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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Development of a finite element based delamination analysis for laminates subject to extension, bending, and torsion

Delamination is a common failure mode of laminated composite materials. This type of failure frequently occurs at the free edges of laminates where singular interlaminar stresses are developed due to the difference in Poisson's ratios between adjacent plies. Typically the delaminations develop between 90 degree plies and adjacent angle plies. Edge delamination has been studied by several investigators using a variety of techniques. Recently, Chan and Ochoa applied the quasi-three-dimensional finite element model to the analysis of a laminate subject to bending, extension, and torsion. This problem is of particular significance relative to the structural integrity of composite helicopter rotors. The task undertaken was to incorporate Chan and Ochoa's formulation into a Raju Q3DG program. The resulting program is capable of modeling extension, bending, and torsional mechanical loadings as well as thermal and hygroscopic loadings. The addition of the torsional and bending loading capability will provide the capability to perform a delamination analysis of a general unsymmetric laminate containing four cracks, each of a different length. The solutions obtained using this program are evaluated by comparing them with solutions from a full three-dimensional finite element solution. This comparison facilitates the assessment of three dimensional affects such as the warping constraint imposed by the load frame grips. It wlso facilitates the evaluation of the external load representation employed in the Q3D formulation. Finally, strain energy release rates computed from the three-dimensional results are compared with those predicted using the quasi-three-dimensional formulation.

Hooper, Steven J.↗
NASA NTRS1990

Damage states in laminated composite three-point bend specimens: An experimental-analytical correlation study

Damage states in laminated composites were studied by considering the model problem of a laminated beam subjected to three-point bending. A combination of experimental and theoretical research techniques was used to correlate the experimental results with the analytical stress distributions. The analytical solution procedure was based on the stress formulation approach of the mathematical theory of elasticity. The solution procedure is capable of calculating the ply-level stresses and beam displacements for any laminated beam of finite length using the generalized plane deformation or plane stress state assumption. Prior to conducting the experimental phase, the results from preliminary analyses were examined. Significant effects in the ply-level stress distributions were seen depending on the fiber orientation, aspect ratio, and whether or not a grouped or interspersed stacking sequence was used. The experimental investigation was conducted to determine the different damage modes in laminated three-point bend specimens. The test matrix consisted of three-point bend specimens of 0 deg unidirectional, cross-ply, and quasi-isotropic stacking sequences. The dependence of the damage initiation loads and ultimate failure loads were studied, and their relation to damage susceptibility and damage tolerance of the mean configuration was discussed. Damage modes were identified by visual inspection of the damaged specimens using an optical microscope. The four fundamental damage mechanisms identified were delaminations, matrix cracking, fiber breakage, and crushing. The correlation study between the experimental results and the analytical results were performed for the midspan deflection, indentation, damage modes, and damage susceptibility.

Starbuck, J. Michael↗
NASA NTRS1939

Stress Concentration Around an Open Circular Hole in a Plate Subjected to Bending Normal to the Plane of the Plate

An aluminum-alloy plate containing an open circular hole of diameter large compared with the thickness of the plate was subjected to bending forces normal to the plane of the plate. Deflection and strain measurements were taken for two different loads. Stress concentrations occurred at the edge of the hole and the maximum stresses were tangential to the hole at the ends of the transverse diameter. The maximum stress at the edge of the hole was 1.59 times the computed stress on the net section and 1.85 times the computed stress in a solid plate of the same dimensions subjected to the same bending forces. The maximum deflections were about 20 percent greater than the corresponding deflection for a solid plate of the same size subjected to the same bending forces. The smallest edge distance was equal to 2-1/2 times the diameter of the hole and the stress concentration on this side of the hole was the same as on the side where the edge distance was about 4-1/2 diameters. A theoretical analysis of the problem shows that, for an aluminum plate of infinite width, the stress concentration at the edge of the hole would be 1.87 times the stress in a solid plate, which is substantially the same relation obtained for the plate tested.

