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Sample records for fuselage acoustic treatment

  1. Laboratory tests on an aircraft fuselage to determine the insertion loss of various acoustic add-on treatments

    NASA Technical Reports Server (NTRS)

    Heitman, K. E.; Mixson, J. S.

    1984-01-01

    This paper describes a laboratory study of add-on acoustic treatments for a propeller-driven light aircraft fuselage. The treatments included: no treatment (i.e., baseline fuselage); a production-type double-wall interior; and various amounts of high density fiberglass added to the baseline fuselage. The sound source was a pneumatic-driver with attached exponential horn, supplied with a broadband signal. Data were acquired at the approximate head positions of the six passenger seats. The results were analyzed on space-averaged narrowband, one-third octave band and overall insertion loss basis. In addition, insertion loss results for the different configurations at specific frequencies representing propeller tone spectra are presented. The propeller tone data includes not only the space-averaged insertion loss, but also the variation of insertion loss at these particular frequencies across the six microphone positions.

  2. 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.

  3. Transport composite fuselage technology: Impact dynamics and acoustic transmission

    NASA Technical Reports Server (NTRS)

    Jackson, A. C.; Balena, F. J.; Labarge, W. L.; Pei, G.; Pitman, W. A.; Wittlin, G.

    1986-01-01

    A program was performed to develop and demonstrate the impact dynamics and acoustic transmission technology for a composite fuselage which meets the design requirements of a 1990 large transport aircraft without substantial weight and cost penalties. The program developed the analytical methodology for the prediction of acoustic transmission behavior of advanced composite stiffened shell structures. The methodology predicted that the interior noise level in a composite fuselage due to turbulent boundary layer will be less than in a comparable aluminum fuselage. The verification of these analyses will be performed by NASA Langley Research Center using a composite fuselage shell fabricated by filament winding. The program also developed analytical methodology for the prediction of the impact dynamics behavior of lower fuselage structure constructed with composite materials. Development tests were performed to demonstrate that the composite structure designed to the same operating load requirement can have at least the same energy absorption capability as aluminum structure.

  4. Investigation of fuselage acoustic treatment for a twin-engine turboprop aircraft in flight and laboratory tests

    NASA Technical Reports Server (NTRS)

    Mixson, J. S.; Oneal, R. L.; Grosveld, F. W.

    1984-01-01

    A flight and laboratory study of sidewall acoustic treatment for cabin noise control is described. In flight, cabin noise levels were measured at six locations with three treatment configurations. Noise levels from narrow-band analysis are reduced to one-third octave format and used to calculate insertion loss, IL, defined as the reduction of interior noise associated with the addition of a treatment. Laboratory tests used a specially constructed structural panel modeled after the propeller plane section of the aircraft sidewall, and acoustic treatments representing those used in flight. Lab measured transmission loss and absorption values were combined using classical acoustic procedures to obtain a prediction of IL. Comparison with IL values measured in flight for the boundary layer component of the noise indicated general agreement.

  5. Distribution of Acoustic Power Spectra for an Isolated Helicopter Fuselage

    NASA Astrophysics Data System (ADS)

    Kusyumov, A. N.; Mikhailov, S. A.; Garipova, L. I.; Batrakov, A. S.; Barakos, G.

    2016-03-01

    The broadband aerodynamic noise can be studied, assuming isotropic flow, turbulence and decay. Proudman's approach allows practical calculations of noise based on CFD solutions of RANS or URANS equations at the stage of post processing and analysis of the solution. Another aspect is the broadband acoustic spectrum and the distribution of acoustic power over a range of frequencies. The acoustic energy spectrum distribution in isotropic turbulence is non monotonic and has a maximum at a certain value of Strouhal number. In the present work the value of acoustic power peak frequency is determined using a prescribed form of acoustic energy spectrum distribution presented in papers by S. Sarkar and M. Y. Hussaini and by G. M. Lilley. CFD modelling of the flow around isolated helicopter fuselage model was considered using the HMB CFD code and the RANS equations.

  6. Acoustically Tailored Composite Rotorcraft Fuselage Panels

    NASA Technical Reports Server (NTRS)

    Hambric, Stephen; Shepherd, Micah; Koudela, Kevin; Wess, Denis; Snider, Royce; May, Carl; Kendrick, Phil; Lee, Edward; Cai, Liang-Wu

    2015-01-01

    A rotorcraft roof sandwich panel has been redesigned to optimize sound power transmission loss (TL) and minimize structure-borne sound for frequencies between 1 and 4 kHz where gear meshing noise from the transmission has the most impact on speech intelligibility. The roof section, framed by a grid of ribs, was originally constructed of a single honeycomb core/composite face sheet panel. The original panel has coincidence frequencies near 700 Hz, leading to poor TL across the frequency range of 1 to 4 kHz. To quiet the panel, the cross section was split into two thinner sandwich subpanels separated by an air gap. The air gap was sized to target the fundamental mass-spring-mass resonance of the double panel system to less than 500 Hz. The panels were designed to withstand structural loading from normal rotorcraft operation, as well as 'man-on-the-roof' static loads experienced during maintenance operations. Thin layers of VHB 9469 viscoelastomer from 3M were also included in the face sheet ply layups, increasing panel damping loss factors from about 0.01 to 0.05. Measurements in the NASA SALT facility show the optimized panel provides 6-11 dB of acoustic transmission loss improvement, and 6-15 dB of structure-borne sound reduction at critical rotorcraft transmission tonal frequencies. Analytic panel TL theory simulates the measured performance quite well. Detailed finite element/boundary element modeling of the baseline panel simulates TL slightly more accurately, and also simulates structure-borne sound well.

  7. Experimental and analytical investigations of fuselage modal characteristics and structural-acoustic coupling

    NASA Technical Reports Server (NTRS)

    Simpson, Myles A.; Mathur, Gopal P.

    1992-01-01

    Measurements conducted on a DC-9 aircraft test section to define the shell and cavity modes of the fuselage, understand its structural-acoustic coupling characteristics, and measure its response to different types of acoustic and vibration excitations are reported. The data were processed to generate spatial plots and wavenumber maps of the shell acceleration and cabin acoustic pressure field. Analysis and interpretation of the spatial plots and wavenumber maps showed that the only structural-acoustic coupling occurred at 105 Hz between the N=2 circumferential structural mode and the (n=2, p=0) circumferential cavity mode. The fuselage response to vibration excitation was found to be dominated by modes whose order increases with frequency.

  8. Nonlinear Acoustic Response of an Aircraft Fuselage Sidewall Structure by a Reduced-Order Analysis

    NASA Technical Reports Server (NTRS)

    Przekop, Adam; Rizzi, Stephen A.; Groen, David S.

    2006-01-01

    A reduced-order nonlinear analysis of a structurally complex aircraft fuselage sidewall panel is undertaken to explore issues associated with application of such analyses to practical structures. Of primary interest is the trade-off between computational efficiency and accuracy. An approach to modal basis selection is offered based upon the modal participation in the linear regime. The nonlinear static response to a uniform pressure loading and nonlinear random response to a uniformly distributed acoustic loading are computed. Comparisons of the static response with a nonlinear static solution in physical degrees-of-freedom demonstrate the efficacy of the approach taken for modal basis selection. Changes in the modal participation as a function of static and random loading levels suggest a means for improvement in the basis selection.

  9. A study of the structural-acoustic response and interior noise levels of fuselage structures

    NASA Technical Reports Server (NTRS)

    Koval, L. R.

    1978-01-01

    Models of both flat and curved fuselage panels were tested for their sound transmission characteristics. The effect of external air flow on transmission loss was simulated in a subsonic wind-tunnel. By numerically evaluating the known equations for field-incidence transmission loss of single-walled panels in a computer program, a comparison of the theory with the test results was made. As a further extension to aircraft fuselage simulation, equations for the field-incidence transmission loss of a double-walled panel were derived. Flow is shown to provide a small increase in transmission loss for a flat panel. Curvature is shown to increase transmission loss for low frequencies, while also providing a sharp decrease in transmission loss at the ring frequency of the cylindrical panel. The field-incidence transmission loss of a double-walled panel was found to be approximately twice that for a single-walled panel, with the addition of dips in the transmission loss at the air gap resonances and at the critical frequency of the internal panel.

  10. Post Treatment of Acoustic Neuroma

    MedlinePlus

    Home What is an AN What is an Acoustic Neuroma? Identifying an AN Symptoms Acoustic Neuroma Keywords Educational Video Pre-Treatment Treatment Options Summary Treatment Options Watch and Wait Radiation Microsurgery Acoustic Neuroma Decision Tree Questions for Your Physician Questions ...

  11. Experimental study using nearfield acoustic holography of sound transmission through fuselage sidewall structures

    NASA Technical Reports Server (NTRS)

    Maynard, J. D.

    1986-01-01

    The reduction of cabin noise in lightweight, propeller-driven aircraft is an especially difficult problem in noise control. Nearfield Acoustic Holography (NAH) was used to determine the mode of vibration and acoustic intensity for panels which differed in: construction (number of stiffening ribs, size of stifening ribs, construction material, and panel surface curvature); boundary support condition (free edge condition or clamped edge condition); and mode of excitation (structural-borne forces or airborne forces). The different samples of aircraft panels are described and the measurement of the natural response frequencies was discussed under various boundary support and excitation conditions. The results of the NAH measurements are presented.

  12. Experimental study using Nearfield Acoustical Holography of sound transmission fuselage sidewall structures

    NASA Technical Reports Server (NTRS)

    Maynard, J. D.

    1983-01-01

    This project involves the development of the Nearfield Acoustic Holography (NAH) technique (in particular its extension from single frequency to wideband noise measurement) and its application in a detailed study of the noise radiation characteristics of several samples of aircraft sidewall panels. With the extensive amount of information provided by the NAH technique, the properties of the sound field radiated by the panels may be correlated with their structure, mounting, and excitation (single frequency or wideband, spatially correlated or uncorrelated, structure-borne). The work accomplished at the beginning of this grant period included: (1) Calibration of the 256 microphone array and test of its accuracy. (2) extension of the facility to permit measurements on wideband noise sources. The extensions incuded the addition of high-speed data acquisition hardware and an array processor, and the development of new software. (3) Installation of motion picture graphics for correlating panel motion with structure, mounting, radiation, etc. (4) Development of new holographic data processing techniques.

  13. 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.

  14. 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. PMID:16018470

  15. Design and performance of duct acoustic treatment

    NASA Technical Reports Server (NTRS)

    Motsinger, R. E.; Kraft, R. E.

    1991-01-01

    The procedure for designing acoustic treatment panels used to line the walls of aircraft engine ducts and for estimating the resulting suppression of turbofan engine duct noise is discussed. This procedure is intended to be used for estimating noise suppression of existing designs or for designing new acoustic treatment panels and duct configurations to achieve desired suppression levels.

  16. 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.

  17. 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.

  18. Some comments on fuselage drag

    NASA Technical Reports Server (NTRS)

    Roskam, J.

    1975-01-01

    The following areas relating to fuselage drag are considered: (1) fuselage fineness - ratio and why and how this can be selected during preliminary design; (2) windshield drag; (3) skin roughness; and (4) research needs in the area of fuselage drag.

  19. Laboratory study of the effects of sidewall treatment, source directivity and temperature on the interior noise of a light aircraft fuselage

    NASA Technical Reports Server (NTRS)

    Heitman, K. E.; Mixson, J. S.

    1986-01-01

    This paper describes a laboratory study of add-on {coustic treatments for a twin-engine, propeller-driven aircraft fuselage. The sound source was a pneumatic-driver, with attached horn to simulate propeller noise distribution, powered by a white noise signal. Treatments included a double-wall, production-line treatment and various fiberglass and lead-vinyl treatments. Insertion losses, space-averaged across six interior microphone positions, were used to evaluate the treatments. In addition, the effects of sound source angle and ambient temperature on interior sound pressure level are presented. The sound source angle is shown to have a significant effect on one-third octave band localized sound pressure level. While changes in ambient temperature are shown to have little effect on one-third octave band localized sound pressure level, the change in narrowband localized sound pressure level may be dramatic.

  20. Home studio acoustic treatments on a budget

    NASA Astrophysics Data System (ADS)

    Haverstick, Gavin A.

    2003-04-01

    Digital technology in the recording industry has evolved and expanded, allowing it to be widely available to the public at a significantly lower cost than in previous years. Due to this fact, numerous home studios are either being built inside or converted from bedrooms, dens, and basements. Hobbyists and part-time musicians that typically do not have the advantage of a large recording budget operate the majority of these home studios. Along with digital equipment, acoustic treatment has become more affordable over the years giving many musicians the ability to write, record, and produce an entire album in the comfort of their own home without having to sacrifice acoustical quality along the way. Three separate case studies were conducted on rooms with various sizes, applications, and budgets. Acoustical treatment such as absorption, diffusion, and bass trapping were implemented to reduce the effects of issues such as flutter echo, excessive reverberation, and bass build-up among others. Reactions and subjective comments from each individual studio owner were gathered and assessed to determine how effective home studios can be on a personal and professional level if accurately treated acoustically.

  1. Acoustic Treatment Design Scaling Methods. Phase 2

    NASA Technical Reports Server (NTRS)

    Clark, L. (Technical Monitor); Parrott, T. (Technical Monitor); Jones, M. (Technical Monitor); Kraft, R. E.; Yu, J.; Kwan, H. W.; Beer, B.; Seybert, A. F.; Tathavadekar, P.

    2003-01-01

    The ability to design, build and test miniaturized acoustic treatment panels on scale model fan rigs representative of full scale engines provides not only cost-savings, but also an opportunity to optimize the treatment by allowing multiple tests. To use scale model treatment as a design tool, the impedance of the sub-scale liner must be known with confidence. This study was aimed at developing impedance measurement methods for high frequencies. A normal incidence impedance tube method that extends the upper frequency range to 25,000 Hz. without grazing flow effects was evaluated. The free field method was investigated as a potential high frequency technique. The potential of the two-microphone in-situ impedance measurement method was evaluated in the presence of grazing flow. Difficulties in achieving the high frequency goals were encountered in all methods. Results of developing a time-domain finite difference resonator impedance model indicated that a re-interpretation of the empirical fluid mechanical models used in the frequency domain model for nonlinear resistance and mass reactance may be required. A scale model treatment design that could be tested on the Universal Propulsion Simulator vehicle was proposed.

  2. 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.

  3. Advanced composite fuselage technology

    NASA Technical Reports Server (NTRS)

    Ilcewicz, Larry B.; Smith, Peter J.; Horton, Ray E.

    1993-01-01

    Boeing's ATCAS program has completed its third year and continues to progress towards a goal to demonstrate composite fuselage technology with cost and weight advantages over aluminum. Work on this program is performed by an integrated team that includes several groups within The Boeing Company, industrial and university subcontractors, and technical support from NASA. During the course of the program, the ATCAS team has continued to perform a critical review of composite developments by recognizing advances in metal fuselage technology. Despite recent material, structural design, and manufacturing advancements for metals, polymeric matrix composite designs studied in ATCAS still project significant cost and weight advantages for future applications. A critical path to demonstrating technology readiness for composite transport fuselage structures was created to summarize ATCAS tasks for Phases A, B, and C. This includes a global schedule and list of technical issues which will be addressed throughout the course of studies. Work performed in ATCAS since the last ACT conference is also summarized. Most activities relate to crown quadrant manufacturing scaleup and performance verification. The former was highlighted by fabricating a curved, 7 ft. by 10 ft. panel, with cocured hat-stiffeners and cobonded J-frames. In building to this scale, process developments were achieved for tow-placed skins, drape formed stiffeners, braided/RTM frames, and panel cure tooling. Over 700 tests and supporting analyses have been performed for crown material and design evaluation, including structural tests that demonstrated limit load requirements for severed stiffener/skin failsafe damage conditions. Analysis of tests for tow-placed hybrid laminates with large damage indicates a tensile fracture toughness that is higher than that observed for advanced aluminum alloys. Additional recent ATCAS achievements include crown supporting technology, keel quadrant design evaluation, and

  4. Flying wings / flying fuselages

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.

    2001-01-01

    The present paper has documented the historical relationships between various classes of all lifting vehicles, which includes the flying wing, all wing, tailless, lifting body, and lifting fuselage. The diversity in vehicle focus was to ensure that all vehicle types that map have contributed to or been influenced by the development of the classical flying wing concept was investigated. The paper has provided context and perspective for present and future aircraft design studies that may employ the all lifting vehicle concept. The paper also demonstrated the benefit of developing an understanding of the past in order to obtain the required knowledge to create future concepts with significantly improved aerodynamic performance.

  5. Inlet total pressure loss due to acoustic wall treatment

    NASA Technical Reports Server (NTRS)

    Miller, B. A.

    1977-01-01

    The effect of diffuser wall acoustic treatment on inlet total pressure loss was experimentally determined. Data were obtained by testing an inlet model with 10 different acoustically treated diffusers differing only in the design of the Helmholtz resonator acoustic treatment. Tests were conducted in a wind tunnel at forward velocities to 41 meters per second for inlet throat Mach numbers of .5 to .8 and angles of attack as high as 50 degrees. Results indicate a pressure loss penalty due to acoustic treatment that increases linearly with the porosity of the acoustic facing sheet. For a surface porosity of 14 percent the total pressure loss was 21 percent greater than that for an untreated inlet.

  6. Unsteady transonic flow calculations for wing-fuselage configurations

    NASA Technical Reports Server (NTRS)

    Batina, J. T.

    1986-01-01

    Unsteady transonic flow calculations are presented for wing-fuselage configurations. Calculations are performed by extending the XTRAN3S unsteady transonic small-disturbance code to allow the treatment of a fuselage. Details of the XTRAN3S fuselage modeling are discussed in the context of the small-disturbance equation. Transonic calculations are presented for three wing-fuselage configurations with leading edge sweep angles ranging from 0 deg to 46.76 deg. Simple bending and torsion modal oscillations of the wing are calculated. Sectional lift and moment coefficients for the wing-alone and wing-fuselage cases are compared and the effects of fuselage aerodynamic interference on the unsteady wing loading are revealed. Tabulated generalized aerodynamic forces used in flutter analyses, indicate small changes in the real in-phase component and as much as a 30% change in the imaginary component when the fuselage is included in the calculation. These changes result in a 2 to 5% increase in total magnitude and a several degree increase in phase.

  7. 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

  8. Ducted fan acoustic radiation including the effects of nonuniform mean flow and acoustic treatment

    NASA Technical Reports Server (NTRS)

    Eversman, Walter; Roy, Indranil Danda

    1993-01-01

    Forward and aft acoustic propagation and radiation from a ducted fan is modeled using a finite element discretization of the acoustic field equations. The fan noise source is introduced as equivalent body forces representing distributed blade loading. The flow in and around the nacelle is assumed to be nonuniform, reflecting the effects of forward flight and flow into the inlet. Refraction due to the fan exit jet shear layer is not represented. Acoustic treatment on the inlet and exhaust duct surfaces provides a mechanism for attenuation. In a region enclosing the fan a pressure formulation is used with the assumption of locally uniform flow. Away from the fan a velocity potential formulation is used and the flow is assumed nonuniform but irrotational. A procedure is developed for matching the two regions by making use of local duct modal amplitudes as transition state variables and determining the amplitudes by enforcing natural boundary conditions at the interface between adjacent regions in which pressure and velocity potential are used. Simple models of rotor alone and rotor/exit guide vane generated noise are used to demonstrate the calculation of the radiated acoustic field and to show the effect of acoustic treatment. The model has been used to assess the success of four techniques for acoustic lining optimization in reducing far field noise.

  9. Advanced technology commercial fuselage structure

    NASA Technical Reports Server (NTRS)

    Ilcewicz, L. B.; Smith, P. J.; Walker, T. H.; Johnson, R. W.

    1991-01-01

    Boeing's program for Advanced Technology Composite Aircraft Structure (ATCAS) has focused on the manufacturing and performance issues associated with a wide body commercial transport fuselage. The primary goal of ATCAS is to demonstrate cost and weight savings over a 1995 aluminum benchmark. A 31 foot section of fuselage directly behind the wing to body intersection was selected for study purposes. This paper summarizes ATCAS contract plans and review progress to date. The six year ATCAS program will study technical issues for crown, side, and keel areas of the fuselage. All structural details in these areas will be included in design studies that incorporate a design build team (DBT) approach. Manufacturing technologies will be developed for concepts deemed by the DBT to have the greatest potential for cost and weight savings. Assembly issues for large, stiff, quadrant panels will receive special attention. Supporting technologies and mechanical tests will concentrate on the major issues identified for fuselage. These include damage tolerance, pressure containment, splices, load redistribution, post-buckled structure, and durability/life. Progress to date includes DBT selection of baseline fuselage concepts; cost and weight comparisons for crown panel designs; initial panel fabrication for manufacturing and structural mechanics research; and toughened material studies related to keel panels. Initial ATCAS studies have shown that NASA's Advanced Composite Technology program goals for cost and weight savings are attainable for composite fuselage.

