NASA Technical Memorandum 107300

_J

Structures Division 1995 Annual Report

ii  ̧

National Aeronautics and

Space Administration

July 1997

Lewis Research Center
Trade names or manufacturers' names are used in this report for identification

only. This usage does not constitute an official endorsement, either expressed

or implied, by the National Aeronautics and Space Administration.
Introduction

The 1995 Annual Report of the Structures

Division reflects the work areas performed by the

Division staff during the 1995 calendar year. This

publication covers 54 separate topics. Its purpose

is to give a brief, but comprehensive review of the

Division's technical accomplishments. As with the

reports for the previous years, the report is

organized topically. The descriptions of the

research reflect work that has been reported in

the open literature during the year. A

bibliography containing 132 citations is provided

to the reader.

The Structures Division comprises a staff of

approximately 68 engineers and scientists plus

administrative and support personnel. The work

areas comprise but are not limited to composite

mechanisms, fatigue, fracture, and dynamics

with emphasis on life and reliability. Additional

work in future years will include mechanical

component technologies and acoustics.

The Division works cooperatively with both

industry and universities to develop the

technology necessary for state-of-the-art

advancement in aeronautical and space

propulsion systems. In the future, propulsion

systems will need to be lighter, to operate at

higher temperatures and to be more reliable in

order to achieve higher performance. Achieving

these goals is complex and challenging. If you

need additional information, please do not

hesitate to contact me or the appropriate Division

staff contact provided in this publication.

Acting Chief, Structures Division

NASA TM-107300 i
Contents

Structures Division

Soft Computing Methods in Design of Superalloys ....................................... 1

Fuzzy and Neural Net Approach to Tolerance Level Uncertainty ............................. 1

Effect of Cyclic Thermal Loads on Fatigue Reliability in Polymer Matrix Composites .............. 1

T/BEST: Technology Benefit Estimator Select Features and Applications ...................... 2

Probabilistic Thermomechanical Fatigue of Polymer Matrix Composites ....................... 4

Prestraining Effects on Fatigue Life ................................................... 5

Structurally Compliant Rocket Engine Combustion Chamber ............................... 6

Kinetics of Cyclic Oxidation and Cracking and Finite Element Analysis of

MA956 and Sapphire/MA956 Composite System ..................................... 7

Large-Displacement Structural Durability Analyses of Simple Bend Specimen Emulating

Rocket Nozzle Liners ........................................................... 8

Effect of Hoop Stress on Ball Bearing Life .............................................. 9

Predicting Rolling-Element Bearing Life .............................................. 10

Structural Mechanics

Effect of Combined Loads on the Durability of a Stiffened Adhesively Bonded

Composite Structure .......................................................... 11

Damage Progression in Bolted Composite Structures .................................... 11

Computational Simulation of Fiber Composite Thin Shell Structures in a

Hygrothermal Environment ..................................................... 12

Optimization Improves Air-Breathing Propulsion Engine Concept ........................... 13

A Novel Boundary Condition Completed the Beltrami-Michell Formulation in Elasticity .......... 15

Active Thermal Distortion Management With Smart Piezoelectric Structures ................... 15

Micromechanics for Woven Composites ............................................... 17

Ballistic Impact Research in Support of HSCT Engine Fan Containment ...................... 17

Monitoring Damage Progression in Ceramic Matrix Composites ............................ 19

High-Temperature Strain Gage and Adhesive System Development and Testing ................ 19

Micromechanics Predict Uniaxial Response and Local Stresses of Composites With Shape

Memory Alloy Fibers .......................................................... 20

Generalized Finite Element Enables Prediction of Global and Local Response of

Multilayered Composite Plates With Embedded Piezoelectric Actuators and Sensors .......... 21

Fatigue and Fracture

A Study to Optimize a Ring-On-Ring Test Fixture for Testing CMC's Under

In-Plane Biaxial Loading Conditions .............................................. 22

Yield Surface Study of Cast Gamma Titanium Alumlnide friAl) ............................. 25

EPM Structural Component Successfully Tested Under Pseudo-Operating Conditions ........... 26

Fatigue Behavior of Haynes 188 Under Thermomechanical, Axial-Torsional Loading ............. 27

Retirement for Cause as an Alternate Means of Managing Component Lives ................... 28

Thermomechanical Deformation and Strain Rate Sensitivity of the Dynamic Strain

Aging Alloy Haynes 188 ....................................................... 29

A Study of Elevated-Temperature Testing Techniques for the Fatigue Behavior of PMC's .......... 31

ii NASA TM-107300
ModelDeterminedforPredictingFatigueLivesof MetalMatrix Composites

Under Mean Stresses ......................................................... 33

Oxidation Embrittlement Observed in SiC/SiC Composites ................................ 34

Structural Dynamics

Rotordynamics on the PC: Transient Analysis With ARDS ................................. 36

A Magnetically Sealed Bearings System .............................................. 38

An Active Magnetic Bearing With High-Temperature Superconductor Coils and

Ferromagnetic Cores .......................................................... 39

Application of Magnetic Bearing to Dynamic Spin Rig .................................... 41

Measurement of Gust Response on a Turbine Cascade ................................... 41

Flutter Analysis of Ducted Rotors ................................................... 42

Fan Blade Deflection Measurement and Analyses Correlation .............................. 43

Space Mechanisms Lessons Learned Study ........................................... 44

Alternative Operating Modes for Magnetic Bearing Control ................................ 46

A New Antiwear Additive for Bearings Used in Spacecraft ................................. 46

Users Guide for ECAP2D: An Euler Aeroelastic Stability Analysis Code for

Oscillating Cascades .......................................................... 48

Users Guide for FPCAS3D: A Three Dimensional Full-Potential Aeroelastic Solver ............... 48

Accelerated Testing of Spacecraft Mechanisms Study .................................... 49

GE Uses NASA Seals to Meet JTAGG Engine Goals ...................................... 50

Structural Integrity

Acousto-Ultrasonics and Tensile Cycle Degradation in Ceramic Matrix Composites ............... 51

Award-Winning CARES/Life Ceramics Durability Evaluation Software: Making

Advanced Technology Accessible .................................................. 52

Integrated Design Software Predicts Creep Life of Monolithic Ceramic Components ............... 53

Thermomechanical Analysis of Ceramic Matrix Composite Components ....................... 54

Modeling the Effective Elastic Behavior of Transversely Cracked Laminated Composite ............ 55

Capability of Thermographic Imaging Defined for Detection of Defects in High-Temperature

Composite Materials ........................................................... 56

Crack Extension Measurement in Brittle Materials ....................................... 57

Experimental Verification of Crack Growth Models for Ceramics ............................. 58

Bibliography ........................................................... 6o

NASA TM-107300 iii

Structures Division

Soft Computing Methods in Design of

Superalloys

In this research, done under contract with the

University of Toledo in Ohio, neural networks and

genetic algorithms were used to design superalloys.

The cyclic oxidation attack parameter Ka was

modeled using a backpropagation neural network

as a function of the superalloy chemistry and the

test temperature. This model was then used in

unison with a genetic algorithm to obtain an

optimized superalloy composition, which resulted

in low values of Ka. The key feature of our approach

was employing the neural network to model the

material properties and genetic algorithms that were

used for optimizing the objective function values

provided as the outputs of the neural network.

Lewis contact: Laszlo Berke (216) 433-5648

Fuzzy and Neural Net Approach to

Tolerance Level Uncertainty

Under a small disadvantaged business (SDB) 8(a)

contract, the practical special case of only tolerance

level input uncertainty was examined for structural

analysis and optimization. Also assumed was that

precise information concerning the distribution of

uncertainty within a small tolerance level was not

known. One can view such a case through fuzzy

techniques and its definitions of "membership

functions" as an alternative to detailed probability

distributions, with accuracy depending on the

quality of available statistical data. The simplest

case is to be concerned only with the possible

tolerance range defined by minimum and maximum

input values and to examine methods for their

efficient mapping into possible minimum and

maximum output values. The obvious approach is

to perform analyses of all combinations of such

input values, but that is efficient for only a very

small number of variables. One can define various

membership functions between the extremes and

perform conventional Monte Carlo experiments

capturing the extreme tails produced. The neural

network models allow a very large number of

samplings. As is well known, the tails increase

asymptotically towards the extreme values

obtainable by exhaustive combinatorial exploration.

This operation is made efficient by first generating

a fast-executing neural network model for the

ranges in question. One may also surmise that the

extreme output values do not belong to some

combination of extreme input values but that some

intermediate input values are also involved in

producing them. When a fast-executing neural

network model is available, optimization iterations

can be run to minimize and maximize output values.

Monte Carlo experimentation also can capture such

cases. Finally, one can apply "interval equation"

techniques of fuzzy sets generated by considering

"alpha cuts" of input membership functions, again

just the extreme values in the simplest case. These

equations can be solved with about twice the

computations required by a conventional solution

and can produce extreme values that closely match

the values generated by the above methods but only

for small (tolerance) differences between minimum

and maximum input values. For large ranges, over

a few percent of mean values, this approach

becomes unreliable. Because of the computational

advantage, this approach warrants further

development to achieve accuracy for larger ranges

of uncertainty. This project is nearing completion.

