground, simulating a total loss of control. The GCAS was designed to take command in such emergencies and bring the aircraft back to level flight. The AFTI F-16 program ended at Dryden on November 4, 1997 after 15 years and over 700 research flights. The USAF continued to fly the aircraft until retiring it to the Air Force Museum on January 9, 2001.

Slender body theory programmed for bodies with arbitrary cross section. [including fuselages

NASA Technical Reports Server (NTRS)

Werner, J.; Krenkel, A. R.

1978-01-01

A computer program developed for determining the subsonic pressure, force, and moment coefficients for a fuselage-type body using slender body theory is described. The program is suitable for determining the angle of attack and sideslipping characteristics of such bodies in the linear range where viscous effects are not predominant. Procedures developed which are capable of treating cross sections with corners or regions of large curvature are outlined.

The Influence of Feedback on the Aeroelastic Behavior of Tilt Proprotor Aircraft Including the Effects of Fuselage Motion

NASA Technical Reports Server (NTRS)

Curtiss, H. C., Jr.; Komatsuzaki, T.; Traybar, J. J.

1979-01-01

The influence of single loop feedbacks to improve the stability of the system are considered. Reduced order dynamic models are employed where appropriate to promote physical insight. The influence of fuselage freedom on the aeroelastic stability, and the influence of the airframe flexibility on the low frequency modes of motion relevant to the stability and control characteristics of the vehicle were examined.

An Application of Cost-Effectiveness Analysis in a Major Defense Acquisition Program: the Decision by the U.S. Department of Defense to Retain the C-17 Transport Aircraft

DTIC Science & Technology

2010-12-01

adding hush kits to dampen the sound of the engines and adding extra material to the fuselage. The cost to restart the C-5 line was estimated at $750...include adding hush kits to dampen the sound of the engines and extra material to the fuselage. A recently completed SAB had recommended that the weep

Vertical drop test of a transport fuselage center section including the wheel wells

NASA Technical Reports Server (NTRS)

Williams, M. S.; Hayduk, R. J.

1983-01-01

A Boeing 707 fuselage section was drop tested to measure structural, seat, and anthropomorphic dummy response to vertical crash loads. The specimen had nominally zero pitch, roll and yaw at impact with a sink speed of 20 ft/sec. Results from this drop test and other drop tests of different transport sections will be used to prepare for a full-scale crash test of a B-720.

76 FR 26962 - Airworthiness Directives; Airbus Model A300 B2-1C, A300 B2-203, A300 B2K-3C, A300-B4-103, A300 B4...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-10

... repetitive detailed inspections for disbonding and cracking of the fuselage inner doubler; eddy current and... detailed inspections for disbonding and cracking of the fuselage inner doubler; eddy current and ultrasonic... or Modification if Necessary (g) For airplanes on which an eddy current inspection of the ``special...

Development of fire test methods for airplane interior materials

NASA Technical Reports Server (NTRS)

Tustin, E. A.

1978-01-01

Fire tests were conducted in a 737 airplane fuselage at NASA-JSC to characterize jet fuel fires in open steel pans (simulating post-crash fire sources and a ruptured airplane fuselage) and to characterize fires in some common combustibles (simulating in-flight fire sources). Design post-crash and in-flight fire source selections were based on these data. Large panels of airplane interior materials were exposed to closely-controlled large scale heating simulations of the two design fire sources in a Boeing fire test facility utilizing a surplused 707 fuselage section. Small samples of the same airplane materials were tested by several laboratory fire test methods. Large scale and laboratory scale data were examined for correlative factors. Published data for dangerous hazard levels in a fire environment were used as the basis for developing a method to select the most desirable material where trade-offs in heat, smoke and gaseous toxicant evolution must be considered.

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Tomek, W. G.; Hall, R. M.; Wahls, R. A.; Luckring, J. M.; Owens, L. R.

2002-01-01

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

Analysis, Design and Optimization of Non-Cylindrical Fuselage for Blended-Wing-Body (BWB) Vehicle

NASA Technical Reports Server (NTRS)

Mukhopadhyay, V.; Sobieszczanski-Sobieski, J.; Kosaka, I.; Quinn, G.; Charpentier, C.