  10. Lightweight acoustic treatments for aerospace applications

    NASA Astrophysics Data System (ADS)

    Naify, Christina Jeanne

    2011-12-01

    Increase in the use of composites for aerospace applications has the benefit of decreased structural weight, but at the cost of decreased acoustic performance. Stiff, lightweight structures (such as composites) are traditionally not ideal for acoustic insulation applications because of high transmission loss at low frequencies. A need has thus arisen for effective sound insulation materials for aerospace and automotive applications with low weight addition. Current approaches, such as the addition of mass law dominated materials (foams) also perform poorly when scaled to small thickness and low density. In this dissertation, methods which reduce sound transmission without adding significant weight are investigated. The methods presented are intended to be integrated into currently used lightweight structures such as honeycomb sandwich panels and to cover a wide range of frequencies. Layering gasses of differing acoustic impedances on a panel substantially reduced the amount of sound energy transmitted through the panel with respect to the panel alone or an equivalent-thickness single species gas layer. The additional transmission loss derives from successive impedance mismatches at the interfaces between gas layers and the resulting inefficient energy transfer. Attachment of additional gas layers increased the transmission loss (TL) by as much as 17 dB at high (>1 kHz) frequencies. The location and ordering of the gasses with respect to the panel were important factors in determining the magnitude of the total TL. Theoretical analysis using a transfer matrix method was used to calculate the frequency dependence of sound transmission for the different configurations tested. The method accurately predicted the relative increases in TL observed with the addition of different gas layer configurations. To address low-frequency sound insulation, membrane-type locally resonant acoustic materials (LRAM) were fabricated, characterized, and analyzed to understand their

  11. 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... Fuselage doors. (a) General. This section applies to fuselage doors, which includes all doors, hatches... of tools to open or close. This also applies to each door or hatch through a pressure...

  12. 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... Fuselage doors. (a) General. This section applies to fuselage doors, which includes all doors, hatches... of tools to open or close. This also applies to each door or hatch through a pressure...

  13. Optimizing acoustical treatment. [structural design criteria for theater

    NASA Technical Reports Server (NTRS)

    Beuran, N.; Ramboiu, S.; Farcas, I.; Halpert, E.

    1974-01-01

    A mathematical linear programming model is presented for optimizing acoustical treatment and interior decoration of concert and other public halls. This method provides the designer with a range of acoustically correct solutions at increased economical efficiency. The mathematical model uses geometrical data about the room, recommended reverberation time values, the architect's choice of given sound absorbing structures and finishing materials. The model permits inclusion of aesthetical considerations about conditioning, proportioning, or, on the contrary, reciprocal exclusion of any classes of material and/or sound absorbing structure.

  14. Fuselage upwash effects on RSRA rotor systems

    NASA Technical Reports Server (NTRS)

    Cowan, J.; Dadone, L.

    1985-01-01

    The effects of RSRA fuselage configurations on rotor performance and loads have been quantified analytically by means of currently available potential flow and rotor analysis. Four configurations of the Rotor Systems Research Aircraft (RSRA) were considered in this study. They were: (1) fuselage alone (conventional helicopter); (2) fuselage with auxiliary propulsion; (3) fuselage with wings (auxiliary lift); and (4) fuselage with both auxiliary lift propulsion. The rotor system investigated was identical to a CH-47D front rotor except that it had four instead of three blades. Two scaled-down versions of the same rotor were also analyzed to determine the effect of rotor scale on the fuselage upwash effects. The flight conditions considered for the upwash study are discussed. The potential flow models for the RSRA configuration, with and without the wings and the auxiliary propulsion system, are presented. The results of fuselage/wing/propulsion system upwash on performance and loads are also presented.

  15. Laboratory test and acoustic analysis of cabin treatment for propfan test assessment aircraft

    NASA Technical Reports Server (NTRS)

    Kuntz, H. L.; Gatineau, R. J.

    1991-01-01

    An aircraft cabin acoustic enclosure, built in support of the Propfan Test Assessment (PTA) program, is described. Helmholtz resonators were attached to the cabin trim panels to increase the sidewall transmission loss (TL). Resonators (448) were located between the trim panels and fuselage shell. In addition, 152 resonators were placed between the enclosure and aircraft floors. The 600 resonators were each tuned to a 235 Hz resonance frequency. After flight testing on the PTA aircraft, the enclosure was tested in the Kelly Johnson R and D Center Acoustics Lab. Laboratory noise reduction (NR) test results are discussed. The enclosure was placed in a Gulfstream 2 fuselage section. Broadband (138 dB overall SPL) and tonal (149 dB overall SPL) excitations were used in the lab. Tonal excitation simulated the propfan flight test excitation. The fundamental tone was stepped in 2 Hz intervals from 225 through 245 Hz. The resonators increase the NR of the cabin walls around the resonance frequency of the resonator array. The effects of flanking, sidewall absorption, cabin adsorption, resonator loading of trim panels, and panel vibrations are presented. Increases in NR of up to 11 dB were measured.

  16. Acoustic hemostasis device for automated treatment of bleeding in limbs

    NASA Astrophysics Data System (ADS)

    Sekins, K. Michael; Zeng, Xiaozheng; Barnes, Stephen; Hopple, Jerry; Kook, John; Moreau-Gobard, Romain; Hsu, Stephen; Ahiekpor-Dravi, Alexis; Lee, Chi-Yin; Ramachandran, Suresh; Maleke, Caroline; Eaton, John; Wong, Keith; Keneman, Scott

    2012-10-01

    A research prototype automated image-guided acoustic hemostasis system for treatment of deep bleeding was developed and tested in limb phantoms. The system incorporated a flexible, conformal acoustic applicator cuff. Electronically steered and focused therapeutic arrays (Tx) populated the cuff to enable dosing from multiple Tx's simultaneously. Similarly, multiple imaging arrays (Ix) were deployed on the cuff to enable 3D compounded images for targeting and treatment monitoring. To affect a lightweight cuff, highly integrated Tx electrical circuitry was implemented, fabric and lightweight structural materials were used, and components were minimized. Novel cuff and Ix and Tx mechanical registration approaches were used to insure targeting accuracy. Two-step automation was implemented: 1) targeting (3D image volume acquisition and stitching, Power and Pulsed Wave Doppler automated bleeder detection, identification of bone, followed by closed-loop iterative Tx beam targeting), and 2) automated dosing (auto-selection of arrays and Tx dosing parameters, power initiation and then monitoring by acoustic thermometry for power shut-off). In final testing the device automatically detected 65% of all bleeders (with various bleeder flow rates). Accurate targeting was achieved in HIFU phantoms with end-dose (30 sec) temperature rise reaching the desired 33-58°C. Automated closed-loop targeting and treatment was demonstrated in separate phantoms.

  17. Fuselage panel noise attenuation by piezoelectric switching control

    NASA Astrophysics Data System (ADS)

    Makihara, Kanjuro; Miyakawa, Takeya; Onoda, Junjiro; Minesugi, Kenji

    2010-08-01

    This paper describes a problem that we encountered in our noise attenuation project and our solution for it. We intend to attenuate low-frequency noise that transmits through aircraft fuselage panels. Our method of noise attenuation is implemented with a piezoelectric semi-active system having a selective switch instead of an active energy-supply system. The semi-active controller is based on the predicted sound pressure distribution obtained from acoustic emission analysis. Experiments and numerical simulations demonstrate that the semi-active method attenuates acoustic levels of not only the simple monochromatic noise but also of broadband noise. We reveal that tuning the electrical parameters in the circuit is the key to effective noise attenuation, to overcome the acoustic excitation problem due to sharp switching actions, as well as to control chattering problems. The results obtained from this investigation provide meaningful insights into designing noise attenuation systems for comfortable aircraft cabin environments.

  18. 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.

  19. Acoustic Wave Treatment For Cellulite—A New Approach

    NASA Astrophysics Data System (ADS)

    Russe-Wilflingseder, Katharina; Russe, Elisabeth

    2010-05-01

    Background and Objectives: Cellulite is a biological caused modification of the female connective tissue. In extracorporeal shockwave therapy (ESWT) pulses are penetrating into the tissue without causing a thermal effect or micro lesions, but leading to a stimulation of tissue metabolism and blood circulation, inducing a natural repair process with cell activation and stem cells proliferation. Recently ESWT treatment showed evidence of remodelling collagen within the dermis and of stimulating microcirculation in fatty tissue. Study Design and Methods: The study was designed to assess acoustic wave treatment for cellulite by comparison treated vs. untreated side (upper-leg and buttock). Each individual served as its own control. 11 females with a BMI less then 30 and an age over 18 years were included. 6 treatments were given weekly with radial acoustic waves. Documentation was done before and 1, 4, 12 weeks after last treatment by standardized photo documentation, relaxed and with muscle contraction, measurement of body weight and circumference of the thigh, pinch test, and evaluation of hormonal status and lifestyle. The efficacy of AWT/EPAT was evaluated before and 1, 4, 12 weeks after last treatment. Patients rated the improvement of cellulite, overall satisfaction and acceptance. The therapist assessed improvement of cellulite, side effects and photo documentation treated vs. untreated side, before vs. after treatment. The blinded investigator evaluated the results using photo documentation right vs. left leg, before vs. after treatment in a frontal, lateral and dorsal view, relaxed and with muscle contraction. Results: The improvement of cellulite at the treated side was rated by patients with 27,3% at week 4 and 12, by the therapist with 34,1% at week 4 and 31,2% at week 12 after the last treatment The blinded investigator could verify an improvement of cellulite in an increasing number of patients with increasing time interval after treatment. No side

  20. Non-waisted fuselage design for supersonic aircraft

    NASA Technical Reports Server (NTRS)

    Hager, James O. (Inventor); Agrawal, Shreekant (Inventor); Antani, Dhamanshu L. (Inventor)

    1999-01-01

    A method for designing a non-waisted fuselage for supersonic wing/fuselage configurations that increases the fuselage volume and improves the supersonic aerodynamic performance compared to a conventional waisted-fuselage configuration. The method entails removing the waisted region of an existing waisted-fuselage configuration by linearly reconstructing cross-sections between the endpoints representing the waisted cross-sectional area portion to create a modified fuselage configuration without waisting. This configuration will have increased fuselage volume and improved supersonic aerodynamic performance. The fuselage camber can then be optimized using non-linear aerodynamic methods to further increase the supersonic aerodynamic performance.

  1. 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 STANDARDS: TRANSPORT CATEGORY AIRPLANES Design and Construction Personnel and Cargo Accommodations § 25.783 Fuselage doors. (a) General....

  2. Advanced Technology Composite Fuselage - Manufacturing

    NASA Technical Reports Server (NTRS)

    Wilden, K. S.; Harris, C. G.; Flynn, B. W.; Gessel, M. G.; Scholz, D. B.; Stawski, S.; Winston, V.

    1997-01-01

    The goal of Boeing's Advanced Technology Composite Aircraft Structures (ATCAS) program is to develop the technology required for cost-and weight-efficient use of composite materials in transport fuselage structure. 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 stringer-stiffened and sandwich skin panels. Circumferential 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 cocured to skin structures. Significant process development efforts included AFP, braiding, RTM, autoclave cure, and core blanket fabrication for both sandwich and stiffened-skin structure. Outer-mold-line and inner-mold-line tooling was developed for sandwich structures and stiffened-skin structure. The effect of design details, process control and tool design on repeatable, dimensionally stable, structure for low cost barrel assembly was assessed. Subcomponent panels representative of crown, keel, and side quadrant panels were fabricated to assess scale-up effects and manufacturing anomalies for full-scale structures. Manufacturing database including time studies, part quality, and manufacturing plans were generated to support the development of designs and analytical models to access cost, structural performance, and dimensional tolerance.

  3. Performance of fuselage pressure structure

    NASA Astrophysics Data System (ADS)

    Maclin, James R.

    1992-07-01

    There are currently more than 1,000 Boeing airplanes around the world over 20 years old. That number is expected to double by the year 1995. With these statistics comes the reality that structural airworthiness will be in the forefront of aviation issues well into the next century. The results of previous and recent test programs Boeing has implemented to study the structural performance of older airplanes relative to pressurized fuselage sections are described. Included in testing were flat panels with multiple site damage (MSD), a full-scale 737 and 2 747s as well as panels representing a 737 and 777, and a generic aircraft in large pressure-test fixtures. Because damage is a normal part of aging, focus is on the degree to which structural integrity is maintained after failure or partial failure of any structural element, including multiple site damage (MSD), and multiple element damage (MED).

  4. Acoustics

    NASA Astrophysics Data System (ADS)

    The acoustics research activities of the DLR fluid-mechanics department (Forschungsbereich Stroemungsmechanik) during 1988 are surveyed and illustrated with extensive diagrams, drawings, graphs, and photographs. Particular attention is given to studies of helicopter rotor noise (high-speed impulsive noise, blade/vortex interaction noise, and main/tail-rotor interaction noise), propeller noise (temperature, angle-of-attack, and nonuniform-flow effects), noise certification, and industrial acoustics (road-vehicle flow noise and airport noise-control installations).

  5. Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program

    NASA Technical Reports Server (NTRS)

    Salikuddin, M.; Kraft, R. E.; Syed, A. a.; Vu, D. D.; Mungur, P.; Langenbrunner, L. E.; Majjigi, R. K.

    2006-01-01

    The objectives of the initial effort (Phase I) of HSR Liner Technology Program, the selection of promising liner concepts, design and fabrication of these concepts for laboratory tests, testing these liners in the laboratory by using impedance tube and flow ducts, and developing empirical impedance/suppression correlation, are successfully completed. Acoustic and aerodynamic criteria for the liner design are established. Based on these criteria several liners are designed. The liner concepts designed and fabricated include Single-Degree-of-Freedom (SDOF), Two-Degree-of-Freedom (2DOF), and Bulk Absorber. Two types of SDOF treatment are fabricated, one with a perforated type face plate and the other with a wiremesh (woven) type faceplate. In addition, special configurations of these concepts are also included in the design. Several treatment panels are designed for parametric study. In these panels the facesheets of different porosity, hole diameter, and sheet thickness are utilized. Several deep panels (i.e., 1 in. deep) are designed and instrumented to measure DC flow resistance and insitu impedance in the presence of grazing flow. Basic components of these panels (i.e., facesheets, bulk materials, etc.) are also procured and tested. The results include DC flow resistance, normal impedance, and insertion loss.

  6. Astronaut Eileen Collins in Full Fuselage Trainer

    NASA Technical Reports Server (NTRS)

    1993-01-01

    Astronaut Eileen M. Collins, pilot for the STS-63 mission, participates in STS-63 training at JSC's Shuttle mockup and integration laboratory. Collins is seated at the pilot's station in the Full Fuselage Trainer (FFT).

  7. Acoustics

    NASA Technical Reports Server (NTRS)

    Goodman, Jerry R.; Grosveld, Ferdinand

    2007-01-01

    The acoustics environment in space operations is important to maintain at manageable levels so that the crewperson can remain safe, functional, effective, and reasonably comfortable. High acoustic levels can produce temporary or permanent hearing loss, or cause other physiological symptoms such as auditory pain, headaches, discomfort, strain in the vocal cords, or fatigue. Noise is defined as undesirable sound. Excessive noise may result in psychological effects such as irritability, inability to concentrate, decrease in productivity, annoyance, errors in judgment, and distraction. A noisy environment can also result in the inability to sleep, or sleep well. Elevated noise levels can affect the ability to communicate, understand what is being said, hear what is going on in the environment, degrade crew performance and operations, and create habitability concerns. Superfluous noise emissions can also create the inability to hear alarms or other important auditory cues such as an equipment malfunctioning. Recent space flight experience, evaluations of the requirements in crew habitable areas, and lessons learned (Goodman 2003; Allen and Goodman 2003; Pilkinton 2003; Grosveld et al. 2003) show the importance of maintaining an acceptable acoustics environment. This is best accomplished by having a high-quality set of limits/requirements early in the program, the "designing in" of acoustics in the development of hardware and systems, and by monitoring, testing and verifying the levels to ensure that they are acceptable.

  8. Enhancement of Focused Ultrasound Treatment by Acoustically Generated Microbubbles

    NASA Astrophysics Data System (ADS)

    Umemura, Shin-ichiro; Yoshizawa, Shin; Takagi, Ryo; Inaba, Yuta; Yasuda, Jun

    2013-07-01

    Microbubbles, whether introduced from outside the body or ultrasonically generated in situ, are known to significantly enhance the biological effects of ultrasound, including the mechanical, thermal, and sonochemical effects. Phase-change nanodroplets, which selectively accumulate in tumor tissue and whose phase changes to microbubbles can be induced by ultrasonic stimulation, have been proposed for high-intensity focused ultrasound (HIFU) tumor treatment with enhanced selectivity and efficiency. In this paper, a purely acoustic approach to generate microbubble clouds in the tissue to be treated is proposed. Short pulses of focused ultrasound with extremely high intensity, named trigger pulses, are used for exposure. They are immediately followed by focused ultrasound for heating with an intensity similar to or less than that of normal HIFU treatment. The localized generation of microbubble clouds by the trigger pulses is observed in a polyarylamide gel by a high-speed camera, and the effectiveness of the generated clouds in accelerating ultrasonically induced thermal coagulation is confirmed in excised chicken breast tissue. The use of second-harmonic superimposed waves as the trigger pulses is also proposed. The highly reproducible initiation of cavitation by waves with the negative peak pressure emphasized and the efficient expansion of the generated microbubble clouds by waves with the positive peak pressure emphasized are also observed by a high-speed camera in partially degassed water.

  9. The Acoustic Voice Quality Index: Toward Improved Treatment Outcomes Assessment in Voice Disorders

    ERIC Educational Resources Information Center

    Maryn, Youri; De Bodt, Marc; Roy, Nelson

    2010-01-01

    Voice practitioners require an objective index of dysphonia severity as a means to reliably track treatment outcomes. To ensure ecological validity however, such a measure should survey both sustained vowels and continuous speech. In an earlier study, a multivariate acoustic model referred to as the Acoustic Voice Quality Index (AVQI), consisting…

  10. 14 CFR 25.856 - Thermal/Acoustic insulation materials.

    Code of Federal Regulations, 2013 CFR

    2013-01-01

    ... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Thermal/Acoustic insulation materials. 25....856 Thermal/Acoustic insulation materials. (a) Thermal/acoustic insulation material installed in the.../acoustic insulation materials (including the means of fastening the materials to the fuselage) installed...

  11. 14 CFR 25.856 - Thermal/Acoustic insulation materials.

    Code of Federal Regulations, 2011 CFR

    2011-01-01

    ... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Thermal/Acoustic insulation materials. 25....856 Thermal/Acoustic insulation materials. (a) Thermal/acoustic insulation material installed in the.../acoustic insulation materials (including the means of fastening the materials to the fuselage) installed...

  12. 14 CFR 25.856 - Thermal/Acoustic insulation materials.

    Code of Federal Regulations, 2014 CFR

    2014-01-01

    ... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Thermal/Acoustic insulation materials. 25....856 Thermal/Acoustic insulation materials. (a) Thermal/acoustic insulation material installed in the.../acoustic insulation materials (including the means of fastening the materials to the fuselage) installed...

  13. 14 CFR 25.856 - Thermal/Acoustic insulation materials.

    Code of Federal Regulations, 2012 CFR

    2012-01-01

    ... 14 Aeronautics and Space 1 2012-01-01 2012-01-01 false Thermal/Acoustic insulation materials. 25....856 Thermal/Acoustic insulation materials. (a) Thermal/acoustic insulation material installed in the.../acoustic insulation materials (including the means of fastening the materials to the fuselage) installed...

  14. 14 CFR 25.856 - Thermal/Acoustic insulation materials.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Thermal/Acoustic insulation materials. 25....856 Thermal/Acoustic insulation materials. (a) Thermal/acoustic insulation material installed in the.../acoustic insulation materials (including the means of fastening the materials to the fuselage) installed...

  15. An electromagnetic finite difference time domain analog treatment of small signal acoustic interactions

    NASA Astrophysics Data System (ADS)

    Kunz, K.; Steich, D.; Lewis, K.; Landrum, C.; Barth, M.

    1994-03-01

    Hyperbolic partial differential equations encompass an extremely important set of physical phenomena including electromagnetics and acoustics. Small amplitude acoustic interactions behave much the same as electromagnetic interactions for longitudinal acoustic waves because of the similar nature of the governing hyperbolic equations. Differences appear when transverse acoustic waves are considered; nonetheless, the strong analogy between the acoustic and electromagnetic phenomena prompted the development of a Finite Difference Time Domain (FDTD) acoustic analog to the existing electromagnetic FDTD technique. The advantages of an acoustic FDTD (AFDTD) code are as follows: (1) boundary condition-free treatment of the acoustic scatterer--only the intrinsic properties of the scatterer's material are needed, no shell treatment or other set of special equations describing the macroscopic behavior of a sheet of material or a junction, etc. are required; this allows completely general geometries and materials in the model. (2) Advanced outer radiation boundary condition analogs--in the electromagnetics arena, highly absorbing outer radiation boundary conditions were developed that can be applied with little modification to the acoustics arena with equal success. (3) A suite of preexisting capabilities related to electromagnetic modeling--this includes automated model generation and interaction visualization as its most important components and is best exemplified by the capabilities of the LLNL generated TSAR electromagnetic FDTD code.

  16. A Requirements-Driven Optimization Method for Acoustic Treatment Design

    NASA Technical Reports Server (NTRS)

    Berton, Jeffrey J.

    2016-01-01

    Acoustic treatment designers have long been able to target specific noise sources inside turbofan engines. Facesheet porosity and cavity depth are key design variables of perforate-over-honeycomb liners that determine levels of noise suppression as well as the frequencies at which suppression occurs. Layers of these structures can be combined to create a robust attenuation spectrum that covers a wide range of frequencies. Looking to the future, rapidly-emerging additive manufacturing technologies are enabling new liners with multiple degrees of freedom, and new adaptive liners with variable impedance are showing promise. More than ever, there is greater flexibility and freedom in liner design. Subject to practical considerations, liner design variables may be manipulated to achieve a target attenuation spectrum. But characteristics of the ideal attenuation spectrum can be difficult to know. Many multidisciplinary system effects govern how engine noise sources contribute to community noise. Given a hardwall fan noise source to be suppressed, and using an analytical certification noise model to compute a community noise measure of merit, the optimal attenuation spectrum can be derived using multidisciplinary systems analysis methods. The subject of this paper is an analytical method that derives the ideal target attenuation spectrum that minimizes noise perceived by observers on the ground.