Lewis contact: Laszlo Berke (216} 433-5648

Effect of Cyclic Thermal Loads on

Fatigue Reliability in Polymer Matrix

Composites

Under the NASA High Speed Research (HSR)

Program, technological solutions that will ensure

the economic viability and environmental

compatibility of a future High Speed Civil Transport

(HSCT) aircraft are currently being sought. Concepts

for both lighter airframe primary structures and

engine structural components are being

investigated. It is envisioned that such objectives

can be achieved through the use of hightemperature composites and other conventional

lighter weight alloys. One of the prime issues for

these structural components is assured long-term

behavior with a specified reliability.

NASA TM-107300 1
An investigation was conducted to develop a

computationalsimulationmethodologythat would

predict the fatigue life (fig. 1), reliability, and

probabflisticlong-termbehaviorof polymermatrix

composites(PMC's).A unified time-, stress-,and

load-dependentMultiFactor Interaction Equation

(MFIE)modeldevelopedattheNASALewisResearch

Center was used to simulate the long-term behavior

of polymer matrix composites.

Although the methodology can be applied to other

types of PMC's, a typical composite system

consisting of graphite fibers in epoxy matrix was

chosen to illustrate the application of the

methodology. The layup of the composite system

was chosen to be [0/-+45/90] s. The cumulative

probability distribution functions (CDF's) for the

fatigue life cycles were computed for different

thermal cycle rates and constant applied stress. The

Composite:

Graphite/epoxy

d

Laminate configuration:

[0/+45/90] s

Ply thickness, 0.127 mm

E

=o

o

=-

--- Amplitude

Time

.--- Amplitude

Mean

Mean

Time

Figure 1.--Therrnomechanical cycfic load.

,. 1.0-

¢-

=e 0.8-

¢0

o. 0.6-

_ 0.4-

= 0.2--

¢,j

,<

0.0

0.3

I I I l I I I

0.4 0.5 0.6 0.7 0.8 0.9 1.0

Ratio of current to final cycles, Nm/Nrnf

Figure 2.--Fatigue life variation resulting from thermomechanical cyclic load. Mean applied load, 50 percent

of ply strength; mean temperature, 65.5 °C; cyclic

temperature 51.7 °C.

laminate strength was evaluated on the basis of

the first ply failure criteria (hereinafter referred to

as laminate strength). First ply failure criteria

assumes that the laminate has failed when any

stress component in a ply exceeds its respective

allowable. Using these CDI_s, a fatigue life cycle

curve for a reliability of 0.999 was obtained (fig. 2).

The results show that at low mechanical cyclic loads

and low thermal cyclic amplitudes, the fatigue life

for 0.999 reliability is most sensitive to matrix

compressive strength, matrix modulus, thermal

expansion coefficient, and ply thickness. However,

at high mechanical cyclic loads and high thermal

cyclic amplitudes, the fatigue life at 0.999 reliability

is more sensitive to the shear strength of the matrix,

the longitudinal fiber modulus, the matrix modulus,

and the ply thickness.

Lewis contact: Chx/stos C. Clmm/s, (216) 433-3252

T/BEST: Technology Benefit Estimator

Select Features and Applications

Progress in aerospace propulsion necessitated an

assessment of the benefits to be gained by

interfacing advanced technologies. These benefits

will be used as guidelines to identify and prioritize

high-payoff research areas, to manage research with

limited resources, and to demonstrate the link

between advanced and basic concepts. An effort was

undertaken at the NASA Lewis Research Center to

develop a formal computational method, T/BEST

(Technology Benefit Estimator), to assess advanced

2 NASA TM-107300
Material 1.2 1.2

BETitanium

E] Graphite epoxy [_+45/90/0]s

0.0 0.0

Weight Cost

Figure 1.--Maurer stock weight and cost estimation for

manufacturing a fan blade.

aerospace technologies and to credibly communicate

the benefits of research. Fibrous composites are

ideal for structural applications such as highperformance aircraft engine blades where the ratios

of high strength to weight and stiffness to weight

are required. These factors, along with the flexibility

to select the composite system and layup and to

favorably orient fiber directions, reduce the

displacements and stresses caused by large

rotational speeds in aircraft engines.

T/BEST can readily evaluate the benefits of utilizing

composites to construct fan and compressor blades

and to update the blade geometry to maximize rotor

efficiency and the blade manufacturing process.

Figure 1 shows the cost benefits of using graphite/

epoxy composites instead of titanium in the

manufacturing of state-of-the-art blades. The cost

required to manufacture composite fan blades is

estimated using T/BEST (fig. 2).

Lewis contact: Christos C. Chamis, (216} 433-3252

Subprocess

• _ 1. Procure material,

cut plies, collate plies

10.6% _ 2. Assemble pressure and

suction halves, inspect

_ 3. Mold, inspect

12.9O/o/(_/ _ _! I [_ 4. Machine root, inspect

__ _ _ 5. Fabricate and bind

root outsert

6. Machine tip

2.1%

_i_:_:_i_ 7. Bond stainless

6.1% _ 8._meh Nickel plate mesh

_9. Mold shelf halves

11.20_ _ 10. Bond shelf moldings

57.1 %

11. Inspect final

Figure 2._ost of manufacturing process of resin matrix composite blades.

NASA TM-107300 3
ProbabilisUc Thermomechanical Fatigue

of Polymer Matrix Composites

Traditional computational approaches for life and

long-term behavior rely on empirical data and are

not generic or unique. Also, those approaches are

not easy to implement in the design procedure in

an effective integrated manner. The focus of ongoing

research at the NASA Lewis Research Center has

been to develop advanced integrated computaUonal

methods and related computer codes to perform a

complete reliability-based assessment of composite

structures. These methods account for uncertainties

in all the constituent properties, fabrication process

variables, and loads to predict probabilistic micro,

ply, laminate and structural responses. These

methods have already been implemented in the

Integrated Probabilistic Assessment of Composite

Structures (IPACS) computer code. The main

objective of the present evaluation was to illustrate

the effectiveness of IPACS to predict the long-term

behavior of composites under combined mechanical

and thermal cyclic loading conditions. A unified

time-, stress-, and load-dependent MultiFactor

Interaction Equation (MFIE) model developed at the

NASA Lewis Research Center was used to simulate

the long-term behavior of polymer matrix

composites. The MFIE model evaluates the

magnitude of the degradation and the properties of

constituent materials at every cycle step at that

temperature which, in turn, is used for

micromechanics and laminate analysis. Possible

impending failure modes are checked at every cycle

6O

oo

d

_= 50

e_

E

m 40

P

30

a.

E

¢)

.2

O

20--

lO I

0.50 0.80

I I I I t

0.55 0.60 0.65 0.70 0.75

Ratio of current to final cycles, Nm/Nmf

Figure 1.--Fatigue life variation for 0.999-reliability of

[0/+45/90]s graphiteepoxy laminate. Ply thickness,

O.127 mm under thermal cyclic load; mean mechanical

load = 0.5 xstatic strength.

.,Ii / / c,,c

2 0.6 / ,' / / temperature

/ " / / amplitude,

>_ ' ,' / / oc

=m 0.4 / / / /----- 15-6

" : / /--- 23.9

0.2 -' ,' / ........ 37.8

/ ,.' / / ---

0.o )_L -___'l___/ t I I I

0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2

Ratio of current to final cycles, Nm/Nmf

Cyclic

• temperature

• .> amplitude,

1.0- _ _ _ oC

"5 _ _ _ 156

I_ O _

o i " o _ _ 23.9

._ 0.8-- = r-o o_ _ 378 o

x "_ "i X I_ _ "

_ ._xc _= _ 51.7 ,.

u= O.6_=EH _x- =E_ E

P w "=_ ¢-_ c = ._o

•_ ,-,43d _la I I_IZ] .__. "6

•1_ > I..

® •[=_- E = J

x_ "6 o

" l_. > Z

<o>0011NININN IN 

Figure 2..--Cumulative distribution function for and sensitivity of fatigue life for 0.999-reliability of [0/+45/90] s

graphiteepoxy laminate. Ply thickness, O.127 mm under

thermal cyclic load; mean mechanical load = 0.5 x static

strength. (a) Cumulative distribution function. (b) Sensitivity of fatique life for 0.999-reliability.

step. The deterministic part of the methodology

has been implemented in the in-house computer

code ICAN (Integrated Composite Analyzer). A

methodology to compute the fatigue life for different

ratios of applied stress to the laminate strength

based on first ply failure criteria combined with

thermal cyclic loads was demonstrated. Degradation

effects due to long-term environmental exposure

and thermomechanical cyclic loads were considered

in the simulation process.

