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At least 505 records · Page 28

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NASA NTRS2019

Experimental and Numerical Analysis of Skin-Stiffener Separation Using a Seven-Point Bend Configuration

Skin-stiffener separation in stiffened composite panels consists of a complex interaction between multiple scales of progressive damage and failure mechanisms. This work used a superposed cohesive element method of modeling resistance curve effects that represents the structural interface of a unidirectional tape skin and a fabric stiffener. Finite element models using cohesive elements with input properties obtained from tape-to-fabric interface characterization tests were compared to experimental results of a stable skin-stiffener separation characterization test using a seven-point bend fixture. This fixture deformed the stiffened panel specimens into a buckled configuration which induced mixed-mode interlaminar stress states at the skin-stiffener interface. The advantage of this fixture was the potential for more stable damage initiation and delamination growth compared to a stringer-stiffened panel under axial compression. Use of the superposed cohesive elements showed promise, but the characterization of interface material properties as input to the cohesive elements remains a crucial component to be investigated to ensure accurate prediction of skin-stiffener separation.

Kosztowny, Cyrus J. R.↗

NASA NTRSDate not supplied

Clearance Sensitivity Mitigation in Small Core Compressors

As future aircraft become lighter and more aerodynamically efficient, thrust requirements will decrease, reducing the core size of the engine. Furthermore, in the pursuit of improved fuel burn, engine overall pressure ratio and bypass ratio will increase, further driving down engine core size. These drivers together mean that the core size, or corrected mass flow rate at the compressor exit, for future single-aisle aircraft applications will shrink below 3.0 lbm/s. Traditionally, this small core compressor size is in the domain of axi-centrifugal designs, machines that are typically less efficient and limited to pressure ratios of ~25 due to stress and thermomechanical fatigue in the centrifugal impeller. In this light, NASA, Pratt & Whitney and the Raytheon Technologies Research Center embarked upon a program to develop technologies to enable an all-axial high-pressure compressor with a core size below 3.0 lbm/s and an overall pressure ratio greater than 50. One of the challenges with an all-axial high-pressure compressor at this core size is the small span at the rear of the compressor. As core size is scaled down, the rotor tip clearances do not scale with span, leading to significant efficiency penalties. This paper documents a numerical and experimental effort to mitigate this penalty through design optimization and technology insertion, enabling an estimated 5 to 10% fuel burn reduction relative to 2020 best-in-class. The experimental program consisted of two builds of a high-speed rig: a baseline build and a second build to insert technology to manage large rotor tip gap/span ratios. The results demonstrated a reduction in the sensitivity of the compressor to large clearance ratios, which would be likely at the end-of-life of a small core application. The test campaign demonstrated that, with the insertion of technology, small core compressors can maintain a similar efficiency to current best-in-class large core size compressors. In addition to advancing the state-of-the-art of technology, the program also advanced the modeling standards for multistage compressors with large clearance-to-span ratios. The validation of this approach is described in the paper.

Axial Compressors↗

NASA NTRSDate not supplied

Clearance Sensitivity Mitigation in Small Core Compressors

As future aircraft become lighter and more aerodynamically efficient, thrust requirements will decrease, reducing the core size of the engine. Furthermore, in the pursuit of improved fuel burn, engine overall pressure ratio and bypass ratio will increase, further driving down engine core size. These drivers together mean that the core size, or corrected mass flow rate at the compressor exit, for future single-aisle aircraft applications will shrink below 3.0 lb/s. Traditionally, this small core compressor size is in the domain of axi-centrifugal designs, machines that are typically less efficient and limited to pressure ratios of ~25 due to stress and thermomechanical fatigue in the centrifugal impeller. In this light, NASA, Pratt & Whitney and the Raytheon Technologies Research Center embarked upon a program to develop technologies to enable an all-axial high-pressure compressor with a core size below 3.0 lb/s and an overall pressure ratio greater than 50. One of the challenges with an all-axial high-pressure compressor at this core size is the small span at the rear of the compressor. As core size is scaled down, the rotor tip clearances do not scale with span, leading to significant efficiency penalties. This paper documents a numerical and experimental effort to mitigate this penalty through design optimization and technology insertion, enabling an estimated 5 to 10% fuel burn reduction relative to 2020 best-in-class. The experimental program consisted of two builds of a high-speed rig: a baseline build and a second build to insert technology to manage large rotor tip gap/span ratios. The results demonstrated a reduction in the sensitivity of the compressor to large clearance ratios, which would be likely at the end-of-life of a small core application. The test campaign demonstrated that, with the insertion of technology, small core compressors can maintain a similar efficiency to current best-in-class large core size compressors. In addition to advancing the state-of-the-art of technology, the program also advanced the modeling standards for multistage compressors with large clearance-to-span ratios. The validation of this approach is described in the paper.