2002-01-01

Initial results of an investigation towards finding an efficient non-cylindrical fuselage configuration for a conceptual blended-wing-body flight vehicle were presented. A simplified 2-D beam column analysis and optimization was performed first. Then a set of detailed finite element models of deep sandwich panel and ribbed shell construction concepts were analyzed and optimized. Generally these concepts with flat surfaces were found to be structurally inefficient to withstand internal pressure and resultant compressive loads simultaneously. Alternatively, a set of multi-bubble fuselage configuration concepts were developed for balancing internal cabin pressure load efficiently, through membrane stress in inner-stiffened shell and inter-cabin walls. An outer-ribbed shell was designed to prevent buckling due to external resultant compressive loads. Initial results from finite element analysis appear to be promising. These concepts should be developed further to exploit their inherent structurally efficiency.

Low-speed stability and control characteristics of a transport model with aft-fuselage-mounted advanced turboprops

NASA Technical Reports Server (NTRS)

Applin, Z. T.; Coe, P. L., Jr.

1986-01-01

A limited experimental investigation was conducted in the Langley 4- by 7-Meter Tunnel to explore the effects of aft-fuselage-mounted advanced turboprop installations on the low-speed stability and control characteristics of a representative transport aircraft in a landing configuration. In general, the experimental results indicate that the longitudinal and lateral-directional stability characteristics for the aft-fuselage-mounted single-rotation tractor and counter-rotation pusher propeller configurations tested during this investigation are acceptable aerodynamically. For the single-rotation tractor configuration, the propeller-induced aerodynamics are significantly influenced by the interaction of the propeller slipstream with the pylon and nacelle. The stability characteristics for the counter-rotation pusher configuration are strongly influenced by propeller normal forces. The longitudinal and directional control effectiveness, engine-out characteristics, and ground effects are also presented. In addition, a tabulated presentation of all aerodynamic data presented in this report is included as an appendix.

Simulator study of flight characteristics of several large, dissimilar, cargo transport airplanes during approach and landing

NASA Technical Reports Server (NTRS)

Grantham, W. D.; Smith, P. M.; Deal, P. L.; Neely, W. R., Jr.

1984-01-01

A six-degree-of-freedom, ground based simulator study is conducted to evaluate the low-speed flight characteristics of four dissimilar cargo transport airplanes. These characteristics are compared with those of a large, present-day (reference) transport configuration similar to the Lockheed C-5A airplane. The four very large transport concepts evaluated consist of single-fuselage, twin-fuselage, triple-fuselage, and span-loader configurations. The primary piloting task is the approach and landing operation. The results of his study indicate that all four concepts evaluated have unsatisfactory longitudinal and lateral directional low speed flight characteristics and that considerable stability and control augmentation would be required to improve these characteristics (handling qualities) to a satisfactory level. Through the use of rate command/attitude hold augmentation in the pitch and roll axes, and the use of several turn-coordination features, the handling qualities of all four large transports simulated are improved appreciably.

Low-speed aerodynamic characteristics from wind-tunnel tests of a large-scale advanced arrow-wing supersonic-cruise transport concept

NASA Technical Reports Server (NTRS)

Smith, P. M.

1978-01-01

Tests have been conducted to extend the existing low speed aerodynamic data base of advanced supersonic-cruise arrow wing configurations. Principle configuration variables included wing leading-edge flap deflection, wing trailing-edge flap deflection, horizontal tail effectiveness, and fuselage forebody strakes. A limited investigation was also conducted to determine the low speed aerodynamic effects due to slotted training-edge flaps. Results of this investigation demonstrate that deflecting the wing leading-edge flaps downward to suppress the wing apex vortices provides improved static longitudinal stability; however, it also results in significantly reduced static directional stability. The use of a selected fuselage forebody strakes is found to be effective in increasing the level of positive static directional stability. Drooping the fuselage nose, which is required for low-speed pilot vision, significantly improves the later-directional trim characteristics.

Ditching Tests of a 1/24-Scale Model of the Lockheed XR60-1 Airplane, TED No. NACA 235

NASA Technical Reports Server (NTRS)

Fisher, Lloyd J.; Cederborg, Gibson A.

