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All sourcesNASA NTRSDOE OSTI
NASA NTRS2000

Crack Turning in Integrally Stiffened Aircraft Structures

Current emphasis in the aircraft industry toward reducing manufacturing cost has created a renewed interest in integrally stiffened structures. Crack turning has been identified as an approach to improve the damage tolerance and fail-safety of this class of structures. A desired behavior is for skin cracks to turn before reaching a stiffener, instead of growing straight through. A crack in a pressurized fuselage encounters high T-stress as it nears the stiffener--a condition favorable to crack turning. Also, the tear resistance of aluminum alloys typically varies with crack orientation, a form of anisotropy that can influence the crack path. The present work addresses these issues with a study of crack turning in two-dimensions, including the effects of both T-stress and fracture anisotropy. Both effects are shown to have relation to the process zone size, an interaction that is central to this study. Following an introduction to the problem, the T-stress effect is studied for a slightly curved semi-infinite crack with a cohesive process zone, yielding a closed form expression for the future crack path in an infinite medium. For a given initial crack tip curvature and tensile T-stress, the crack path instability is found to increase with process zone size. Fracture orthotropy is treated using a simple function to interpolate between the two principal fracture resistance values in two-dimensions. An extension to three-dimensions interpolates between the six principal values of fracture resistance. Also discussed is the transition between mode I and mode II fracture in metals. For isotropic materials, there is evidence that the crack seeks out a direction of either local symmetry (pure mode I) or local asymmetry (pure mode II) growth. For orthotropic materials the favored states are not pure modal, and have mode mixity that is a function of crack orientation.

Pettit, Richard Glen↗
NASA NTRS2000

IC Engine Applications of Carbon-Carbon

Many of the properties of carbon-carbon make it an ideal material for reciprocating materials of intermittent combustion (IC) engines. Recent diesel engine tests, shown herein, indicate that the thermal and mechanical properties of carbon-carbon are adequate for piston applications, However, reducing the manufacturing costs and providing long term oxidation protection are still issues that need to be addressed.

Northam, G. Burton↗
NASA NTRS1984

Big Savings from Smart Motors

Chesebrough-Pond's operates 32 plants across the nation and in those plants are more than 10,000 electric motors. In an effort to cut down on waste of electrical power used by these motors, Chesebrough organized a Corporate Advanced Technology Group to devise ways of improving productivity and cut manufacturing costs. Chesebrough used NASA's Marshall Space Flight Center's Power Factor Controller technology as a departure point for development of their own computerized motor controller that enables motors to operate at maximum efficiency regardless of the motor's applications or operating condition.

Source record↗
NASA NTRS1980

Loudspeaker Performance Aid

Many manufacturers of loudspeakers are now using a magnetic liquid cooling agent known as ferrofluid. Commercialized by Ferrofluids Corporation, ferrofluid is a liquid material in which sub-microscopic particles of iron oxide are permanently suspended. Injected into the voice coil segment of speaker system, magnetic liquid serves as superior heat transfer medium for cooling the voice coil, thus substantially increasing the system's ability to handle higher power levels and decreasing chance of speaker failure. Ferrofluid offers several additional advantages which add up to improved speaker performance, lower manufacturing costs and fewer rejects.

Source record↗
NASA NTRS1981

Laser Balancing

Mechanical Technology, Incorporated developed a fully automatic laser machining process that allows more precise balancing removes metal faster, eliminates excess metal removal and other operator induced inaccuracies, and provides significant reduction in balancing time. Manufacturing costs are reduced as a result.