Axial Compressors↗

DOE OSTI2024

Multidimensional modeling of fuel-cladding friction in an LWR fuel rod

Solving finite element problems with friction often increases the level of difficulty to obtain properly converged solutions. This type of challenge becomes more salient when advanced, possibly multiscale, material models are employed to capture the thermomechanical behavior of fuel and cladding materials. Here the present work details our recent developments in a nuclear fuel performance finite element code for the systematic consideration of friction in nuclear reactor finite element simulations. We show the application of friction and its effects on the mechanics of light-water reactor rods accounting for various fuel constitutive modeling techniques, model dimensionalities, pellet assumed geometries, and power conditions. In particular, we focus on the fuel rod mechanical behavior as it relates to fuel constitutive models, sensitivity to the coefficient of friction, pellet states of stress, and rod elongation. We discuss the trade-offs between the various multidimensional modeling options and highlight the relevance of frictional effects in the prediction of the fuel rod deformation and interfacial stresses. To relate our modeling results with actual reactor operation, simulations including frictional effects are compared with fuel rod elongation experimental data and some challenges for carrying out a full validation of the axial mechanics are discussed.

42 ENGINEERING↗

NASA NTRS1992

Theory of anisotropic thin-walled closed-section beams with hygrothermal effects

A variationally consistent theory is developed in order to derive the governing equations of anisotropic thin-walled beams with closed sections subjected to axial load, torsion, bending and hygrothermal loads. The theory is based on an asymptotical analysis of two-dimensional shell theory. Closed-form expressions for the beam stiffness coefficients, equivalent hygrothermal loads, stress and displacement fields are provided. The influence of material anisotropy and hygrothermal coefficients on the displacement field is identified. The analysis is applied to predict the response of composite beams under thermomechanical loading.

Badir, Ashraf M.↗

NASA NTRS1993

Turbulent structure of a wingtip vortex in the near field

The turbulent rollup of a vortex generated by a rectangular wing has been investigated. Extensive mean and turbulence measurements of the flowfield on a wingtip and in the near field have been completed. Velocity fluctuation measurements show that the near-field core is not laminar. A large axial velocity excess was found to exist in the core of the vortex. A momentum balance in the near-field of the wingtip showed that the magnitude of the core Reynolds-stress gradient terms are the same order as the largest terms in the governing equations. Navier-Stokes computations of the identical configuration, including wind tunnel walls and using measured inflow and outflow boundary conditions, reproduced many of the features of the experiment. Inherent limitations of the Baldwin-Barth turbulence model combined with limited grid resolution caused the computed vortex core to be more diffuse than desired. The momentum balance also demonstrated that the level of numerically generated false diffusion in the vortex core is relatively high.

Zilliac, Gregory G.↗

NASA NTRS1947

Effects of sweepback on boundary layer and separation

Following the law of stress adopted in the Navier-Stokes equations, the configuration of the viscous flow in planes at right angles to the axis of an infinite cylinder is found to be independent of the axial motion of the cylinder. In the limiting case of a yawed or swept wing of very high aspect ratio, certain boundary-layer and separation phenomena are thus determined independently by the crosswise component of velocity.