  17. Analytical prediction of the interior noise for cylindrical models of aircraft fuselages for prescribed exterior noise fields. Phase 2: Models for sidewall trim, stiffened structures and cabin acoustics with floor partition

    NASA Technical Reports Server (NTRS)

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

    1982-01-01

    An airplane interior noise prediction model is developed to determine the important parameters associated with sound transmission into the interiors of airplanes, and to identify apropriate noise control methods. Models for stiffened structures, and cabin acoustics with floor partition are developed. Validation studies are undertaken using three test articles: a ring stringer stiffened cylinder, an unstiffened cylinder with floor partition, and ring stringer stiffened cylinder with floor partition and sidewall trim. The noise reductions of the three test articles are computed using the heoretical models and compared to measured values. A statistical analysis of the comparison data indicates that there is no bias in the predictions although a substantial random error exists so that a discrepancy of more than five or six dB can be expected for about one out of three predictions.

  18. Full scale GLARE fuselage panel tests

    NASA Technical Reports Server (NTRS)

    Vercammen, Roland W. A.; Ottens, Harold H.

    1996-01-01

    A GLARE fuselage panel, representative of the crown section of the Fokker 100 fuselage in front of the wing, was tested. The panels were loaded by air pressure resulting in tangential stress in the panel by axial loading, representative of both the cabin pressure and the fuselage bending due to taxiing and gust loading. A fatigue test, simulating 180000 flights, followed by static tests were performed. The panel was loaded to failure at 1.32 ultimate load. The test set-up, the uniform strain distribution of the panel, and the fatigue loads applied at high test frequency are described. The use of GLARE leads to a substantial weight reduction without affecting the fatigue static strength.

  19. On the Strength of Box Type Fuselages

    NASA Technical Reports Server (NTRS)

    Mathar, J

    1929-01-01

    The present investigation relates to a box-type fuselage with sides consisting of thin smooth sheet metal, stiffened by longitudinal members riveted to the flanged channel-section bulkheads or transverse frames and to the semicircular corrugated corner stiffenings. The results obtained in this particular case can be applied to a great number of similar structures.

  20. Astronaut Eileen Collins in Full Fuselage Trainer

    NASA Technical Reports Server (NTRS)

    1993-01-01

    Astronaut Eileen M. Collins, pilot for the STS-63 mission, participates in STS-63 training at JSC's Shuttle mockup and integration laboratory. Collins is seated at the pilot's station in the Full Fuselage Trainer (FFT) (48403-4); Collins looks out the aft flight deck window in the Shuttle mockup trainer (48405).

  1. The Aerodynamic Aspect of Wing-fuselage Fillets

    NASA Technical Reports Server (NTRS)

    Muttray, H

    1935-01-01

    Model tests prove the feasibility of enhancing the aerodynamic qualities of wing-fuselage fillets by appropriate design of fuselage and wing roots. Abrupt changes from maximum fuselage height to wing chord must be avoided and every longitudinal section of fuselage and wing roots must be so faired and arranged as to preserve the original lift distribution of the continuous wing. Adapting the fuselage to the curvilinear circulation of the wing affords further improvement. The polars of such arrangements are almost the same as those of the "wing alone," thus voiding the superiority of the high-wing type airplane known with conventional design.

  2. 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.

  3. 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.

  4. PRSEUS Acoustic Panel Fabrication

    NASA Technical Reports Server (NTRS)

    Nicolette, Velicki; Yovanof, Nicolette P.; Baraja, Jaime; Mathur, Gopal; Thrash, Patrick; 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 with regards to acoustic response. A PRSEUS panel was designed and fabricated and provided to NASA-LaRC for acoustic response testing in the Structural Acoustics Loads and Transmission (SALT) facility). Preliminary assessments of the sound transmission characteristics of a PRSEUS panel subjected to a representative Hybrid Wing Body (HWB) operating environment were completed for the NASA Environmentally Responsible Aviation (ERA) Program.

  5. Acoustic and vibration response of a structure with added noise control treatment under various excitations.

    PubMed

    Rhazi, Dilal; Atalla, Noureddine

    2014-02-01

    The evaluation of the acoustic performance of noise control treatments is of great importance in many engineering applications, e.g., aircraft, automotive, and building acoustics applications. Numerical methods such as finite- and boundary elements allow for the study of complex structures with added noise control treatment. However, these methods are computationally expensive when used for complex structures. At an early stage of the acoustic trim design process, many industries look for simple and easy to use tools that provide sufficient physical insight that can help to formulate design criteria. The paper presents a simple and tractable approach for the acoustic design of noise control treatments. It presents and compares two transfer matrix-based methods to investigate the vibroacoustic behavior of noise control treatments. The first is based on a modal approach, while the second is based on wave-number space decomposition. In addition to the classical rain-on-the-roof and diffuse acoustic field excitations, the paper also addresses turbulent boundary layer and point source (monopole) excitations. Various examples are presented and compared to a finite element calculation to validate the methodology and to confirm its relevance along with its limitations. PMID:25234878

  6. 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.

  7. 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.

  8. 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.

  9. Composite fuselage crown panel manufacturing technology

    NASA Technical Reports Server (NTRS)

    Willden, Kurtis; Metschan, S.; Grant, C.; Brown, T.

    1992-01-01

    Commercial fuselage structures contain significant challenges in attempting to save manufacturing costs with advanced composite technology. Assembly issues, material costs, and fabrication of elements with complex geometry are each expected to drive the cost of composite fuselage structures. Boeing's efforts under the NASA ACT program have pursued key technologies for low-cost, large crown panel fabrication. An intricate bond panel design and manufacturing concepts were selected based on the efforts of the Design Build Team (DBT). The manufacturing processes selected for the intricate bond design include multiple large panel fabrication with the Advanced Tow Placement (ATP) process, innovative cure tooling concepts, resin transfer molding of long fuselage frames, and utilization of low-cost material forms. The process optimization for final design/manufacturing configuration included factory simulations and hardware demonstrations. These efforts and other optimization tasks were instrumental in reducing cost by 18 percent and weight by 45 percent relative to an aluminum baseline. The qualitative and quantitative results of the manufacturing demonstrations were used to assess manufacturing risks and technology readiness.

  10. Advanced Technology Composite Fuselage: Program Overview

    NASA Technical Reports Server (NTRS)

    Ilcewicz, L. B.; Smith, P. J.; Hanson, C. T.; Walker, T. H.; Metschan, S. L.; Mabson, G. E.; Wilden, K. S.; Flynn, B. W.; Scholz, D. B.; Polland, D. R.; Fredrikson, H. G.; Olson, J. T.; Backman, B. F.

    1997-01-01

    The Advanced Technology Composite Aircraft Structures (ATCAS) program has studied transport fuselage structure with a large potential reduction in the total direct operating costs for wide-body commercial transports. The baseline fuselage section was divided into four 'quadrants', crown, keel, and sides, gaining the manufacturing cost advantage possible with larger panels. Key processes found to have savings potential include (1) skins laminated by automatic fiber placement, (2) braided frames using resin transfer molding, and (3) panel bond technology that minimized mechanical fastening. The cost and weight of the baseline fuselage barrel was updated to complete Phase B of the program. An assessment of the former, which included labor, material, and tooling costs, was performed with the help of design cost models. Crown, keel, and side quadrant cost distributions illustrate the importance of panel design configuration, area, and other structural details. Composite sandwich panel designs were found to have the greatest cost savings potential for most quadrants. Key technical findings are summarized as an introduction to the other contractor reports documenting Phase A and B work completed in functional areas. The current program status in resolving critical technical issues is also highlighted.

  11. Acoustic testing of a supersonic tip speed fan with acoustic treatment and rotor casting slots. Quiet engine program scale model fan C

    NASA Technical Reports Server (NTRS)

    Kazin, S. B.

    1973-01-01

    Acoustic tests were conducted on a high tip speed (1550 ft/sec, 472.44 m/sec) single stage fan with varying amounts of wall acoustic treatment and with circumferential slots over the rotor blade tips. The slots were also tested with acoustic treatment placed behind the slots. The wall treatment results show that the inlet treatment is more effective at high fan speeds and aft duct treatment is more effective at low fan speeds. Maximum PNL's on a 200-foot (60.96 m) sideline show the untreated slots to have increased the rear radiated noise at approach. However, when the treatment was added to the slots inlet radiated noise was decreased, resulting in little change relative to the solid casing on an EPNL basis.

  12. Drag of Exposed Fittings and Surface Irregularities on Airplane Fuselages

    NASA Technical Reports Server (NTRS)

    Wood, Donald H

    1928-01-01

    Measurements of drag were made on fittings taken from a typical fuselage to determine whether the difference between the observed full size fuselage drag and model fuselage drag could be attributed to the effects of fittings and surface irregularities found on the full size fuselage and not on the model. There are wide variations in the drag coefficients for the different fittings. In general those which protrude little from the surface or are well streamlined show very low and almost negligible drag. The measurements show, however, that a large part of the difference between model and full scale test results may be attributed to these fittings.

  13. An unsteady helicopter rotor-fuselage aerodynamic interaction analysis

    NASA Technical Reports Server (NTRS)

    Lorber, Peter F.; Egolf, T. Alan

    1990-01-01

    A computational method has been developed to treat the unsteady aerodynamic interaction between a helicopter rotor, wake, and fuselage. Two existing codes, a lifting line-prescribed wake rotor analysis and a source panel fuselage analysis, were modified and coupled to allow prediction of unsteady fuselage pressures and airloads. A prescribed displacement technique was developed to position the rotor wake about the fuselage. Also coupled into the method were optional blade dynamics or rigid blade performance analyses to set the rotor operating conditions. Sensitivity studies were performed to determine the influence of the wake and fuselage geometry on the computational results. Solutions were computed for an ellipsoidal fuselage and a four bladed rotor at several advance ratios, using both the classical helix and the generalized distorted wake model. Results are presented that describe the induced velocities, pressures, and airloads on the fuselage and the induced velocities and bound circulation at the rotor. The ability to treat arbitrary geometries was demonstrated using a simulated helicopter fuselage. Initial computations were made to simulate the geometry of an experimental rotor-fuselage interaction study performed at the Georgia Institute of Technology.

  14. Acoustic Treatment Design Scaling Methods. Volume 2; Advanced Treatment Impedance Models for High Frequency Ranges

    NASA Technical Reports Server (NTRS)

    Kraft, R. E.; Yu, J.; Kwan, H. W.

    1999-01-01

    The primary purpose of this study is to develop improved models for the acoustic impedance of treatment panels at high frequencies, for application to subscale treatment designs. Effects that cause significant deviation of the impedance from simple geometric scaling are examined in detail, an improved high-frequency impedance model is developed, and the improved model is correlated with high-frequency impedance measurements. Only single-degree-of-freedom honeycomb sandwich resonator panels with either perforated sheet or "linear" wiremesh faceplates are considered. The objective is to understand those effects that cause the simple single-degree-of- freedom resonator panels to deviate at the higher-scaled frequency from the impedance that would be obtained at the corresponding full-scale frequency. This will allow the subscale panel to be designed to achieve a specified impedance spectrum over at least a limited range of frequencies. An advanced impedance prediction model has been developed that accounts for some of the known effects at high frequency that have previously been ignored as a small source of error for full-scale frequency ranges.

  15. Acoustic Treatment Design Scaling Methods. Volume 3; Test Plans, Hardware, Results, and Evaluation

    NASA Technical Reports Server (NTRS)

    Yu, J.; Kwan, H. W.; Echternach, D. K.; Kraft, R. E.; Syed, A. A.

    1999-01-01

    The ability to design, build, and test miniaturized acoustic treatment panels on scale-model fan rigs representative of the full-scale engine provides not only a cost-savings, but an opportunity to optimize the treatment by allowing tests of different designs. To be able to use scale model treatment as a full-scale design tool, it is necessary that the designer be able to reliably translate the scale model design and performance to an equivalent full-scale design. The primary objective of the study presented in this volume of the final report was to conduct laboratory tests to evaluate liner acoustic properties and validate advanced treatment impedance models. These laboratory tests include DC flow resistance measurements, normal incidence impedance measurements, DC flow and impedance measurements in the presence of grazing flow, and in-duct liner attenuation as well as modal measurements. Test panels were fabricated at three different scale factors (i.e., full-scale, half-scale, and one-fifth scale) to support laboratory acoustic testing. The panel configurations include single-degree-of-freedom (SDOF) perforated sandwich panels, SDOF linear (wire mesh) liners, and double-degree-of-freedom (DDOF) linear acoustic panels.

  16. Study of utilization of advanced composites in fuselage structures of large transports

    NASA Technical Reports Server (NTRS)

    Jackson, A. C.; Campion, M. C.; Pei, G.

    1984-01-01

    The effort required by the transport aircraft manufacturers to support the introduction of advanced composite materials into the fuselage structure of future commercial and military transport aircraft is investigated. Technology issues, potential benefits to military life cycle costs and commercial operating costs, and development plans are examined. The most urgent technology issues defined are impact dynamics, acoustic transmission, pressure containment and damage tolerance, post-buckling, cutouts, and joints and splices. A technology demonstration program is defined and a rough cost and schedule identified. The fabrication and test of a full-scale fuselage barrel section is presented. Commercial and military benefits are identified. Fuselage structure weight savings from use of advanced composites are 16.4 percent for the commercial and 21.8 percent for the military. For the all-composite airplanes the savings are 26 percent and 29 percent, respectively. Commercial/operating costs are reduced by 5 percent for the all-composite airplane and military life cycle costs by 10 percent.

  17. SU-E-T-536: Inhomogeneity Correction in Planning of Gamma Knife Treatments for Acoustic Schwannoma

    SciTech Connect

    Lu, L; Gupta, N; Hessler, J; Liu, A; Weldon, M; McGregor, J; Ammirati, M; Guiou, M; Xia, F; Grecula, J

    2014-06-01

    Purpose: To find out the dose difference on targets and organs at risk for the treatment of acoustic schwannoma if the inhomogeneity correction (Convolution algorithm) is applied. Methods: Images of patients treated for acoustic schwannoma with Gamma Knife using TMR 10 algorithm were retrieved from database and replanned with Convolution and TMR 10 algorithm respectively. These patients were treated using a preplan scheme in following: (1) Before the actual treatment day, using the MRI image that was taken without a head frame on the patient's skull, a pre-treatment plan was made based on the default skull coordinates in the Gamma Knife treatment planning system (LGP); (2) then on treatment day, a head frame was placed on the patient's skull, and a CT image was taken. The CT image with head frame was registered and fused with the completed preplan; (3) the treatment plan was finalized and the treatment was delivered. To find out the dosimetry impact of inhomogeneity correction, we used the retrieved CT images to replan the treatment using Convolution algorithm in LGP software version 10.1.1. The dose distributions and the dose volume histograms for targets and OARs were compared for these two dose calculation algorithms. Results: The dose calculated with the Convolution algorithm in general is slightly lower than the one from TMR 10 around the boney area. The effect from the inhomogeneity correction is observable but not significant, and varies with the location of the tumor. Conclusion: Inhomogeneity correction slightly improve the dose accuracy for acoustic schwannoma Gamma Knife treatments although the correction may not be very significant. Our Result provides evidence for dose prescription adjustment to treat acoustic schwannoma. The actual clinical outcome of switching from using TMR10 to using Convolution needs to be further investigated.

  18. Quiet Clean Short-Haul Experimental Engine (QCSEE): Acoustic treatment development and design

    NASA Technical Reports Server (NTRS)

    Clemons, A.

    1979-01-01

    Acoustic treatment designs for the quiet clean short-haul experimental engines are defined. The procedures used in the development of each noise-source suppressor device are presented and discussed in detail. A complete description of all treatment concepts considered and the test facilities utilized in obtaining background data used in treatment development are also described. Additional supporting investigations that are complementary to the treatment development work are presented. The expected suppression results for each treatment configuration are given in terms of delta SPL versus frequency and in terms of delta PNdB.

  19. Evaluation of Isolated Fuselage and Rotor-Fuselage Interaction Using CFD

    NASA Technical Reports Server (NTRS)

    Renaud, Thomas; O'Brien, David; Smith, Marilyn; Potsdam, Mark

    2004-01-01

    The US Army Aeroflightdynamics Directorate (AFDD), the French Office National d'Etudes et de Recherches Aerospatiales (ONERA) and the Georgia Institute of Technology (GIT) are working under the United States/France Memorandum of Agreement on Helicopter Aeromechanics to study rotorcraft aeromechanics issues of interest to both nations. As a task under this agreement, a comparative study of the Dauphin 365N helicopter has been undertaken to analyze the capabilities and weaknesses of state-of-the-art computational fluid dynamics (CFD) codes, with the aim of fuselage performance prediction and investigation of rotor-fuselage interaction. Three CFD flow solvers applied on three meshes provide similar results in terms of pressure coefficient. Force predictions vary somewhat. This paper presents details on the grid sensitivity and the low Mach number preconditioning influence. The importance of taking into account the wind tunnel strut and the rotor hub is shown. The pressure coefficients along top and bottom centerlines of the fuselage are in good agreement with the experiment except in the area aft of the hub. There remains a discrepancy between the computed forces and the experimental data due in part to modeling inaccuracies. Rotor- fuselage interactions are performed using uniform and non-uniform actuator disk models in order to simulate the rotor downwash.

  20. Acoustic treatment of the NASA Langley 4- by 7-meter tunnel: A feasibility study

    NASA Technical Reports Server (NTRS)

    Yu, J. C.; Abrahamson, A. L.

    1986-01-01

    A feasibility study for upgrading the NASA Langley 4- by 7-Meter Tunnel so that it may be used for aeroacoustic research related to helicopters is described. The requirements for noise research leading to the design of the next generation of helicopters impose a set of acoustic test criteria that no existing wind tunnel in the United States can presently meet. Included in this feasibility study are the following considerations: (1) an evaluation of general wind-tunnel requirements and desired tunnel background noise levels for helicopter aeroacoustic research; (2) an assessment of the present acoustic environment for testing model rotors; (3) a diagnostic investigation of tunnel background noise sources and paths; (4) acoustic treatment options for tunnel background noise reduction and a trade-off study between these options; (5) an engineering feasibility assessment of the selected option; and (6) an integrated analysis of study components and recommendations of treatment for an approach to meet the tunnel background noise reduction goal. It is concluded that the Langley 4- by 7-Meter Tunnel is a fundamentally suitable facility for helicopter aeroacoustic research. It is also concluded that acoustic treatment of this facility for meeting the required tunnel background noise goal can be accomplished technically at reasonable risk and cost.

  1. Advanced fiber placement of composite fuselage structures

    NASA Technical Reports Server (NTRS)

    Anderson, Robert L.; Grant, Carroll G.

    1991-01-01

    The Hercules/NASA Advanced Composite Technology (ACT) program will demonstrate the low cost potential of the automated fiber placement process. The Hercules fiber placement machine was developed for cost effective production of composite aircraft structures. The process uses a low cost prepreg tow material form and achieves equivalent laminate properties to structures fabricated with prepreg tape layup. Fiber placement demonstrations planned for the Hercules/NASA program include fabrication of stiffened test panels which represent crown, keel, and window belt segments of a typical transport aircraft fuselage.

  2. Voice acoustic measures of depression severity and treatment response collected via interactive voice response (IVR) technology

    PubMed Central

    Mundt, James C.; Snyder, Peter J.; Cannizzaro, Michael S.; Chappie, Kara; Geralts, Dayna S.

    2011-01-01

    Efforts to develop more effective depression treatments are limited by assessment methods that rely on patient-reported or clinician judgments of symptom severity. Depression also affects speech. Research suggests several objective voice acoustic measures affected by depression can be obtained reliably over the telephone. Thirty-five physician-referred patients beginning treatment for depression were assessed weekly, using standard depression severity measures, during a six-week observational study. Speech samples were also obtained over the telephone each week using an IVR system to automate data collection. Several voice acoustic measures correlated significantly with depression severity. Patients responding to treatment had significantly greater pitch variability, paused less while speaking, and spoke faster than at baseline. Patients not responding to treatment did not show similar changes. Telephone standardization for obtaining voice data was identified as a critical factor influencing the reliability and quality of speech data. This study replicates and extends previous research with a larger sample of patients assessing clinical change associated with treatment. The feasibility of obtaining voice acoustic measures reflecting depression severity and response to treatment using computer-automated telephone data collection techniques is also established. Insight and guidance for future research needs are also identified. PMID:21253440

  3. 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

  4. Effects of long-chord acoustically treated stator vanes on fan noise. 2: Effect of acoustical treatment

    NASA Technical Reports Server (NTRS)

    Dittmar, J. H.; Scott, J. N.; Leonard, B. R.; Stakolich, E. G.

    1976-01-01

    A set of long chord stator vanes was designed to replace the vanes in an existing fan stage. The long chord stator vanes consisted of a turning section and axial extension pieces, all of which incorporated acoustic damping material. The long chord stator vanes were tested in two lengths, with the long version giving more noise reduction than the short, primarily because of the additional lining material. The noise reduction achieved with the acoustically treated long chord stator vanes was compared with the reduction achieved by an acoustically treated exhaust splitter. The long chord stator was at least as good as the splitter as a method for incorporating acoustic lining material. In addition, comparing an acoustic three ring inlet and an acoustic wall-only inlet discloses that the wall-only inlet could be used in an engine where the noise reduction requirements are not too stringent.