Dumont, C↗
NASA NTRS1954

Rapid estimation of bending frequencies of rotating beams

A procedure is presented in the form of charts which permits the rapid estimation of the natural bending frequencies of helicopter rotor blades, both rotating and nonrotating. Since the approach is based on Southwell's equation, an evaluation of the method with regard to such things as higher modes, blade offset, and variable mass and stiffness distributions is also given. The evaluation shows that when nonrotating beam bending modes are used, Southwell's equation yields reasonably accurate bending frequencies for rotating helicopter blades. Example comparisons of frequencies estimated using the charts with values given by the manufacturer for several actual blades show that the simplified procedure yields good practical results.

Robert T Yntema↗
NASA NTRS1945

Bending and shear stresses developed by the instantaneous arrest of the root of a cantilever beam rotating with constant angular velocity about a transverse axis through the root

A theoretical investigation was made of the behavior of a cantilever beam in rotational motion about a transverse axis through the root determining the stresses, the deflections, and the accelerations that occur in the beam as a result of the arrest of motion. The equations for bending and shear stress reveal that, at a given percentage of the distance from root to tip and at a given trip velocity, the bending stresses for a particular mode are independent of the length of the beam and the shear stresses vary inversely with the length. When examined with respect to a given angular velocity instead of a given tip velocity, the equations reveal that the bending stress is proportional to the length of the beam whereas the shear stress is independent of the length. Sufficient experimental verification of the theory has previously been given in connection with another problem of the same type.

BEAMS - VIBRATION↗
NASA NTRS1939

Experimental and analytical investigation of a monocoque wing model loaded in bending

Bending tests with transverse loads and with pure bending were undertaken on a double-web monocoque wing model in order to establish the relation between the state of stress and the results from the elementary bending theory. The longitudinal stresses in the stiffeners were measured with tensiometers and the shear stresses in the sheet were calculated from them. The measurements were made at both moderate loads with no buckles in the covering and at loads with which the critical stress in the individual panels was exceeded. For the comparison, the wing skin was considered as stiffened sheet according to the shear panel scheme. In this way, the statistically indeterminate calculation was confirmed by the test results.

Schapitz, E↗
NASA NTRS1993

Numerical and experimental investigation of the bending response of thin-walled composite cylinders

A numerical and experimental investigation of the bending behavior of six eight-ply graphite-epoxy circular cylinders is presented. Bending is induced by applying a known end-rotation to each end of the cylinders, analogous to a beam in bending. The cylinders have a nominal radius of 6 inches, a length-to-radius ratio of 2 and 5, and a radius-to-thickness ratio of approximately 160. A (+/- 45/0/90)S quasi-isotropic layup and two orthotropic layups, (+/- 45/0 sub 2)S and (+/- 45/90 sub 2)S, are studied. A geometrically nonlinear special-purpose analysis, based on Donnell's nonlinear shell equations, is developed to study the prebuckling responses and gain insight into the effects of non-ideal boundary conditions and initial geometric imperfections. A geometrically nonlinear finite element analysis is utilized to compare with the prebuckling solutions of the special-purpose analysis and to study the buckling and post buckling responses of both geometrically perfect and imperfect cylinders. The imperfect cylinder geometries are represented by an analytical approximation of the measured shape imperfections. Extensive experimental data are obtained from quasi-static tests of the cylinders using a test fixture specifically designed for the present investigation. A description of the test fixture is included. The experimental data are compared to predictions for both perfect and imperfect cylinder geometries. Prebuckling results are presented in the form of displacement and strain profiles. Buckling end-rotations, moments, and strains are reported, and predicted mode shapes are presented. Observed and predicted moment vs. end-rotation relations, deflection patterns, and strain profiles are illustrated for the post buckling responses. It is found that a geometrically nonlinear boundary layer behavior characterizes the prebuckling responses. The boundary layer behavior is sensitive to laminate orthotropy, cylinder geometry, initial geometric imperfections, applied end-rotation, and non-ideal boundary conditions. Buckling end-rotations, strains, and moments are influenced by laminate orthotropy and initial geometric imperfections. Measured buckling results correlate well with predictions for the geometrically imperfect specimens. The postbuckling analyses predict equilibrium paths with a number of scallop-shaped branches that correspond to unique deflection patterns. The observed postbuckling deflection patterns and measured strain profiles show striking similarities to the predictions in some cases. Ultimate failure of the cylinders is attributed to an interlaminar shear failure mode along the nodal lines of the postbuckling deflection patterns.