Jones, Robert T↗

NASA NTRS1947

Effects of sweepback on boundary layer and separation

Following a law of stress adopted in the Navier-Stokes equations, the configuration of the viscous flow in planes at right angles to the axis of an infinite cylinder is found to be independent of the axial motion of the cylinder. In the limiting case of a yawed or swept wing of very high aspect ratio, certain boundary-layer and separation phenomena are thus determined independently by the crosswise component of velocity. It follows that the effect of sweepback is to increase the area of stable laminar flow and to decrease the lift coefficient at which flow separation occurs.

Jones, Robert T↗

NASA NTRS1996

Characterization of the Dynamic Material Properties of Magnetostrictive Terfenol-D

A major limitation in use of electromagnetic and/or magnetomechanical models for design of Terfenol-D actuators is the lack of reliable material property data for Terfenol-D. In particular data on the performance of Terfenol-D as employed in a transducer, operating under real world dynamic conditions is needed. To provide this information, Terfenol-D rod properties need to be measured under as run prestressed and magnetically biased states. Using a Terfenol-D actuator, the following properties can be measured and/or calculated: mechanical quality factor, speed of sound in the material, the resonant frequency, the anti-resonant frequency, two magnetic permeabilities (one at constant stress and one at constant strain), two Young's moduli (one at constant amplitude applied magnetic field and one at constant amplitude magnetic flux density in the material), the magnetomechanical coupling, and the axial strain coefficient. The development of the material properties measurements and calculations is based on the model of low signal, linear, magnetostriction from Clark, the linear transduction equations for a transducer from Hunt, and a one degree of freedom mechanical model of the transducer. The electrical impedance and admittance mobility loops are used to determine the resonant, anti-resonant, and half power point frequencies. The rest of the material properties indicated above can then be calculated using these frequencies, acceleration from an accelerometer mounted on the actuator arm, and readily measurable transducer and Terfenol-D rod parameters.

Calkins, Frederick T.↗

NASA NTRS1967

Analysis of stability-critical orthotropic cylinders subjected to axial compression

Analytical procedure for determining critical buckling loads of orthotropic cylinders subjected to axial compression loading has been defined. Three modes of instability have been considered - general instability, local instability caused by panel and interframe buckling, and local instability caused by yielding and crippling in areas of stress concentration.

Finley, R. L.↗

NASA NTRS2003

Optimization of Elastically Tailored Tow-Placed Plates with Holes

Elastic stiffness tailoring of laminated composite panels by allowing the fibers to curve within the plane of the laminate is a design concept that has been demonstrated to be both beneficial and practical. The objective of the present paper is to demonstrate the effectiveness of stiffness tailoring through the use of curvilinear fibers to reduce stress concentrations around the hole and improve the load carrying capability of panels. Preliminary panel designs that are to be manufactured and tested were determined through design studies for flat plates without holes under axial compression using an optimization program. These candidate designs were then analyzed with finite element models that accurately reflect the test conditions and geometries in order to decide upon the final designs for manufacture and testing. An advanced tow-placement machine is used to manufacture the test panels with varying fiber orientation angles. A total of six large panels measuring three feet by six feet, each of which is used to produce four specimens with or without holes, are fabricated. The panels were machined into specimens with holes and tested at NASA Langley Research Center. Buckling response and failure of panels without holes and with two different hole dimensions are presented. Buckling and failure loads of tow-steered specimens are significantly greater than the buckling and failure loads of traditional straight-fiber specimens.

Jegley, Dawn C.↗

NASA NTRS1990

On the design and development of a miniature ceramic gimbal bearing

A review is made of a program to develop ceramic gimbal bearings for a miniaturized missile guidance system requiring nonmagnetic properties and higher load capacity than possible with conventional AISI 440C stainless steel bearings. A new gimbal design concept is described which utilizes the compressive strength and nonmagnetic properties of silicon nitride (Si3N4) ceramics for the gimbal bearing. Considerable manufacturing development has occurred in the last 5 years making ceramic bearings a viable option in the gimbal design phase. A preliminary study into the feasibility of the proposed design is summarized. Finite element analysis of the brittle ceramic bearing components under thermal stress and high acceleration loading were conducted to ensure the components will not fail catastrophically in service. Finite element analysis was also used to optimize the adhesive joint design. Bearing torque tests run at various axial loads indicate that the average running torque of ceramic bearings varies with load similarly to that of conventional steel bearings.