4 NASA TM-107300
ApplicationofIPACSwasillustratedbyconsidering

a [0/+-45/90] s graphite fiber/epoxy matrix

composite.The fatiguelife cycleswere computed

for differentthermal cyclesandfor differentratios

r of the magnitude of applied stress cycles to the

laminate strength cycles based on first ply failure

criteria. These curves can be used to assess the

fatigue llfe of a component subjected to mechanical

cyclic loading for a given reliability (fig. 1).

Cumulative probability distribution functions for

mechanical fatigue due to different cyclic stress

magnitudes and the respective sensitivity factors

are shown in figure 2. Note that results similar to

those in figures 1 and 2 can also be developed for

different stress ratios but with constant

temperatures.

Lewis contact: ChHstos C. Chain/s, {216) 433-3252

Prestraining Effects on Fatigue Life

It is generally acknowledged that the fatigue life of

metals is influenced by prior prestraining. What is

not well understood is the nature and extent of the

influence. One school of thought views prior

deformation as being beneficial because it imparts

cold working that invariably raises the yield and

ultimate tensile strengths of metals. These in turn

ce-

?.

1o-I

10-2

....... I ' ' ' .... I .... '''

Prestrain,

percent

• -2

A 2

[] 5

O 10

_ Fatigue life l

[]_I) 0 _1

t

. , J , lllll . , , ,,,LI i i . . ,,,,t

104 105 106

Cyclic life, Nf

103

Figure 1.---Fatigue results for prestrained specimens

compared with baseline data for Inconel 718 superalloy.

can raise the fatigue strength. This view arises from

prior deformation being imposed as a result of

rolling, swaging, forging, and other mechanical

metal-working operations wherein the applied

compressive deformation is not aligned with the

direction of subsequent fatigue loading. The

opposing view that prior deformation is damaging

to ensuing fatigue resistance has its origins in

deformations aligned with the direction of fatigue

loading. Losses of ductility, flattening of voids into

sharp discontinuities with higher stress

concentration factors, and actual initiation of

microcracks are among the mechanisms for

imparting damage. An experimental program was

conducted to study the effects of both tensile and

compressive axial prestrains on the subsequent

fatigue life of the nickel-base alloy Inconel 718. All

testing was at room temperature. Completely

reversed, axial strain-controlled testing was

employed. Baseline tests with zero prestrain were

conducted over the range of 103 to 106 cycles to

failure. Additional specimens were prestrained 2,

5, and 10 percent in tension and 2 percent in

compression prior to conducting fatigue tests at

the baseline loading conditions. No beneficial effects

were observed, although compressive prestrains

were less detrimental than tensile. The general

trend of results was for tensile prestrain to cause

the greatest life losses at the highest lives (fig. 1).

Three fatigue damage llfing models were used to

predict the experimental results: the Linear

Damage Rule, the Linear Strain/Life Fraction Rule,

and the Nonlinear Damage Curve Approach. The

NASA TM- 107300 5
I08

Q

UJ

nO.

c 106

6

m

Q)

.__

E

0

IE

Prestrain,

percent

• -2

A 2

[] 5 •

0 10

..... Factors of 2 ,..-"

on fatigue life --_''/"_'I ..

_,,_ "°_ _,o_. '_S

102 J J

103 104 105 106

Remaining fatigue life, (n2)OBS

Figure 2._Remaining fatigue fife predicted by nonlinear

Damage Curve Approach.

Smith Watson-Topper parameter was used to

estimate the effects of any fatigue mean stresses

present. The most accurate predictions of fatigue

life following prestrain were provided by the

Nonlinear Damage Curve Approach (fig. 2).

Lewis contacts: 8reermmesh Kallttx/, (216} 433-6727 and

Gary R. Halford, (216) 433-3265

Structurally Compliant Rocket Engine

Combustion Chamber

A new, structurally compliant design for a rocket

engine combustion chamber was validated through

a series of analyses and experiments. Subscale,

tubular channel, rocket nozzle chambers were

cyclically tested and analytically evaluated. Cyclic

lives were determined to have a potential for a

tenfold increase over those of rectangular channel

designs, the current state of the art. Greater

structural compliance in the circumferential

direction gave rise to lower thermal strains during

hot firing, resulting in lower thermal strain

ratcheting and longer predicted fatigue lives

(fig. I). Thermal, structural, and durability analyses

of the combustion-chamber design, involving cyclic

I

I

r

I

I

i

I

I

I

I

Figure 1._Computed deformed shapes of conventional rectangular and compliant tubular

channels after first five successive cycles of

hot firing (IO00X).

6 NASA TM-107300
2

E

._=

etJ: 1 ITubular

BB Rectangular

0 1 2 3 4 5

Cycles

Figure 2._alculated thinning of hot-gas wall of conventional rectangular channel and compliant tubular

channel configurations.

temperatures, strains, and low-cycle fatigue lives,

have corroborated the experimental observations

(fig. 2).

Lewis contacts: Robert Jankovsky, (216) 433-7515,

Vinod Arya, (216) 433-2816, and Gary R. Halford, (216)

433-3265

Kinetics of Cyclic Oxidation and

Cracking and Finite Element Analysis

of MA956 and Sapphire/MA956 Composite System

Sapphire fiber-reinforced MA956 composites hold

promise for significant weight savings and increased

high-temperature structural capability compared

with unreinforced MA956. As part of an overall

assessment of the high-temperature characteristics

of this material system, cyclic oxidation behavior

was studied at 1093 and 1204 °C. Initially, both

sets of coupons exhibited parabolic oxidation

kinetics (fig. 1). Later, monolithic MA956 exhibited

spallation and a linear weight loss whereas the

composite showed a linear weight gain without

spaUation. Weight loss of the monolithic MA956

resulted from the linking of a multiplicity of

Unequibiaxial straining of oxide layer

c 0.010 c 0.010

o0o8 o .,

. o .. ex

ooo, .x I ;

'_ 0.002 _ _ '_ 0.002 =_o o

E 0.000 E 0.000

0000000000000

00000000

MA956/sapphire plate MA956/plate

Figure 1._aximum mechanical strains in oxide layer of composite

(MA956/sapphire) and matrix (MA956) plate (continuum analysis).

Equibiaxial straining of oxide layer

NASA TM-107300 7
0.00581 ,

Effective

mechanical

strain,

m/m

0.00817

.00738

.00660

.00581

.00503

.00424

.00346

z

.00267 |

x _Y Maximum strain in upperoxide .00189

layer is 0.00817 along A-A'.

Figure 2.--Effective mechanical strain distribution in

MA956/sapphire composite unit cell (micromechanical analysis).

randomly oriented and closely spaced surface

cracks that facilitated ready spallation. By contrast,

cracking of the composite's oxide layer was

nonintersecting and was aligned nominally parallel

with the orientation of the subsurface reinforcing

fibers. The oxidative lifetime of monolithic MA956

was projected from the observed oxidation kinetics.

Linear-elastic, finite element continuum, and

micromechanics analyses were performed on

coupons of the monolithic and composite materials.

Results of the analyses qualitatively agreed well

with the observed oxide cracking and spaUation

behavior of both the MA956 and the sapphire/

MA956 composite coupons (fig. 2).

Lewis contacts: Gary R. Halford, (216) 433-3265 tnd Vinod

K. Arya, (216} 433-2616

Large-Displacement Structural Durabil/ty Analyses of Simple Bend Specimen

Emulating Rocket Nozzle Liners

Large-displacement elastic and elastic-plastic finite

element stress-strain analyses using an updated

Lagrangian formulation were performed on an

OFHC (oxygen-free high-conductivity) copper plate

specimen. The plate specimen is intended for lowcost experiments that emulate the most important

thermomechanical loading and failure modes of a

more complex rocket nozzle. The plate, which is

loaded in bending at 593 °C, contains a centrally

located and internally pressurized channel (fig. 1).

The cyclic crack initiation lives were estimated using

the results from the analyses and isothermal straincontrolled low-cycle fatigue data for OFHC copper.

The table presents a comparison of the predicted

and experimental cyclic lives and shows that an

elastic analysis predicts by a factor greater than 4

a longer cyclic life than that observed in

experiments. The results from elastic-plastic

analysis for the plate bend specimen, however,

predicted a cyclic life in close agreement with that

of experiment, thus justifying the need for the more

rigorous stress-strain analysis.

Figure 1._Deforrned channel of plate specimen

emulating the "bulging" effect.