Fuchs, J. P.↗
NASA NTRS1994

Compliance measurements of chevron notched four point bend specimen

The experimental stress intensity factors for various chevron notched four point bend specimens are presented. The experimental compliance is verified using the analytical solution for a straight through crack four point bend specimen and the boundary integral equation method for one chevron geometry. Excellent agreement is obtained between the experimental and analytical results. In this report, stress intensity factors, loading displacements and crack mouth opening displacements are reported for different crack lengths and different chevron geometries, under four point bend loading condition.

Calomino, Anthony↗
NASA NTRS1993

Effects of cutouts on the behavior of symmetric composite laminates subjected to bending and twisting loads

Composite structures are used for a wide variety of aerospace applications. Practical structures contain cutouts and these structures are subjected to in-plane and out-of-plane loading conditions. Structurally efficient designs for composite structures require a thorough understanding of the effects of cutouts on the response of composite plates subjected to inplane or out-of-plane loadings. Most investigations of the behavior of composite plates with cutouts have considered in-plane loadings only. Out-of-plane loadings suchas bending or twisting have received very limited attention. The response of homogeneous plates (e.g., isotropic or orthotropic plates) subjected to bending or twisting moments has been studied analytically. These analyses are for infinite plates and neglect finite-plate effects. Recently, analytical and experimental studies were conducted to determine the effects of cutouts on the response of laminated composite plates subjected to bending moments. No analytical or experimental results are currently available for the effects of cutouts on the response of composite laminates subjected to twisting moments.

Prasad, C. B.↗
NASA NTRS1996

Investigation of Two-Phase Flows in Piping Bends and Elbows

An experimental investigation of the hydrodynamic characteristics of two-phase R-113 flow has been carried out. Straight tube pressure drop data, as a function of mass flow rate (mass flux) and flow quality has been obtained using the Two-Phase Flow Test Facility located in the Advanced Thermal Laboratories of the Crew and Thermal Systems Division at the Lyndon B. Johnson Space Center. Additionally, after successfully obtaining the straight tube pressure drop data, the test facility was modified in order to obtain pressure drop data for the flow of two-phase R-113 through 180 deg piping bends. Inherent instabilities of the test facility prevented the successful acquisition of pressure drop data through the piping bends. The experimental straight tube data will be presented and compared with existing predictive correlations in an attempt to gain insight into the utility of such correlations as the basis for developing design criteria. A discussion of the instabilities which rendered successful acquisition of the piping bend data will be presented and suggestions will be made for eliminating these system tendencies. Finally, recommendations for future investigations, based on successful reconfiguration of the test facility, will be made.

Duncan, Allen B.↗
NASA NTRS1959

Evaluation of Several Approximate Methods for Calculating the Symmetrical Bending-Moment Response of Flexible Airplanes to Isotropic Atmospheric Turbulence

Several approximate procedures for calculating the bending-moment response of flexible airplanes to continuous isotropic turbulence are presented and evaluated. The modal methods (the mode-displacement and force-summation methods) and a matrix method (segmented-wing method) are considered. These approximate procedures are applied to a simplified airplane for which an exact solution to the equation of motion can be obtained. The simplified airplane consists of a uniform beam with a concentrated fuselage mass at the center. Airplane motions are limited to vertical rigid-body translation and symmetrical wing bending deflections. Output power spectra of wing bending moments based on the exact transfer-function solutions are used as a basis for the evaluation of the approximate methods. It is shown that the force-summation and the matrix methods give satisfactory accuracy and that the mode-displacement method gives unsatisfactory accuracy.