Hanson, Robert A.↗

NASA NTRS1990

Analyses of composite fuselage structure under various loading conditions

With the characteristics of high strength/weight and stiffness/weight ratios in advanced composite materials, composite plate structures have been successfully applied to secondary load-carrying structural components in the aerospace industry for the past two decades. Recently, filament wound composite shells are being considered for design of primary fuselage structures. To implement the design the structural response under various loading conditions needs to be predicted accurately. Therefore, it is essential to establish a rigorous analytical solution in the area of composite laminated shells. A closed form solution is presented that predicts the response of a composite shell subjected to internal pressure, axial tension, bending, and torsion. The material of the shell is assumed to be general cylindrically anisotropic. Based on the theory of cylindrical anisotropic elasticity coupled partial differential governing equations are developed using Lehknitskii's stress function approach. The general expressions for the stresses and displacements in the composite cylinders under these loading conditions is discussed. Three examples (45) off-axis unidirectional, (45/45) unsymmetric, and (45/45)(sub s) symmetric angle-ply fiber-reinforced laminated shells are shown to illustrate the effect of radius-to-thickness ratios, coupling, and stacking sequence.

Yuan, Fuh-Gwo↗

NASA NTRS2003

Failure Criterion For Isotropic Time Dependent Materials Which Accounts for Multi-Axial Loading

The Space Shuttle's Reusable Solid Rocket Motor (RSRM) nozzle program has recently conducted testing to characterize the effects of multi-axial loading, temperature and time on the failure characteristics of TIGA321, EA913NA, EA946 (three filled epoxy adhesives). From the test data a "Multi-Axial, Temperature, and Time Dependent" or MATT failure criterion was developed. It is shown that this criterion simplifies, for constant load and constant load rate conditions, into a form that can be easily used for stress analysis. Failure for TIGA321 and EA913NA are characterized below their glass transition temperature. Failure for EA946 is characterized for conditions that pass through its glass transition. The MATT failure criterion is shown to be accurate for a wide range of conditions for these adhesives.

Richardson, D. E.↗

NASA NTRS2007

A Semi-Analytical Method for Determining the Energy Release Rate of Cracks in Adhesively-Bonded Single-Lap Composite Joints

A semi-analytical method for determining the strain energy release rate due to a prescribed interface crack in an adhesively-bonded, single-lap composite joint subjected to axial tension is presented. The field equations in terms of displacements within the joint are formulated by using first-order shear deformable, laminated plate theory together with kinematic relations and force equilibrium conditions. The stress distributions for the adherends and adhesive are determined after the appropriate boundary and loading conditions are applied and the equations for the field displacements are solved. Based on the adhesive stress distributions, the forces at the crack tip are obtained and the strain energy release rate of the crack is determined by using the virtual crack closure technique (VCCT). Additionally, the test specimen geometry from both the ASTM D3165 and D1002 test standards are utilized during the derivation of the field equations in order to correlate analytical models with future test results. The system of second-order differential field equations is solved to provide the adherend and adhesive stress response using the symbolic computation tool, Maple 9. Finite element analyses using J-integral as well as VCCT were performed to verify the developed analytical model. The finite element analyses were conducted using the commercial finite element analysis software ABAQUS. The results determined using the analytical method correlated well with the results from the finite element analyses.

Yang, Charles↗

NASA NTRS1937

Stress Analysis of Beams with Shear Deformation of the Flanges

This report discusses the fundamental action of shear deformation of the flanges on the basis of simplifying assumptions. The theory is developed to the point of giving analytical solutions for simple cases of beams and of skin-stringer panels under axial load. Strain-gage tests on a tension panel and on a beam corresponding to these simple cases are described and the results are compared with analytical results. For wing beams, an approximate method of applying the theory is given. As an alternative, the construction of a mechanical analyzer is advocated.

