Thermomechanical fatigue (TMF) crack growth of Alloy 718 when the cycle includes long high-temperature tensile dwells is often accelerated because sustained-load crack growth occurs during the dwells and a temperature affected zone (TAZ) develops ahead of a crack tip that reduces fatigue crack growth resistance upon application of a cycle after the long dwell. This paper presents new TMF and sustained-load tests that aim to study and characterize the TAZ development and growth ahead of semi-elliptical and single-edge cracks. The influence of constraint on crack growth is assessed. This investigation focuses on a coarse-grained (CG) Alloy 718 at 650 degrees C. By comparing the crack growth behavior between a fine-grained and CG microstructure after a long tensile dwell, a modification of the TAZ growth prediction model that includes the influence of grain size is proposed. Using this improved model for TAZ growth and acceleration, the TMF crack growth prediction of surface cracks is improved. Temperature affected zone (TAZ) develops during tensile dwells in Alloy 718. The extent of TAZ is reduced when grains are larger. The extent of TAZ is also dependent on the constraint near the crack tip. A new grain size dependent TAZ growth model improves TMF predictions.

Isothermal and thermomechanical fatigue crack growth tests up to and above the maximum usage temperature for Inconel 718, nominally, 650°C, were executed to characterize the impact of time spent at elevated temperature on the fatigue crack growth rate. Various spectra were tested which showed tensile holds executed at a spectrum’s maximum load or immediately following an increase in load increased the fatigue crack growth rate. Additionally, a series of tests were executed to evaluate the impact of the stress intensity, duration, and temperature of a tensile hold on fatigue crack growth rate. The concept of a thermally affected zone to describe an area ahead of the crack tip weakened by the tensile hold that accelerates crack propagation is presented. As the stress intensity, duration, or temperature of a tensile hold increase, the temperature affected zone increases in size leading to an increase in fatigue crack growth rate for subsequent cycling.

Cracks starting at surfaces will grow under fatigue loading conditions both along the surface and in the thickness directions of the component geometry. In those cases where the crack grows through the thickness, the fatigue crack may transition to a corresponding through crack geometry. While the fatigue crack growth behaviour of both the surface flaws and complete through cracks are well understood, the method for modelling the process by which they transition from one to the other is not. This paper seeks to bring greater clarity and understanding to the transition process by implementing a transition method and developing the associated codes and equations to do so based on careful consideration of boundary conditions, experimental data, and finite element simulations.

Hot sections in turbine engines are subjected to large variations in temperatures and mechanical/thermal loadings. As such, accurate predictions of fatigue crack growth must account for many physical phenomena: temperature dependent crack growth behavior, load interaction history effects, time at temperature effects, temperature history effects, and the effects of stress/temperature/time on the materials. Through extensive experimental work on superalloys, a very definite “temperature history effect” on the resulting crack growth behavior has been identified and modeled. This work also identified a Temperature Affected Zone (TAZ) that occurs ahead of the crack tip and affects subsequent crack growth rates. The size of the TAZ is dependent on temperature, hold time, and stress state. Measurements of the TAZ were made under various conditions. The changes that occur in this TAZ are a combination of oxidation and material microstructure evolution. Various simplified “hot section” engine spectra (changing temperatures and stress levels) were tested to determine resulting crack growth behavior. Correlation between the experiments and model predictions were good and generally conservative.

Accurate characterization and understanding of the fatigue crack growth behaviour of components in jet turbine engines is critical for successfully using a damage tolerant design method to maximise safety and efficiency. The hot section components experience changing loads and temperatures, and hence, fatigue crack growth rates are typically studied under thermomechanical loading. One question that remains unclear is the role of the compressive holds that are often part of an aircraft loading-temperature spectrum. This experimental study was undertaken to investigate a turbine disk alloy, Inconel 718, subjected to different cycling and temperature profiles considering different lengths of hot compressive holds to determine its effect on the fatigue crack growth rate. It was found that the addition of a compressive hold at temperatures from 650 to 725 degrees C has no significant impact on the fatigue crack growth rate when compared with a cycle without a compressive hold. Fractographic analysis shows that crack growth is primarily transgranular in all cases studied suggesting that grain boundary oxidation, often observed during hot tensile holds, is insignificant.

ABSTRACTThis research looked at the effect of crystallographic orientation and temperature on the fatigue crack growth rate and the resulting fracture surface morphology in PWA1484 single crystal superalloy. Two groups of single edge notch tension specimens, one group with controlled secondary orientations and one group with uncontrolled secondary orientation, were tested at temperatures from 649°C to 982°C at R‐ratios of 0.1 and 0.7. It was found that the effect of temperature on the crack growth rate becomes more pronounced as the crack driving force increases while the secondary orientation and R‐ratio effects on the crack growth rate increase with increasing temperature. Two types of crack surface morphology were seen during fractography. The first was a precipitate avoidance (γ′ avoidance) morphology that was rolling but still predominantly flat when observed on a larger scale. In <001> primary oriented specimens, this fracture mode tended to follow the precipitate/matrix faces (microscopically cubic) while macroscopically staying essentially normal to the applied loads. The second mode was a form of cleavage (γ′ shearing) and occurred predominantly on octahedral crystallographic planes.