EXPERIMENTAL AND PREDICTED CYCLIC LIVES OF

PLATE BEND SPECIMEN

Analysis Maximum total Life,

effective strain range, cycles

percent

Large elastic 1.1 500

Large elastic-plastic 1.7 190

Experiment - - - 109

Lewis contacts: Gary R. Halford, (216) 433-3265 and Vinod

K. Arya, (218) 433-2816

8 NASA TM-107300
Effect of Hoop Stress on Ball Bearing Life

To prevent motion of the inner race of a ball or roller

bearing around a shaft, designers have been

specifying extremely tight interference fits between

the inner race and the shaft when it is not practical

to provide a keyway of locknut arrangement. The

interference fit is usually based on the anticipated

growth of the shaft and the bearing under the most

severe operating conditions. These conditions

sometimes exist only for short periods in the

machine's operating cycle. Nevertheless, it is an

extremely important design consideration for both

safety of operation and maintainability. In recent

years, some engineers have noticed that bearings

with tighter than usual press (or interference) fits

may have shorter field lives than anticipated or

calculated. The failure mechanism is usually

classical rolling-element (subsurface) fatigue.

A finite element analysis (FEA) of a generic,

dimensionally normalized inner race of an angularcontact ball bearing was performed under varying

conditions of speed and press (or interference) fit

of the inner-race bore on a journal. The FEA results

at the ball-race contact were used to derive an

equation from which was obtained the radius of

an equivalent cylindrical bearing race with the same

or similar hoop stress. The radius of the equivalent

cylinder was used to obtain a generalized closedform approximation of the hoop stresses at the ballinner-race contact in an angular-contact ball

bearing. A life analysis was performed on both a

45- and a 120-mm-bore, angular-contact ball

bearing. The predicted lives with and without hoop

stress were compared with experimental endurance

results obtained at 12 000 and 25 000 rpm with

the 120-mm-bore ball bearing (fig. 1). A life factor

equation based on hoop stress is presented. A

threshold hoop stress below which, for a given

bearing steel, hoop stress would be ineffective in

decreasing life is suggested.

Lewis contact: Erwin V, Zaretsky, (216) 433-3241

Experimental Theoretical

L10 life I-10 life 5.5x1_ 24x103

-- • 12 000 rpm Without hoop stress

2.4--_ • 25 000 rpm With hoop stress

5.0 -- 22 ,, ,.

4.5 -- 20 -- ,, _,, _ Inner-race

', \ \, _ speed,

_ ',, \ \, _ rpm ._ 4.0 -- 18

-- z 12 000

-'_3.5 -- 1_16I_ _"_"_ "_'i iiii!_",

'_ '_ ', ,_,

,- 2 14 '_ "_,

F- 3.0 -- _ _ "

12 _ ",,

2.5 -- ', "'-,

% %% 1o

2.0 -- , ,,

1.5-- I i I i I L I il I l I t I t I

6

2 4 6 8 10 20 40 60 80

LI0 life, billions of inner-race revolutions

Figure 1.--_ornparison of theoretical Llolife with and without hoop stress and experimental LIO

life at a thrust load of 22 241 N (5000 Ibf) for 120-turn-bore, angular-contact ball bearings.

Contact angle, 24°; race temperature, 218 °C (425 °F); bulk oil tempereture, 191 °C (375 °F);

material, VIM-VAR AISI M-50 steel; inner race, forged; material hardness, Rc 63; difference between hardness of rolling elements and inner race, O;lubricant, tetraester MIL-4.-23699; life

factors, a2 (material and processing), a3 (operating conditions).

NASA TM-107300 9
Predicting Rolling-Element Bearing Life

Nearly five decades have passed since G. Lundberg

and A. Palmgren published their life theory in 1947

and 1952 and it was adopted as an ANSI/ABMA

and ISO standard in 1950 and 1953. Subsequently,

many variations and deviations from their life theory

have been proposed, the most recent being that of

E. Ioannides and T.A. Harris in 1985. A critical

analysis was performed comparing the results of

different life theories and discussing their

implications in the design and analysis of rollingelement bearings. Variations in the stress-life

relation and in the critical stress related to bearing

life are discussed using stress fields obtained from

a three-dimensional, finite element analysis (FEA)

of a ball in a nonconforming race under varying

load (fig. I). The results showed that for a ninth

power stress-life exponent, the Lundberg-Palmgren

theory best predicts life as exhibited by most airmelted bearing steels. For a 12th power relation

reflected by modern vacuum-processed bearing

steels, a Zaretsky-modifled Weibull equation is

superior. The assumption of a fatigue-limiting stress

as advocated by Ioannides and Harris distorts the

stress-life exponent and overpredicts life.

Normal load Lewis contact: Erwin V. Zaretaky. (216) 433-3241

all-race model

1.0

_ 0.8 solution

o'61

P

0.4

N

_ o.2_

o.o t

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6

Normalized depth below groove surface, Z/b

Figure 1._C,omparison of finite-element stress analysis (FEA) with closed-form solution for principal stresses and

_45 shear stress as a function of depth below groove surface. Ball diameter, 12.7 mm (0.5 in.); groove conformity,

52 percent; normal load, 623 N (140 IbO;maximum Hertz stress, 1.4 GPa (200 ksi); semimajor contact axis, a,

1.3 mm (0.05 in.); semiminor contact axis, b, 0.16 mm (0.006 in.).

10 NASA TM-107300
Structural Mechanics

Figure 1._Stiffness composite panel finite element

model.

Effect of Combined Loads on the DurabUity of a Stiffened Adhesively Bonded

Composite Structure

The progressive fracture and damage tolerance

characteristics of a stiffened graphite/epoxy panel

(fig. 1) were investigated via computational

simulation. An integrated computer code,

CODSTRAN, was utilized for the simulation of

composite structural degradation under loading.

Damage initiation, growth accumulation, and

propagation to structural fracture are included in

the simulation. The results indicated that damage

initiation and progression significantly effect

structural behavior under loading. The damage

versus the applied bending load is plotted in figure 2. For an applied load equal to 88 percent of

0.60 --

0.48--

_Q.0.36 --

0.24 -

r_ 0.12 --

i

o.oo I I

0.50 0.60 0.70 0.80 0.90 1.00

Normalized bending load

Figure 2.--Damage initiation and growth under bending

load.

the failure load (1.1354 kN-m or 10.05 kips-in.),

the fracture propagates very quickly and the

panel collapses at a load equal to 1.1354 kN-m

{I0.05 kips-in.).

Lewis contact: Pascal K. Gotsis, (216) 433-3331

Damage Progression in Bolted Composite Structures

Structural durability, damage tolerance, and

progressive fracture characteristics of bolted

graphite/epoxy composite laminates (fig. 1) are

evaluated via computational simulation. An

integrated computer code, CODSTRAN, was used

I 8 in. ___.

_"_ - (20.32 cm)

i 2 in.

='_(5.08 cm): 4

4 in.

(10.16 cm)

I

!

Figure 1..--Bolted laminate finite element model.

Loading

for the simulation of structural degradation under

loading. Damage initiation, growth, accumulation,

and propagation to fracture were included in the

simulation. The results showed the damage

progression sequence and structural fracture

resistance during different degradation stages. The

damage versus the applied load is plotted in

figure 2. Because of matrix fracture, damage

initiation began at a load of 17.792 kN (4 kips). The

damage growth accelerated very fast at 51.152 kN

(11.5 kips), involving more damage in the matrix

and fracture in the fibers. Finally, fracture in the

bolted laminate occurred at 191.26 kN (43 kips).

The results indicate that damage progression

characteristics are not very sensitive to the laminate

width or to the bolt space in the investigated range.

NASA TM-107300 11
36

28--32___-- Structural_

8--

J Damage growth

4 -- /_ Damage initiation

0 -----Y_ I I I I

0 10 20 30 40 50

Load, kips

I I I I I

0.00 44.48 88,96 133.45 177.93

Load, kN

Figure 2.---Damage progression for bolted laminate.

24

G)

=o

m 20 D.

em16 ¢0

E

c0

a 12

An optimal value of the bolt spacing-to-diameter

ratio of 2 gives the best ultimate fracture load of

approximately 2.3 MN/m.

Lewis contact: Pascal K. Gotsis. (216) 433-3331

/

/

/

/

/

/

¢

/

/

/

/

/

/

- L _

L/N=4

R/t = 33.3

R = 25.4 cm

W-o/o/o]

Computational Simulation of Fiber

Composite Thin Shell Structures in a

Hygrothermal Environment

A computational simulation of fiber composite thin

shell structures subjected to static loads in a

hygrothermal environment was performed using

CSTEM, a three-dimensional finite element

computer code. The graphite/epoxy fiber composite

cylindrical shell structure is composed of a [+0] 2

angle-ply laminate (fig. 1). The simulation of the

processing was taken into account. The results

showed that (I) the ply angle 0 significantly

influenced the global and local behavior of the fiber

composite thin shell (fig. 2); (2) the angle of twist

Thermal

load

I

rocessing _,

Time

Figure 1._Geometry, materials, and processing of laminated cylinder. Fiber volume ratio, 55 percent; moisture,

2 percent; fibers, 7"300;matrix, epoxy with high strength; angle-ply laminate [_0]_" curing temperature Tg of the

matrix equal to 215.5 °C.

12 NASA TM-107300
o,

'%

.o

o,

o.

o.