Bennett, Floyd V.↗
NASA NTRS1999

An Experimental Study of a Stitched Composite with a Notch Subjected to Combined Bending and Tension Loading

A series of tests was conducted to measure the strength of stitched carbon/epoxy composites containing through-thickness damage in the form of a crack-like notch. The specimens were subjected to three types of loading: pure bending, pure tension, and combined bending and tension loads. Measurements of applied loads, strains near crack tips, and crack opening displacements (COD) were monitored in all tests. The transverse displacement at the center of the specimen was measured using a Linear Variable Differential Transformer (LVDT). The experimental data showed that the outer surface of the pure tension specimen failed at approximately 6,000 microstrain, while in combined bending and tension loads the measured tensile strains reached 10,000 microstrain.

Palmer, Susan O.↗
NASA NTRS1999

Elastic Moduli of Pyrolytic Boron Nitride Measured Using 3-Point Bending and Ultrasonic Testing

Three-point bending and ultrasonic testing were performed on a flat plate of PBN. In the bending experiment, the deformation mechanism was believed to be shear between the pyrolytic layers, which yielded a shear modulus, c (sub 44), of 2.60 plus or minus .31 GPa. Calculations based on the longitudinal and shear wave velocity measurements yielded values of 0.341 plus or minus 0.006 for Poisson's ratio, 10.34 plus or minus .30 GPa for the elastic modulus (c (sub 33)), and 3.85 plus or minus 0.02 GPa for the shear modulus (c (sub 44)). Since free basal dislocations have been reported to affect the value of c (sub 44) found using ultrasonic methods, the value from the bending experiment was assumed to be the more accurate value.

Kaforey, M. L.↗
NASA NTRS1961

Flight Investigation of Effects of Selected Operating Conditions on the Bending and Torsional Moments Encountered by a Helicopter Rotor Blade

Flight tests have been conducted with a single-rotor helicopter to determine the effects of partial-power descents with forward speed, high-speed level turns, pull-outs from autorotation, and high-forward-speed high-rotor-speed autorotation on the flapwise bending and torsional moments of the rotor blade. One blade of the helicopter was equipped at 14 percent and 40 percent of the blade radius with strain gages calibrated to measure moments rather than stresses. The results indicate that the maximum moments encountered in partial-power descents with forward speed tend to be generally reduced from the maximum moments encountered during partid-power descents at zero forward speed. High-speed level turns and pull-outs from auto-rotation caused retreating-blade stall which produced torsional moments (values up to 2,400 inch-pounds). at the 14-percent-radius station that were as large as those encountered during the previous investigations of retreating-blade stall (values up t o 2,500 inch-pounds). High-forward- speed high-rotor-speed autorotation produced flapwise bending moments (values up to 7,200 inch-pounds) at the 40-percent-radius station which were as large as the flapwise bending moments (values up to 7,800 inch-pounds) a t the 14-percent-radius station encountered during partial - power vertical descents. The results of the present investigation (tip-speed ratios up to 0.325 and an unaccelerated level-flight mean lift coefficient of about 0.6), in combination with the related results of at zero forward speed produce the largest rotor-blade vibratory moments. However, inasmuch as these large moments occur only during 1 percent of the cycles and 88 percent of the cycles are at moment values less than 70 percent of these maximum values in partial-power descents, other conditions, such as high-speed flight where the large moments are combined with large percentages of time spent,must not be neglected in any rotor-blade service-life assessment.

