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At least 145 records · Page 8
All sourcesNASA NTRSDOE OSTI
NASA NTRS1952

An engineering method for estimating notch-size effect in fatigue tests on steel

Neuber's proposed method of calculating a practical factor of stress concentration for parts containing notches of arbitrary size depends on the knowledge of a "new material constant" which can be established only indirectly. In this paper, the new constant has been evaluated for a large variety of steels from fatigue tests reported in the literature, attention being confined to stresses near the endurance limit. Reasonably satisfactory results were obtained with the assumption that the constant depends only on the tensile strength of the steel. Even in cases where the notches were cracks of which only the depth was known, reasonably satisfactory agreement was found between calculated and experimental factors. It is also shown that the material constant can be used in an empirical formula to estimate the size effect on unnotched specimens tested in bending fatigue.

Kuhn, Paul↗
NASA NTRS1994

The Role of Crack Formation in Chevron-Notched Four-Point Bend Specimens

The failure sequence following crack formation in a chevron-notched four-point bend 1 specimen is examined in a parametric study using the Bluhm slice synthesis model. Premature failure resulting from crack formation forces which exceed those required to propagate a crack beyond alpha (min) is examined together with the critical crack length and critical crack front length. An energy based approach is used to establish factors which forecast the tendency of such premature failure due to crack formation for any selected chevron-notched geometry. A comparative study reveals that, for constant values of alpha (1) and alpha (0), the dimensionless beam compliance and stress intensity factor are essentially independent of specimen width and thickness. The chevron tip position, alpha (0) has its primary effect on the force required to initiate a sharp crack. Small values for alpha (0) maximize the stable region length, however, the premature failure tendency is also high for smaller alpha (0) values. Improvements in premature failure resistance can be realized for larger values of alpha (0) with only a minor reduction in the stable region length. The stable region length is also maximized for larger chevron based positions, alpha (1) but the chance for premature failure is also raised. Smaller base positions improve the premature failure resistance with only minor decreases in the stable region length. Chevron geometries having a good balance of premature failure resistance, stable region length, and crack front length are 0.20 less than or equal to alpha (0) is less than or equal to 0.30 and 0.70 is less than or equal to alpha (1) is less than or equal to 0.80.

Calomino, Anthony M.↗
NASA NTRS2004

Notch Antennas

Notch antennas, also known as the tapered slot antenna (TSA), have been the topics of research for decades. TSA has demonstrated multi-octave bandwidth, moderate gain (7 to 10 dB), and symmetric E- and H- plane beam patterns and can be used for many different applications. This chapter summarizes the research activities on notch antennas over the past decade with emphasis on their most recent advances and applications. This chapter begins with some discussions on the designs of single TSA; then follows with detailed discussions of issues associated with TSA designs and performance characteristics. To conclude the chapter, some recent developments in TSA arrays and their applications are highlighted.

Lee, Richard Q.↗
NASA NTRS2011

Structural Efficiency and Behavior of Pristine and Notched Stitched Structure

Two driving factors in aircraft panel design are structural efficiency and response to in-service damage. Stitching through the thickness can improve both of these considerations. Combining stitching with a post-buckling design approach can provide additional improvements. The buckling behavior of stitched structure is considered since lighter structures can be achieved if local skin buckling is allowed to occur at less than design ultimate load. Through-the-thickness stitching can suppress delamination between skin and flange, thereby allowing the structure to reliably carry load into the postbuckling range. Hat-stiffened and rod-stiffened panels in which the skin and flanges were stitched together through-the-thickness prior to curing are considered through experiment and analysis. In both types of panels no mechanical fasteners were used for the assembly. Specimens were loaded to failure in axial compression. In this study all specimens buckled in the skin between the stiffeners and continued to carry load. In addition, the behavior of panels with a severed stringer or notch are considered. Failure loads and strain distributions in the notched panel are compared to those in the unnotched panel.