Source record↗
NASA NTRS2003

Residual Strength Analyses of Monolithic Structures

Finite-element fracture simulation methodology predicts the residual strength of damaged aircraft structures. The methodology uses the critical crack-tip-opening-angle (CTOA) fracture criterion to characterize the fracture behavior of the material. The CTOA fracture criterion assumes that stable crack growth occurs when the crack-tip angle reaches a constant critical value. The use of the CTOA criterion requires an elastic- plastic, finite-element analysis. The critical CTOA value is determined by simulating fracture behavior in laboratory specimens, such as a compact specimen, to obtain the angle that best fits the observed test behavior. The critical CTOA value appears to be independent of loading, crack length, and in-plane dimensions. However, it is a function of material thickness and local crack-front constraint. Modeling the local constraint requires either a three-dimensional analysis or a two-dimensional analysis with an approximation to account for the constraint effects. In recent times as the aircraft industry is leaning towards monolithic structures with the intention of reducing part count and manufacturing cost, there has been a consistent effort at NASA Langley to extend critical CTOA based numerical methodology in the analysis of integrally-stiffened panels.In this regard, a series of fracture tests were conducted on both flat and curved aluminum alloy integrally-stiffened panels. These flat panels were subjected to uniaxial tension and during the test, applied load-crack extension, out-of-plane displacements and local deformations around the crack tip region were measured. Compact and middle-crack tension specimens were tested to determine the critical angle (wc) using three-dimensional code (ZIP3D) and the plane-strain core height (hJ using two-dimensional code (STAGS). These values were then used in the STAGS analysis to predict the fracture behavior of the integrally-stiffened panels. The analyses modeled stable tearing, buckling, and crack branching at the integral stiffener using different values of critical CTOA for different material thicknesses and orientation. Comparisons were made between measured and predicted load-crack extension, out-of-plane displacements and local deformations around the crack tip region. Simultaneously, three-dimensional capabilities to model crack branching and to monitor stable crack growth of multiple cracks in a large thick integrally-stiffened flat panels were implemented in three-dimensional finite element code (ZIP3D) and tested by analyzing the integrally-stiffened panels tested at Alcoa. The residual strength of the panels predicted from STAGS and ZP3D code compared very well with experimental data. In recent times, STAGS software has been updated with new features and now one can have combinations of solid and shell elements in the residual strength analysis of integrally-stiffened panels.

Forth, Scott↗
NASA NTRS2003

Optimal Design and Damage Tolerance Verification of an Isogrid Structure for Helicopter Application

A composite isogrid panel design for application to a rotorcraft fuselage is presented. An optimum panel design for the lower fuselage of the rotorcraft that is subjected to combined in-plane compression and shear loads was generated using a design tool that utilizes a smeared-stiffener theory in conjunction with a genetic algorithm. A design feature was introduced along the edges of the panel that facilitates introduction of loads into the isogrid panel without producing undesirable local bending gradients. A low-cost manufacturing method for the isogrid panel that incorporates these design details is also presented. Axial compression tests were conducted on the undamaged and low-speed impact damaged panels to demonstrate the damage tolerance of this isogrid panel. A combined loading test fixture was designed and utilized that allowed simultaneous application of compression and shear loads to the test specimen. Results from finite element analyses are presented for the isogrid panel designs and these results are compared with experimental results. This study illustrates the isogrid concept to be a viable candidate for application to the helicopter lower fuselage structure.

Baker, Donald J.↗
NASA NTRS1999

Fabrication Of Carbon-Boron Reinforced Dry Polymer Matrix Composite Tape

Future generation aerospace vehicles will require specialized hybrid material forms for component structure fabrication. For this reason, high temperature composite prepregs in both dry and wet forms are being developed at NASA Langley Research Center (LaRC). In an attempt to improve compressive properties of carbon fiber reinforced composites, a hybrid carbon-boron tape was developed and used to fabricate composite laminates which were subsequently cut into flexural and compression specimens and tested. The hybrid material, given the designation HYCARB, was fabricated by modifying a previously developed process for the manufacture of dry polymer matrix composite (PMC) tape at LaRC. In this work, boron fibers were processed with IM7/LaRC(TradeMark)IAX poly(amide acid) solution-coated prepreg to form a dry hybrid tape for Automated Tow Placement (ATP). Boron fibers were encapsulated between two (2) layers of reduced volatile, low fiber areal weight poly(amide acid) solution-coated prepreg. The hybrid prepreg was then fully imidized and consolidated into a dry tape suitable for ATP. The fabrication of a hybrid boron material form for tow placement aids in the reduction of the overall manufacturing cost of boron reinforced composites, while realizing the improved compression strengths. Composite specimens were press-molded from the hybrid material and exhibited excellent mechanical properties.