  5. Vibration Damping Via Acoustic Treatment Attached To Vehicle Body Panels

    NASA Astrophysics Data System (ADS)

    Gambino, Carlo

    Currently, in the automotive industry, the control of noise and vibration is the subject of much research, oriented towards the creation of innovative solutions to improve the comfort of the vehicle and to reduce its cost and weight. This thesis fits into this particular framework, as it aims to investigate the possibility of integrating the functions of sound absorptioninsulation and vibration damping in a unique component. At present the bituminous viscoelastic treatments, which are bonded to the car body panels, take charge of the vibration damping, while the sound absorption and insulation is obtained by means of the poroacoustic treatments. The solution proposed here consists of employing porous materials to perform both these functions, thus allowing the partial or complete removal of the viscoelastic damping treatments from the car body. This should decrease the weight of the vehicle, reducing fuel consumption and emissions, and it might also benefit production costs.

  6. Acoustic Treatment Design Scaling Methods. Volume 1; Overview, Results, and Recommendations

    NASA Technical Reports Server (NTRS)

    Kraft, R. E.; Yu, J.

    1999-01-01

    Scale model fan rigs that simulate new generation ultra-high-bypass engines at about 1/5-scale are achieving increased importance as development vehicles for the design of low-noise aircraft engines. Testing at small scale allows the tests to be performed in existing anechoic wind tunnels, which provides an accurate simulation of the important effects of aircraft forward motion on the noise generation. The ability to design, build, and test miniaturized acoustic treatment panels on scale model fan rigs representative of the fullscale engine provides not only a cost-savings, but an opportunity to optimize the treatment by allowing tests of different designs. The primary objective of this study was to develop methods that will allow scale model fan rigs to be successfully used as acoustic treatment design tools. The study focuses on finding methods to extend the upper limit of the frequency range of impedance prediction models and acoustic impedance measurement methods for subscale treatment liner designs, and confirm the predictions by correlation with measured data. This phase of the program had as a goal doubling the upper limit of impedance measurement from 6 kHz to 12 kHz. The program utilizes combined analytical and experimental methods to achieve the objectives.

  7. Changes in acoustic emission peaks in precipitation strengthened alloys with heat treatment

    SciTech Connect

    Heiple, C.R.; Carpenter, S.H.

    1983-01-01

    Acoustic emission was measured during tensile deformation in a number of precipitation-strengthened alloys as a function of prior heat treatment. The alloys tested included 7075, 6061, and 2219 aluminum; a modified A-286 stainless steel (JBK-75) and an experimental beryllium-containing stainless steel; and Incoloy 903. A rms voltage peak was observed in all the alloys near the onset of plastic flow, and a second peak was usually observed in 7075, 2219, and Incoloy 903 at plastic strains greater than 1%. Some evidence of a second peak was also observed in 6061 aluminum. Changes with heat treatment in the stress and strain at which the second peak occurred were consistent with the peak arising from the fracture of inclusions. The shifts in the location of the peak were in a direction so as to make the stress on the inclusions at the second peak relatively insensitive to prior heat treatment. The amplitude distributions of acoustic emission signals were also consistent with this interpretation. The strain at which the first acoustic emission peak occurred also varied with heat treatment, but the dependence of peak location on prior aging was different for the various alloys.

  8. 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.

  9. 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.

  10. Design-Oriented Analysis of Aircraft Fuselage Structures

    NASA Technical Reports Server (NTRS)

    Giles, Gary L.

    1998-01-01

    A design-oriented analysis capability for aircraft fuselage structures that utilizes equivalent plate methodology is described. This new capability is implemented as an addition to the existing wing analysis procedure in the Equivalent Laminated Plate Solution (ELAPS) computer code. The wing and fuselage analyses are combined to model entire airframes. The paper focuses on the fuselage model definition, the associated analytical formulation and the approach used to couple the wing and fuselage analyses. The modeling approach used to minimize the amount of preparation of input data by the user and to facilitate the making of design changes is described. The fuselage analysis is based on ring and shell equations but the procedure is formulated to be analogous to that used for plates in order to take advantage of the existing code in ELAPS. Connector springs are used to couple the wing and fuselage models. Typical fuselage analysis results are presented for two analytical models. Results for a ring-stiffened cylinder model are compared with results from conventional finite-element analyses to assess the accuracy of this new analysis capability. The connection of plate and ring segments is demonstrated using a second model that is representative of the wing structure for a channel-wing aircraft configuration.

  11. 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.

  12. Aero Commander in flight - Upswept fuselage study

    NASA Technical Reports Server (NTRS)

    1965-01-01

    The NASA Flight Research Center's Aero Commander 680F is shown in flight with tufts attached to the side and bottom sections of the aircraft. These were placed on the aircraft for a Upswept Fuselage Study to see if the flow separated on the aft section of a small aircraft for comparison of data acquired from a large cargo-type aircraft with an upswept aft section. The photo of the tufts demonstrates that the flow is attached with no turbulence present. (Note the straight lines of tufts). The Aero Commander was used both for support and as a research aircraft. Among other uses, it was flown to outlying dry lakebeds, used as emergency landing sites, before X-15 flights. It could reach the lakebeds quickly and land on the hard-packed surfaces to ensure they were not soft from rainfall or some other cause. Between 1964 and 1966, the Flight Research Center used the aircraft in the Aviation Safety and Operating Problems Program to evaluate the aerodynamics of various light aircraft and to define possible technological improvements. The Aero Commander left what had become the Dryden Flight Research Center on March 14, 1979, and was transferred to the Customs Air Branch in San Diego. The Aero Commander 680F (N6297), built by the Aero Commander Company of Bethany, Oklahoma, is a pressurized five-place aircraft that is powered by two 380-horsepower reciprocating engines built by Lycoming Company. The fuselage length is 24.2 feet with a wing span of 35.98 feet.

  13. Composite fuselage shell structures research at NASA Langley Research Center

    NASA Technical Reports Server (NTRS)

    Starnes, James H., Jr.; Shuart, Mark J.

    1992-01-01

    Fuselage structures for transport aircraft represent a significant percentage of both the weight and the cost of these aircraft primary structures. Composite materials offer the potential for reducing both the weight and the cost of transport fuselage structures, but only limited studies of the response and failure of composite fuselage structures have been conducted for transport aircraft. The behavior of these important primary structures must be understood, and the structural mechanics methodology for analyzing and designing these complex stiffened shell structures must be validated in the laboratory. The effects of local gradients and discontinuities on fuselage shell behavior and the effects of local damage on pressure containment must be thoroughly understood before composite fuselage structures can be used for commercial aircraft. This paper describes the research being conducted and planned at NASA LaRC to help understand the critical behavior or composite fuselage structures and to validate the structural mechanics methodology being developed for stiffened composite fuselage shell structure subjected to combined internal pressure and mechanical loads. Stiffened shell and curved stiffened panel designs are currently being developed and analyzed, and these designs will be fabricated and then tested at Langley to study critical fuselage shell behavior and to validate structural analysis and design methodology. The research includes studies of the effects of combined internal pressure and mechanical loads on nonlinear stiffened panel and shell behavior, the effects of cutouts and other gradient-producing discontinuities on composite shell response, and the effects of local damage on pressure containment and residual strength. Scaling laws are being developed that relate full-scale and subscale behavior of composite fuselage shells. Failure mechanisms are being identified and advanced designs will be developed based on what is learned from early results from

  14. 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.

  15. Numerical computation of transonic flow about wing-fuselage configurations on a vector computer

    NASA Technical Reports Server (NTRS)

    Thomas, S. D.; Holst, T. L.

    1983-01-01

    The transonic wing analysis code TWING, which uses the AF2 relaxation algorithm, has been vectorized to run on the Cray-1S computer. Vectorization of this code improved computational efficiency over that of the CDC 7600 computer by factors of 11 to 13. The improvement compares favorably with the prediction of a theoretical performance model. A convenient generalization now permits the treatment of rudimentary wing-fuselage combinations. Flow predictions for a transport configuration in both isolated-wing and wing-fuselage modes show the expected trends in shock strength and position when compared with wind-tunnel results. An isolated fighter wing is examined in terms of execution time on three different computers and in comparison with experimental data. The computational fluid dynamics code produced during this study is a careful union of an efficient three-dimensional, transonic, numerical algorithm and the vector features presently available on modern computers.

  16. Efficient modeling of flat and homogeneous acoustic treatments for vibroacoustic finite element analysis. Direct field formulations

    NASA Astrophysics Data System (ADS)

    Alimonti, L.; Atalla, N.

    2016-04-01

    This paper is concerned with the development of a simplified model for noise control treatments to speed up finite element analysis in vibroacoustic applications. The methodology relies on the assumption that the acoustic treatment is flat and homogeneous. Moreover, its finite lateral extent is neglected. This hypothesis is justified by short wavelength and large dissipation, which suggest that the reflected field emanating from the acoustic treatment lateral boundaries does not substantially affect its dynamic response. Under these circumstances, the response of the noise control treatment can be formally obtained by means of convolution integrals involving simple analytical kernels (i.e. Green functions). Such fundamental solutions can be computed efficiently by the transfer matrix method. However, some arbitrariness arises in the formulation of the mathematical model, resulting in different baffling conditions at the two ends of the treatment to be considered. Thus, the paper investigates the possibility of different formulations (i.e. baffling conditions) within the same hybrid finite element-transfer matrix framework, seeking for the best strategy in terms of tradeoff between efficiency and accuracy. Numerical examples are provided to show strengths and limitations of the proposed methodology.

  17. 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.

  18. 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.

  19. Comparison between design and installed acoustic characteristics of NASA Lewis 9- by 15-foot low-speed wind tunnel acoustic treatment

    NASA Technical Reports Server (NTRS)

    Dahl, Milo D.; Woodward, Richard P.

    1990-01-01

    The test section of the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel was acoustically treated to allow the measurement of sound under simulated free-field conditions. The treatment was designed for high sound absorption at frequencies above 250 Hz and for withstanding the environmental conditions in the test section. In order to achieve the design requirements, a fibrous, bulk-absorber material was packed into removable panel sections. Each section was divided into two equal-depth layers packed with material to different bulk densities. The lower density was next to the facing of the treatment. The facing consisted of a perforated plate and screening material layered together. Sample tests for normal-incidence acoustic absorption were also conducted in an impedance tube to provide data to aid in the treatment design. Tests with no airflow, involving the measurement of the absorptive properties of the treatment installed in the 9- by 15-foot wind tunnel test section, combined the use of time-delay spectrometry with a previously established free-field measurement method. This new application of time-delay spectrometry enabled these free-field measurements to be made in nonanechoic conditions. The results showed that the installed acoustic treatment had absorption coefficients greater than 0.95 over the frequency range 250 Hz to 4 kHz. The measurements in the wind tunnel were in good agreement with both the analytical prediction and the impedance tube test data.

  20. Acoustically accessible window determination for ultrasound mediated treatment of glycogen storage disease type Ia patients

    NASA Astrophysics Data System (ADS)

    Wang, Shutao; Raju, Balasundar I.; Leyvi, Evgeniy; Weinstein, David A.; Seip, Ralf

    2012-10-01

    Glycogen storage disease type Ia (GSDIa) is caused by an inherited single-gene defect resulting in an impaired glycogen to glucose conversion pathway. Targeted ultrasound mediated delivery (USMD) of plasmid DNA (pDNA) to liver in conjunction with microbubbles may provide a potential treatment for GSDIa patients. As the success of USMD treatments is largely dependent on the accessibility of the targeted tissue by the focused ultrasound beam, this study presents a quantitative approach to determine the acoustically accessible liver volume in GSDIa patients. Models of focused ultrasound beam profiles for transducers of varying aperture and focal lengths were applied to abdomen models reconstructed from suitable CT and MRI images. Transducer manipulations (simulating USMD treatment procedures) were implemented via transducer translations and rotations with the intent of targeting and exposing the entire liver to ultrasound. Results indicate that acoustically accessible liver volumes can be as large as 50% of the entire liver volume for GSDIa patients and on average 3 times larger compared to a healthy adult group due to GSDIa patients' increased liver size. Detailed descriptions of the evaluation algorithm, transducer-and abdomen models are presented, together with implications for USMD treatments of GSDIa patients and transducer designs for USMD applications.

  1. 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.

  2. 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.

  3. Crack curving in a ductile pressurized fuselage

    NASA Astrophysics Data System (ADS)

    Lam, Paul W.

    Moire interferometry was used to study crack tip displacement fields of a biaxially loaded cruciform type 0.8mm thick 2024-T3 aluminum specimen with various tearstrap reinforcement configurations: Unreinforced, Bonded, Bonded+Riveted, and Machined Pad-up. A program was developed using the commercially available code Matlab to derive strain, stress, and integral parameters from the experimental displacements. An FEM model of the crack tip area, with experimental displacements as boundary conditions, was used to validate FEM calculations of crack tip parameters. The results indicate that T*-integral parameter reaches a value of approximately 120 MPa-m0.5 during stable crack propagation which agrees with previously published values for straight cracks in the same material. The approximate computation method employed in this study uses a partial contour around the crack tip that neglects the contribution from the portion behind the crack tip where there is significant unloading. Strain distributions around the crack tip were obtained from experimental displacements and indicate that Maximum Principal Strain or Equivalent Strain can predict the direction of crack propagation, and is generally comparable with predictions using the Erdogan-Sih and Kosai-Ramulu-Kobayashi criteria. The biaxial tests to failure showed that the Machined Pad-up specimen carried the highest load, with the Bonded specimen next, at 78% of the Machined Pad-up value. The Bonded+Riveted specimen carried a lower load than the Bonded, at 67% of the Machined Pad-up value, which was the same as that carried by the Unreinforced specimen. The tearstraps of the bonded specimens remained intact after the specimen failed while the integrally machined reinforcement broke with the specimen. FEM studies were also made of skin flapping in typical Narrow and Wide-body fuselage sections, both containing the same crack path from a full-scale fatigue test of a Narrow-body fuselage. Results indicate that the

  4. Classification of fatigue cracking data in a simulated aircraft fuselage using a self-organizing map

    SciTech Connect

    Marsden, M.L.; Hill, E.V.K.

    1994-12-31

    Many aircraft are being flown beyond their design lifespans and have therefore fallen victim to fatigue cracking. In some cases, such as the 1988 Aloha Airlines 737-200 incident, catastrophic fatigue growth has caused the loss of life. Acoustic emission (AE) nondestructive testing has been used to detect and classical the growth of fatigue cracks in complex structures, such as aircraft fuselages and wings since as early as 1979. In order to simulate an aircraft fuselage undergoing pressurization cycle fatigue, a test was developed in which a thin-walled aluminum pressure vessel was instrumented with AE sensors and cyclically fatigued to promote crack growth at a stress concentration built into the vessel. The AE data acquisition system. extracted the six AE parameters - amplitude, counts, duration, energy, risetime, and count-to-peak from each of the sensor signals. One-third of these parameter data sets were used to tram a Kohonen self-organizing map (SOM) neural network. The remaining data sets were used to test the SOM. The SOM output is a two-dimensional map with similar input data sets located at similar coordinates on the map. Because the continuous AE parameter data are grouped into discrete bands or intervals, e.g., all the events having amplitudes between 51.00 dB and 51.99 dB are classified as 51 dB events, the initial SOM output showed no distinct clustering. However, when the output was transformed into three-dimensions, with the third dimension being the frequency of occurrence of each two-dimensional coordinate, several distinct peaks were evident. These peaks correspond to the three AE source in the vessel: metal rubbing, rivet fretting, and fatigue cracking. Thus, the three-dimensional SOM was able to unambiguously classify fatigue crack growth events in a simulated aircraft fuselage structure.

  5. 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.

  6. Characterization of Retrogression and Re-Aging Heat Treatment of AA7075-T6 Using Nonlinear Acoustics and Eddy Current

    SciTech Connect

    Ananthula, Rajeshwar; Ko, Ray T.; Sathish, Shamachary; Blodgett, Mark

    2004-02-26

    Nonlinear acoustic parameter and eddy current methods have been utilized to characterize the heat treatment process of retrogression and re-aging of aluminum 7075-T6. The results of nonlinear acoustic parameter measurements show two distinct peaks at 30 minutes and 45 minutes of retrogression time. The phase of the through-thickness eddy current signal shows a minimum at 42 minutes of retrogression time. Application of combined methods for identifying the optimized properties in the material is discussed.

  7. Influence of magnetic and acoustic treatment of superplasticizer solutions on the properties of portland cement concretes

    NASA Astrophysics Data System (ADS)

    Belous, N. Kh.; Azharonok, V. V.; Rodtsevich, S. P.; Koshevar, V. D.; Goncharik, S. V.; Chubrik, N. I.; Orlovich, A. I.; Rubannik, V. V.

    2012-05-01

    We have investigated the influence of the regimes of high-frequency magnetic-impulse and acoustic action on the physicochemical properties of water solutions of polycarboxylate superplasticizers and technological indices of fine concretes plasticized by them. The dependences of technological properties of concretes on the concentration of water solutions of the superplasticizers, the content of impurity ions in the water used for dilution, and the conditions of acousto-radiowave treatment have been determined. The regimes of activation of superplasticizer solutions, which permit increasing the mobility and keeping quality of concrete and solution mixes tempered with water and the density and strength of fine concretes formed from them, have been established.

  8. 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.

  9. 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

  10. Composite Structure Modeling and Analysis of Advanced Aircraft Fuselage Concepts

    NASA Technical Reports Server (NTRS)

    Mukhopadhyay, Vivek; Sorokach, Michael R.

    2015-01-01

    NASA Environmentally Responsible Aviation (ERA) project and the Boeing Company are collabrating to advance the unitized damage arresting composite airframe technology with application to the Hybrid-Wing-Body (HWB) aircraft. The testing of a HWB fuselage section with Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) construction is presently being conducted at NASA Langley. Based on lessons learned from previous HWB structural design studies, improved finite-element models (FEM) of the HWB multi-bay and bulkhead assembly are developed to evaluate the performance of the PRSEUS construction. In order to assess the comparative weight reduction benefits of the PRSEUS technology, conventional cylindrical skin-stringer-frame models of a cylindrical and a double-bubble section fuselage concepts are developed. Stress analysis with design cabin-pressure load and scenario based case studies are conducted for design improvement in each case. Alternate analysis with stitched composite hat-stringers and C-frames are also presented, in addition to the foam-core sandwich frame and pultruded rod-stringer construction. The FEM structural stress, strain and weights are computed and compared for relative weight/strength benefit assessment. The structural analysis and specific weight comparison of these stitched composite advanced aircraft fuselage concepts demonstrated that the pressurized HWB fuselage section assembly can be structurally as efficient as the conventional cylindrical fuselage section with composite stringer-frame and PRSEUS construction, and significantly better than the conventional aluminum construction and the double-bubble section concept.

  11. An experimental study of the effects of water repellant treatment on the acoustic properties of Kevlar

    NASA Technical Reports Server (NTRS)

    Smith, C. D.; Parrott, T. L.

    1978-01-01

    The treatment consisted of immersing samples of Kevlar in a solution of distilled water and Zepel. The samples were then drained, dried in a circulating over, and cured. Flow resistance tests showed approximately one percent decrease in flow resistance of the samples. Also there was a density increase of about three percent. It was found that the treatment caused a change in the texture of the samples. There were significant changes in the acoustic properties of the treated Kevlar over the frequency range 0.5 to 3.5 kHz. In general it was found that the propagation constant and characteristic impedance increased with increasing frequency. The real and imaginary components of the propagation constant for the treated Kevlar exhibited a decrease of 8 to 12 percent relative to that for the untreated Kevlar at the higher frequencies. The magnitude of the reactance component of the characteristic impedance decreased by about 40 percent at the higher frequencies.

  12. Potential efficacy of early treatment of acute acoustic trauma with steroids and piracetam after gunshot noise.

    PubMed

    Psillas, George; Pavlidis, Pavlos; Karvelis, Ioannis; Kekes, George; Vital, Victor; Constantinidis, John

    2008-12-01

    The purpose of this randomized study was to evaluate the early effect of the treatment of acute acoustic trauma (AAT) with steroids and piracetam in a sample of 52 young soldiers who were exposed to intense gunfire noise (G3 rifle). These patients were divided into three groups: (1) group A (20 patients) in which the treatment began within the first hour after the AAT, (2) group B (17 patients) in which the treatment started more than 1 h later and less than 16 h after the AAT and (3) group C (15 patients) in which the treatment began after 24 h or more. One month after the treatment onset, 36 (69%) patients of all the groups showed hearing improvement (complete-partial recovery) in the pure tone audiometry. The greater number of patients who showed complete recovery after AAT was noted in group A (65%) compared to group B (23.5%) and C (13.3%). Moreover, in group A, in the final audiogram, the averaged hearing threshold was statistically better (P < 0.001) than that of groups B and C. In spite of the lack of control group, our data demonstrated the possible effectiveness of the immediate onset of treatment of AAT. PMID:18463885

  13. Evaluation of the Acoustic Measurement Capability of the NASA Langley V/STOL Wind Tunnel Open Test Section with Acoustically Absorbent Ceiling and Floor Treatments

    NASA Technical Reports Server (NTRS)

    Theobald, M. A.

    1978-01-01

    The single source location used for helicopter model studies was utilized in a study to determine the distances and directions upstream of the model accurate at which measurements of the direct acoustic field could be obtained. The method used was to measure the decrease of sound pressure levels with distance from a noise source and thereby determine the Hall radius as a function of frequency and direction. Test arrangements and procedures are described. Graphs show the normalized sound pressure level versus distance curves for the glass fiber floor treatment and for the foam floor treatment.

  14. Mapping acoustic emissions from hydraulic fracture treatments using coherent array processing: Concept

    SciTech Connect

    Harris, D.B.; Sherwood, R.J.; Jarpe, S.P.; Harben, P.E.

