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NASA NTRS1999

Residual Strength Pressure Tests and Nonlinear Analyses of Stringer- and Frame-Stiffened Aluminum Fuselage Panels with Longitudinal Cracks

The results of residual strength pressure tests and nonlinear analyses of stringer- and frame-stiffened aluminum fuselage panels with longitudinal cracks are presented. Two types of damage are considered: a longitudinal crack located midway between stringers, and a longitudinal crack adjacent to a stringer and along a row of fasteners in a lap joint that has multiple-site damage (MSD). In both cases, the longitudinal crack is centered on a severed frame. The panels are subjected to internal pressure plus axial tension loads. The axial tension loads are equivalent to a bulkhead pressure load. Nonlinear elastic-plastic residual strength analyses of the fuselage panels are conducted using a finite element program and the crack-tip-opening-angle (CTOA) fracture criterion. Predicted crack growth and residual strength results from nonlinear analyses of the stiffened fuselage panels are compared with experimental measurements and observations. Both the test and analysis results indicate that the presence of MSD affects crack growth stability and reduces the residual strength of stiffened fuselage shells with long cracks.

Young, Richard D.↗

NASA NTRS1999

Aeroelastic Analysis of a Trimmed Generic Hypersonic Vehicle

The aeroelastic equations of motion governing a hypersonic vehicle in free flight are derived. The equations of motion for a translating and rotating flexible body using Lagrange's equations in terms of quasi-coordinates are presented. These equations are simplified for the case of a vehicle with pitch and plunge rigid body degrees of freedom and small elastic displacements. The displacements are approximated by a truncated series of the unrestrained mode shapes, which are obtained using equivalent plate theory. Subsequently, the nonlinear equations of motion are linearized about the trim state, which is obtained using a rigid body trim model and steady hypersonic aerodynamics. The appropriate flutter derivatives are calculated from piston theory. Results describing mode shapes, trim behavior, and aeroelastic stability of a generic hypersonic vehicle are presented.

Nydick, I.↗

NASA NTRS1979

Temperature Measurement Aid

NASA's Ames Research Center has designed a simple but medically important device--one which holds temperature probes, called thermistors, to a person's skin without affecting the characteristics of the skin segment being measured. The device improves the accuracy of skin surface temperature measurements, valuable data in health evaluation. The need for such a device was recognized in the course of life science experiments at Ames. In earlier methods, the sensing head of the temperature probe was affixed to the patient's skin by tape or elastic bands. This created a heat variance which altered skin temperature readings. The Ames-developed thermistor holder is a plastic ring with tab extensions, shown in the upper photo on the chest, arm and leg of the patient undergoing examination. The ring holds the sensing head of the temperature probe and provides firm, constant pressure between the skin and the probe. The tabs help stabilize the ring and provide attachment points for the fastening tape or bands, which do not directly touch the sensor. With this new tool, it is possible to determine more accurately the physiological effects of strenuous exercise, particularly on the treadmill. The holder is commercially available from Yellow Springs Instrument Company, Inc., Yellow Springs, Ohio, which is producing the device under a NASA patent license.

Source record↗

NASA NTRS1980

Composite wall concept for high temperature turbine shrouds: Survey of low modulus strain isolator materials

Plasma sprayed yttria stabilized zirconium oxide turbine seal specimens, incorporating various low modulus porous metal strain isolator pads between the zirconium oxide and a dense metal substrate, were subjected to cyclic thermal shock testing. Specimens that had a low modulus pad composed of sintered FeNiCrAlY fibermetal survived 1000 thermal shock cycles without spalling of the ceramic. A figure of merit for the low modulus pad materials taking into consideration the elastic modulus, thermal conductivity, strength, and oxidation resistance of the pad was proposed, and showed reasonable agreement with the thermal shock results. A potential surface distress problem on the zirconium oxide, associated with nonuniform temperature distribution and rapid stress relaxation was identified. One approach to solving the surface distress problem through application of laser surface fusion of the zirconium oxide layer showed some promise, but improvements in the laser surface fusion process are necessary to prevent process associated damage to the ceramic.