1948-01-01

The ditching characteristics of the Lockheed XR60-1 airplane were determined by tests of a 1/24-scale dynamic model in calm water at the Langley tank no. 2 monorail. Various landing attitudes, flap settings, speeds, and conditions of damager were investigated. The ditching behavior was evaluated from recordings of decelerations, length of runs, and motions of the model. Scale-strength bottoms and simulated crumpled bottoms were used to reproduce probable damage to the fuselage. It was concluded that the airplane should be ditched at a landing attitude of about 5 deg with flaps full down. At this attitude, the maximum longitudinal deceleration should not exceed 2g and the landing run will be bout three fuselage lengths. Damage to the fuselage will not be excessive and will be greatest near the point of initial contact with the water.

A comparison of the aerodynamic characteristics at transonic speeds of four wing-fuselage configurations as determined from different test techniques, 4 October 1960

NASA Technical Reports Server (NTRS)

Donlan, C. J.; Myers, B. C., II; Mattson, A. T.

1976-01-01

The high speed aerodynamic characteristics of a family of four wing-fuselage configurations of 0, 35, 45, and 60 deg sweepback were determined from transonic bump model tests that were conducted in the Langley high speed 7 by 10 foot tunnel; sting supported model tests were conducted in the Langley 8 foot high speed tunnel and in the Langley high speed 7 by 10 foot tunnel, and rocket model tests were conducted by the Langley Pilotless Aircraft Research Division. A complementary study of the effect of Mach number gradients and streamline curvature on bump results is also included. The qualitative data obtained from the various test facilities for the wing-fuselage configurations were in essential agreement as regards the relative effects of sweepback and Mach number except for drag at zero lift. Quantitatively, important differences were present.

Over-the-wing propeller

NASA Technical Reports Server (NTRS)

Johnson, Joseph L., Jr. (Inventor); White, E. Richard (Inventor)

1986-01-01

This invention is an aircraft with a system for increasing the lift drag ratio over a broad range of operating conditions. The system positions the engines and nacelles over the wing in such a position that gains in propeller efficiency is achieved simultaneously with increases in wing lift and a reduction in wing drag. Adverse structural and torsional effects on the wings are avoided by fuselage mounted pylons which attach to the upper portion of the fuselage aft of the wings. Similarly, pylon-wing interference is eliminated by moving the pylons to the fuselage. Further gains are achieved by locating the pylon surface area aft of the aircraft center of gravity, thereby augmenting both directional and longitudinal stability. This augmentation has the further effect of reducing the size, weight and drag of empennage components. The combination of design changes results in improved cruise performance and increased climb performance while reducing fuel consumption and drag and weight penalties.

Gemini: A long-range cargo transport

NASA Technical Reports Server (NTRS)

1994-01-01

The proposed Gemini, a long-range cargo transport, is designed as a high capacity, dedicated cargo transporter of 8'x8'x20' inter-modal containers, and long-range design. These requirements will result in a design that is larger than any existing aircraft. Due to the size, a conventional configuration would result in an aircraft unable to operate economically at existing airports. It is necessary to design for a minimum possible empty weight, wingspan, and landing gear track. After considering both a single fuselage biplane and a double fuselage biplane configuration, the design team choose the double fuselage biplane configuration. Both of these configuration choices result in a reduced wing root bending moment and subsequently in substantial savings in the wing weight. An overall decrease in the weight of the airplane, its systems, and fuel will be a direct result of the wing weight savings.

Results of a Cyclic Load Test of an RB-47E Airplane

NASA Technical Reports Server (NTRS)

Huston, Wilber B.

1959-01-01

Results of a cyclic load test made by NASA on an EB-47E airplane are given. The test reported on is for one of three B-47 airplanes in a test program set up by the U. S. Air Force to evaluate the effect of wing structural reinforcements on fatigue life. As a result of crack development in the upper fuselage longerons of the other two airplanes in the program, a longeron and fuselage skin modification was incorporated early in the test. Fuselage strain-gage measurements made before and after the longeron modification and wing strain-gage measurements made only after wing reinforcement are summarized. The history of crack development and repair is given in detail. Testing was terminated one sequence short of the planned end of the program with the occurrence of a major crack in the lower right wing skin.