,--[+20] 2

/

///F- [+30]2

/

//////,-- [+10]2

[:L-6012

[+70]2

L/R=6

R/t = 33.3 [+90]2 .......... .o....

"....°..

0.2 0.4 0.6 0.8 1.0

Longitudinal distance, x/L

Figure 2._Radial displacements along axial line that passes through point of contact with

concentrated load as function of normalized longitudinal distance xlL. Fiber volume ratio,

55 percent; fibers, moisture, 2 percent; 7"300; matrix, epoxy with high strength.

was influenced significantly for 0_<0_<10 and

80_<0<90, whereas it was influenced moderately for

10_<0_<80; (3) the global results obtained using the

three-dimensional finite element method were in

reasonably good agreement with those obtained

using an elementary beam theory.

Lewis contact: Pascal K. Gotsis, (216) 433-3331

Optimization Improves Air-Breathing

Propulsion Engine Concept

The optimization of air-breathing propulsion

engines with constraints specified on the discharge

temperature, pressure ratios, mixer Mach number,

compressor and fan maximum speeds and surge

margins has been cast as a nonlinear optimization

problem. Engine thrust is the merit function, and

bypass ratios, R-values of the fans, and fuel flow

are typical active design variables. Because of the

large number of mission points, diverse constraint

types, and overall ill-conditioning of the design

space, the most reliable optimizer available in

CometBoards (which has 1 dozen, state-of-the-art

optimizers) would provide feasible optimum results

only for a portion of the flight envelope. Therefore,

a solution strategy that soft couples the engine

performance analyzer NEPP with the CometBoards

design tool was developed. Some of the unique

features of CometBoards (cascade strategy, variable

and constraint formulations, and scaling devised

for difficult multidisciplinary applications) could

successfully optimize the performance of both

NASA TM-107300 13
subsonicand supersonicenginesover their flight

envelopes. The combined CometBoards and NEPP

strategy converged to the same global optimum

solution even when it was initiated from different

design points. Typically, when NEPP was used

alone, many cycles of manual interventions were

necessary to extract an acceptable solution because

the optimization scheme was unreliable. The

reliable and robust combined tool eliminated

manual interventions and it was much easier to

use, adding value to the engine cycle analysis.

Solutions for two numerical examples, a waverotor-topped subsonic engine and a mixed-flow

turbofan engine design, illustrate the capability of

the combined design tool. A wave-rotor-topped

engine with four ports was optimized for a

47-mission flight envelope. So that the benefit of

wave-rotor topping could be assessed, most

baseline variables and constraints were declared

passive, whereas the rotational speed, heat added,

and fuel flow were declared active variables. The

important active constraints included limits on the

temperature and on the maximum speeds and

surge margins of all the compressors. Engine thrust

was the merit function. The cascade strategy

successfully solved the design (with 47 subproblems, one for each mission point). For the

mission at Mach 0.1 and a 5000-ft altitude,

figure 1 shows the convergence of the two-optimizer

cascade strategy (sequential quadratic programming (SQP) followed by the method of feasible

directions (FD)). The SQP optimizer produced an

infeasible design at a 67 060.87-Ib thrust. At a

thrust of 66 901.28 Ib, the second FD optimizer

produced a feasible optimum design that was

verified graphically. Engine thrust was the merit

function for the optimization of a mixed-flow

supersonic engine with constraints specified on the

maximum compressor speed, acceptable surge

margins, discharge temperatures, pressure ratios,

and mixer Mach number. The bypass ratio, mixer

pressure balance, fuel flow, and R-values for fan

and compressor were active design variables. The

122-mission flight envelope required the solution

of 122 optimization subproblems. The combined tool

successfully solved the engine problem producing

the same global solution even when it was initiated

from different points. Figure 2 depicts the engine

solutions, which were normalized with respect to

NEPP answers. The combined tool was superior for

most of the 122 missions, successfully solving the

122 mission mixed-flow turbofan engine design

optimization problem.

Lewis contacts: Dale A. Hopkins, (216) 433-3260 and 6urya

N. Patnalk, (216) 962-3135

7.75x104

_.B7.50

_ 7.25

_ 7.00

E

.E_6.75

o

-=: _

t'_' _

O, O'J

6.50 ____ SQP-----_i_,-_ -- FD =

6.25 I I i I I l

0 2 4 6 8 10

Number of design iterations

Figure 1.---Convergence of cascade solution for a

wave-retor-topped engine.

I

12

2

S

E

o

._N

o

Z

1.00

Solution

1.10

I CometBoards

NEPP optimizer

1.05 _

0.95 1 I [ I I

0 25 50 75 100 125

Number of mission points

Figure 2.reValue-added benefit in design of mixed-flow

turbofan engine.

14 NASA TM-107300
A Novel Boundary Condition Completed

the Beltrami-Michell Formulation in

Elasticity

In elasticity, the method of forces, wherein stresses

are considered the primary unknowns, is known

as the Beltrami-Michell Formulation. This method

can solve only stress boundary value problems; it

cannot handle the more prevalent displacement or

mixed-boundary-value problems of elasticity.

Therefore, the classical method, which has

restricted application, could not become a true

alternative to the Navler's displacement method,

which can solve all three types of boundary value

problems. The restriction was alleviated by

augmenting the classical formulation with a novel

set of conditions identified as the boundary

compatibility conditions. This new method, which

completes the classical force formulation, has been

termed the Completed Beltrami-Michell

Formulation. The completed formulation can solve

general elasticity problems with stress,

displacement, and mixed-boundary conditions in

terms of stresses as the primary unknowns. The

a

t Xl x 2 x 3 x4

a ] _t I II

t t t t t t t t t t t t , t,,t

I- L -_- L q_, -t 1 2 3

(a)

e_

Z

6 -20

o

o

_ .40

o

Z

-60

lllllllll

' , ! I I I I I l

I I I I I I I I I

fX._ .j

I|illllll

Figure L--Analysis of composite cylindrical shell. (a) Composite cylindrical shell. (b) Normal force distribution along

x-axis.

completed formulation is derived from the stationary

condition of the variational functional of the

Integrated Force Method. In the completed BeltramiMichell formulation, stresses for kinematically

stable structures can be obtained without any

reference to displacements either in the field or on

the boundary. Displacements, ff required, can be

back calculated from stresses.

The solution obtained through the completed

Beltraml-Michell formulation for a composite shell

with mixed-boundary conditions is shown in figure I. The composite shell, made of two different

materials, is subjected to thermomechanical loads.

Notice the peak response and the discontinuity at

the shell interface.

Lewis contacts: Dale A. Hopkins, (216) 433-3260 and 6urya

N. Patnalk, (216) 962-3135

Active Thermal Distortion Management

With Smart Piezoelectric Structures

The development of smart piezoelectric structures

offers great potential for their use in advanced

aerospace applications. By combining the

traditional performance advantages of composite

materials with the inherent capability of

piezoelectric materials to sense and adapt to their

environment, piezoelectric composite materials can

provide dramatic advantages in the development of

smart structures with the capability to actively

manage adverse thermomechanical conditions. To

investigate potential applications for active thermal

distortion management, a thermal analysis

capability was incorporated in a refined laminate

theory and a finite element formulation developed

for both beam and plate elements. A numerical

study was performed on a cantilevered [453/-453]

graphite/epoxy plate with attached piezoceramic

patches. The application of a thermal gradient, with

the piezoceramic patches grounded, resulted in the

combined bending and twisting deformation shown

in figure 1 (also shown is the original undeformed

geometry). Through the application of active voltages

with varying polarities, the bending and twisting

deformation of the plate can be controlled either

individually or simultaneously. The resulting

deformation from inhibiting both the bending and

twisting behavior simultaneously is shown in

figure 2. These results demonstrate the potential

NASA TM-107300 15
o033t 0.00

-0.67 0.00

Vj.<_._---__----____ 020

-_.oo_ _ _ 0.40

0.40 0.80 0.80

y/b 0.80 1.00

1.00

Figure l.--lnitial thermal bending and twisting deformation of [4531-453/plate

under a thermal gradient.

1.00--

0.67

E

0,00

0.20

0.40

0.60

y/b 0.80

0.00

0.20

0.40

0.60 x/a

0.80

1.00

1.00

Figure 2.DActive compensation of thermal bending and twisting deformation of

[4531-453/plate under a thermal gradient.

16 NASA TM-107300
application of smart piezoelectric structures for

active thermal distortion management of propulsion

components and space structures.

Lewis contacts: Ho-Jun Lee, {216) 433-3316 and Dimitris

A. Saravanos, (216) 962-3211

Micromechanics for Woven Composites

There is a growing interest in the use of woven

composites for structural applications.