    1991-09-01

    Hydraulic fracturing is a widely-used well completion technique for enhancing the recovery of gas and oil in low-permeability formations. Hydraulic fracturing consists of pumping fluids into a well under high pressure (1000--5000 psi) to wedge-open and extend a fracture into the producing formation. The fracture acts as a conduit for gas and oil to flow back to the well, significantly increasing communication with larger volumes of the producing formation. A considerable amount of research has been conducted on the use of acoustic (microseismic) emission to delineate fracture growth. The use of transient signals to map the location of discrete sites of emission along fractures has been the focus of most research on methods for delineating fractures. These methods depend upon timing the arrival of compressional (P) or shear (S) waves from discrete fracturing events at one or more clamped geophones in the treatment well or in adjacent monitoring wells. Using a propagation model, the arrival times are used to estimate the distance from each sensor to the fracturing event. Coherent processing methods appear to have sufficient resolution in the 75 to 200 Hz band to delineate the extent of fractures induced by hydraulic fracturing. The medium velocity structure must be known with a 10% accuracy or better and no major discontinuities should be undetected. For best results, the receiving array must be positioned directly opposite the perforations (same depths) at a horizontal range of 200 to 400 feet from the region to be imaged. Sources of acoustic emission may be detectable down to a single-sensor SNR of 0.25 or somewhat less. These conclusions are limited by the assumptions of this study: good coupling to the formation, acoustic propagation, and accurate knowledge of the velocity structure.

  15. Surface grid generation for wing-fuselage bodies

    NASA Technical Reports Server (NTRS)

    Smith, R. E.; Kudlinski, R. A.; Pitts, J. I.

    1984-01-01

    In the application of finite-difference methods to obtain numerical solutions of viscous compressible fluid flow about wing-fuselage bodies, it is advantageous to transform the governing equations to an idealized boundary-fitted coordinate system. The advantages are reduced computational complexity and added accuracy in the application of boundary conditions. The solution process requires that a grid be superimposed on the physical solution domain which corresponds to a uniform grid on a rectangular computational domain (uniform rectangular parallel-epiped). Grid generation is the determination of a one to one relationship between grid points in the physical domain and grid points in the computational domain. A technique for computing wing-fuselage surface grids using the Harris geometry and software for smooth-surface representation is described. Grid spacing control concepts which govern the relationship between the wing-fuselage surface and the computational grid are also presented.

  16. Helicopter fuselage drag - combined computational fluid dynamics and experimental studies

    NASA Astrophysics Data System (ADS)

    Batrakov, A.; Kusyumov, A.; Mikhailov, S.; Pakhov, V.; Sungatullin, A.; Valeev, M.; Zherekhov, V.; Barakos, G.

    2015-06-01

    In this paper, wind tunnel experiments are combined with Computational Fluid Dynamics (CFD) aiming to analyze the aerodynamics of realistic fuselage configurations. A development model of the ANSAT aircraft and an early model of the AKTAI light helicopter were employed. Both models were tested at the subsonic wind tunnel of KNRTU-KAI for a range of Reynolds numbers and pitch and yaw angles. The force balance measurements were complemented by particle image velocimetry (PIV) investigations for the cases where the experimental force measurements showed substantial unsteadiness. The CFD results were found to be in fair agreement with the test data and revealed some flow separation at the rear of the fuselages. Once confidence on the CFD method was established, further modifications were introduced to the ANSAT-like fuselage model to demonstrate drag reduction via small shape changes.

  17. Treatment of murine tumors using acoustic droplet vaporization-enhanced high intensity focused ultrasound

    NASA Astrophysics Data System (ADS)

    Zhu, Meili; Jiang, Lixing; Fabiilli, Mario L.; Zhang, Aili; Fowlkes, J. Brian; Xu, Lisa X.

    2013-09-01

    High intensity focused ultrasound (HIFU) can be applied focally and noninvasively to thermally ablate solid tumors. Long treatment times are typically required for large tumors, which can expose patients to certain risks while potentially decreasing the therapeutic efficacy of the treatment. Acoustic droplet vaporization (ADV) is a promising modality that can enhance the efficacy of tumor treatment using HIFU. In this study, the therapeutic effects of combined HIFU and ADV was evaluated in mice bearing subcutaneously-implanted 4T1 tumors. Histological examination showed that the combination of HIFU and ADV generated a mean necrotic area in the tumor that was 2.9-fold larger than with HIFU alone. A significant enhancement of necrosis was found in the periphery of the tumor, where the blood supply was abundant. Seven days after treatment, the tumors treated with combined HIFU and ADV were 30-fold smaller in volume than tumors treated with HIFU alone. The study demonstrates the potential advantage of combining HIFU and ADV in tumor treatment.

  18. Acoustic Characteristics of Various Treatment Panel Designs Specific to HSCT Mixer-Ejector Application

    NASA Technical Reports Server (NTRS)

    Salikuddin, M.; Kinzie, K.; Vu, D. D.; Langenbrunner, L. E.; Szczepkowski, G. T.

    2006-01-01

    The development process of liner design methodology is described in several reports. The results of the initial effort of concept development, screening, laboratory testing of various liner concepts, and preliminary correlation (generic data) are presented in a report Acoustic Characteristics of Various Treatment Panel Designs for HSCT Ejector Liner Acoustic Technology Development Program. The second phase of laboratory test results of more practical concepts and their data correlations are presented in this report (product specific). In particular, this report contains normal incidence impedance measurements of several liner types in both a static rig and in a high temperature flow duct rig. The flow duct rig allows for temperatures up to 400 F with a grazing flow up to Mach 0.8. Measurements of impedance, DC flow resistance, and in the flow rig cases, impact of the liner on boundary layer profiles are documented. In addition to liner rig tests, a limited number of tests were made on liners installed in a mixer-Ejector nozzle to confirm the performance of the liner prediction in an installed configuration.

  19. 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.

  20. 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

  1. 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

  2. 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

  3. 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

  4. 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

  5. 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.

  6. Effect of hyperbaric oxygen treatment on permanent threshold shift in acoustic trauma among rats.

    PubMed

    Kuokkanen, J; Virkkala, J; Zhai, S; Ylikoski, J

    1997-01-01

    Impulse noise from firearms is a common cause of acute acoustic trauma (AAT). Recently hyperbaric oxygen treatment has become available in many hospitals treating AAT. We exposed 39 Wistar rats to intense impulse noise of 60 shots from the assault rifle (162 dB SPL). After the exposure 15 animals were given hyperbaric oxygen treatment (HBO) by 10 treatment cycles of 90 minutes 100% oxygen in 0.25 MPa, one treatment cycle per day. Four weeks after the exposure the hearing thresholds were measured with auditory brainstem response audiometry at frequencies of 1.0, 2.0, 4.0, 6.0, 8.0 and 10.0 kHz. Characteristics for the resulting noise-induced hearing loss were large variations in its severity not only between animals, but also between the ears of a single animal. The largest permanent threshold shifts were found at 6.0, 8.0 and 10.0 kHz. Most of the HBO-treated animals showed less threshold shift than the non-treated animals. The difference between the HBO group and the control group was only slightly statistically significant (p = 0.067). PMID:9288276

  7. Numerical Investigation of Rotorcraft Fuselage Drag Reduction Using Active Flow Control

    NASA Technical Reports Server (NTRS)

    Allan, Brian G.; Schaeffler, Norman W.

    2011-01-01

    The effectiveness of unsteady zero-net-mass-flux jets for fuselage drag reduction was evaluated numerically on a generic rotorcraft fuselage in forward flight with a rotor. Previous efforts have shown significant fuselage drag reduction using flow control for an isolated fuselage by experiment and numerical simulation. This work will evaluate a flow control strategy, that was originally developed on an isolated fuselage, in a more relevant environment that includes the effects of a rotor. Evaluation of different slot heights and jet velocity ratios were performed. Direct comparisons between an isolated fuselage and rotor/fuselage simulations were made showing similar flow control performance at a -3deg fuselage angle-of-attack condition. However, this was not the case for a -5deg angle-of-attack condition where the performance between the isolated fuselage and rotor/fuselage were different. The fuselage flow control resulted in a 17% drag reduction for a peak C(sub mu) of 0.0069 in a forward flight simulation where mu = 0:35 and CT/sigma = 0:08. The CFD flow control results also predicted a favorable 22% reduction of the fuselage download at this same condition, which can have beneficial compounding effects on the overall performance of the vehicle. This numerical investigation was performed in order to provide guidance for a future 1/3 scale wind tunnel experiment to be performed at the NASA 14-by 22-Foot Subsonic Tunnel.

  8. 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... Requirements § 27.549 Fuselage, landing gear, and rotor pylon structures. (a) Each fuselage, landing gear, and... accelerated flight and landing conditions, including engine torque. (Secs. 604, 605, 72 Stat. 778, 49...

  9. 14 CFR 27.549 - Fuselage, landing gear, and rotor pylon structures.

    Code of Federal Regulations, 2010 CFR

    2010-01-01

    ... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Fuselage, landing gear, and rotor pylon... Requirements § 27.549 Fuselage, landing gear, and rotor pylon structures. (a) Each fuselage, landing gear, and... accelerated flight and landing conditions, including engine torque. (Secs. 604, 605, 72 Stat. 778, 49...

  10. 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.

  11. Aircraft interior noise prediction using a structural-acoustic analogy in NASTRAN modal synthesis

    NASA Technical Reports Server (NTRS)

    Grosveld, Ferdinand W.; Sullivan, Brenda M.; Marulo, Francesco

    1988-01-01

    The noise induced inside a cylindrical fuselage model by shaker excitation is investigated theoretically and experimentally. The NASTRAN modal-synthesis program is used in the theoretical analysis, and the predictions are compared with experimental measurements in extensive graphs. Good general agreement is obtained, but the need for further refinements to account for acoustic-cavity damping and structural-acoustic interaction is indicated.

  12. 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

  13. Manufacturing scale-up of composite fuselage crown panels

    NASA Technical Reports Server (NTRS)

    Willden, Kurtis; Gessel, M.; Grant, Carroll G.; Brown, T.

    1993-01-01

    The goal of the Boeing effort under the NASA ACT program is to reduce manufacturing costs of composite fuselage structure. Materials, fabrication of complex subcomponents and assembly issues are expected to drive the costs of composite fuselage structure. Several manufacturing concepts for the crown section of the fuselage were evaluated through the efforts of a Design Build Team (DBT). A skin-stringer-frame intricate bond design that required no fasteners for the panel assembly was selected for further manufacturing demonstrations. The manufacturing processes selected for the intricate bond design include Advanced Tow Placement (ATP) for multiple skin fabrication, resin transfer molding (RTM) of fuselage frames, innovative cure tooling, and utilization of low-cost material forms. Optimization of these processes for final design/manufacturing configuration was evaluated through the fabrication of several intricate bond panels. Panels up to 7 ft. by 10 ft. in size were fabricated to simulate half scale production parts. The qualitative and quantitative results of these manufacturing demonstrations were used to assess manufacturing risks and technology readiness for production.

  14. 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

  15. 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

  16. 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.

  17. RWF rotor-wake-fuselage code software reference guide

    NASA Technical Reports Server (NTRS)

    Berry, John D.

    1991-01-01

    The RWF (Rotor-Wake-Fuselage) code was developed from first principles to compute the aerodynamics associated with the complex flow field of helicopter configurations. The code is sized for a single, multi-bladed main rotor and any configuration of non-lifting fuselage. The mathematical model for the RWF code is based on the integration of the momentum equations and Green's theorem. The unknowns in the problem are the strengths of prescribed singularity distributions on the boundaries of the flow. For the body (fuselage) a surface of constant strength source panels is used. For the rotor blades and rotor wake a surface of constant strength doublet panels is used. The mean camber line of the rotor airfoil is partitioned into surface panels. The no-flow boundary condition at the panel centroids is modified at each azimuthal step to account for rotor blade cyclic pitch variation. The geometry of the rotor wake is computers at each time step of the solution. The code produces rotor and fuselage surface pressures, as well as the complex geometry of the evolving rotor wake.

  18. 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

  19. An evaluation of proposed acoustic treatments for the NASA LaRC 4 x 7 meter wind tunnel

    NASA Technical Reports Server (NTRS)

    Abrahamson, A. L.

    1985-01-01

    The NASA LaRC 4 x 7 Meter Wind Tunnel is an existing facility specially designed for powered low speed (V/STOL) testing of large scale fixed wing and rotorcraft models. The enhancement of the facility for scale model acoustic testing is examined. The results are critically reviewed and comparisons are drawn with a similar wind tunnel (the DNW Facility in the Netherlands). Discrepancies observed in the comparison stimulated a theoretical investigation using the acoustic finite element ADAM System, of the ways in which noise propagating around the tunnel circuit radiates into the open test section. The reasons for the discrepancies noted above are clarified and assists in the selection of acoustic treatment options for the facility.

  20. Risk of a second cancer from scattered radiation in acoustic neuroma treatment

    NASA Astrophysics Data System (ADS)

    Yoon, Myonggeun; Lee, Hyunho; Sung, Jiwon; Shin, Dongoh; Park, Sungho; Chung, Weon Kuu; Jahng, Geon-Ho; Kim, Dong Wook

    2014-06-01

    The present study aimed to compare the risk of a secondary cancer from scattered and leakage doses in patients receiving intensity-modulated radiotherapy (IMRT), volumetric modulated arc therapy (VMAT), and stereotactic radiosurgery (SRS). Four acoustic neuroma patients were treated with IMRT, VMAT, or SRS. Their excess relative risk (ERR), excess absolute risk (EAR), and lifetime attributable risk (LAR) of a secondary cancer were estimated using the corresponding secondary doses measured at various organs by using radio-photoluminescence glass dosimeters (RPLGD) placed inside a humanoid phantom. When a prescription dose was delivered in the planning target volume of the 4 patients, the average organ equivalent doses (OED) at the thyroid, lung, liver, bowel, bladder, prostate (or ovary), and rectum were 14.6, 1.7, 0.9, 0.8, 0.6, 0.6, and 0.6 cGy, respectively, for IMRT whereas they were 19.1, 1.8, 2.0, 0.6, 0.4, 0.4, and 0.4 cGy, respectively, for VMAT, and 22.8, 4.6, 1.4, 0.7, 0.5, 0.5, and 0.5 cGy, respectively, for SRS. The OED decreased as the distance from the primary beam increased. The thyroid received the highest OED compared to other organs. A lifetime attributable risk evaluation estimated that more than 0.03% of acoustic neuroma (AN) patients would get radiation-induced cancer within 20 years of receiving radiation therapy. The organ with the highest radiation-induced cancer risk after radiation treatment for AN was the thyroid. We found that the LAR could be increased by the transmitted dose from the primary beam. No modality-specific difference in radiation-induced cancer risk was observed in our study.

  1. SU-E-T-208: Incidence Cancer Risk From the Radiation Treatment for Acoustic Neuroma Patient

    SciTech Connect

    Kim, D; Chung, W; Shin, D; Yoon, M

    2014-06-01

    Purpose: The present study aimed to compare the incidence risk of a secondary cancer from therapeutic doses in patients receiving intensitymodulated radiotherapy (IMRT), volumetric modulated arc therapy (VMAT), and stereotactic radiosurgery (SRS). Methods: Four acoustic neuroma patients were treated with IMRT, VMAT, or SRS. Their incidnece excess relative risk (ERR), excess absolute risk (EAR), and lifetime attributable risk (LAR) were estimated using the corresponding therapeutic doses measured at various organs by radio-photoluminescence glass dosimeters (RPLGD) placed inside a humanoid phantom. Results: When a prescription dose was delivered in the planning target volume of the 4 patients, the average organ equivalent doses (OED) at the thyroid, lung, normal liver, colon, bladder, prostate (or ovary), and rectum were measured. The OED decreased as the distance from the primary beam increased. The thyroid received the highest OED compared to other organs. A LAR were estimated that more than 0.03% of AN patients would get radiation-induced cancer. Conclusion: The tyroid was highest radiation-induced cancer risk after radiation treatment for AN. We found that LAR can be increased by the transmitted dose from the primary beam. No modality-specific difference in radiation-induced cancer risk was observed in our study.

  2. 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.

  3. 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.

  4. Algebraic grid generation for wing-fuselage bodies

    NASA Technical Reports Server (NTRS)

    Smith, R. E.; Everton, E. L.; Kudlinski, R. A.

    1984-01-01

    An algebraic procedure for the generation of boundary-fitted grids about wing-fuselage configurations is presented. A wing-fuselage configuration is specified by cross sections and mathematically represented by Coons' patches. A configuration is divided into sections so that several grid blocks that either adjoin each other or partially overlap each other can be generated, and each grid has six surfaces that map into a computational cube. Grids are first determined on the six boundary surfaces and then in the interior. Grid curves that are on the surface of the configuration are derived using plane-patch intersections, and single-valued functions relating approximate arc lengths along the curves to computational coordinates define the distribution of grid points. The two-boundary technique and transfinite interpolation are used to determine the boundary surface grids that are not on the configuration, and transfinite interpolation with linear blending functions is used to determine the interior grids.

  5. Algebraic grid generation about wing-fuselage bodies

    NASA Technical Reports Server (NTRS)

    Smith, R.E.; Kudlinski, R. A.; Everton, E. L.; Wiese, M. R.

    1987-01-01

    An algebraic procedure for generating boundary-fitted grids about wing-fuselage configurations is presented. A wing-fuselage configuration consists of two aircraft components specified by cross sections and mathematically represented by Coons' patches. Several grid blocks are constructed to cover the entire region surrounding the configuration, and each grid block maps into a computational cube. Grid points are first determined on the six boundary surfaces of a block and then in the interior. Grid points on the surface of the configuration are derived from the intersection of planes with the Coons' patch definition. Approximate arc length distributions along the resulting grid curves concentrate and disperse grid points. The two-boundary technique and transfinite interpolation are used to determine grid points on the remaining boundary surfaces and block interiors.

  6. 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.

  7. 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.

  8. Aerodynamic analysis of a helicopter fuselage with rotating rotor head

    NASA Astrophysics Data System (ADS)

    Reß, R.; Grawunder, M.; Breitsamter, Ch.

    2015-06-01

    The present paper describes results of wind tunnel experiments obtained during a research programme aimed at drag reduction of the fuselage of a twin engine light helicopter configuration. A 1 : 5 scale model of a helicopter fuselage including a rotating rotor head and landing gear was investigated in the low-speed wind tunnel A of Technische Universität a München (TUM). The modelled parts of the helicopter induce approxiu mately 80% of the total parasite drag thus forming a major potential for shape optimizations. The present paper compares results of force and moment measurements of a baseline configuration and modified variants with an emphasis on the aerodynamic drag, lift, and yawing moment coefficients.

  9. 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.

  10. Design of fuselage shapes for natural laminar flow

    NASA Technical Reports Server (NTRS)

    Dodbele, S. S.; Vandam, C. P.; Vijgen, P. M. H. W.

    1986-01-01

    Recent technological advances in airplane construction techniques and materials allow for the production of aerodynamic surfaces without significant waviness and roughness, permitting long runs of natural laminar flow (NLF). The present research effort seeks to refine and validate computational design tools for use in the design of axisymmetric and nonaxisymmetric natural-laminar-flow bodies. The principal task of the investigation involves fuselage body shaping using a computational design procedure. Analytical methods were refined and exploratory calculations conducted to predict laminar boundary-layer on selected body shapes. Using a low-order surface-singularity aerodynamic analysis program, pressure distribution, boundary-layer development, transition location and drag coefficient have been obtained for a number of body shapes including a representative business-aircraft fuselage. Extensive runs of laminar flow were predicted in regions of favorable pressure gradient on smooth body surfaces. A computational design procedure was developed to obtain a body shape with minimum drag having large extent of NLF.

  11. Design of fuselage shapes for natural laminar flow

    NASA Astrophysics Data System (ADS)

    Dodbele, S. S.; Vandam, C. P.; Vijgen, P. M. H. W.

    1986-03-01

    Recent technological advances in airplane construction techniques and materials allow for the production of aerodynamic surfaces without significant waviness and roughness, permitting long runs of natural laminar flow (NLF). The present research effort seeks to refine and validate computational design tools for use in the design of axisymmetric and nonaxisymmetric natural-laminar-flow bodies. The principal task of the investigation involves fuselage body shaping using a computational design procedure. Analytical methods were refined and exploratory calculations conducted to predict laminar boundary-layer on selected body shapes. Using a low-order surface-singularity aerodynamic analysis program, pressure distribution, boundary-layer development, transition location and drag coefficient have been obtained for a number of body shapes including a representative business-aircraft fuselage. Extensive runs of laminar flow were predicted in regions of favorable pressure gradient on smooth body surfaces. A computational design procedure was developed to obtain a body shape with minimum drag having large extent of NLF.

  12. 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.

  13. 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

  14. 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

  15. 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

  16. 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

  17. 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.

  18. Skin, Stringer, and Fastener Loads in Buckled Fuselage Panels

    NASA Technical Reports Server (NTRS)

    Young, Richard D.; Rose, Cheryl A.; Starnes, James H., Jr.

    2001-01-01

    The results of a numerical study to assess the effect of skin buckling on the internal load distribution in a stiffened fuselage panel, with and without longitudinal cracks, are presented. In addition, the impact of changes in the internal loads on the fatigue life and residual strength of a fuselage panel is assessed. A generic narrow-body fuselage panel is considered. The entire panel is modeled using shell elements and considerable detail is included to represent the geometric-nonlinear response of the buckled skin, cross section deformation of the stiffening components, and details of the skin-string attachment with discrete fasteners. Results are presented for a fixed internal pressure and various combinations of axial tension or compression loads. Results illustrating the effect of skin buckling on the stress distribution in the skin and stringer, and fastener loads are presented. Results are presented for the pristine structure, and for cases where damage is introduced in the form of a longitudinal crack adjacent to the stringer, or failed fastener elements. The results indicate that axial compression loads and skin buckling can have a significant effect on the circumferential stress in the skin, and fastener loads, which will influence damage initiation, and a comparable effect on stress intensity factors for cases with cracks. The effects on stress intensity factors will influence damage propagation rates and the residual strength of the panel.