Belvin, Harry L.↗
NASA NTRS2004

High Temperature VARTM with LaRC Polyimides

Recent work at NASA Langley Research Center (LaRC) has concentrated on developing new polyimide resin systems for advanced aerospace applications that can be processed without the use of an autoclave. Polyimide composites are very attractive for applications that require a high strength to weight ratio and thermal stability. Vacuum assisted resin transfer molding (VARTM) has shown potential to reduce the manufacturing cost of composite structures. In VARTM, the fibrous preform is infiltrated on a rigid tool surface contained beneath a flexible vacuum bag. Both resin injection and fiber compaction are achieved under pressures of 101.3 KPa or less. Recent studies have demonstrated the feasibility of the VARTM process for fabrication of void free structures utilizing epoxy resin systems with fiber volume fractions approaching 60%. In this work, the VARTM process has been extended to the fabrication of composite panels from polyimide systems developed at the Langley Research Center. This work has focused on processing LARC(trademark) PETI-8 (Langley Research Center Phenylethynyl Terminated Imide- 8), an aromatic polyimide based on 3,3',4,4' -biphenyltetracarboxylic dianhydride, a 50:50 molar ratio of 3,4'-oxydianiline and 1,3-bis(3-aminophenoxy)benzene, with 4-phenylethynylphthalic anhydride as the endcapping agent. Various molecular weight versions were investigated to determine their feasibility of being processed by VARTM at elevated temperatures. An injection temperature of approximately 280 C was required to achieve the necessary viscosity (<5 Poise) for flow at VARTM pressures. Laminate quality and initial mechanical properties are presented for LARC(trademark) PETI-8 and 6k IM7 uniweave fabric.

Cano, Roberto J.↗
NASA NTRS2004

Residual Strength Characterization of a Curved Integrally-Stiffened Panel

Over the years, Finite-element fracture simulation methodology has been very well established at NASA Langley to predict the residual strength of damaged aircraft structures. This methodology has been experimentally verified at NASA Langley for structures ranging from laboratory coupons up to full-scale built-up structural components with single and multiple-site damage cracking. The methodology uses the critical crack-tip-opening-angle (CTOA) fracture criterion to characterize the fracture behavior of the material. The CTOA fracture criterion assumes that stable crack growth occurs when the crack-tip angle reaches a constant critical value. The use of the CTOA criterion requires an elastic-plastic, finite-element analysis. The critical CTOA value is determined by simulating fracture behavior in laboratory specimens, such as a compact specimen, to obtain the angle that best fits the observed test behavior. The critical CTOA value appears to be independent of loading, crack length, and in-plane dimensions. However, it is a function of material thickness and local crack-front constraint. Modeling the local constraint requires either a three-dimensional analysis or a two-dimensional analysis with an approximation to account for the constraint effects. In recent times as the aircraft industry is leaning towards monolithic structures with the intension of reducing part count and manufacturing cost, there has been a consistent effort at NASA Langley to extend critical CTOA based numerical methodology in the analysis of integrally-stiffened panels. In this regard, a series of fracture tests were conducted on curved aluminum-alloy integrally-stiffened panels. These curved panels were subjected to uniaxial tension and pressure loading. During the test, applied load-crack extension, out-of-plane displacements and local deformations around the crack tip region were measured. Compact and middle-crack tension specimens were tested to determine the critical angle (psi(sub c) using three-dimensional code (ZIP3D) and the plane-strain core height (h(sub c)) using two-dimensional code (STAGS).

Seshadri, B. R.↗
NASA NTRS2005

Evaluation of a Compression-Loaded-Stitched-Multi-Bay Fuselage Panel With Barely Visible Impact Damage