Bill, R. C.↗

NASA NTRS1991

Passive dynamic controllers for nonlinear mechanical systems

A methodology for model-independant controller design for controlling large angular motion of multi-body dynamic systems is outlined. The controlled system may consist of rigid and flexible components that undergo large rigid body motion and small elastic deformations. Control forces/torques are applied to drive the system and at the same time suppress the vibration due to flexibility of the components. The proposed controller consists of passive second-order systems which may be designed with little knowledge of the system parameter, even if the controlled system is nonlinear. Under rather general assumptions, the passive design assures that the closed loop system has guaranteed stability properties. Unlike positive real controller design, stabilization can be accomplished without direct velocity feedback. In addition, the second-order passive design allows dynamic feedback controllers with considerable freedom to tune for desired system response, and to avoid actuator saturation. After developing the basic mathematical formulation of the design methodology, simulation results are presented to illustrate the proposed approach to a flexible six-degree-of-freedom manipulator.

Juang, Jer-Nan↗

NASA NTRS2019

Preliminary Scale-Resolving Simulations of Laminar-To-Turbulent Transition in Swept-Wing Flow

Swept wings and control surfaces are common elements of modern aircraft, and it has been shown both experimentally and theoretically that laminar-to-turbulent transition of the three-dimensional boundary layer that develops over them is highly sensitive to surface roughness. Numerous studies have been conducted on the effect of discrete roughness elements or distributed roughness elements on swept flow transition, however so far limited computational effort has been dedicated to the study of transition over swept wings with randomly distributed micron-sized roughness. In the present work, we set up to reproduce the extensive experimental data base generated by Dagenhart et al for the infinite swept wing NLF(2)-0415. To this purpose, we perform scale-resolving simulations of flow transition over smooth and rough surfaces using a high-order space-time spectral-element Discontinuous-Galerkin solver. Different types of surface roughnesses are implemented by elastically deforming the original mesh. The study shows that the experimental results cannot be accounted for by a perfectly smooth wing and reveals a strong sensitivity of the transition process to the representation of the surface roughness. The crossflow patterns and transition location approach those measured for some of the surface profiles, however a correlation between the wavenumber spectrum of the surface, grid resolution and boundary layer stability is yet to be established.

Denison, Marie↗

NASA NTRS1939

A method of estimating the critical buckling load for structural members

The relations between load on the structure and rotation of a joint can be used to estimate the lowest critical load after the equation for neutral stability has been tested for three assumed critical loads, each of which is less than the lowest critical load. The solutions of six simple problems are included to illustrate the application of the method of estimating critical loads and to reveal certain characteristics of the method that should be known by the practical engineer using it. Four of these problems are concerned with members that lie in the elastic, or long-column, range. The other two problems are concerned with members that lie in the short-column range.

Lundquist, Eugene E↗

NASA NTRS1991

Passive dynamic controllers for non-linear mechanical systems

A methodology for model-independent controller design for controlling large angular motion of multi-body dynamic systems is outlined.The controlled system may consist of rigid and flexible components that undergo large rigid body motion and small elastic deformations. Control forces/torques are applied to drive the system, and at the same time suppress the vibrations due to flexibility of the components. The proposed controller consists of passive second-order systems which may be designed with little knowledge of the system parameters, even if the controlled system is non-linear. Under rather general assumptions, the passive design assures that the closed loop system has guaranteed stability properties. Unlike positive real controller design, stabilization can be accomplished without direct velocity feedback. In addition, the second-order passive design allows dynamic feedback controllers with considerable freedom to tune for desired system response, and to avoid actuator saturation. After developing the basic mathematical formulation of the design methodology, simulation results are presented to illustrate the proposed approach applied to a flexible six-degree-of-freedom manipulator.

Juang, Jer-Nan↗

NASA NTRS1959

Transonic Flutter Investigation of Models of T-Tail of Blackburn NA-39 Airplane

A transonic flutter investigation has been made of models of the T-tail of the Blackburn NA-39 airplane. The models were dynamically and elastically scaled from measured airplane data in accordance with criteria which include a flutter safety margin. The investigation was made in the Langley transonic blowdown tunnel and covered a Mach number range from 0.73 to 1.09 at simulated altitudes extending to below sea level. The results of the investigation indicated that, if differences between the measured model and scaled airplane properties are disregarded, the airplane with the normal value of stabilizer pitching stiffness should have a stiffness margin of safety of at least 32 percent at all Mach numbers and altitudes within the flight boundary. However, the airplane with the emergency value of stabilizer pitching stiffness would not have the required margin of safety from symmetrical flutter at Mach numbers greater than about 0.85 at low altitudes. First-order corrections for some differences between the measured model and scaled airplane properties indicated that the airplane with the normal value of stabilizer pitching stiffness would still have an adequate margin of safety from flutter and that the flutter safety margin for the airplane with the emergency value of stabilizer pitching stiffness would be changed from inadequate to adequate. However, the validity of the corrections is questionable.