Kuhn, Paul↗

NASA NTRS1997

Nonlinear Dynamic Analysis of Disordered Bladed-Disk Assemblies

In a effort to address current needs for efficient, air propulsion systems, we have developed some new analytical predictive tools for understanding and alleviating aircraft engine instabilities which have led to accelerated high cycle fatigue and catastrophic failures of these machines during flight. A frequent cause of failure in Jets engines is excessive resonant vibrations and stall flutter instabilities. The likelihood of these phenomena is reduced when designers employ the analytical models we have developed. These prediction models will ultimately increase the nation's competitiveness in producing high performance Jets engines with enhanced operability, energy economy, and safety. The objectives of our current threads of research in the final year are directed along two lines. First, we want to improve the current state of blade stress and aeromechanical reduced-ordered modeling of high bypass engine fans, Specifically, a new reduced-order iterative redesign tool for passively controlling the mechanical authority of shroudless, wide chord, laminated composite transonic bypass engine fans has been developed. Second, we aim to advance current understanding of aeromechanical feedback control of dynamic flow instabilities in axial flow compressors. A systematic theoretical evaluation of several approaches to aeromechanical feedback control of rotating stall in axial compressors has been conducted. Attached are abstracts of two .papers under preparation for the 1998 ASME Turbo Expo in Stockholm, Sweden sponsored under Grant No. NAG3-1571. Our goals during the final year under Grant No. NAG3-1571 is to enhance NASA's capabilities of forced response of turbomachines (such as NASA FREPS). We with continue our development of the reduced-ordered, three-dimensional component synthesis models for aeromechanical evaluation of integrated bladeddisk assemblies (i.e., the disk, non-identical bladeing etc.). We will complete our development of component systems design optimization strategies for specified vibratory stresses and increased fatigue life prediction of assembly components, and for specified frequency margins on the Campbell diagrams of turbomachines. Finally, we will integrate the developed codes with NASA's turbomachinery aeromechanics prediction capability (such as NASA FREPS).

McGee, Oliver G., III↗

DOE OSTI2026

A comparison of wrought and powder metallurgical FeCrAl claddings under simulated LWR accident transients

Iron-chromium-aluminum (FeCrAl) alloys are potential accident tolerant fuel (ATF) cladding candidates for light-water reactors but are difficult to fabricate as thin-walled tubes via conventional cast-and-wrought routes. Powder metallurgy (PM) offers a manufacturing alternative with improved compositional control, but its transient accident performance has not been directly benchmarked against wrought variants. This study evaluates the burst behavior of commercially developed PM-processed FeCrAl alloys, PM-C26M (Fe-12Cr-6Al-2Mo) and the high precipitate density FA-SMT (Fe-22Cr-5Al-3Mo), under simulated light-water reactor accident transient conditions. Burst testing was conducted using the Severe Accident Test Station with heating rates of 5 °C/s and 50 °C/s and internal pressures ranging from 25 MPa to 100 MPa. PM-C26M reproduced wrought C26M burst behavior within 7–37 °C across the stress range, indicating that PM processing does not compromise transient strength. FA-SMT exhibited markedly higher burst temperatures and reduced heating-rate sensitivity, consistent with its engineered precipitate strengthening. FA-SMT rupture exhibited axial "unzipping" rather than the lateral tearing characteristic of PM- and wrought C26M. Post-test EBSD and fractography indicate that this behavior is strongly correlated with strain-gated intergranular void nucleation associated with the dense precipitate architecture of FA-SMT, a response absent in the comparatively clean PM-C26M matrix and consistent with rupture morphologies reported for oxide-dispersion strengthened (ODS) FeCrAl of similar base-matrix chemistry to PM-C26M. These findings highlight the potential of powder metallurgy as a viable fabrication route for ATF claddings from an accident performance standpoint.

Bell, Sam [ORNL] (ORCID:0000000251905657)↗