Reinforcement of these composites is accomplished

by various processes such as weaving, braiding, or

knitting. Woven composites, in particular, are

constructed by weaving two fiber tows into each

other to form a layer. The interlacing of fiber bundles

has several advantages such as increasing the

intra- and interlaminar strength, providing greater

damage tolerance, and making it possible to

produce near net shapes for thick structural

components. The design methodologies for such

composites must account for geometrical as well

as processing parameters. A micromechanics-based

technique that is used in combination with

micromechanics-based analysis codes was

developed to analyze such composites. The

intention of the present work was to develop a

methodology that could make use of the capabilities

that already exist for the analysis of laminated

continuous fiber-reinforced composite mechanics

codes developed in-house at the NASA Lewis

Research Center and yet account for fiber waviness

and fiber-end distributions through the thickness.

Because this technique is based on simplified

micromechanics equations, it is inherently more

efficient than numerical analysis techniques and

TABLE L---SiC/SiC COMPOSITE PROPERTIES

[Fiber volume ratio, 0.4; void, 0.2.]

Property Micromechanics Measured

prediction

Density, g/cm3

Exx, GPa

F__, GPa

G_y, GPa

Gyz, GPa

Vxy

V_z

2.45

214.4

193

80

79.3

.18

2

2.5

214

0.17

400 r --"0--" Micromechanics based

/

& Data

200

100

0 I I I I

0 5 10 15 20

Porosity, percent

Figure 1.--ln-plane modulus versus porosity of SiC/SiC

composite. Fiber volume ratio, 0.4.

"o

o

I

25

provides a very detailed response. The properties

for a generic graphite/epoxy plain-weave textile

composite and a SiC/SIC plain-weave composite

were generated using the above technique, ICAN

(Integrated Composite Analyzer), and CEMCAN

(Ceramic Matrix Composite Analyzer) computer

codes. Limited verification with a set of detailed

three-dimensional finite element analyses as well

as with available experimental data shows good

agreement (table I). Fabrication-related issues and

environmental degradation effects can be easily

taken into account using this micromechanicsbased technique (fig. 1). Work is underway to

incorporate this model into the existing composite

mechanics codes and to develop capabilities that

will allow prediction of the composite properties

and/or response of three dimensionally reinforced

fabric composites.

Lewis contacts: Bubodh Mltad, (216) 433-3261 and Pappu

Murthy. (216) 433-3332

NASA TM-107300 17
Ballistic Impact Research in Support of

HSCT Engine Fan Containment

Ballistic impact tests and analyses are being

conducted at the NASA Lewis Research Center to

support the development of a lightweight, hightemperature, engine fan containment system for the

High Speed Civil Transport initiative. During 1995,

materials and concepts were evaluated by

conducting impact tests on fiat 2- by 2-ft panels

and structures in the Structures Division's Ballistic

Impact Facility. Disk-shaped projectiles (4.5 in.

(11.43 cm) in diam by 0.75 in. (1.90 cm) thick)

weighing 2 Ib (0.09 kg) were accelerated to speeds

in excess of 1200 ft/sec and were impacted on the

test articles. A number of different concepts and

materials were found to be viable candidates for

reducing the weight of a baseline fan containment

system.

Computer analyses were conducted to support these

tests and the large subcomponent tests at the

Figure 1.mTypical progression of fan blade impact on subcomponent panel

Figure 2.---Ballistic impact facility showing 6-in. (15.24-cm) gun.

18 NASA TM-107300
Universityof DaytonResearchInstitute.TheLSTC

DYNAthree-dimensionalexplicitfiniteelementcode

wasusedto designtheexperimentsandto predict

their outcomes•Test conditions and test article

configurationsweredeterminedbasedin part on

thefinite elementpredictions.Figure1depictsthe

progression of a fan blade impacting a subcomponentpanel.

TheBallisticImpactFacility{fig.2)hasseveralguns

with barrelsrangingfrom 0.0625to 8 in. (0.1588

to20.32 cm) in diameter and is capable of shooting

projectiles at velocities in excess of the speed of

sound. The facility is supported by high-speed data

acquisition, video and film photography, and

modeling with DYNA.

Lewis contacts: J. Michael Pereira, (216) 433-6738 and

Matt Melis, (216) 433-3322

Monitoring Damage Progression in

Ceramic Matrix Composites

Ceramic matrix composites (CMC's) are being

evaluated for use in the engine for the High Speed

Civil Transport. There is a need to understand how

these materials respond to extreme mechanical and

thermal conditions. Traditional nondestructive

techniques have had limited success in evaluating

these materials. A unique vibration technique was

used on a model CMC material to obtain a

specimen's natural frequency and damping value.

3.5 D

3.0 --

2.5--

G)

¢_ 2.0 --

=Z

0

1.5

O3 ¢0 1.0

0

0.5

r

.e

oO

I ,,,.,Ill I ,t,,,,,I , ,L,,ll,l t ,I,,,I,I , ,L,.,L,I

100 101 102 103 104

Number of fatigue cycles

Figure 1.--Vibrational NDE loss factor versus mechanical

cycles for SiCSiC [451-45] composite.

After various mechanical loading cycles at room

temperature, measurements were conducted on an

enhanced SiC/SIC CVI (chemical vapor infiltration)

material with three layup configurations and two

stress levels. The frequency decreased dramatically

after the first mechanical loading cycle and then

leveled off after successive cycles. The loss factor,

twice the damping value, increased dramatically

after the first mechanical cycle and then slowly

increased, as shown in figure l. Gripping of the

sample had negligible effect on either frequency or

loss factor. The greater the stress, the greater the

decrease in frequency. The modulus calculated from

the frequency response compared well with the

direct measurement. The changes in frequency and

loss were associated with damage in the material.

The conclusion was that the most significant

damage occurred in the first loading cycle. This

vibration technique tracked modulus changes and

damage progression in CMC's.

Lewis contacts: Christopher Rabzak, (216) 977-1238 and

J. Michael Perelra, (216) 433-6738

High-Temperature Strain Gage and

Adhesive System Development and

Testing

Polymer-matrix-based composites continue to be

evaluated for high-temperature applications in

engine components. A unique strain gage and

adhesive system was developed for the mechanical

property measurement of polymers and polymer

composites at elevated temperatures. This system

overcomes some of the problems encountered in

using commercial strain gages and adhesives. An

important limitation with typical commercial strain

gage adhesives is that they require a postcure at

temperatures substantially higher than the

maximum test temperature. The exposure of the

specimen to this temperature may affect

subsequent results; in some cases, the postcure

temperature may be higher than the glass transition

temperature of the polymer. In addition, although

typical commercial strain gages can be used for

short times at temperatures up to 370 °C (700 °F),

their long-term use is limited to 230 °C (450 °F).

This precludes their use in testing some hightemperature polyimides near their maximum

temperature capability. The system that was

developed consists of a nonencapsulated, unbacked

gage grid that is bonded directly to the polymer

NASA TM-107300 19
0.16

-'- Extensometer _.,_

°oo12

0.08

_0.04

0 0 5' 10' 15' 20' 2--5 30

Load, Ib

Figure 1._omparison of elevated-ternpereture strain

gage and extensorneter on PMR-15 specimen at 315 °C

(600 °F).

using a specially formulated adhesive. The gage is

applied after the specimen has been cured but prior

to the normal postcure cycle. The adhesive was

formulated to cure under the specimen postcure

conditions. Special handling, mounting, and

electrical connection procedures were developed;

a fixture was designed to calibrate each strain gage

after its application on the specimen. At elevated

temperatures, a variety of tests was conducted to

determine the performance characteristics of the

gages on PMR- 15 neat resin specimens. These tests

included static tension and compression, thermal

exposure, and creep. Figure I compares the gage

and an extensometer on PMR-15 at 315 °C

(600 °F). The gage and adhesive system performed

within normal strain gage specifications at

temperatures up to 315 °C (600 °F).

Lewis contact: Christopher Rab=ak, (216) 977-1238 and

J. Michael Pereira, (216) 433-6738

Micromechanics Predict Uniaxial

Response and Local Stresses of Composites With Shape Memory Alloy Fibers

Shape memory alloys (SMA's) have found

increasing applications as candidate smart

materials and components of adaptive structures.

One potential application of shape memory alloys

is their use as fibrous sensors or actuators within

a composite material system. The development of

fibrous shape memory (SM) composites wiU provide

weight savings, directionality in actuation and

sensing, and some control over the resultant SM

effect. However, to analyze the complicated

thermomechanical response of unidirectional

composite systems with continuous shape memory

alloy fibers, the development of new micromechanical models is required. Such a

micromechanical method based on a combination

of the three-concentric-cylinder model (fibers,

matrix, and equivalent shape memory composite)

with the multicell micromechanical method was

developed. The micromechanics can predict the

uniaxial response and the local stresses in the

composite when the fibers undergo martensitic or

reverse-phase transformations as a result of

applied mechanical or thermal loads.

Applications on candidate shape memory

composite systems with Nitlnol fibers (-30 percent)

in an epoxy matrix illustrated that these

composites may exhibit a unique longitudinal

response under isothermal loading and unloading

conditions (fig. 1). The capability of shape memory

composites to induce very high longitudinal strains

upon the application of mechanical stress (fig. I)

or with a variation of temperature was also

demonstrated. The ranges of stress, strain, and

temperature where such actuation may occur may

be also predicted. The evolution of microstresses

in the composite during stress- or temperatureca 400 o.