The experimental results from a stitched VaRTM carbon-epoxy composite panel tested under uni-axial compression loading are presented along with nonlinear finite element analysis prediction of the response. The curved panel is divided by frames and stringers into six bays with a column of three bays along the compressive loading direction. The frames are supported at the frame ends to resist out-of-plane translation. Back-to-back strain gages are used to record the strain and displacement transducers were used to record the out-of-plane displacements. In addition a full-field-displacement measurement technique that utilizes a camera-based-stereo-vision system was used to record the displacements. The panel was loaded to 1.5 times the predicted initial buckling load (1st bay buckling load, P(sub er) from the nonlinear finite element analysis and then was removed from the test machine for impact testing. After impacting with 20 ft-lbs of energy using a spherical impactor to produce barely visible damage the panel was loaded in compression until failure. The buckling load of the first bay to buckle was 97% of the buckling load before impact. The stitching constrained the impact damage from growing during the loading to failure. Impact damage had very little overall effect on panel stiffness. Panel stiffness measured by the full-field-displacement technique indicated a 13% loss in stiffness after impact. The panel failed at 1.64 times the first panel buckling load. The barely visible impact damage did not grow noticeably as the panel failed by global instability due to stringer-web terminations at the frame locations. The predictions from the nonlinear analysis of the finite element modeling of the entire specimen were very effective in the capture of the initial buckling and global behavior of the panel. In addition, the prediction highlighted the weakness of the panel under compression due to stringer web terminations. Both the test results and the nonlinear predictions serve to reinforce the severe penalty in structural integrity caused by the low cost manufacturing technique to terminate the stringer webs, and demonstrates the importance of this type of sub-component testing and high fidelity failure analysis in the design of a composite fuselage.

Baker, Donald J.↗
NASA NTRS2000

Erosion Coatings for High-Temperature Polymer Composites: A Collaborative Project With Allison Advanced Development Company

The advantages of replacing metals in aircraft turbine engines with high-temperature polymer matrix composites (PMC's) include weight savings accompanied by strength improvements, reduced part count, and lower manufacturing costs. Successfully integrating high-temperature PMC's into turbine engines requires several long-term characteristics. Resistance to surface erosion is one rarely reported property of PMC's in engine applications because PMC's are generally softer than metals and their erosion resistance suffers. Airflow rates in stationary turbine engine components typically exceed 2.3 kg/sec at elevated temperatures and pressures. In engine applications, as shown in the following photos, the survivability of PMC components is clearly a concern, especially when engine and component life-cycle requirements become longer. Although very few publications regarding the performance of erosion coatings on PMC's are available particularly in high-temperature applications the use of erosion-resistant coatings to significantly reduce wear on metallic substrates is well documented. In this study initiated by the NASA Glenn Research Center at Lewis Field, a low-cost (less than $140/kg) graphite-fiber-reinforced T650 35/PMR 15 sheet-molding compound was investigated with various coatings. This sheet-molding compound has been compression molded into many structurally complicated components, such as shrouds for gas turbine inlet housings and gearboxes. Erosion coatings developed for PMC s in this study consisted of a two-layered system: a bondcoat sprayed onto a cleaned PMC surface, followed by an erosion-resistant, hard topcoat sprayed onto the bondcoat as shown in following photomicrograph. Six erosion coating systems were evaluated for their ability to withstand harsh thermal cycles, erosion resistance (ASTM G76 83 "Standard Practice for Conducting Erosion Tests by Solid Particle Impingement Using Gas Jets") using Al2O3, and adhesion to the graphite fiber polyimide composite (ASTM D 4541 95 "Pull Off Strength of Coatings"). Glenn and Allison Advanced Development Company collaborated to optimize erosion coatings for gas turbine fan and compressor applications. All the coating systems survived aggressive thermal cycling without spalling. During erosion tests (see the final photo), the most promising coating systems tested had Cr3C2-NiCr and WC-Co as the hard topcoats. In all cases, these coating systems performed significantly better than that with a TiN hard topcoat. When material depth (thickness) loss is considered, the Cr3C2-NiCr and WC-Co coating systems provided, on average, an erosion resistance 8.5 times greater than that for the uncoated PMR 15/T650 35 composite. Similarly, Cr3C2-NiCr and WC-Co coating systems adhered to the PMC substrate during tensile tests significantly better than systems containing a TiN topcoat. Differences in topcoats of Cr3C2-NiCr and WC-Co were determined by considering issues such as cost and environmental impact. The preferred erosion-resistant coating system for PMR 15/T650 35 has WC-Co as the hard topcoat. This system provides the following benefits in comparison to the coating system with Cr3C2-NiCr topcoat: lower powder material cost (15 to 20 percent), environmentally friendly materials (Cr3C2-NiCr is hazardous), and higher deposition yield (10 to 15 percent), which results in less waste.