An approach was developed to predict the thermo-mechanical fatigue crack growth rates under typical gas turbine engine spectrum loading conditions. The material studied in the development of this model was a polycrystalline superalloy, Inconel 100. Load interaction effects were determined to have a major effect on the crack growth life. A yield zone load interaction life prediction model was modified to include temperature dependent properties. Multiple overload effects were included in the model to incorporate enhanced retardation compared to single overload retardation behavior. Temperature interaction effects were included and proved to be very important because of the wide temperature ranges to which turbine engine components are subjected. The effects of oxidation and temperature changes were accounted for in the model by accelerating crack growth in regions that had been previously affected by elevated temperatures. Experimental data of isolated, first order effects were used to calibrate and verify the model. Temperature dependent mechanical properties were determined and were essential in the model’s development. Parametric studies were performed using this model to assess the sensitivity of specific crack growth variables on life predictions.

Abstract A study was conducted to explore some of the load and temperature interaction effects on the fatigue crack growth rate (FCGR) of polycrystalline superalloy IN100. Load interaction testing in the form of single overloads was performed at 316°C and 649°C. Temperature interaction testing was performed by cycling between 316°C and 649°C in blocks of 1, 10, and 100 cycles. After compiling a database of constant temperature, constant amplitude FCGR for IN100, fatigue crack growth predictions assuming no load or temperature interactions were made. Experimental fatigue crack propagation data were then compared with these predictions to assess interaction effects. The fracture mechanisms observed during interaction testing using a scanning electron microscope were compared with the mechanisms present during constant temperature, constant amplitude testing. Overload interaction testing led to full crack retardation at 2.0 × overloads for both 316°C and 649°C testing. Overloading by 1.6 × at both temperatures led to retarded crack growth, whereas 1.3 × overloads at 649°C created accelerated crack growth and at 316°C the crack growth was retarded. One block alternating temperature interaction testing grew significantly faster than the non-interaction prediction, while 10 block alternating temperature interaction testing also grew faster but not to the same extent. One hundred block alternating testing grew slower than non-interaction predictions. Possible explanations for the interaction effects responsible for the observed crack growth acceleration and retardation are discussed.

ABSTRACTA study is undertaken to investigate the fatigue crack growth rate properties of polycrystalline IN100 through the identification of crack growth mechanisms as a function of temperature, frequency and ΔK. An additional goal is to determine the stress free activation energy of IN100. Constant amplitude, load controlled tests are performed at room temperature (22 °C), 316 °C, 482 °C and 649 °C under two different loading frequencies of 20 and 0.33 Hz. These specimens are then analysed via scanning electron microscopy (SEM) to determine failure mechanisms. SEM shows that, as temperature increased from room temperature to 649 °C, the fracture mechanism transitions from transgranular to intergranular. The fracture mechanism is shown to transition from intergranular to transgranular at elevated temperatures as da/dN increases as a result of growing ΔK. Scanning electron microscopy shows that, as frequency decreases from 20 to 0.33 Hz at 649 °C, the fracture mechanism transitions from transgranular to intergranular.

This paper investigates the tensile and fatigue properties of a newly developed fibre metal laminate (FML) manufactured using the vacuum assisted resin transfer moulding (VARTM) method. This manufacturing method allows the glass fibre reinforced epoxy and 2024-T3 aluminium FML to be prepared at lower cost than conventionally manufactured FMLs. However, in order for the resin to infiltrate the FML, the metal sheets need to be perforated. These perforation holes act as crack initiators and reduce the FML's performance. Tension and fatigue test results of three different designs are reported and compared to mechanical property predictions. Additionally, single sheet Al alloy specimens were tested in order to analyse the influence of the drilling method.

This article presents the effects of several variables on the damage progression within a mechanically fastened graphite/epoxy composite joint. The variables included the composite lay-up, loading configuration (single shear vs. double shear), R-value, stress level, and damage mechanisms observed in each specimen. In situ X-ray of the individual laminates recorded the extent of damage, mostly longitudinal splitting and bearing delamination, as a function of the cycle count. The following lay-ups were investigated: [04/45/03/90/0]s, [45/90/−45/02/45/02/−45/0]s, [±5/65/(±5)2/−65/±5]s, and [±5/65/(±5)2/−65/5/65]s. All of these lay-ups are considered to be ‘hard’ lay-ups, much stiffer in the 0° direction than in the 90° direction. The stress levels at which detectable damage develops was determined. The researchers chose to apply 50,000 cycles at each stress level. Once damage was detected, the stress level was typically raised to 17.25 MPa (2.5 ksi). Another 50,000 cycles was then applied until the bolt hole diameter elongated by 10% of its original length. The damage length vs. stress level is plotted as a way to compare damage onset stresses and growth as a function of lay-up and stress ratio. The new ‘non-traditional’ lay-ups are shown to offer some unique advantages over traditional lay-ups.