Arrow-wing supersonic cruise aircraft structural design concepts evaluation. Volume 4: Sections 15 through 21

NASA Technical Reports Server (NTRS)

Sakata, I. F.; Davis, G. W.

1975-01-01

The analyses performed to provide structural mass estimates for the arrow wing supersonic cruise aircraft are presented. To realize the full potential for structural mass reduction, a spectrum of approaches for the wing and fuselage primary structure design were investigated. The objective was: (1) to assess the relative merits of various structural arrangements, concepts, and materials; (2) to select the structural approach best suited for the Mach 2.7 environment; and (3) to provide construction details and structural mass estimates based on in-depth structural design studies. Production costs, propulsion-airframe integration, and advanced technology assessment are included.

On-line structural health and fire monitoring of a composite personal aircraft using an FBG sensing system

NASA Astrophysics Data System (ADS)

Chandler, K.; Ferguson, S.; Graver, T.; Csipkes, A.; Mendez, A.

2008-03-01

We report in this paper on the design and development of a novel on-line structural health monitoring and fire detection system based on an array of optical fiber Bragg grating (FBG) sensors and interrogation system installed on a new, precommercial compact aircraft. A combined total of 17 FBG sensors - strain, temperature and high-temperature - were installed at critical locations in an around the wings, fuselage and engine compartment of a prototype, Comp Air CA 12 all-composite, ten-passenger personal airplane powered by a 1,650 hp turbine engine. The sensors are interrogated online and in real time by a swept laser FBG interrogator (Micron Optics sm125-700) mounted on board the plane. Sensors readings are then combined with the plane's avionics system and displayed on the pilot's aviation control panel. This system represents the first of its kind in commercial, small frame, airplanes and a first for optical fiber sensors.

Study of Surface Roughness and Cutting force in machining for 6068 Aluminium alloy

NASA Astrophysics Data System (ADS)

Purushothaman, D.; Kaushik Yanamundra, Krishna; Krishnan, Gokul; Perisamy, C.

2018-04-01

Metal matrix composites, in particular, Aluminium Hybrid Composites are gaining increasing attention for applications in air and land because of their superior strength to weight ratio, density and high temperature resistance. Aluminium alloys are being used for a wide range of applications in Aerospace and Automobile industries, to name a few. The Aluminium Alloy 6068 has been used as the specimen. It is mainly composed of Aluminium (93.22 - 97.6 %), Magnesium (0.60 - 1.2 %), Silicon (0.60 - 1.4 %) and Bismuth (0.60 - 1.1 %). Aluminium 6068 is widely used for manufacturing aircraft structures, fuselages and wings. It is also extensively used in fabricating automobile parts such as wheel spacers. In this study, tests for the measurement of surface roughness and cutting force has been carried out on the specimen, the results evaluated and conclusions are drawn. Also the simulation of the same is carried out in a commercial FE software – ABAQUS.

Vibroacoustic Response Data of Stiffened Panels and Cylinders

NASA Technical Reports Server (NTRS)

Cabell, Randolph; Klos, Jake; Buehrle, Ralph; Schiller, Noah

2008-01-01

NASA has collected vibroacoustic response data on a variety of complex, aerospace structures to support research into numerical modeling of such structures. This data is being made available to the modeling community to promote the development and validation of analysis methods for these types of structures. Existing data from two structures is described, as well as plans for a data set from a third structure. The first structure is a 1.22 m by 1.22 m stiffened aluminum panel, typical of a commercial aircraft sidewall section. The second is an enclosed, stiffened aluminum cylinder, approximately 3.66 m long and 1.22 m in diameter, constructed to resemble a small aircraft fuselage with no windows and a periodic structure. The third structure is a filament-wound composite cylinder with composite stiffeners. Numerous combinations of excitation and response variables were measured on the structures, including: shaker excitation; diffuse acoustic field; velocity response from a laser vibrometer; intensity scans; and point acceleration.