Jones, George W., Jr.↗

NASA NTRS1973

Nonlinear vibrations of rectangular plates.

A finite-difference method is developed to determine the large amplitude dynamic responses of thin elastic plates subjected to uniform pressure pulse-type loads. Four different sets of boundary conditions are considered. Some specific problems are solved. The results are compared with approximate solutions obtained by Yamaki (1961). The numerical method presented provides an accurate and efficient approximate solution to the problem, and should be useful as a check on other approximate methods. The grid-size and the time-step necessary for obtaining numerical stability depend on the particular problem. For many cases the method converges rapidly and a rather large grid-size and time-step is adequate.

Bayles, D. J.↗

NASA NTRS1978

The dynamics and control of large flexible space structures. Part A: Discrete model and modal control

Attitude control techniques for the pointing and stabilization of very large, inherently flexible spacecraft systems were investigated. The attitude dynamics and control of a long, homogeneous flexible beam whose center of mass is assumed to follow a circular orbit was analyzed. First order effects of gravity gradient were included. A mathematical model which describes the system rotations and deflections within the orbital plane was developed by treating the beam as a number of discretized mass particles connected by massless, elastic structural elements. The uncontrolled dynamics of the system are simulated and, in addition, the effects of the control devices were considered. The concept of distributed modal control, which provides a means for controlling a system mode independently of all other modes, was examined. The effect of varying the number of modes in the model as well as the number and location of the control devices were also considered.

Bainum, P. M.↗

NASA NTRS2005

Buckling of Carbon Nanotube-Reinforced Polymer Laminated Composite Materials Subjected to Axial Compression and Shear Loadings

A multi-scale method to predict the stiffness and stability properties of carbon nanotube-reinforced laminates has been developed. This method is used in the prediction of the buckling behavior of laminated carbon nanotube-polyethylene composites formed by stacking layers of carbon nanotube-reinforced polymer with the nanotube alignment axes of each layer oriented in different directions. Linking of intrinsic, nanoscale-material definitions to finite scale-structural properties is achieved via a hierarchical approach in which the elastic properties of the reinforced layers are predicted by an equivalent continuum modeling technique. Solutions for infinitely long symmetrically laminated nanotube-reinforced laminates with simply-supported or clamped edges subjected to axial compression and shear loadings are presented. The study focuses on the influence of nanotube volume fraction, length, orientation, and functionalization on finite-scale laminate response. Results indicate that for the selected laminate configurations considered in this study, angle-ply laminates composed of aligned, non-functionalized carbon nanotube-reinforced lamina exhibit the greatest buckling resistance with 1% nanotube volume fraction of 450 nm uniformly-distributed carbon nanotubes. In addition, hybrid laminates were considered by varying either the volume fraction or nanotube length through-the-thickness of a quasi-isotropic laminate. The ratio of buckling load-to-nanotube weight percent for the hybrid laminates considered indicate the potential for increasing the buckling efficiency of nanotube-reinforced laminates by optimizing nanotube size and proportion with respect to laminate configuration.

Riddick, J. C.↗

NASA NTRS1974

Interfaces in directionally solidified eutectics

Consideration of the interface between the phases in directionally solidified eutectic structures with respect to its morphology, crystallography, stability, and overall response to internal and external stress fields. Yielding behavior in such eutectic structures is shown to be controlled by several possible mechanisms. Two that can be related to the interface in directionally-solidified eutectic alloys are: (1) the interaction of slip dislocations with interface dislocations of the semicoherent boundaries, and (2) restrictions of allowable deformation mechanisms in two contiguous phases due to orientation effects. Evidence found thus far for elastic-plastic eutectic composites indicates that interfaces are capable of diverting fatigue cracks parallel to the loading axis.

Hertzberg, R. W.↗

NASA NTRS1975

Astronautic Structures Manual, Volume 3

This document (Volumes I, II, and III) presents a compilation of industry-wide methods in aerospace strength analysis that can be carried out by hand, that are general enough in scope to cover most structures encountered, and that are sophisticated enough to give accurate estimates of the actual strength expected. It provides analysis techniques for the elastic and inelastic stress ranges. It serves not only as a catalog of methods not usually available, but also as a reference source for the background of the methods themselves. An overview of the manual is as follows: Section A is a general introduction of methods used and includes sections on loads, combined stresses, and interaction curves; Section B is devoted to methods of strength analysis; Section C is devoted to the topic of structural stability; Section D is on thermal stresses; Section E is on fatigue and fracture mechanics; Section F is on composites; Section G is on rotating machinery; and Section H is on statistics. These three volumes supersede Volumes I and II, NASA TM X-60041 and NASA TM X-60042, respectively.