Ludi, LeRoy H.↗
NASA NTRS2002

Composite Bending Box Section Modal Vibration Fault Detection

One of the primary concerns with Composite construction in critical structures such as wings and stabilizers is that hidden faults and cracks can develop operationally. In the real world, catastrophic sudden failure can result from these undetected faults in composite structures. Vibration data incorporating a broad frequency modal approach, could detect significant changes prior to failure. The purpose of this report is to investigate the usefulness of frequency mode testing before and after bending and torsion loading on a composite bending Box Test section. This test article is representative of construction techniques being developed for the recent NASA Blended Wing Body Low Speed Vehicle Project. The Box section represents the construction technique on the proposed blended wing aircraft. Modal testing using an impact hammer provides an frequency fingerprint before and after bending and torsional loading. If a significant structural discontinuity develops, the vibration response is expected to change. The limitations of the data will be evaluated for future use as a non-destructive in-situ method of assessing hidden damage in similarly constructed composite wing assemblies. Modal vibration fault detection sensitivity to band-width, location and axis will be investigated. Do the sensor accelerometers need to be near the fault and or in the same axis? The response data used in this report was recorded at 17 locations using tri-axial accelerometers. The modal tests were conducted following 5 independent loading conditions before load to failure and 2 following load to failure over a period of 6 weeks. Redundant data was used to minimize effects from uncontrolled variables which could lead to incorrect interpretations. It will be shown that vibrational modes detected failure at many locations when skin de-bonding failures occurred near the center section. Important considerations are the axis selected and frequency range.

Werlink, Rudy↗
NASA NTRS2004

Ethnic Differences in Bending Stiffness of the Ulna and Tibia

There is considerable information about the variations in bone mass associated with different opportunity to compare a mechanical property of bone in young college women of Caucasian, Hispanic and Asian descent who gave informed consent to participate in an exercise study. The subjects were sedentary, in good health, eumenorrheic, non-smokers and had body mass indices (BMI) less than 30. Measurements acquired were body weight, kg, and height, cm, calcaneal and wrist bone density, g/square cm (PIXI, Lunar GE) and bending stiffness (EI, Nm(exp 2)) in the ulna and tibia. E1 was determined non-invasively with an instrument called the Mechanical Response Tissue Analyzer (MRTA) that delivers a vibratory stimulus to the center of the ulna or tibia and analyzes the response curve based on the equation E1 = k(sub b) L(exp 3)/48 where k, is lateral bending stiffness, L is the length of the bone, E is Young's modulus of elasticity and I, the bending moment of inertia. The error of the test (CV) based on measurements of an aluminum rod with a known E1 was 4.8%, of calcaneal BMD, 0.54%, and of wrist bone density, 3.45%.

Arnaud, S. B.↗
NASA NTRS1998

Experimental Observations of a Stitched Composite with a Notch Subjected to Combined Bending and Tension Loading

A series of tests was conducted to support development of an analytical model for predicting the failure strains of stitched warp-knit carbon/epoxy composite materials with through-thicknesss damage in the form of a crack-like notch. Measurements of strain near notch tips, crack opening displacement (COD), and applied load were monitored in all tests. The out-of-plane displacement at the center of the notch was also measured when the specimen was subjected to bending. Three types of loading were applied: pure bending, pure tension, and combined bending and tension.

Palmer, Susan O.↗
NASA NTRS2006

Analytical and Numerical Results for an Adhesively Bonded Joint Subjected to Pure Bending

A one-dimensional, semi-analytical methodology that was previously developed for evaluating adhesively bonded joints composed of anisotropic adherends and adhesives that exhibit inelastic material behavior is further verified in the present paper. A summary of the first-order differential equations and applied joint loading used to determine the adhesive response from the methodology are also presented. The method was previously verified against a variety of single-lap joint configurations from the literature that subjected the joints to cases of axial tension and pure bending. Using the same joint configuration and applied bending load presented in a study by Yang, the finite element analysis software ABAQUS was used to further verify the semi-analytical method. Linear static ABAQUS results are presented for two models, one with a coarse and one with a fine element meshing, that were used to verify convergence of the finite element analyses. Close agreement between the finite element results and the semi-analytical methodology were determined for both the shear and normal stress responses of the adhesive bondline. Thus, the semi-analytical methodology was successfully verified using the ABAQUS finite element software and a single-lap joint configuration subjected to pure bending.

Smeltzer, Stanley S., III↗