Total fatigue life is traditionally composed of the time to crack initiation plus the time for the initiated crack to grow to a critical crack size. Fracture mechanics does reasonably well in predicting the growth portion but there is still a lot of uncertainty about the definition of an initiated crack and scatter associated with the number of cycles to “initiation”. This paper will review some of the history, logic and uses of the Equivalent Initial Flaw Size (EIFS) approach to total life prediction. In short, this is a method where found cracks are analytically grown backwards to time equal zero (time or cycles) to determine an initial flaw, referred to as an EIFS. By growing a number of found cracks back to time equal zero a distribution of EIFS can be established. Example of establishing this distribution are given for the C-130 aircraft with a 7075 aluminum structure and for gas powered turbine blades made of directional solidified super-alloys.

Fiber metal laminates (FMLs) are multi-component materials utilizing metals, fibers and matrix resins. Tailoring their properties is readily achievable by varying one or more of these components. Established FMLs like GLARE utilize aluminum foils, glass fibers and epoxy matrices and are manufactured using an autoclave. Two new processes for manufacturing FMLs using vacuum assisted resin transfer molding (VARTM) have been developed at the NASA Langley Research Center (LaRC). A description of these processes and the resulting FMLs are presented.

The room temperature tensile and fatigue response of non-perforated and perforated titanium for laminar flow control application was investigated both experimentally and analytically. Results showed that multiple perforations did not affect the tensile response, but did reduce the fatigue life. A two dimensional finite element stress analysis was used to determine that the stress fields from adjacent perforations did not influence one another. The stress fields around the holes did not overlap one another, allowing the materials to be modeled as a plate with a center hole. Fatigue life was predicted using an equivalent MW flow size approach to relate the experimental results to microstructural features of the titanium. Predictions using flaw sizes ranging from 1 to 15 microns correlated within a factor of 2 with the experimental results by using a flow stress of 260 MPa. By using two different flow stresses in the crack closure model and correcting for plasticity, the experimental results were bounded by the predictions for high applied stresses. Further analysis of the complex geometry of the perforations and the local material chemistry is needed to further understand the fatigue behavior of the perforated titanium.

This experimental exploratory study investigates the effects of embedding a ‘barrier’ layer within a graphite/epoxy composite during manufacture on reducing the permeability after thermal cycling and low-velocity impact events. The baseline composite material was an eight-ply graphite/epoxy system. Barrier layer candidates include aluminized Mylar®, aluminum foil, and two β-Ti 15-3 films. Cryogenic thermal cycling was performed on the hybrid composites and control composites to determine if the interleaved composites exhibited reduced permeability after thermal cycling as compared to the controls. Drop-weight impact tests were performed on the hybrids and controls to determine the effect of interleaving on the critical impact energy of the graphite/epoxy composites. The results of this research suggest that the addition of an embedded barrier layer can increase a graphite/epoxy composite’s resistance to low-velocity impacts that cause permeation. This research indicates that hybrid composites are promising materials for applications where extreme temperatures and moderate impact loads are experienced.

The mode I critical strain energy release rates, GIC, of two polymer matrix composites were experimentally measured at temperatures ranging from -196°C to 160°C. The two composite materials investigated in this study were IM7/PETI-5 and IM7/977-2. Double cantilever beam specimens were manufactured with a °[0°7/ ± 3°/0°7] lay-up. The experimental results showed that GI had a strong dependence on temperature above 25°C and minimal dependence on temperature below 25°C. In addition, it was found that IM7/PETI-5 was significantly tougher than IM7/977-2 at all the temperatures investigated. These materials can be considered suitable, in the future, for manufacturing cryogenic fuel tanks in space applications; therefore, understanding critical composite properties such as the interlaminar toughness at such temperatures is very important.

New polyimide matrix composite materials are leading candidates for aerospace structural applications due to their high strength to weight ratio and excellent mechanical properties at elevated temperatures. The high fatigue resistance of these composites often results in the bolts being the weak link of a structure. Aircraft-quality bolts made of 4340 steel with a minimum UTS = 1241 MPa (180 ksi) were tested in three-point bend fatigue-Two life prediction methodologies were accessed for bending stress: S-N curves and fracture mechanics. The tensile S-N curve from the Mil-Handbook-5 conservatively predicts the bending fatigue life and run-out stress. Crack growth data, in the form of da/dN versus Delta K, from the Damage Tolerant Design Handbook was converted to a versus N data using five geometric correction factors. None of the five correction factors accurately predict crack growth, but all five correction factors did conservatively predict crack growth.

The goal of this research was to evaluate possible damage suppression in multi-directional notched laminates through the use of slightly off-axis longitudinal plies. These `non-traditional' laminates were compared to laminates with 0° longitudinal plies for quasi-isotropic and hard lay-ups, for open hole tension, filled hole tension, and single-shear bearing loadings. In-situ radiographic inspections were performed to evaluate damage initiation, progression, and suppression. In notched tension, the non-traditional laminates were less strong and stiff than their traditional counterparts, but X-ray images clearly showed the suppression of longitudinal splitting and delamination. Under single-shear bearing loads, the non-traditional laminates demonstrated increased bearing resistance compared to traditional laminates, and the non-traditional laminates were less prone to damage in the longitudinal plies.