:E

0

¢,1

E

0

0

._c

tO

--I

3OO

200

100

0

-100

-0.02

C D

0.00 0.02 0.04 0.06 0.08 0.10 0.12

Axial strain

Figure L--Predicted isothermal loading-unloading stressstrain cycle of 30-percent-fiber-volume-ratio nitinoll

epoxy composite (40 °C). Segment AB, loading with

fibers in parent (austenite) phase (6 = 0); segment BC,

loading with fibers undergoing transformation to martensitic phase (0<6<1); segment DE, unloading with fully

transformed fibers (6 = 1); segment EF, unloading

with fibers undergoing reverse phase transformation.

20 NASA TM-107300
induced phase transformations can be also

simulated.In summary,the researchhasenabled

themodelingof the complex,yet promising,active

responseofSMcompositesandhasoutlinedsome

of their unique advantagesas activematerialsin

themanagementofthermaldistortionsin advanced

propulsioncomponents.

Lewis contacts: Dimitris A_ Saravanos, (216) 962-3211 and

Dale Hopkins, (216) 433-3260

Generalized Finite Element Enables

Prediction of Global and Local Response

of Multilayered Composite Plates With

Embedded Piezoelectric Actuators and

Sensors

The development of a new class of smart composite

materials and adaptive structures with sensory and

active capabilities is expected to further improve

the performance and reliability of aeronautical

structural systems. However, such an effort requires

the development of admissible mechanics entailing

capabilities to model the unified electromechanical

response of sensory/active structures as well as

the evolution of the complicated stress-strain fields

in smart composites and the interfacial phenomena

between the embedded microdevices and the

passive composite plies.

Recent developments in generalized mechanics for

smart composite laminates with embedded

piezoelectric actuators and sensors have

culminated in the development of a specialty fournode finite element for adaptive piezoelectric plates.

The finite element provides unique features not

found elsewhere, including the capabilities to

model (1) the global response of structures with

embedded actuators and/or sensors; (2) the local

through-the-thickness response (intralaminar and

interlaminar stresses, interfacial phenomena

between structure and piezoelectric devices, etc.),

(31 a wide spectrum of thicknesses, ranging from

very thin to very thick structures; and (4)

continuous or discrete piezoelectric actuators and

sensors.

1.0 _ i i Solution I

_l ]Finite i Exact element

) e _rx ] _rx

0.. ! + _y i ...... _y

i ] ,

•_ i"_ -<_"_" _ F"--" _.. - ,_"+"-+'"i

! -'-..- i

°, ! "'t z

4.5 k '

/

P_i 0° I 900 l 0° I PZT

i I i I , I i , i i

0.0 0.2 0.4 0.6 0.8 1.0

Thickness, z/h

Figure 1.--In-plane modal stresses for fundamental mode in thick [0/90/0] square plate

with surface-bonded piezoceramic sensors and comparisons with exact solution.

Spanthickness ratio, 4.

NASA TM-107300 21
0.05

0.04

0.01

© Measured

Predicted

0

0

0

0.2 0.4 0.6 0.8 1.0

Axial distance, x/L

Figure 2.--Predicted bending deformations of [01+_45]s cantilever plate beam induced by

opposite-polarity electric field applied on shown surface-bonded actuators. Measurements reported by Crawley and Lazarus (1989).

Numerous case studies have provided analytical

and experimental verification of the finite element

and the prototype software. Figure 1 illustrates one

of the extreme capabilities of the mechanics: the

excellent prediction of local axial stresses in a thick

composite plate with surface-bonded piezoceramic

(PZT-4) sensory layers. A typical simulation of the

global deformation of a laminated cantilever plate

with discrete piezoelectric actuators is shown in

figure 2. In summary, the evaluations have

demonstrated the versatility of the finite element

to analyze and thus to facilitate the formal

development of adaptive piezoelectric components

for advanced propulsion systems.

Lewis contacts: Dim/tris A. 8aravano,,. (216} 962-3211 and

Dale Hopkins. (216} 433-3260

Fatigue and Fracture

A Study to Optimize a Ring-On-Ring Test

Fixture for Testing CMC's Under In-Plane

Biaxial Loading Conditions

A study was conducted to optimize a ring-on-rlng

test fixture to investigate the applicability of

composite materials to key engine components used

in high-speed aircraft. These components operate

under severe and highly complex loading conditions

and are continuously exposed to high temperatures

and aggressive environments that lead to cracking

and other structural deficiencies. The ability to

reliably design and life-test these components

depends on the generation of a comprehensive data

base for composite materials under complex loading

conditions. Equibiaxial bending, a condition that

occurs in many components, is a through-thickness

thermal gradient. Because no experimental results

are available in the open literature for equibiaxial

bending failure of fiber-reinforced ceramic

composites, a test method was proposed. Analytical

verification of the testing setup is the first step to

develop a standard test procedure.

22 NASA TM-107300
Ring-on-ring fixture and specimen

z

Finite element model ring-on-ring

fixture (2219 nodes and

2040 elements)

z

Y

Figure l.nFinite element model of ring-on-ring biaxial flexure specimen.

fil

I11

Ill

Ill

ill

I

I

= I I !

, !

I

: : ;

I

= ,

i

I

" ii'

: II:

i i i

NASA TM- 107300 23
500

40O

3OO

200

e=

o. 100 :E

_ o

-1 O0

Stress Load,

kN

+- Radial 0.845

...... Radial 2.05

x--- Radial 3.225

o - - Radial 4.50

[]-- Radial 5.80

o--- Hoop 0.845

A- Hoop 2.05

m ...... Hoop 3.225

--- Hoop 4.50

......... Hoop 5.80

-200 f

-3O0

--400

0.000

l I I

0.005 0.010 0.015

Distance, m

!

[3

t

0.020

Figure 2.--Contact stresses at ring-specimen interface.

I

0.025

A finite element analysis was conducted on a biaxialflexure specimen ring-on-ring setup (fig. I) to

evaluate the stress-straln field for monolithic and

ceramic matrix composite materials. The evaluation

included a comparison of the finite element results

with a closed-form solution and experimental data.

Effects of the disk geometry and the material

characteristics on the resulting stress-strain field

and the contact stresses were explored (fig. 2),

modeled, and discussed. The analysis considered

two-dimensional axisymetric calculations as well

as a three-dimensional anisotropic solution.

Another objective of this work was to further

promote the adoption of the disk and related test

procedures as a standard biaxial flexure specimen.

Lewis contacts: Ali Abdul-Aziz, |216) 433-6729 and

Dave Brewer, (216) 433-3304

24 NASA TM-107300
Yield Surface Study of Cast Gamma

Titanium Aluminide (TiAI)

Cast gamma TiAl has been identified for advanced

aeropropulsion components where high stiffness

and low weight at high temperatures are required.

These components are geometrically complex and

are composed of relatively thin (1.0- to 2.0-mm)

structural elements. The typical TiAI microstructure

has grain sizes ranging from 0.5 to 1.0 mm. The

combination of thin structural members and large

grains can provide a potential for nonhomogeneous

material behavior (e.g., anisotropy). In an attempt

to investigate this potential, a series of yield surface

tests were conducted at room temperature on two

cast TiAl materials (Ti-48Al-2Cr-2Nb and

proprietary XD).

The tests were conducted under in-plane biaxial

loading conditions at room temperature. The

cruciform specimen used in these tests has a test

section 95 by 95 mm in area and 2 mm in thickness

(fig. I). For both materials, an equivalent inelastic

strain limit of 150 microstrain was used to define

the yield surface. This small strain limit was

measured using several strain gages mounted in

the test section. A final equiaxed tensile test was

performed for each specimen after establishing the

yield surface.

As illustrated in figure 2, the XD material exhibited

a typical polycrystalline material yield surface. In

contrast, the Ti48AI-2Cr-2Nb (48--2-2) TiAI yield

surface indicated a preferential yield in the Y-loading

axis. This yielding behavior is typical of a material

that has a texture. A macroetching technique

revealed flow patterns in the grain structure of the

48-2-2 material. A study is currently investigating

the relationship of this flow pattern to the observed

texturing effect.

Customers: General Electric Aircraft Engines and

Pratt & Whitney

Lewis contacts: Paul A. Bartclotta, (216) 433-3338,

Peter Kantzos, (216) 433-5202, and David L. Krause,

(216) 433-5465

Figure 1.mFailed in-plane biaxial specimen of cast TiAI.

100 ' ' ' '

Failure stress, ksi

80 • XD (65)

O 48-2-2 (60)

6O XD Yield

40 surface _."_lf

--6O

I i i i

, @____@ 0

48-2-2 Yield

surface Reference yield

-80 surface (45 ksi)

-100 ........