Jegley, Dawn C.↗
NASA NTRS2012

Oxidation and the Effects of High Temperature Exposures on Notched Fatigue Life of an Advanced Powder Metallurgy Disk Superalloy

Oxidation and the effects of high temperature exposures on notched fatigue life were considered for a powder metallurgy processed supersolvus heat-treated ME3 disk superalloy. The isothermal static oxidation response at 704 C, 760 C, and 815 C was consistent with other chromia forming nickel-based superalloys: a TiO2-Cr2O3 external oxide formed with a branched Al2O3 internal subscale that extended into a recrystallized - dissolution layer. These surface changes can potentially impact disk durability, making layer growth rates important. Growth of the external scales and dissolution layers followed a cubic rate law, while Al2O3 subscales followed a parabolic rate law. Cr- rich M23C6 carbides at the grain boundaries dissolved to help sustain Cr2O3 growth to depths about 12 times thicker than the scale. The effect of prior exposures was examined through notched low cycle fatigue tests performed to failure in air at 704 C. Prior exposures led to pronounced debits of up to 99 % in fatigue life, where fatigue life decreased inversely with exposure time. Exposures that produced roughly equivalent 1 m thick external scales at the various isotherms showed statistically equivalent fatigue lives, establishing that surface damage drives fatigue debit, not exposure temperature. Fractographic evaluation indicated the failure mode for the pre-exposed specimens involved surface crack initiations that shifted with exposure from predominately single intergranular initiations with transgranular propagation to multi-initiations from the cracked external oxide with intergranular propagation. Weakened grain boundaries at the surface resulting from the M23C6 carbide dissolution are partially responsible for the intergranular cracking. Removing the scale and subscale while leaving a layer where M23C6 carbides were dissolved did not lead to a significant fatigue life improvement, however, also removing the M23C6 carbide dissolution layer led to nearly full recovery of life, with a transgranular initiation typical to that observed in unexposed specimens.

Sudbrack, Chantal K.↗
NASA NTRS2015

An Analysis Methodology to Predict Damage Propagation in Notched Composite Fuselage Structures

A new methodology is proposed for predicting damage propagation in notched composite skin-stiffened structures. The proposed approach considers the interaction of damage propagation in the skin with delamination of the stiffener in order to assess the damage containment behavior of the structure. The damage propagating from a notch within a thin fiber-reinforced polymer skin is idealized as a through-the-thickness cohesive crack. The cohesive law for the through-the-thickness crack is characterized using a compact tension test. The approach was applied to a full-scale pultruded rod stitched efficient unitized structure (PRSEUS) concept fuselage panel that was tested recently. The stitched skin/stringer interfaces, a key feature of the PRSEUS concept, were modeled to assess the effectiveness of the stitching at containing damage propagation. Comparison between the predicted and observed damage extension shows acceptable agreement throughout loading. These results indicate that the model can represent accurately the complex interactions between a through-the-thickness crack in the skin and delamination between the skin and stringer. The model is used to demonstrate that the skin-stiffener interface toughness is critical to damage containment capability

Bergan, A.↗
NASA NTRS2016

Fatigue Analysis of Notched Laminates: A Time-Efficient Macro-Mechanical Approach

A coupled transversely isotropic deformation and damage fatigue model is implemented within the finite element method and was utilized along with a static progressive damage model to predict the fatigue life, stiffness degradation as a function of number of cycles, and post-fatigue tension and compression response of notched, multidirectional laminates. Initially, the material parameters for the fatigue model were obtained utilizing micromechanics simulations and the provided [0], [90] and [plus or minus 45] experimental composite laminate S-N (stress-cycle) data. Within the fatigue damage model, the transverse and shear properties of the plies were degraded with an isotropic scalar damage variable. The damage in the longitudinal (fiber) ply direction was suppressed, and only the strength of the fiber was degraded as a function of fatigue cycles. A maximum strain criterion was used to capture the failure in each element, and once this criterion was satisfied, the longitudinal stiffness of the element was decreased by a factor of 10 (sup 4). The resulting, degraded properties were then used to calculate the new stress state. This procedure was repeated until final failure of the composite laminate was achieved or a specified number of cycles reached. For post-fatigue tension and compression behavior, four internal state variables were used to control the damage and failure. The predictive capability of the above-mentioned approach was assessed by performing blind predictions of the notched multidirectional IM7/977-3 composite laminates response under fatigue and post-fatigue tensile and compressive loading, followed by a recalibration phase. Although three different multidirectional laminates were analyzed in the course of this study, only detailed results (i.e., stiffness degradation and post-fatigue stress-train curves as well as damage evolution states for a single laminate ([30/60/90/minus 30/minus 60] (sub 2s)) are discussed in detail here.