Sutter, James K.↗
NASA NTRS2001

Fan Flutter Analysis Capability Enhanced

The trend in the design of advanced transonic fans for aircraft engines has been toward the use of complex high-aspect-ratio blade geometries with a larger number of blades and higher loading. In addition, integrally bladed disks or blisks are being considered in fan designs for their potential to reduce manufacturing costs, weight, and complexity by eliminating attachments. With such design trends, there is an increased possibility within the operating region of part-speed stall flutter (self-excited vibrations) that is exacerbated by the reduced structural damping of blisk fans. To verify the aeroelastic soundness of the design, the NASA Glenn Research Center is developing and validating an accurate aeroelastic prediction and analysis capability. Recently, this capability was enhanced significantly as described here.

Bakhle, Milind A.↗
NASA NTRS2003

Superalloy Lattice Block Developed for Use in Lightweight, High-Temperature Structures

Successful development of advanced gas turbine engines for aircraft will require lightweight, high-temperature components. Currently titanium-aluminum- (TiAl) based alloys are envisioned for such applications because of their lower density (~4 g/cm3) in comparison to superalloys (~8.5 g/cm3), which have been utilized for hot turbine engine parts for over 50 years. However, a recently developed concept (lattice block) by JAMCORP, Inc., of Willmington, Massachusetts, would allow lightweight, high-temperature structures to be directly fabricated from superalloys and, thus, take advantage of their well-known, characterized properties. In its simplest state, lattice block is composed of thin ligaments arranged in a three dimensional triangulated trusslike configuration that forms a structurally rigid panel. Because lattice block can be fabricated by casting, correctly sized hardware is produced with little or no machining; thus very low cost manufacturing is possible. Together, the NASA Glenn Research Center and JAMCORP have extended their lattice block methodology for lower melting materials, such as Al alloys, to demonstrate that investment casting of superalloy lattice block is possible. This effort required advances in lattice block pattern design and assembly, higher temperature mold materials and mold fabrication technology, and foundry practice suitable for superalloys (ref. 1). Lattice block panels have been cast from two different Ni-base superalloys: IN 718, which is the most commonly utilized superalloy and retains its strength up to 650 C; and MAR M247, which possesses excellent mechanical properties to at least 1100 C. In addition to the open-cell lattice block geometry, same-sized lattice block panels containing a thin (~1-mm-thick) solid face on one side have also been cast from both superalloys. The elevated-temperature mechanical properties of the open cell and face-sheeted superalloy lattice block panels are currently being examined, and the microstructure is being characterized in terms of casting defects. In addition, a small study (ref. 3) is being undertaken with GE Aircraft Engines to determine the suitability of superalloy lattice block for engine components.

Hebsur, Mohan G.↗
NASA NTRS2003

Thin-Film Air-Mass-Flow Sensor of Improved Design Developed

Researchers at the NASA Glenn Research Center have developed a new air-mass-flow sensor to solve the problems of existing mass flow sensor designs. NASA's design consists of thin-film resistors in a Wheatstone bridge arrangement. The resistors are fabricated on a thin, constant-thickness airfoil to minimize disturbance to the airflow being measured. The following photograph shows one of NASA s prototype sensors. In comparison to other air-mass-flow sensor designs, NASA s thin-film sensor is much more robust than hot wires, causes less airflow disturbance than pitot tubes, is more accurate than vane anemometers, and is much simpler to operate than thermocouple rakes. NASA s thin-film air-mass-flow sensor works by converting the temperature difference seen at each leg of the thin-film Wheatstone bridge into a mass-flow rate. The following figure shows a schematic of this sensor with air flowing around it. The sensor operates as follows: current is applied to the bridge, which increases its temperature. If there is no flow, all the arms are heated equally, the bridge remains in balance, and there is no signal. If there is flow, the air passing over the upstream legs of the bridge reduces the temperature of the upstream legs and that leads to reduced electrical resistance for those legs. After the air has picked up heat from the upstream legs, it continues and passes over the downstream legs of the bridge. The heated air raises the temperature of these legs, increasing their electrical resistance. The resistance difference between the upstream and downstream legs unbalances the bridge, causing a voltage difference that can be amplified and calibrated to the airflow rate. Separate sensors mounted on the airfoil measure the temperature of the airflow, which is used to complete the calculation for the mass of air passing by the sensor. A current application for air-mass-flow sensors is as part of the intake system for an internal combustion engine. A mass-flow sensor is used to provide accurate information about the amount of air entering the engine so that the amount of fuel can be adjusted to give the most efficient combustion. The ideal mass-flow sensor would be a rugged design that minimizes the disturbance to the flow stream and provides an accurate reading of both smooth and turbulent flows; NASA's design satisfies these requirements better than any existing design. Most of the mass-flow sensors used today are the hot wire variety. Hot wires can be fragile and cannot accurately measure a turbulent or reversing flow, which is often encountered in an intake manifold. Other types of mass-flow sensors include pitot tubes, vane anemometers, and thermocouple rakes-all of which suffer from some type of performance problem. Because it solves these performance problems while maintaining a simple design that lends itself to low-cost manufacturing techniques, NASA s thin-film resistance temperature detector air-mass-flow sensor should lead to more widespread use of mass-flow sensors.