  19. Full-scale testing and analysis of fuselage structure

    NASA Technical Reports Server (NTRS)

    Miller, M.; Gruber, M. L.; Wilkins, K. E.; Worden, R. E.

    1994-01-01

    This paper presents recent results from a program in the Boeing Commercial Airplane Group to study the behavior of cracks in fuselage structures. The goal of this program is to improve methods for analyzing crack growth and residual strength in pressurized fuselages, thus improving new airplane designs and optimizing the required structural inspections for current models. The program consists of full-scale experimental testing of pressurized fuselage panels in both wide-body and narrow-body fixtures and finite element analyses to predict the results. The finite element analyses are geometrically nonlinear with material and fastener nonlinearity included on a case-by-case basis. The analysis results are compared with the strain gage, crack growth, and residual strength data from the experimental program. Most of the studies reported in this paper concern the behavior of single or multiple cracks in the lap joints of narrow-body airplanes (such as 727 and 737 commercial jets). The phenomenon where the crack trajectory is curved creating a 'flap' and resulting in a controlled decompression is discussed.

  20. Cabin fuselage structural design with engine installation and control system

    NASA Technical Reports Server (NTRS)

    Balakrishnan, Tanapaal; Bishop, Mike; Gumus, Ilker; Gussy, Joel; Triggs, Mike

    1994-01-01

    Design requirements for the cabin, cabin system, flight controls, engine installation, and wing-fuselage interface that provide adequate interior volume for occupant seating, cabin ingress and egress, and safety are presented. The fuselage structure must be sufficient to meet the loadings specified in the appropriate sections of Federal Aviation Regulation Part 23. The critical structure must provide a safe life of 10(exp 6) load cycles and 10,000 operational mission cycles. The cabin seating and controls must provide adjustment to account for various pilot physiques and to aid in maintenance and operation of the aircraft. Seats and doors shall not bind or lockup under normal operation. Cabin systems such as heating and ventilation, electrical, lighting, intercom, and avionics must be included in the design. The control system will consist of ailerons, elevator, and rudders. The system must provide required deflections with a combination of push rods, bell cranks, pulleys, and linkages. The system will be free from slack and provide smooth operation without binding. Environmental considerations include variations in temperature and atmospheric pressure, protection against sand, dust, rain, humidity, ice, snow, salt/fog atmosphere, wind and gusts, and shock and vibration. The following design goals were set to meet the requirements of the statement of work: safety, performance, manufacturing and cost. To prevent the engine from penetrating the passenger area in the event of a crash was the primary safety concern. Weight and the fuselage aerodynamics were the primary performance concerns. Commonality and ease of manufacturing were major considerations to reduce cost.

  1. Advanced Technology Composite Fuselage - Repair and Damage Assessment Supporting Maintenance

    NASA Technical Reports Server (NTRS)

    Flynn, B. W.; Bodine, J. B.; Dopker, B.; Finn, S. R.; Griess, K. H.; Hanson, C. T.; Harris, C. G.; Nelson, K. M.; Walker, T. H.; Kennedy, T. C.; Nahan, M. F.

    1997-01-01

    Under the NASA-sponsored contracts for Advanced Technology Composite Aircraft Structures (ATCAS) and Materials Development Omnibus Contract (MDOC), Boeing is studying the technologies associated with the application of composite materials to commercial transport fuselage structure. Included in the study is the incorporation of maintainability and repairability requirements of composite primary structure into the design. This contractor report describes activities performed to address maintenance issues in composite fuselage applications. A key aspect of the study was the development of a maintenance philosophy which included consideration of maintenance issues early in the design cycle, multiple repair options, and airline participation in design trades. Fuselage design evaluations considered trade-offs between structural weight, damage resistance/tolerance (repair frequency), and inspection burdens. Analysis methods were developed to assess structural residual strength in the presence of damage, and to evaluate repair design concepts. Repair designs were created with a focus on mechanically fastened concepts for skin/stringer structure and bonded concepts for sandwich structure. Both a large crown (skintstringer) and keel (sandwich) panel were repaired. A compression test of the keel panel indicated the demonstrated repairs recovered ultimate load capability. In conjunction with the design and manufacturing developments, inspection methods were investigated for their potential to evaluate damaged structure and verify the integrity of completed repairs.

  2. Rotorcraft Fuselage Flow Control Using Plasma Streamwise Vortex Generators

    NASA Astrophysics Data System (ADS)

    Coleman, Dustin; Thomas, Flint

    2012-11-01

    Active flow control, in the form of dielectric barrier discharge (DBD) plasma actuators, is applied to a NASA ROBIN-mod7 generic rotorcraft fuselage model. The model is considered in what would be a typical cruise position i.e. a nose down position at α = -5° . This configuration gives rise to a massive 3-D flow separation over the aft ramp section of the fuselage, characterized by two counter-rotating, streamwise vortices. The control objective is to minimize these concentrated vortices by means of flush fuselage-mounted plasma streamwise vortex generators (PSVGs), and consequently, reduce the form drag of the vehicle. Experiments were conducted at freestream Mach and Reynolds numbers of M∞ = 0 . 12 and ReL = 2 . 65 million, respectively. Aerodynamic loads under both natural and controlled conditions were acquired through use of an ATI Mini40 6-component force sensor. The pressure field on the ramp section was monitored by a 128 count static pressure array. Likewise, the flow field was captured by time-resolved PIV wake surveys. Results are compared with previous studies that utilized active flow control by way of pulsed jets or combustion actuators. This work is supported under NASA Cooperative Agreement NNX10AM32G.

  3. High Fidelity Failure Analysis for a Composite Fuselage Section

    NASA Technical Reports Server (NTRS)

    Li, Jain; Davila, Carlos G.; Chen, Tzi-Kang

    2001-01-01

    A high fidelity delamination failure analysis was developed by combining a local failure analysis with a global full-scale finite element structural analysis to address strength and delamination failure in a single package. The methodology was demonstrated through a local three-dimensional pull-off failure analysis and a geometrically nonlinear structural analysis of a five-foot composite helicopter fuselage section. Pull-off specimens were used to identify potential debonding failure of co-cured skin-stringer/frame fuselage structures. An investigation of the failed pull-off specimens was performed to determine the location of the failure initiation. Three-dimensional strain energy release rate analysis indicates that the delamination initiation and growth is controlled by Mode 1 opening mode. The bending moment at the delamination tip was identified as the crucial factor controlling the failure. The geometrically nonlinear structural analysis of a five-foot composite fuselage section was performed using a detailed finite element model. Loads were applied along the periphery of the subcomponent using displacement fields generated from solutions of a full-vehicle model.

  4. 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.

  5. 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.

  6. Full-scale testing and analysis of fuselage structure

    NASA Astrophysics Data System (ADS)

    Miller, M.; Gruber, M. L.; Wilkins, K. E.; Worden, R. E.

    1994-09-01

    This paper presents recent results from a program in the Boeing Commercial Airplane Group to study the behavior of cracks in fuselage structures. The goal of this program is to improve methods for analyzing crack growth and residual strength in pressurized fuselages, thus improving new airplane designs and optimizing the required structural inspections for current models. The program consists of full-scale experimental testing of pressurized fuselage panels in both wide-body and narrow-body fixtures and finite element analyses to predict the results. The finite element analyses are geometrically nonlinear with material and fastener nonlinearity included on a case-by-case basis. The analysis results are compared with the strain gage, crack growth, and residual strength data from the experimental program. Most of the studies reported in this paper concern the behavior of single or multiple cracks in the lap joints of narrow-body airplanes (such as 727 and 737 commercial jets). The phenomenon where the crack trajectory is curved creating a 'flap' and resulting in a controlled decompression is discussed.

  7. 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.

  8. 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.

  9. Fabrication of the V-22 composite AFT fuselage using automated fiber placement

    NASA Technical Reports Server (NTRS)

    Pinckney, Robert L.

    1991-01-01

    Boeing Helicopters and its subcontractors are working together under an Air Force Wright Research and Development Center (WRDC)-Manufacturing-Technology Large-Composite Primary Structure Fuselage program to develop and demonstrate new manufacturing techniques for producing composite fuselage skin and frame structures. Three sets of aft fuselage skins and frames have been fabricated and assembled, and substantial reductions in fabrication and assembly costs demonstrated.

  10. 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.

  11. 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.

  12. Bubbles trapped at the coupling surface of the treatment head significantly reduce acoustic energy delivered in shock wave lithotripsy

    NASA Astrophysics Data System (ADS)

    Pishchalnikov, Yuri A.; McAteer, James A.; Pishchalnikova, Irina V.; Beard, Spencer; Williams, James C.; Bailey, Michael R.

    2006-05-01

    The coupling efficiency of a "dry head" electromagnetic lithotripter (Dornier Compact Delta) was studied in vitro. A fiber-optic probe hydrophone (FOPH-500) was positioned in a test tank filled with degassed water. The tank was coupled through a semi-transparent latex membrane to the water-filled cushion of the lithotripter head, so that bubbles (air pockets) trapped between the two coupling surfaces could be easily observed and photographed. When gel was applied to both the latex membrane and the water cushion, numerous bubbles (some several millimeters in diameter) could be seen at the coupling interface. Hydrophone measurements in the geometric focus of the lithotripter showed that the acoustic pressure could be two times lower when bubbles were present than when they were manually removed. In our in vitro design, trapped bubbles could be easily observed and therefore removed from the acoustic path. However, during patient treatment with a dry-head lithotripter one cannot see whether bubbles are trapped against the skin. This study provides a demonstration of the dramatic effect that trapped bubbles can have on the amount of acoustic energy actually delivered for treatment.

  13. 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.

  14. Acoustic neuroma

    MedlinePlus

    Vestibular schwannoma; Tumor - acoustic; Cerebellopontine angle tumor; Angle tumor ... Acoustic neuromas have been linked with the genetic disorder neurofibromatosis type 2 (NF2). Acoustic neuromas are uncommon.

  15. New wrinkles on black hole perturbations: Numerical treatment of acoustic and gravitational waves

    NASA Astrophysics Data System (ADS)

    Tenyotkin, Valery

    2009-06-01

    This thesis develops two main topics. A full relativistic calculation of quasinormal modes of an acoustic black hole is carried out. The acoustic black hole is formed by a perfect, inviscid, relativistic, ideal gas that is spherically accreting onto a Schwarzschild black hole. The second major part is the calculation of sourceless vector (electromagnetic) and tensor (gravitational) covariant field evolution equations for perturbations on a Schwarzschild background using the relatively recent [Special characters omitted.] decomposition method. Scattering calculations are carried out in Schwarzschild coordinates for electromagnetic and gravitational cases as validation of the method and the derived equations.

  16. Alleviation of fuselage form drag using vortex flows: Final report

    SciTech Connect

    Wortman, A.

    1987-09-15

    The concept of using vortex generators to reduce the fuselage form drag of transport aircraft combines the outflow from the plane of symmetry which is induced by the rotational component of the vortex flow with the energization of the boundary layer to reduce the momentum thickness and to delay or eliminate flow separation. This idea was first advanced by the author in 1981. Under a DOE grant, the concept was validated in wind tunnel tests of approximately 1:17 scale models of fuselages of Boeing 747 and Lockheed C-5 aircraft. The search for the minimum drag involved three vortex generator configurations with three sizes of each in six locations clustered in the aft regions of the fuselages at the beginning of the tail upsweep. The local Reynolds number, which is referred to the length of boundary layer run from the nose, was approximately 10{sup 7} so that a fully developed turbulent boundary layer was present. Vortex generator planforms ranged from swept tapered, through swept straight, to swept reverse tapered wings whose semi-spans ranged from 50% to 125% of the local boundary layer thickness. Pitch angles of the vortex generators were varied by inboard actuators under the control of an external proportional digital radio controller. It was found that certain combinations of vortex generator parameters increased drag. However, with certain configurations, locations, and pitch angles of vortex generators, the highest drag reductions were 3% for the 747 and about 6% for the C-5, thus confirming the arguments that effectiveness increases with the rate of upsweep of the tail. Greatest gains in performance are therefore expected on aft loading military transports. 10 refs., 11 figs., 1 tab.

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

    NASA Technical Reports Server (NTRS)

    Schaefler, 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

  18. Experimental study of noise reduction for an unstiffened cylindrical model of an airplane fuselage

    NASA Technical Reports Server (NTRS)

    Willis, C. M.; Daniels, E. F.

    1981-01-01

    Noise reduction measurements were made for a simplified model of an airplane fuselage consisting of an unstiffened aluminum cylinder 0.5 m in diameter by 1.2 m long with a 1.6-mm-thick wall. Noise reduction was first measured with a reverberant field pink-noise load on the cylinder exterior. Next, noise reduction was measured by using a propeller to provide a more realistic noise load on the cylinder. Structural resonance frequencies and acoustic reverberation times for the cylinder interior volume were also measured. Comparison of data from the relatively simple test using reverberant-field noise with data from the more complex propeller-noise tests indicates some similarity in both the overall noise reduction and the spectral distribution. However, all of the test parameters investigated (propeller speed, blade pitch, and tip clearance) had some effect on the noise-reduction spectra. Thus, the amount of noise reduction achieved appears to be somewhat dependent upon the spectral and spatial characteristics of the flight conditions. Information is also presented on cyclinder resonance frequencies, damping, and characteristics of propeller-noise loads.

  19. 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.

  20. Closeup view of the aft fuselage of the Orbiter Discovery ...

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

    Close-up view of the aft fuselage of the Orbiter Discovery on the starboard side looking forward. This view is of the attach surface for the Orbiter Maneuvering System/Reaction Control System (OMS/RCS) Pod. The OMS/RCS pods are removed for processing and reconditioning at another facility. This view was taken from a service platform in 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

  1. Aeroelastic Analysis of Aircraft: Wing and Wing/Fuselage Configurations

    NASA Technical Reports Server (NTRS)

    Chen, H. H.; Chang, K. C.; Tzong, T.; Cebeci, T.

    1997-01-01

    A previously developed interface method for coupling aerodynamics and structures is used to evaluate the aeroelastic effects for an advanced transport wing at cruise and under-cruise conditions. The calculated results are compared with wind tunnel test data. The capability of the interface method is also investigated for an MD-90 wing/fuselage configuration. In addition, an aircraft trim analysis is described and applied to wing configurations. The accuracy of turbulence models based on the algebraic eddy viscosity formulation of Cebeci and Smith is studied for airfoil flows at low Mach numbers by using methods based on the solutions of the boundary-layer and Navier-Stokes equations.

  2. Closeup view of the aft fuselage looking forward along the ...

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

    Close-up view of the aft fuselage looking forward along the approximate centerline of the Orbiter Discovery looking at the expansion nozzles of the Space Shuttle Main Engines (SSME) and the Orbiter Maneuvering System. Also in the view is the orbiter's body flap with a protective covering over the High-temperature Reusable Surface Insulation tiles on the surface facing the SSMEs. 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

  3. Closeup view of the underside of the forward fuselage of ...

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

    Close-up view of the underside of the forward fuselage of the Orbiter Discovery looking at the nose landing-gear and into the landing-gear well. The vehicle is elevated and supported by jack stands attached to the hoist attach points and the rear External Tank attach points on the propellant disconnect plate assemblies. This photo 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

  4. 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.

  5. 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

  6. Finite element modeling of acoustic wave propagation and energy deposition in bone during extracorporeal shock wave treatment

    NASA Astrophysics Data System (ADS)

    Wang, Xiaofeng; Matula, Thomas J.; Ma, Yong; Liu, Zheng; Tu, Juan; Guo, Xiasheng; Zhang, Dong

    2013-06-01

    It is well known that extracorporeal shock wave treatment is capable of providing a non-surgical and relatively pain free alternative treatment modality for patients suffering from musculoskeletal disorders but do not respond well to conservative treatments. The major objective of current work is to investigate how the shock wave (SW) field would change if a bony structure exists in the path of the acoustic wave. Here, a model of finite element method (FEM) was developed based on linear elasticity and acoustic propagation equations to examine SW propagation and deflection near a mimic musculoskeletal bone. High-speed photography experiments were performed to record cavitation bubbles generated in SW field with the presence of mimic bone. By comparing experimental and simulated results, the effectiveness of FEM model could be verified and strain energy distributions in the bone were also predicted according to numerical simulations. The results show that (1) the SW field will be deflected with the presence of bony structure and varying deflection angles can be observed as the bone shifted up in the z-direction relative to SW geometric focus (F2 focus); (2) SW deflection angels predicted by the FEM model agree well with experimental results obtained from high-speed photographs; and (3) temporal evolutions of strain energy distribution in the bone can also be evaluated based on FEM model, with varied vertical distance between F2 focus and intended target point on the bone surface. The present studies indicate that, by combining MRI/CT scans and FEM modeling work, it is possible to better understand SW propagation characteristics and energy deposition in musculoskeletal structure during extracorporeal shock wave treatment, which is important for standardizing the treatment dosage, optimizing treatment protocols, and even providing patient-specific treatment guidance in clinic.

  7. A fuselage/tank structure study for actively cooled hypersonic cruise vehicles: Aircraft design evaluation

    NASA Technical Reports Server (NTRS)

    Nobe, T.

    1975-01-01

    The effects of fuselage cross sections and structural members on the performance of hypersonic cruise aircraft are evaluated. Representative fuselage/tank area structure was analyzed for strength, stability, fatigue and fracture mechanics. Various thermodynamic and structural tradeoffs were conducted to refine the conceptual designs with the primary objective of minimizing weight and maximizing aircraft range.

  8. Local design optimization for composite transport fuselage crown panels

    NASA Technical Reports Server (NTRS)

    Swanson, G. D.; Ilcewicz, L. B.; Walker, T. H.; Graesser, D.; Tuttle, M.; Zabinsky, Z.

    1992-01-01

    Composite transport fuselage crown panel design and manufacturing plans were optimized to have projected cost and weight savings of 18 and 45 percent, respectively. These savings are close to those quoted as overall NASA Advanced Composite Technology (ACT) program goals. Three local optimization tasks were found to influence the cost and weight of fuselage crown panels. The effects are summarized of each task and the task associated with a design cost model is described in detail. Studies were performed to evaluate the relationship between manufacturing cost and design details. A design tool was developed to aid in these studies. The development of the design tool included combining cost and performance constraints with a random search optimization algorithm. The resulting software was used in a series of optimization studies that evaluated the sensitivity of design variables, guidelines, criteria, and material selection on cost. The effect of blending adjacent design points in a full scale panel subjected to changing load distributions and local variations was shown to be important. Technical issues and directions for future work were identified.

  9. Local design optimization for composite transport fuselage crown panels

    NASA Technical Reports Server (NTRS)

    Swanson, G. D.; Ilcewicz, L. B.; Walker, T. H.; Graesser, D.; Tuttle, M.; Zabinsky, Z.

    1992-01-01

    Composite transport fuselage crown panel design and manufacturing plans were optimized to have projected cost and weight savings of 18 percent and 45 percent, respectively. These savings are close to those quoted as overall NASA ACT program goals. Three local optimization tasks were found to influence the cost and weight of fuselage crown panels. This paper summarizes the effect of each task and describes in detail the task associated with a design cost model. Studies were performed to evaluate the relationship between manufacturing cost and design details. A design tool was developed to aid in these investigations. The development of the design tool included combining cost and performance constraints with a random search optimization algorithm. The resulting software was used in a series of optimization studies that evaluated the sensitivity of design variables, guidelines, criteria, and material selection on cost. The effect of blending adjacent design points in a full scale panel subjected to changing load distributions and local variations was shown to be important. Technical issues and directions for future work were identified.

  10. Algebraic grid generation about wing-fuselage bodies

    NASA Technical Reports Server (NTRS)

    Smith, R. E.

    1986-01-01

    An algebraic procedure for the generation of boundary-fitted grids about wing-fuselage configurations is presented. A wing-fuselage configuration is specified by cross sections and mathematically represented by Coons' patches. A configuration is divided into sections so that several grid blocks that either adjoin each other or partially overlap each other can be generated. Each grid has six exterior surfaces that map into a computational cube. Grids are first determined on the six boundary surfaces and then in the interior. Grid curves that are on the surface of the configuration are derived from the intersection of planes with the Coons' patch definition. Single-valued functions relating approximate arc lengths along the grid curves to a computational coordinate define the distribution of grid points. The two-boundary technique and transfinite interpolation are used to determine the boundary surface grids that are not on the configuration, and transfinite interpolation with linear blending functions is used to determine the interior grid.

  11. Vertical Drop Test of a YS-11 Fuselage Section

    NASA Astrophysics Data System (ADS)

    Minegishi, Masakatsu; Kumakura, Ikuo; Iwasaki, Kazuo; Shoji, Hirokazu; Yoshimoto, Norio; Terada, Hiroyuki; Sashikuma, Hirofumi; Isoe, Akira; Yamaoka, Toshihiro; Katayama, Noriaki; Hayashi, Toru; Akaso, Tetsuya

    The Structures and Materials Research Center of the National Aerospace Laboratory of Japan (NAL) and Kawasaki Heavy Industories, Ltd. (KHI) conducted a vertical drop test of a fuselage section cut from a NAMIC YS-11 transport airplane at NAL vertical drop test facility in December 2001. The main objectives of this program were to obtain background data for aircraft cabin safety by drop test of a full-scale fuselage section and to develop computational method for crash simulation. The test article including seats and anthropomorphic test dummies was dropped to a rigid impact surface at a velocity of 6.1 m/s (20 ft/s). The test condition and result were considered to be severe but potentially survivable. A finite element model of this test article was also developed using the explicit nonlinear transient-dynamic analysis code, LS-DYNA3D. An outline of analytical method and comparison of analysis result with drop test data are presented in this paper.