Source record↗

NASA NTRS1982

Coupled bending-bending-torsion flutter of a mistuned cascade with nonuniform blades

A set of aeroelastic equations describing the motion of an arbitrarily mistuned cascade with flexible, pretwisted, nonuniform blades is developed using an extended Hamilton's principle. The derivation of the equations has its basis in the geometric nonlinear theory of elasticity in which the elongations and shears are negligible compared to unity. A general expression for foreshortening of a blade is derived and is explicity used in the formulation. The blade aerodynamic loading in the subsonic and supersonic flow regimes is obtained from two dimensional, unsteady, cascade theories. The aerodynamic, inertial and structural coupling between the bending (in two planes) and torsional motions of the blade is included. The equations are used to investigate the aeroelastic stability and to quantify the effect of frequency mistuning on flutter in turbofans. Results indicate that a moderate amount of intentional mistuning has enough potential to alleviate flutter problems in unshrouded, high aspect ratio turbofans.

Kaza, K. R. V.↗

NASA NTRS1985

Optimization of cascade blade mistuning. II - Global optimum and numerical optimization

The values of the mistuning which yield the most stable eigenvectors are analytically determined, using the simplified equations of motion which were developed in Part I of this work. It is shown that random mistunings, if large enough, may lead to the maximal stability, whereas the alternate mistunings cannot. The problem of obtaining maximum stability for minimal mistuning is formulated, based on numerical optimization techniques. Several local minima are obtained using different starting mistuning vectors. The starting vectors which lead to the global minimum are identified. It is analytically shown that all minima appear in multiplicities which are equal to the number of compressor blades. The effect of mistuning on the flutter speed is studied using both an optimum mistuning vector and an alternate mistuning vector. Effects of mistunings in elastic axis locations are shown to have a negligible effect on the eigenvalues. Finally, it is shown that any general two-dimensional bending-torsion system can be reduced to an equivalent uncoupled torsional system.

Nissim, E.↗

NASA NTRS2004

Evaluation of a CFD Method for Aerodynamic Database Development using the Hyper-X Stack Configuration

A computational fluid dynamic (CFD) study is performed on the Hyper-X (X-43A) Launch Vehicle stack configuration in support of the aerodynamic database generation in the transonic to hypersonic flow regime. The main aim of the study is the evaluation of a CFD method that can be used to support aerodynamic database development for similar future configurations. The CFD method uses the NASA Langley Research Center developed TetrUSS software, which is based on tetrahedral, unstructured grids. The Navier-Stokes computational method is first evaluated against a set of wind tunnel test data to gain confidence in the code s application to hypersonic Mach number flows. The evaluation includes comparison of the longitudinal stability derivatives on the complete stack configuration (which includes the X-43A/Hyper-X Research Vehicle, the launch vehicle and an adapter connecting the two), detailed surface pressure distributions at selected locations on the stack body and component (rudder, elevons) forces and moments. The CFD method is further used to predict the stack aerodynamic performance at flow conditions where no experimental data is available as well as for component loads for mechanical design and aero-elastic analyses. An excellent match between the computed and the test data over a range of flow conditions provides a computational tool that may be used for future similar hypersonic configurations with confidence.

Parikh, Paresh↗

NASA NTRS1975

Sensitivity of Hingeless Rotor Blade Flap-lag Stability in Hover to Analytical Modelling Assumptions

Prediction of flap-lag stability using a single bending mode for each degree-of-freedom is examined in the case in which the bending modes are assumed to be the same in the flap and lag directions and are independent of pitch angle and stiffness distribution. It is shown that this model gives results analogous to those obtained by Ormiston employing a rigid blade model with the blade and hub stiffness represented by springs in the limiting cases of the elastic coupling parameter R = 0 and 1. For intermediate values of R the results are shown to be quite different. The mode shape assumptions are shown to result in what is referred to as the parallel spring model in contrast to Ormiston's model which is referred to as a series spring model. The similarities and differences between these two models are developed in some details. The differences between these two models are examined for various typical rotor blade characteristics. Other aspects of the sensitivity of this problem are also considered.

Curtiss, H. C., Jr.↗