Fralick, Gustave C.↗
NASA NTRS2003

Active Blade Vibration Control Being Developed and Tested

Gas turbine engines are currently being designed to have increased performance, lower weight and manufacturing costs, and higher reliability. Consequently, turbomachinery components, such as turbine and compressor blades, have designs that are susceptible to new vibration problems and eventual in-service failure due to high-cycle fatigue. To address this problem, researchers at the NASA Glenn Research Center are developing and testing innovative active blade vibration control concepts. Preliminary results of using an active blade vibration control system, involving a rotor supported by an active magnetic bearing in Glenn's Dynamic Spin Rig, indicate promising results (see the photograph). Active blade vibration control was achieved using feedback of blade strain gauge signals within the magnetic bearing control loop. The vibration amplitude was reduced substantially (see the graphs). Also, vibration amplitude amplification was demonstrated; this could be used to enhance structural mode identification, if desired. These results were for a nonrotating two-bladed disk. Tests for rotating blades are planned. Current and future active blade vibration control research is planned to use a fully magnetically suspended rotor and smart materials. For the fully magnetically suspended rotor work, three magnetic bearings (two radial and one axial) will be used as actuators instead of one magnetic bearing. This will allow additional degrees of freedom to be used for control. For the smart materials work, control effectors located on and off the blade will be considered. Piezoelectric materials will be considered for on-the-blade actuation, and actuator placement on a stator vane, or other nearby structure, will be investigated for off-the-blade actuation. Initial work will focus on determining the feasibility of these methods by performing basic analysis and simple experiments involving feedback control.

Johnson, Dexter↗
NASA NTRS2006

Current Research Activities in Drive System Technology in Support of the NASA Rotorcraft Program

Drive system technology is a key area for improving rotorcraft performance, noise/vibration reduction, and reducing operational and manufacturing costs. An overview of current research areas that support the NASA Rotorcraft Program will be provided. Work in drive system technology is mainly focused within three research areas: advanced components, thermal behavior/emergency lubrication system operation, and diagnostics/prognostics (also known as Health and Usage Monitoring Systems (HUMS)). Current research activities in each of these activities will be presented. Also, an overview of the conceptual drive system requirements and possible arrangements for the Heavy Lift Rotorcraft program will be reviewed.

Handschuh, Robert F.↗
NASA NTRS2005

NASA Heavy Lift Rotorcraft Systems Investigation

The NASA Heavy Lift Rotorcraft Systems Investigation examined in depth several rotorcraft configurations for large civil transport, designed to meet the technology goals of the NASA Vehicle Systems Program. The investigation identified the Large Civil Tiltrotor as the configuration with the best potential to meet the technology goals. The design presented was economically competitive, with the potential for substantial impact on the air transportation system. The keys to achieving a competitive aircraft were low drag airframe and low disk loading rotors; structural weight reduction, for both airframe and rotors; drive system weight reduction; improved engine efficiency; low maintenance design; and manufacturing cost comparable to fixed-wing aircraft. Risk reduction plans were developed to provide the strategic direction to support a heavy-lift rotorcraft development. The following high risk areas were identified for heavy lift rotorcraft: high torque, light weight drive system; high performance, structurally efficient rotor/wing system; low noise aircraft; and super-integrated vehicle management system.

Johnson, Wayne↗