  12. 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.

  13. Optimization of Sandwich Composites Fuselages Under Flight Loads

    NASA Astrophysics Data System (ADS)

    Yuan, Chongxin; Bergsma, Otto; Koussios, Sotiris; Zu, Lei; Beukers, Adriaan

    2012-02-01

    The sandwich composites fuselages appear to be a promising choice for the future aircrafts because of their structural efficiency and functional integration advantages. However, the design of sandwich composites is more complex than other structures because of many involved variables. In this paper, the fuselage is designed as a sandwich composites cylinder, and its structural optimization using the finite element method (FEM) is outlined to obtain the minimum weight. The constraints include structural stability and the composites failure criteria. In order to get a verification baseline for the FEM analysis, the stability of sandwich structures is studied and the optimal design is performed based on the analytical formulae. Then, the predicted buckling loads and the optimization results obtained from a FEM model are compared with that from the analytical formulas, and a good agreement is achieved. A detailed parametric optimal design for the sandwich composites cylinder is conducted. The optimization method used here includes two steps: the minimization of the layer thickness followed by tailoring of the fiber orientation. The factors comprise layer number, fiber orientation, core thickness, frame dimension and spacing. Results show that the two-step optimization is an effective method for the sandwich composites and the foam sandwich cylinder with core thickness of 5 mm and frame pitch of 0.5 m exhibits the minimum weight.

  14. Numerical Investigation of a Fuselage Boundary Layer Ingestion Propulsion Concept

    NASA Technical Reports Server (NTRS)

    Elmiligui, Alaa A.; Fredericks, William J.; Guynn, Mark D.; Campbell, Richard L.

    2013-01-01

    In the present study, a numerical assessment of the performance of fuselage boundary layer ingestion (BLI) propulsion techniques was conducted. This study is an initial investigation into coupling the aerodynamics of the fuselage with a BLI propulsion system to determine if there is sufficient potential to warrant further investigation of this concept. Numerical simulations of flow around baseline, Boundary Layer Controlled (BLC), and propelled boundary layer controlled airships were performed. Computed results showed good agreement with wind tunnel data and previous numerical studies. Numerical simulations and sensitivity analysis were then conducted on four BLI configurations. The two design variables selected for the parametric study of the new configurations were the inlet area and the inlet to exit area ratio. Current results show that BLI propulsors may offer power savings of up to 85% over the baseline configuration. These interim results include the simplifying assumption that inlet ram drag is negligible and therefore likely overstate the reduction in power. It has been found that inlet ram drag is not negligible and should be included in future analysis.

  15. 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.

  16. Study on utilization of advanced composites in fuselage structures of large transports

    NASA Technical Reports Server (NTRS)

    Johnson, R. W.; Thomson, L. W.; Wilson, R. D.

    1985-01-01

    The potential for utilizing advanced composites in fuselage structures of large transports was assessed. Six fuselage design concepts were selected and evaluated in terms of structural performance, weight, and manufacturing development and costs. Two concepts were selected that merit further consideration for composite fuselage application. These concepts are: (1) a full depth honeycomb design with no stringers, and (2) an I section stringer stiffened laminate skin design. Weight reductions due to applying composites to the fuselages of commercial and military transports were calculated. The benefits of applying composites to a fleet of military transports were determined. Significant technology issues pertinent to composite fuselage structures were identified and evaluated. Program plans for resolving the technology issues were developed.

  17. Impact of fuselage incidence on the supersonic aerodynamics of two fighter configurations

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    The results of experimental and theoretical investigations into the effect of fuselage upwash on fighter aircraft wing performance are reported. Wind tunnel trials were performed on 4 percent scale models of two supersonic fighters. The trials were run at Mach 1.6-2.0, an Re of 2,000,000 and at angles of attack (AOA) of -4 to 20 deg. Measurements were made of lift, drag and pitching moments. Two linearized theory supersonic aerodynamic prediction codes, PAN AIR and the SDAS lift analysis, were used to predict the same aerodynamic coefficients. The fuselage AOA augmented the lift and pitching moment at 0, 2 and 5 deg. The contribution mainly arose from the fuselage-induced upwash. The PAN AIR code gave superior data for the fuselage aerodynamics and effects, although both codes accurately predicted the overall lift and moment increments due to the fuselage AOA.

  18. Acoustic evaluation of the Helmholtz resonator treatment in the NASA Lewis 8- by 6-foot supersonic wind tunnel

    NASA Technical Reports Server (NTRS)

    Heidelberg, Laurence J.; Gordon, Elliot B.

    1989-01-01

    The acoustic consequences of sealing the Helmholtz resonators of the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel (8x6 SWT) were experimentally evaluated. This resonator sealing was proposed in order to avoid entrapment of hydrogen during tests of advanced hydrogen-fueled engines. The resonators were designed to absorb energy in the 4- to 20-Hz range; thus, this investigation is primarily concerned with infrasound. Limited internal and external noise measurements were made at tunnel Mach numbers ranging from 0.5 to 2.0. Although the resonators were part of the acoustic treatment installed because of a community noise problem their sealing did not seem to indicate a reoccurrence of the problem would result. Two factors were key to this conclusion: (1) A large bulk treatment muffler downstream of the resonators was able to make up for much of the attenuation originally provided by the resonators, and (2) there was no noise source in the tunnel test section. The previous community noise problem occurred when a large ramjet was tested in an open-loop tunnel configuration. If a propulsion system producing high noise levels at frequencies of less than 10 Hz were tested, the conclusion on community noise would have to be reevaluated.

  19. Comparison of Different Measurement Technologies for the In-Flight Assessment of Radiated Acoustic Intensity

    NASA Technical Reports Server (NTRS)

    Klos, Jacob; Palumbo, Daniel L.; Buehrle, Ralph D.; Williams, Earl G.; Valdivia, Nicolas; Herdic, Peter C.; Sklanka, Bernard

    2005-01-01

    A series of tests was planned and conducted in the Interior Noise Test Facility at Boeing Field, on the NASA Aries 757 flight research aircraft, and in the Structural Acoustic Loads and Transmission Facility at NASA Langley Research Center. These tests were designed to answer several questions concerning the use of array methods in flight. One focus of the tests was determining whether and to what extent array methods could be used to identify the effects of an acoustical treatment applied to a limited portion of an aircraft fuselage. Another focus of the tests was to verify that the arrays could be used to localize and quantify a known source purposely placed in front of the arrays. Thus the issues related to backside sources and flanking paths present in the complicated sound field were addressed during these tests. These issues were addressed through the use of reference transducers, both accelerometers mounted to the fuselage and microphones in the cabin, that were used to correlate the pressure holograms. measured by the microphone arrays using either SVD methods or partial coherence methods. This correlation analysis accepts only energy that is coherent with the sources sensed by the reference transducers, allowing a noise control engineer to only identify and study those vibratory sources of interest. The remainder of this paper will present a detailed description of the test setups that were used in this test sequence and typical results of the NAH/IBEM analysis used to reconstruct the sound fields. Also, a comparison of data obtained in the laboratory environments and during flights of the 757 aircraft will be made.

  20. Effect of fuselage upwash on the supersonic longitudinal aerodynamic characteristics of 2 fighter configurations

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    An experimental and theoretical investigation of fuselage incidence effects on two fighter aircraft models, which differed in wing planform only, has been conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.6, 1.8, and 2.0. Results were obtained on the two models at fuselage incidence angles of 0 deg, 2 deg, and 5 deg. The fuselage geometry included two side-mounted, flow-through, half-axisymmetric inlets and twin vertical tails. The two planforms tested were cranked wings with 70 deg/66 deg and 70 deg/30 deg leading-edge sweep angles. Experimental data showed that fuselage incidence resulted in positive increments in configuration lift and pitching moment; most of the lift increment can be attributed to the fuselage-induced upwash acting on the wing and most of the pitching-moment increment is due to the fuselage. Theoretical analysis indicates that linear-theory methods can adequately predict the overall configuration forces and moments resulting from fuselage upwash, but a higher order surface-panel method (PAN AIR) more accurately predicted the distribution of forces and resulting moments between the components.

  1. NOTE: Initial evaluation of acoustic reflectors for the preservation of sensitive abdominal skin areas during MRgFUS treatment

    NASA Astrophysics Data System (ADS)

    Gorny, Krzysztof R.; Chen, Shigao; Hangiandreou, Nicholas J.; Hesley, Gina K.; Woodrum, David A.; Brown, Douglas L.; Felmlee, Joel P.

    2009-04-01

    During MR-guided focused ultrasound (MRgFUS) treatments of uterine fibroids using ExAblate®2000 (InSightec, Haifa, Israel), individual tissue ablations are performed extracorporeally through the patient's abdomen using an annular array FUS transducer embedded within the MR table. Ultrasound intensities in the near field are below therapeutic levels and, under normal conditions, heating of the patient skin is minimal. However, increased absorption of ultrasound energy within sensitive skin areas or areas with differing acoustic properties, such as scars, may lead to skin burns and therefore these areas must be kept outside the near field of the FUS beam. Depending on their location and size the sensitive areas may either obstruct parts of the fibroid from being treated or prevent the entire MRgFUS treatment altogether. The purpose of this work is to evaluate acoustic reflector materials that can be applied to protect skin and the underlying sensitive areas. Reflection coefficients of cork (0.88) and foam (0.91) based materials were evaluated with a hydrophone. An ExAblate 2000 MRgFUS system was used to simulate clinical treatment with discs of reflector materials placed in a near field underneath a gel phantom. MR thermometry was used to monitor temperature elevations as well as the integrity of the focal spot. The phantom measurements showed acoustic shadow zones behind the reflectors with zone depths changing between 7 and 27 mm, for reflector disc diameters increasing from 10 to 30 mm (40 mm diameter discs completely blocked the FUS beam at the depth evaluated). The effects on thermal lesions due to the presence of the reflectors in the FUS beam were found to diminish with decreasing disc diameter and increasing sonication depth. For a 20 mm diameter disc and beyond 50 mm sonication depth, thermal lesions were minimally affected by the presence of the disc. No heating was observed on the skin side of the foam reflectors, as confirmed by measurements performed

  2. A simple analytic treatment of linear growth of structure with baryon acoustic oscillations

    NASA Astrophysics Data System (ADS)

    Slepian, Zachary; Eisenstein, Daniel J.

    2016-03-01

    In linear perturbation theory, all information about the growth of structure is contained in the Green's function, or equivalently, transfer function. These functions are generally computed using numerical codes or by phenomenological fitting formula anchored in accurate analytic results in the limits of large and small scale. Here, we present a framework for analytically solving all scales, in particular the intermediate scales relevant for the baryon acoustic oscillations (BAO). We solve for the Green's function and transfer function using spherically averaged overdensities and the approximation that the density of the coupled baryon-photon fluid is constant interior to the sound horizon.

  3. 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 as the last Space Shuttle Main Engine is being removed, it can be seen on the left side of the image frame. Note that one of the Orbiter Maneuvering System/ Reaction Control System has been removed while one of them remains. Also note that the body flap, below the engine positions has a protective covering to prevent damage to the High-temperature Reusable Surface Insulation tiles. 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

  4. 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 as the last Space Shuttle Main Engine is being removed, it can be seen on the right side of the image frame. Note that one of the Orbiter Maneuvering System/ Reaction Control System has been removed while one of them remains. Also note that the body flap, below the engine positions has a protective covering to prevent damage to the High-temperature Reusable Surface Insulation tiles. 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

  5. Computational investigation of slot blowing for fuselage forebody flow control

    NASA Technical Reports Server (NTRS)

    Murman, Scott M.; Rizk, Yehia M.; Schiff, Lewis B.; Cummings, Russell M.

    1992-01-01

    This paper presents a computational investigation of a tangential slot blowing concept for generating lateral control forces on an aircraft fuselage forebody. The effects of varying both the jet width and jet exit velocity for a fixed location slot are analyzed. This work is aimed at aiding researchers in designing future experimental and computational models of tangential slot blowing. The primary influence on the resulting side force of the forebody is seen to be the jet mass flow rate. This influence is sensitive to different combinations of slot widths and jet velocities over the range of variables considered. Both an actuator plane and an overset grid technique are used to model the tangential slot. The overset method successfully resolves the details of the actual slot geometry, extending the generality of the numerical method. The actuator plane concept predicts side forces similar to those produced by resolving the actual slot geometry.

  6. Acoustic measurements of F-4E aircraft operating in hush house, NSN 4920-02-070-2721

    NASA Astrophysics Data System (ADS)

    Miller, V. R.; Plzak, G. A.; Chinn, J. M.

    1981-09-01

    The primary purpose of this test program was to measure the acoustic environment in the hush house facility located at Kelly Air Force Base, Texas, during operation of the F-4E aircraft to ensure that aircraft structural acoustic design limits were not exceeded. The acoustic measurements showed that sonic fatigue problems are anticipated with the F-4E aircraft aft fuselage structure during operation in the hush house. The measured acoustic levels were less than those measured in an F-4E aircraft water cooled hush house at Hill AFB in the lower frequencies, but were increased over that measured during ground run up on some areas of the aircraft. It was recommended that the acoustic loads measured in this program should be specified in the structural design criteria for aircraft which will be subjected to hush house operation or defining requirements for associated equipment. Recommendations were also made to increase the fatigue life of the aft fuselage.

  7. 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.

  8. Development of pressure containment and damage tolerance technology for composite fuselage structures in large transport aircraft

    NASA Technical Reports Server (NTRS)

    Smith, P. J.; Thomson, L. W.; Wilson, R. D.

    1986-01-01

    NASA sponsored composites research and development programs were set in place to develop the critical engineering technologies in large transport aircraft structures. This NASA-Boeing program focused on the critical issues of damage tolerance and pressure containment generic to the fuselage structure of large pressurized aircraft. Skin-stringer and honeycomb sandwich composite fuselage shell designs were evaluated to resolve these issues. Analyses were developed to model the structural response of the fuselage shell designs, and a development test program evaluated the selected design configurations to appropriate load conditions.

  9. 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.

  10. Experimental and numerical analyses of laminar boundary-layer flow stability over an aircraft fuselage forebody

    NASA Technical Reports Server (NTRS)

    Vijgen, Paul M. H. W.; Holmes, Bruce J.

    1987-01-01

    Fuelled by a need to reduce viscous drag of airframes, significant advances have been made in the last decade to design lifting surface geometries with considerable amounts of laminar flow. In contrast to the present understanding of practical limits for natural laminar flow over lifting surfaces, limited experimental results are available examining applicability of natural laminar flow over axisymmetric and nonaxisymmetric fuselage shapes at relevantly high length Reynolds numbers. The drag benefits attainable by realizing laminar flow over nonlifting aircraft components such as fuselages and nacelles are shown. A flight experiment to investigate transition location and transition mode over the forward fuselage of a light twin engine propeller driven airplane is examined.

  11. Effect of low-level laser treatment on cochlea hair-cell recovery after acute acoustic trauma

    NASA Astrophysics Data System (ADS)

    Rhee, Chung-Ku; Bahk, Chan Woong; Kim, Se Hyung; Ahn, Jin-Chul; Jung, Jae Yun; Chung, Phil-Sang; Suh, Myung-Whan

    2012-06-01

    We investigated the effect of low-level laser radiation on rescuing hair cells of the cochlea after acute acoustic trauma and hearing loss. Nine rats were exposed to noise. Starting the following day, the left ears (NL ears) of the rats were irradiated at an energy output of 100 to 165 mW/cm2 for 60 min for 12 days in a row. The right ears (N ears) were considered as the control group. Frequency-specific hearing levels were measured before the noise exposure and also after the 1st, 3rd to 5th, 8th to 10th and 12th irradiations. After the 12th treatment, hair cells were observed using a scanning electron microscope. Compared to initial hearing levels at all frequencies, thresholds increased markedly after noise exposure. After the 12th irradiation, hearing threshold was significantly lower for the NL ears compared to the N ears. When observed using an electron microscope, the number of hair cells in the middle turn of the NL ears was significantly larger than that of the N ears. Our findings suggest that low-level laser irradiation promotes recovery of hearing thresholds after acute acoustic trauma.

  12. Spinning mode sound propagation in ducts with acoustic treatment and sheared flow

    NASA Technical Reports Server (NTRS)

    Rice, E. J.

    1975-01-01

    The propagation of spinning mode sound was considered for a cylindrical duct with sheared steady flow. Calculations concentrated on the determination of the wall optimum acoustic impedance and the maximum possible attenuation. Both the least attenuated and higher radial modes for spinning lobe patterns were considered. A parametric study was conducted over a wide range of Mach numbers, spinning lobe numbers, sound frequency, and boundary layer thickness. A correlation equation was developed from theoretical considerations starting with the thin boundary layer approximation of Eversman. This correlation agrees well with the more exact calculations for inlets and provides a single boundary layer refraction parameter which determines the change in optimum wall impedance due to refraction effects.

  13. 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.

  14. 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.

  15. Payload specialists Baudry and Chretien in the Shuttle full fuselage trainer

    NASA Technical Reports Server (NTRS)

    1984-01-01

    Payload specialists Patrick Baudry (left) and Jean Loup Chretien are seated at the commander and pilot stations on the Shuttle full fuselage trainer. In this view they are looking at the camera over the backs of the shuttle seats.

  16. STS 51-E crew is briefed on the Shuttle full fuselage trainer

    NASA Technical Reports Server (NTRS)

    1984-01-01

    STS 51-E crew is briefed on the Shuttle full fuselage trainer. Astronauts Dave Griggs (foreground), Jean Loup Chretien (behind Griggs) and Jeff Hoffman are being shown the workings of the trainer by flight instructors.

  17. Vertical Drop Test of a YS-11 Fuselage Section (Part 2)

    NASA Astrophysics Data System (ADS)

    Iwasaki, Kazuo; Kumakura, Ikuo; Minegishi, Masakutsu; Shoji, Hirokazu; Yoshimoto, Norio; Miyaki, Hiromitsu; Terada, Hiroyuki; Isoe, Akira; Yamaoka, Toshihiro; Katayama, Noriaki; Hayashi, Toru; Akaso, Tetsuya; Kosaka, Hideyuki

    The Structures and Materials Research Center of the National Aerospace Laboratory of Japan (NAL) and Kawasaki Heavy Industries, Ltd. (KHI) conducted the 2nd vertical drop test of a fuselage section cut from a NAMC YS-11 transport airplane in July 2002. The main objective of this test program was to obtain background data for aircraft cabin safety by drop test of a full-scale fuselage section and to develop computational tool for crash simulation of aircraft fuselage structure. The test article including seats and anthropomorphic test dummies was dropped to a rigid impact surface by free-fall method at a velocity of 7.6m/s (25ft/s). The impact environment and the resultant response of the fuselage structure and the passenger dummies were considered to be severe but potentially survivable. A description of the results of the 1st drop test and the 2nd drop test is presented in this paper.

  18. Application of a design-build-team approach to low cost and weight composite fuselage structure

    NASA Technical Reports Server (NTRS)

    Ilcewicz, L. B.; Walker, T. H.; Willden, K. S.; Swanson, G. D.; Truslove, G.; Metschan, S. L.; Pfahl, C. L.

    1991-01-01

    Relationships between manufacturing costs and design details must be understood to promote the application of advanced composite technologies to transport fuselage structures. A team approach, integrating the disciplines responsible for aircraft structural design and manufacturing, was developed to perform cost and weight trade studies for a twenty-foot diameter aft fuselage section. Baseline composite design and manufacturing concepts were selected for large quadrant panels in crown, side, and keel areas of the fuselage section. The associated technical issues were also identified. Detailed evaluation of crown panels indicated the potential for large weight savings and costs competitive with aluminum technology in the 1995 timeframe. Different processes and material forms were selected for the various elements that comprise the fuselage structure. Additional cost and weight savings potential was estimated for future advancements.

  19. Coupled rotor-flexible fuselage vibration reduction using open loop higher harmonic control

    NASA Technical Reports Server (NTRS)

    Papavassiliou, I.; Friedmann, P. P.; Venkatesan, C.

    1991-01-01

    A fundamental study of vibration prediction and vibration reduction in helicopters using active controls was performed. The nonlinear equations of motion for a coupled rotor/flexible fuselage system have been derived using computer algebra on a special purpose symbolic computer facility. The trim state and vibratory response of the helicopter are obtained in a single pass by applying the harmonic balance technique and simultaneously satisfying the trim and the vibratory response of the helicopter for all rotor and fuselage degrees of freedom. The influence of the fuselage flexibility on the vibratory response is studied. It is shown that the conventional single frequency higher harmonic control is capable of reducing either the hub loads or only the fuselage vibrations but not both simultaneously. It is demonstrated that for simultaneous reduction of hub shears and fuselae vibrations a new scheme called multiple higher harmonic control is required.

  20. 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.

  1. Low-Speed Aerodynamic Characteristics of a Fuselage Model with Various Arrangements of Elongated Lift Jets

    NASA Technical Reports Server (NTRS)

    Vogler, R. D.; Goodson, K. W.

    1973-01-01

    Data were obtained for a round jet located on the center of the bottom of a fuselage and for elongated slots separated spanwise by distances of 0.8 and 1.2 of the fuselage width. The effect of yawing the slots, inclining the jets laterally, and combining slot yaw with jet inclination was determined. Data were obtained in and out of ground effect through a range of effective velocity ratios and through a range of sideslip angles.

  2. Full-scale engine tests of bulk absorber acoustic inlet treatment

    NASA Technical Reports Server (NTRS)

    Heidelberg, L. J.; Homyak, L.