Methodologies for Combined Loads Tests Using a Multi-Actuator Test Machine

NASA Technical Reports Server (NTRS)

Rouse, Marshall

2013-01-01

The NASA Langley COmbined Loads Test System (COLTS) Facility was designed to accommodate a range of fuselage structures and wing sections and subject them to both quasistatic and cyclic loading conditions. Structural tests have been conducted in COLTS that address structural integrity issues of metallic and fiber reinforced composite aerospace structures in support of NASA Programs (i.e. the Aircraft Structural Integrity (ASIP) Program, High-Speed-Research program and the Supersonic Project, NASA Engineering and Safety Center (NESC) Composite Crew Module Project, and the Environmentally Responsible Aviation Program),. This paper presents experimental results for curved panels subjected to mechanical and internal pressure loads using a D-box test fixture. Also, results are presented that describe use of a checkout beam for development of testing procedures for a combined mechanical and pressure loading test of a Multi-bay box. The Multi-bay box test will be used to experimentally verify the structural performance of the Multi-bay box in support of the Environmentally Responsible Aviation Project at NASA Langley.

POSTOP: Postbuckled open-stiffener optimum panels-theory and capability

NASA Technical Reports Server (NTRS)

Dickson, J. N.; Biggers, S. B.

1984-01-01

The computer program POSTOP was developed to serve as an aid in the analysis and sizing of stiffened composite panels that are loaded in the postbuckling regime. A comprehensive set of analysis routines was coupled to a widely used optimization program to produce this sizing code. POSTOP is intended for the preliminary design of metal or composite panels with open-section stiffeners, subjected to multiple combined biaxial compression (or tension), shear and normal pressure load cases. Longitudinal compression, however, is assumed to be the dominant loading. Temperature, initial bow eccentricity and load eccentricity effects are included. The panel geometry is assumed to be repetitive over several bays in the longitudinal (stiffener) direction as well as in the transverse direction. Analytical routines are included to compute panel stiffnesses, strains, local and panel buckling loads, and skin/stiffener interface stresses. The resulting program is applicable to stiffened panels as commonly used in fuselage, wing, or empennage structures. The analysis procedures and rationale for the assumptions used therein are described in detail.

High Energy Wide Area Blunt Impact on Composite Aircraft Structures

NASA Astrophysics Data System (ADS)

DeFrancisci, Gabriela K.

The largest source of damage to commercial aircraft is caused by accidental contact with ground service equipment (GSE). The cylindrical bumper typically found on GSE distributes the impact load over a large contact area, possibly spanning multiple internal structural elements (frame bays) of a stiffened-skin fuselage. This type of impact can lead to damage that is widespread and difficult to detect visually. To address this problem, monolithic composite panels of various size and complexity have been modeled and tested quasi-statically and dynamically. The experimental observations have established that detectability is dependent on the impact location and immediately-adjacent internal structure of the panel, as well as the impactor geometry and total deformation of the panel. A methodology to model and predict damage caused by wide area blunt impact events was established, which was then applied to more general cases that were not tested in order to better understand the nature of this type of impact event and how it relates to the final damage state and visual detectability.

Effects of intra- and inter-laminar resin content on the mechanical properties of toughened composite materials

NASA Technical Reports Server (NTRS)

Grande, Dodd H.; Ilcewicz, Larry B.; Avery, William B.; Bascom, Willard D.

1991-01-01

Composite materials having multiphase toughened matrix systems and laminate architectures characterized by resin-rich interlaminar layers (RIL) have been the subject of much recent attention. Such materials are likely to find applications in thick compressively loaded structures such as the keel area of commercial aircraft fuselages. The effects of resin content and its interlaminar and intralaminar distribution on mechanical properties were investigated with test and analysis of two carbon-epoxy systems. The RIL was found to reduce the in situ strengthening effect for matrix cracking in laminates. Mode 2 fracture toughness was found to increase with increasing RIL thickness over the range investigated, and Mode 1 interlaminar toughness was negligibly affected. Compressive failure strains were found to increase with increasing resin content for specimens having no damage, holes, and impact damage. Analytical tools for predicting matrix cracking of off-axis plies and damage tolerance in compression after impact (CAI) were successfully applied to materials with RIL.