Composite Fatigue analysis↗
NASA NTRS2018

Validation of a Mesoscale Fiber Kinking Model Through Test and Analysis of Double Edge Notch Compression Specimens

Tests and analyses were conducted on a series of geometrically-scaled double edge notch compression specimens to validate the capability of the fiber-kinking model in the NASA continuum damage mechanics code for progressive damage analysis, CompDam. The tests and analyses focused on the fiber-kinking damage mechanism, which is of critical importance for predicting accurately longitudinal compression failure in carbon fiber reinforced polymer laminates. The fiber-kinking model in CompDam was augmented with a new capability to represent a hardening response once a kink band has fully formed, enabling prediction of band broadening. Correlation of the test and analysis results showed agreement in stiffness and strength with less than 10% error. The analysis predicts the same sequence of events leading to ultimate failure that was found in the test. The overall excellent correlation in terms of stiffness, strength, and failure process validates the capability of the model for predicting longitudinal compression failure in notched laminates with cross-ply layups.

Bergan, Andrew C.↗
NASA NTRS2020

Progressive Failure Analysis Correlation with Notched Composite Laminate Test Data

Testing of small and intermediate laminate panels with notches was performed and Progressive Failure Analyses (PFA) models were developed for aid in test planning and for correlation with the test data. Two progressive damage failure models were included in the study: the commercially available Abaqus built-in damage model and COmplete STress Reduction (COSTR) damage model developed at NASA Langley Research Center (LaRC). The finite elements models used for the analysis were developed using shell elements. The pre-test PFA results obtained from the two damage models were compared to test data. Then a post-test PFA with updated material properties based on additional available material property data was executed with the COSTR damage model and compared to the test data. The panels tested in compression all exhibited less scatter in the failure load with self-similar failure behavior. The panels tested in tension exhibited more scatter in the failure load with a failure mode that included delamination of plies. However, a self-similar crack path was also observed as the overall failure mode for all panels tested. When comparing test data to the pre-test analysis results, both damage models with shell element models were considered adequate for predicting the behavior and failure load of both the small and intermediate panels when tested in compression. However, for the tension load cases, larger differences were observed between the test and analysis results. When comparing test data to the post-test analysis performed using the COSTR damage model, better correlation was observed. However, large discrepancies were still observed for the tension load cases. Consequently, a higher fidelity finite element model including solid elements for sub-laminates and the modeling of cohesive layers between sub-laminate layers is recommended for tension loading of notched composite laminates.

Damage Model,Progressive Failure Analysis,Open Hol↗
NASA NTRSDate not supplied

Experimental Observations of Damage States in Unnotched and Notched 3D Orthogonal Woven Coupons Loaded in Tension

This paper describes a series of unnotched and notched tensile tests on two 3D orthogonal woven architectures. Dog bone specimens were used for the unnotched tests, and open-hole specimens were used for the notched tests. For each architecture and test type, two specimens were loaded under monotonic displacement control to failure and two specimens were incrementally loaded to failure in five steps. X-ray computed tomography (CT) was used to characterize the woven architecture and defects on each specimen prior to testing. Digital image correlation (DIC) images were obtained from both sides of the specimen throughout the loading and unloading. Acoustic emission (AE) data were also continuously acquired throughout the loading. For the incremental tests, X-ray CT was used to document the damage progression after each load increment. The damage progression was analyzed based on the results from these three techniques. The results for both architectures and test specimens followed a pattern where matrix cracks formed at the surface perpendicular to the tensile load direction. New surface cracks formed continuously throughout the loading and cracks extended in depth and width. In general, fiber failure was not detected until specimen failure.

Wade C Jackson↗
NASA NTRSDate not supplied

Experimental Observations of Damage States in Unnotched and Notched 3D Orthogonal Woven Coupons Loaded in Tension

This paper describes a series of unnotched and notched tensile tests on two 3D orthogonal woven architectures loaded in the warp direction. Dogbone specimens were used for the unnotched tests, and open-hole specimens were used for the notched tests. For each architecture and test type, two specimens were loaded under monotonic displacement control to failure, and two specimens were incrementally loaded to failure in five steps. X-ray computed tomography (CT) was used to characterize the woven architecture and defects on each specimen prior to testing. Digital image correlation (DIC) images were obtained from both sides of the specimen throughout the loading and unloading. For the incremental tests, X-ray CT was used to document the damage progression after each load increment. The results for both architectures and test specimens followed a similar pattern where matrix cracks formed perpendicular to the load direction and were primarily located between the stacks of warp tows. Two-piece failure path always occurred perpendicular to the loading direction, along the through-thickness segments of the Z-tows. The combination of stress concentrations due to the specimen geometry, weave geometry, and cut tows created stress concentrations that appear to initiate failure.

Woven materials↗