    1979-01-01

    Three different densities of Keviar bulk absorber fan inlet treatment were tested on a YF 102 turbofan engine. This bulk absorber material may have potential for flight application. Farfield noise measurements were made and the attenuation properties of the three treatment densities were compared. In addition, the best bulk treatment was compared to the best single degree of freedom, SDOF (honeycomb and perforated cover sheet) treatment from another investigation. Although the density was varied over a large range, (3 to 1) the effect on attenuation was small. The highest density treatment, 11.8 lb/cubic ft., had a somewhat broader attenuation bandwidth. The comparison of the best bulk and SDOF treatments showed the bulk to have a much greater attenuation bandwidth. At the design frequency both types of treatment had almost equal performance.

  3. Full-scale engine tests of bulk absorber acoustic inlet treatment

    NASA Technical Reports Server (NTRS)

    Heidelberg, L. J.; Homyak, L.

    1979-01-01

    Three different densities of Kevlar bulk absorber fan inlet treatment were tested on a YF 102 turbofan engine. This bulk absorber material may have potential for flight application. Far-field noise measurements were made and the attenuation properties of the three treatment densities were compared. In addition the best bulk treatment was compared to the best single degree of freedom, SDOF (honeycomb and perforated cover sheet) treatment from another investigation. Although the density was varied over a large range, (3 to 1) the effect on attenuation was small. The highest density treatment, 11.8 lb/cu ft, had a somewhat broader attenuation bandwidth. The comparison of the best bulk and SDOF treatments showed the bulk to have a greater attenuation bandwidth. At the design frequency both types of treatment had almost equal performance.

  4. 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.

  5. A Study of the Utilization of Advanced Composites in Fuselage Structures of Commercial Aircraft

    NASA Technical Reports Server (NTRS)

    Watts, D. J.; Sumida, P. T.; Bunin, B. L.; Janicki, G. S.; Walker, J. V.; Fox, B. R.

    1985-01-01

    A study was conducted to define the technology and data needed to support the introduction of advanced composites in the future production of fuselage structure in large transport aircraft. Fuselage structures of six candidate airplanes were evaluated for the baseline component. The MD-100 was selected on the basis of its representation of 1990s fuselage structure, an available data base, its impact on the schedule and cost of the development program, and its availability and suitability for flight service evaluation. Acceptance criteria were defined, technology issues were identified, and a composite fuselage technology development plan, including full-scale tests, was identified. The plan was based on composite materials to be available in the mid to late 1980s. Program resources required to develop composite fuselage technology are estimated at a rough order of magnitude to be 877 man-years exclusive of the bird strike and impact dynamic test components. A conceptual composite fuselage was designed, retaining the basic MD-100 structural arrangement for doors, windows, wing, wheel wells, cockpit enclosure, major bulkheads, etc., resulting in a 32 percent weight savings.

  6. 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.

  7. A comparison of two approaches to the treatment of chronic cough: perceptual, acoustic, and electroglottographic outcomes.

    PubMed

    Vertigan, Anne E; Theodoros, Deborah G; Winkworth, Alison L; Gibson, Peter G

    2008-09-01

    Voice problems have been reported to occur in association with chronic cough (CC) and can interfere with quality of life. Voice symptoms can improve following behavioral intervention for CC that persists despite medical management; however, formal measures of voice changes have not been reported. The aim of this study was to measure the changes in perceptual, acoustic, and electroglottographic voice characteristics after a SPEech Pathology Intervention Program for CHronic Cough (SPEICH-C) compared to a Healthy Lifestyle Education intervention program (HLE control). Eighty-two participants with CC that was refractory to medical management were randomly allocated to receive either the SPEICH-C or an HLE control. Participants in the SPEICH-C group demonstrated a significant reduction in perceptual ratings of breathy, rough, strain, and glottal fry qualities (P<0.001) in comparison to the HLE control group. There was a significant improvement between pre- and postintervention maximum phonation time, jitter, and harmonic-to-noise ratio values in the SPEICH-C group; however, the magnitude of change was not significantly different between groups. There was no significant change in fundamental frequency, standard deviation of fundamental frequency, phonation range, or closed phase of vocal fold vibration after intervention for either group. These results demonstrated that SPEICH-C can improve perceptual aspects of voice quality suggesting that dysphonia may be a fundamental characteristic of CC. PMID:17485195

  8. Thermal and acoustic problems on root canal treatment with different lasers

    NASA Astrophysics Data System (ADS)

    Ertl, Thomas P.; Benthin, Hartmut; Mueller, Gerhard J.

    1994-12-01

    Side effects of root canal preparation with lasers such as the generation of acoustic shockwaves and heat transfer were investigated. Shockwaves may cause disintegration of root hard substance and too high temperatures may damage the periodontium. Three types of pulsed lasers with different ablation characteristics were chosen for the study. (1) Excimer laser 308 nm/120 ns. (2) Er:YSGG laser 2.78 micrometers /500 microsecond(s) . (3) Nd:YAG laser 1.06 micrometers /180 microsecond(s) . Delivery systems for all lasers were quartz fibers with 400 micrometers core diameter. Canals were pretreated up to size 40 to obtain a comparable root canal shape. The teeth were positioned with the root in chicken egg protein as a heatsink during the laser operation. Shockwaves were measured with a needle hydrophone and visualization of the ablation process was made with high speed flashlamp photography. Temperatures were measured with a fiberoptic device. Results show that lasers with medium pulse length, operating at wavelength highly absorbed by hard and soft tissue, caused minimum side effects. The ablation process with lasers emitting at a low absorbed wavelength rapidly shifts from an initial heat transfer at the beginning of preparation to a noncontrollable ablation and temperature rise when carbonization occurs in the canal. Very short pulsed lasers such as excimer lasers cause stronger shock waves than lasers with a pulse length in the microsecond(s) region. One can conclude that Er:YSGG lasers offer the best ratio between efficiency and side effects.

  9. In situ processing methods for composite fuselage sandwich structures

    NASA Technical Reports Server (NTRS)

    Saatchi, Hossein; Durako, Bill; Reynolds, Dick; Dost, Ernest; Willden, Kurtis

    1993-01-01

    Conventional sandwich structure fabrication methods are labor intensive and high in cost. A low cost method is needed to produce lightweight sandwich structures. Sundstrand has developed a series of in situ composite fabrication methods in which the raw materials (skin and core materials) are placed in a closed mold, and the component is produced in one heating cycle. Internal pressure is generated by chemical agents during the thermal cycles, which consolidates the skins and produces the foam core. The finished part is a net-shape composite sandwich structure with skins and a foamed core. The in situ process reduces cost by eliminating several secondary operations that are used in conventional fabrication methods. Further, a strong molecular bond is produced between the core and skin, which eliminates adhesive bonding and prevents a weak bond section in the sandwich structure. In this investigation, we evaluated the feasibility of the in situ process using thermoset materials currently under consideration for commercial airplane fuselage applications, such as keel sections. The materials used were Hercules 855340 toughened epoxy resin in both liquid and powder forms, and 3M Scotchply PR500 resin, manufactured by 3M Corporation, in powder form. We successfully foamed these resins and produced experimental panels with AS-4/855340 Hercules prepreg skins. Chopped fibers were added to the core to increase performance of the foam. Mechanical property testing on these panels showed properties competitive with other foams. Additional experiments are required to optimize the in situ foam core sandwiches for specific properties and applications.

  10. Interference drag in a simulated wing-fuselage juncture

    NASA Technical Reports Server (NTRS)

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

    1984-01-01

    The interference drag in a wing fuselage juncture as simulated by a flat plate and a body of constant thickness having a 1.5:1 elliptical leading edge is evaluated experimentally. The experimental measurements consist of mean velocity data taken with a hot wire at a streamwise location corresponding to 16 body widths downstream of the body leading edge. From these data, the interference drag is determined by calculating the total momentum deficit (momentum area) in the juncture and also in the two dimensional turbulent boundary layers on the flat plate and body at locations sufficiently far from the juncture flow effect. The interference drag caused by the juncture drag as measured at this particular streamwise station is -3% of the total drag due to the flat plate and body boundary layers in isolation. If the body is considered to be a wing having a chord and span equal to 16 body widths, the interference drag due to the juncture is only -1% of the frictional drag of one surface of such a wing.

  11. Closeup view of the aft fuselage of the Orbiter Discovery ...

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

    Close-up view of the aft fuselage of the Orbiter Discovery looking at the thrust structure that supports the Space Shuttle Main Engines (SSMEs). In this view, SSME number two position is on the left and SSME number three position is on the right. The thrust structure transfers the forces produce by the engines into and through the airframe of the orbiter. The thrust structure includes the SSMEs load reaction truss structure, engine interface fittings and the hydraulic-actuator support structure. The propellant feed lines are the plugged and capped orifices within the engine bays. Note that SSME position two is rotated ninety degrees from position three and one. This was needed to enable enough clearance for the engines to fit and gimbal. Note in engine bay three is a clear view of the actuators that control the gambling of that engine. This view was taken from a service platform in 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

  12. Residual Strength Prediction of Fuselage Structures with Multiple Site Damage

    NASA Technical Reports Server (NTRS)

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

    1999-01-01

    This paper summarizes recent results on simulating full-scale pressure tests of wide body, lap-jointed fuselage panels with multiple site damage (MSD). The crack tip opening angle (CTOA) fracture criterion and the FRANC3D/STAGS software program were used to analyze stable crack growth under conditions of general yielding. The link-up of multiple cracks and residual strength of damaged structures were predicted. Elastic-plastic finite element analysis based on the von Mises yield criterion and incremental flow theory with small strain assumption was used. A global-local modeling procedure was employed in the numerical analyses. Stress distributions from the numerical simulations are compared with strain gage measurements. Analysis results show that accurate representation of the load transfer through the rivets is crucial for the model to predict the stress distribution accurately. Predicted crack growth and residual strength are compared with test data. Observed and predicted results both indicate that the occurrence of small MSD cracks substantially reduces the residual strength. Modeling fatigue closure is essential to capture the fracture behavior during the early stable crack growth. Breakage of a tear strap can have a major influence on residual strength prediction.

  13. Mechanisms of active control in cylindrical fuselage structures

    NASA Technical Reports Server (NTRS)

    Silcox, R. J.; Lester, H. C.; Fuller, C. R.

    1987-01-01

    This paper summarizes ongoing efforts to understand and exploit active control techniques for low frequency noise suppression in aerospace applications. Analytical models are utilized in an effort to understand the mechanisms that govern noise transmission into acoustic spaces enclosed by lightweight structures and to examine the results of experimental implementations of active control schemes. Emphasis is placed on attaining global noise reductions using a minimum number of actuators rather than localized control over many subregions. This program has demonstrated the effect of synchrophasing and interface modal filtering, in limiting the modal density within the acoustic space, and how strong reactive effects may occur in two dimensional geometries. Finally, the performance of active control systems utilizing acoustic and vibration actuators is evaluated. Suppressions of 10 to 30 dB are demonstrated in practice, and performance is discussed in relation to the physical mechanisms and parameters of the system.

  14. Treatment of acoustic fluid-structure interaction by Localized Lagrange Multipliers and comparison to alternative interface coupling methods.

    SciTech Connect

    Felippa, Carlos A.; Sprague, Michael A.; Ross, Michael R.; Park, K. C.

    2008-11-01

    This paper is a sequel on the topic of localized Lagrange multipliers (LLM) for applications of fluid-structure interaction (FSI) between finite-element models of an acoustic fluid and an elastic structure. The prequel paper formulated the spatial-discretization methods, the LLM interface treatment, the time-marching partitioned analysis procedures, and the application to 1D benchmark problems. Here, we expand on formulation aspects required for successful application to more realistic 2D and 3D problems. Additional topics include duality relations at the fluid-structure interface, partitioned vibration analysis, reduced-order modeling, handling of curved interface surfaces, and comparison of LLM with other coupling methods. Emphasis is given to non-matching fluid-structure meshes. We present benchmark examples that illustrate the benefits and drawbacks of competing interface treatments. Realistic application problems involving the seismic response of two existing dams are considered. These include 2D modal analyses of the Koyna gravity dam, transient-response analyses of that dam with and without reduced-order modeling, incorporation of nonlinear cavitation effects, and the 3D transient-response analysis of the Morrow Point arch dam.

  15. 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

  16. Effects of forward velocity and acoustic treatment on inlet fan noise

    NASA Technical Reports Server (NTRS)

    Feiler, C. E.; Merriman, J. E.

    1974-01-01

    Flyover and static noise data from several engines are presented that show inlet fan noise measured in flight can be lower than that projected from static tests for some engines. The differences between flight and static measurements appear greatest when the fan fundamental tone due to rotor-stator interaction or to the rotor-alone field is below cutoff. Data from engine and fan tests involving inlet treatment on the walls only are presented that show the attenuation from this treatment is substantially larger than expected from previous theories or flow duct experience. Data showing noise shielding effects due to the location of the engine on the airplane are also presented. These observations suggest that multiringed inlets may not be necessary to achieve the desired noise reduction in many applications.

  17. Treatment of a sloping fluid-solid interface and sediment layering with the seismo-acoustic parabolic equation.

    PubMed

    Collins, Michael D; Siegmann, William L

    2015-01-01

    The parabolic equation method is extended to handle problems in seismo-acoustics that have multiple fluid and solid layers, continuous depth dependence within layers, and sloping interfaces between layers. The medium is approximated in terms of a series of range-independent regions, and a single-scattering approximation is used to compute transmitted fields across the vertical interfaces between regions. The approach is implemented in terms of a set of dependent variables that is well suited to piecewise continuous depth dependence in the elastic parameters, but one of the fluid-solid interface conditions in that formulation involves a second derivative that complicates the treatment of sloping interfaces. This issue is resolved by using a non-centered, four-point difference formula for the second derivative. The approach is implemented using a matrix decomposition that is efficient when the parameters of the medium have a general dependence within the upper layers of the sediment but only depend on depth in the water column and deep within the sediment. PMID:25618077

  18. Twenty years' experience in the treatment of acoustic neuromas with fractionated radiotherapy: A review of 45 cases

    SciTech Connect

    Maire, Jean-Philippe . E-mail: jean-philippe.maire@chu-bordeaux.fr; Huchet, Aymeri; Milbeo, Yann; Darrouzet, Vincent; Causse, Nicole; Celerier, Denis; Liguoro, Dominique; Bebear, Jean-Pierre

    2006-09-01

    Purpose: To evaluate very long-term results of fractionated radiotherapy (FRT) of acoustic neuromas (AN). Methods and Materials: From January 1986 to January 2004, FRT was performed in 45 consecutive patients (46 AN). Indications were as follows: poor general condition contraindicating surgery, hearing preservation in bilateral neuromas, partial resection, nonsurgical recurrence. A 3-field to 5-field technique with static beams was used. A mean total dose of 51 Gy was given (1.80 Gy/fraction). The median tumor diameter was 31 mm (range, 11-55 mm). The median follow-up from FRT was 80 months (range, 4-227 months). Results: The particularity of our series consists of a very long-term follow-up of FRT given to selected patients. Nineteen patients died, two with progressive disease, and 17 from non-AN causes. A serviceable level of hearing was preserved in 7/9 hearing patients. No patient had facial or trigeminal neuropathy. Tumor shrinkage was observed in 27 (59%) and stable disease in 16 (35%). Tumor progression occurred in three patients, 12 to 15 months after FRT. Two additional tumors recurred after shrinkage 20 and 216 months after treatment and were operated on. Actuarial local tumor control rates at 5 and 15 years were 86%. For the patient who had a tumor recurrence at 216 months, histologic examination documented transformation to a low-grade malignant peripheral nerve sheath tumor. Conclusion: Very long-term efficacy of FRT is well documented in this series. However, our results suggest that malignant transformation can occur many years after FRT so we advocate caution when using this treatment for young patients.

  19. Acoustic Treatment Design Scaling Methods. Volume 5; Analytical and Experimental Data Correlation

    NASA Technical Reports Server (NTRS)

    Chien, W. E.; Kraft, R. E.; Syed, A. A.

    1999-01-01

    The primary purpose of the study presented in this volume is to present the results and data analysis of in-duct transmission loss measurements. Transmission loss testing was performed on full-scale, 1/2-scale, and 115-scale treatment panel samples. The objective of the study was to compare predicted and measured transmission loss for full-scale and subscale panels in an attempt to evaluate the variations in suppression between full- and subscale panels which were ostensibly of equivalent design. Generally, the results indicated an unsatisfactory agreement between measurement and prediction, even for full-scale. This was attributable to difficulties encountered in obtaining sufficiently accurate test results, even with extraordinary care in calibrating the instrumentation and performing the test. Test difficulties precluded the ability to make measurements at frequencies high enough to be representative of subscale liners. It is concluded that transmission loss measurements without ducts and data acquisition facilities specifically designed to operate with the precision and complexity required for high subscale frequency ranges are inadequate for evaluation of subscale treatment effects.

  20. Analysis, design, and test of acoustic treatment in a laboratory inlet duct

    NASA Technical Reports Server (NTRS)

    Kraft, R. E.; Motsinger, R. E.; Gauden, W. H.; Link, J. F.

    1979-01-01

    A suppression prediction program based on the method of modal analysis for spinning mode propagation in a circular duct was used in the analytical design of optimized, multielement, Kevlar bulk-absorber treatment configurations for an inlet duct. The NASA-Langley ANRL anechoic chamber using the spinning mode synthesizer as a sound source was used to obtain in-duct spinning mode measurements, radial mode measurements, and far-field traverses, as well as aerodynamic measurements. The measured suppression values were compared to predicted values, using the in-duct, forward-traveling, radial-mode content as the source for the prediction. The performance of the treatment panels was evaluated from the predicted and measured data. Although experimental difficulties were encountered at the design condition, sufficient information was obtained to confirm the expectation that it is the panel impedance components which are critical to suppression at a single frequency, not the particular construction materials. The agreement obtained between measurement and prediction indicates that the analytical program can be used as an accurate, reliable, and useful design tool.

  1. Development of Rene' 41 honeycomb structure as an integral cryogenic tankage/fuselage concept for future space transportation systems

    NASA Technical Reports Server (NTRS)

    Shideler, J. J.; Swegle, A. R.; Fields, R. A.

    1982-01-01

    The status of the structural development of an integral cryogenic-tankage/hot-fuselage concept for future space transportation systems (STS) is discussed. The concept consists of a honeycomb sandwich structure which serves the combined functions of containment of cryogenic fuel, support of vehicle loads, and thermal protection from an entry heating environment. The inner face sheet is exposed to a cryogenic (LH2) temperature of -423 F during boost; and the outer face sheet, which is slotted to reduce thermal stress, is exposed to a maximum temperature of 1400 F during a high altitude, gliding entry. A fabrication process for a Rene' 41 honeycomb sandwich panel with a core density less than 1 percent was developed which is consistent with desirable heat treatment processes for high strength.

  2. Acoustic Neuroma

    MedlinePlus

    An acoustic neuroma is a benign tumor that develops on the nerve that connects the ear to the brain. ... can press against the brain, becoming life-threatening. Acoustic neuroma can be difficult to diagnose, because the ...

  3. 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.

  4. Flow fields and aerodynamic characteristics for hypersonic missiles with mid-fuselage inlets

    NASA Technical Reports Server (NTRS)

    Hunt, J. L.; Johnston, P. J.; Riebe, G. D.

    1983-01-01

    A study was made to quantify forebody flow fields and to evaluate aerodynamic performance trends on a matrix of fuselage shapes for the mid-inlet/bolt-on-engine class of hypersonic airbreathing missiles for the Navy's vertical box launcher. The study indicated that inlet mass flow and pressure recovery can be increased by cambering the nose and increasing the width of the fuselage at both Mach 4 acceleration and Mach 6 cruise conditions. Aerodynamic trim predictions show that the drag at zero lift at Mach 4 decreases while the L/D max at Mach 6 increases with the nose camber, although these tendencies reverse with increasing width of maximum fuselage cross section.

  5. 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.

  6. 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.

  7. Leeward centerline and side fuselage entry heating predictions for the space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Helms, V. T., III

    1983-01-01

    Heat transfer data measured along the leeward centerline and on the side fuselage of the Space Shuttle orbiter during STS-2 and STS-3 are compared with predictions of empirical heating techniques derived from wind tunnel tests. Steps required to extrapolate an existing leeward centerline theory to flight conditions are described. Generally favorable comparisons from Mach 24 down to approximately Mach 7 for both flights are presented. The side fuselage impingement heating method is currently under development, but some preliminary results are available. The method is briefly described and compared with wind tunnel and flight measurements. Side heating predictions are given for an STS-3 trajectory point near Mach 10 showing good agreement with flight data. There is evidence of embedded vortices emanating from the side fuselage impingement line which significantly enhance local heating rates at both wind tunnel and flight conditions.

  8. Acoustic Seal

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M. (Inventor)

    2006-01-01

    The invention relates to a sealing device having an acoustic resonator. The acoustic resonator is adapted to create acoustic waveforms to generate a sealing pressure barrier blocking fluid flow from a high pressure area to a lower pressure area. The sealing device permits noncontacting sealing operation. The sealing device may include a resonant-macrosonic-synthesis (RMS) resonator.

  9. Acoustic seal

    NASA Technical Reports Server (NTRS)

    Steinetz, Bruce M. (Inventor)

    2006-01-01

    The invention relates to a sealing device having an acoustic resonator. The acoustic resonator is adapted to create acoustic waveforms to generate a sealing pressure barrier blocking fluid flow from a high pressure area to a lower pressure area. The sealing device permits noncontacting sealing operation. The sealing device may include a resonant-macrosonic-synthesis (RMS) resonator.

  10. 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