-100-80 --60 -40 -20 0 20 40 60 80 100

_xx, ksi

Figure 2.--150-microstrain yield surfaces for 48-2-2 and

XD TiAI at room temperature.

NASA TM-107300 25
Figure 1.--Mechanical loading fixture with demonstration component, loading ram, universal

support pivots, and sliding rockers.

EPM Structural Component Successfully Tested Under Pseudo-Operating

Conditions

A fabrication feasibility demonstration component

for the Enabling Propulsion Materials (EPM)

program was evaluated under prototypical engine

loading conditions at the Structural Benchmark

Test Facility at the NASA Lewis Research Center.

The purpose of this test was to verify EPM casting,

joining, coating, and life-prediction methods. The

component was fabricated by Joining two large

superalloy cast sections of an exhaust nozzle flap.

These two sections were then Joined together using

electron beam welding techniques developed in the

EPM program. After the joints were inspected, the

component was coated with an oxidation-resistant

barrier coating and was sent to Lewis for testing.

A special test fixture (fig. 1) was designed and built

at Lewis to produce a biaxial bending condition

similar to the loading condition this part would

encounter in engine operation. Several finite

element analyses were conducted to validate the

mechanical test method. A floating furnace was then

designed to provide prototypical thermal profiles

in the component. An isothermal low-cycle fatigue

(LCF) test was used to evaluate the component. A

cyclic load of 13 kN (max) to I kN (min.) at a

frequency of I Hz was employed. Component failure

was defined as a 30-percent increase in compliance.

Based on this definition, the LCF life of this

component was 35 000 cycles.

As predicted, a fatigue crack was initiated in the

welded joint associated with the high-stress

location. The local temperature at the failure site

was 425 °C. The predicted life of the component

was 18 000 cycles. This prediction was based on

several liflng methods developed for conventional

superaUoys. As shown in figure 2, using average

material properties (solid circle symbol) would

predict lives that were not very conservative (54 000

cycles) and were more representative of those of an

unwelded component. To account for the welded

26 NASA TM-107300
¢0

e_

o

106

• Unwelded IN939 800 °F ,'/

(using average material ," J ,_

properties) jl/j

• Welded IN939 800 °F ,"/,'_

minium properties ," J ,"

105 -- (using Kt = 1.5) ,"X""

104 -- jJ'" J'"

/ / SJ_ /p/Sf/

103 ]

103 104 105 106

Predicted life

Figure 2.mLife prediction of demonstration component

showing predicted life for welded and unwelded component. Component life, 35 000 cycles.

joint and its unknown properties, minimum parent

material properties and a stress concentration

factor Kt of 1.5 were used as knockdown factors.

Thus, the predicted life of the welded component

was 18 000 cycles (solid square symbol). The Lewis

life-prediction method was within a factor of 2 in

life from the actual demonstration component life.

Note that specifics on test conditions, material

details, and engine operation conditions have been

omitted because of Limited Exclusive Rights

restrictions.

Customers: General Electric Aircraft Engines and

Pratt & Whitney

Lewis contacts: Paul A. Bartolotta, (216) 433-3338,

Gary R. Halford, (216} 433-3265, and David L. Krause,

(216) 433-5465

Fatigue Behavior of Haynes 188 Under

Thermomechanical, Axial-Torsional

Loading

It has been shown that multiaxial loading and

thermomechanical loading can cause changes,

relative to uniaxial isothermal loading, in the

deformation and damage accumulation displayed

by a material. Seldom have both multiaxial and

thermomechanical loading been combined in a

single experiment. The results of axial-torsional,

thermomechanical fatigue experiments on thinwalled tubular specimens fabricated from the

wrought cobalt-base superalloy Haynes 188 are

shown in figure 1. The fatigue test matrix consists

of uniaxial, thermally in-phase (TIP, maximum

temperature coincides with the maximum

mechanical strain) and thermally out-of-phase (TOP,

maximum temperature coincides with the minimum

mechanical strain) experiments; a torsional

thermomechanical experiment; a mechanically inphase (maximum axial mechanical strain occurs

at the same point in time as the maximum torsional

strain), thermally in-phase experiment (MIPTIP); a

mechanically in-phase, thermally out-of-phase

experiment (MIPTOP); a mechanically out-of-phase

(maximum axial mechanical strain occurs when the

torsional strain passes through zero), thermally inphase experiment (MOPTIP); and a mechanically

out-of-phase, thermally out-of-phase experiment

(MOPTOP). In all the experiments, the imposed

maximum and minimum temperatures and axial

and torsional strain ranges were the same. Four

multiaxial life models were used to predict the

fatigue lives of these experiments. Using fatigue life

parameters derived from previous axial and

torsional experiments at the peak temperature, it

was found that the Von Mises equivalent strain

range and modified multiaxiality factor approaches

predicted the thermomechanical fatigue lives

somewhat better than both the Smith-WatsonTopper and the Fatemi-Socie-Kurath parameters.

In general, the largest deviations from the

experimentaUy observed lives were exhibited by the

Fatemi-Socie-Kurath model.

Lewis contacts: Peter J. Bonacuse, (216) 433-3309 and

6reeramesh Kalluri, (216} 463-6727

NASA TM-107300 27
105 105

104

._u

o. 103

102

102

oo°°'°

Factors of _..-"_ ..-'"

2 on life _...-'"

..'"-_"_.'"'_ TorsionalTIP

.-"" / ."'" [] AxialTIP

V ""1_'_-_ "'" _ AxialTOP

V ..'" _I_ .-'- V MIPTIP

.-_*,_,."'" L_ MIPTOP

-'"/..--'" T M° LP_

/. ,-'" • MOPTOP (a)

• ' L L I I1' i I = F I I P I I I I II

103 104 105

Observed life

-= 104

"o

._o

_- 103

102

102

.,'°"

Factors of ,.-_ .-i-"""""'_

2 on life _...-_..'_f....-""

.-'_'_.-'"_ Torsional TIP

.._..,"'-/Yf,"'" D Axial TIP

V .-'I_ ,-'" C: Axial TOP

_7..,•'_.J _ ,-"" V MIPTIP

,'_"/_'_" ,-"" A MIPTOP

,'"/ .-'" •MOPTIP

'_ .-'" , • MOPTOP(c )

103 104 105

Observed life

105

_104

nD

.L2

a. 103

105

oo''*

Factors of ...'"'_

2 on life _ .-'"

."'"/.'_.'"'_ TorsionalTIP

..Y" _ •.Y" [] Axial TIP

w ..._,./v.- {_ Axial TOP

Y .."" _ .."" _ MIPTIP

..

,--/ /,_" am • MOPTIP

/ ..-"- • MOPTOP (b)

"*'; i I I I] I I i ' I III ; , I I i I I

103 104 105

Observed life

_104

lo

.2

1O

P103 a.

102 102

102 102

oo•'°

Factors of ..-'"'_

2 on ----life .."- _..-"" _I(

• .

...7....... n Axia,T";

..-'I._/ .,'" (_ AxialTOP

,_,.-"+.,,_J.-"" _ MIPTIP

..'V/_..-"" A MIPTOP

..-_ .L,._" • MOPTIP

103 104 105

Observed life

Figure 1._Observed versus predicted fatigue lives for various fatigue fife models. Therrnomechanical, axial-torsional

fatigue of Haynes 188. (a) Von Mises equivalent strain range model (b) Smith-Watson-Topper model. (c) Modified

muitiaxiality factor approach. (d) Fatemi-Socie-Kurath model

Retirement for Cause as an Alternate

Means of Managing Component Lives

In the current budgetary environment, fielded

equipment will often be used beyond its design life.

To avoid the large cost of replacing critical rotating

parts as they reach their safe-life limits, a retirement

for cause (RFC) program may prove to be a costeffective and safe alternative. Studies indicate that

a full 80 percent of parts replaced at low-cycle

fatigue (LCF) calculated safe-life limits have at least

a full order of magnitude remaining fatigue life. The

Air Force has embraced RFC and currently

successfully uses it to manage parts life for several

of their gas turbine engines. RFC involves periodic

nondestructive evaluation (NDE) to access the

damage state (whether or not detectable cracks

exist) of components (fig. I). Those components that

do not have detectable cracks are returned to

service. This approach allows low-life parts to be

detected and discarded before they can cause an

incident, and high-llfe parts to be used to their full

potential. Although there are initial and ongoing

costs associated with the procurement of the

inspection equipment and the performance of the

inspections, these costs have been more than offset

by the savings in replacement parts. Basic to an

RFC program is the calculation of crack growth

rates under the expected service loads (mechanical

and thermal). The results are used to define safeuse intervals between required (NDE) inspections.

The starting crack size for the fracture mechanics

analysis is a flaw that is just below the detection

limit of the NDE technique employed. Crack growth

in ductile materials is sensitive to loading sequence

in that large-amplitude load excursions in the early

28 NASA TM-107300
0.8

0.7

0.6

E 0.5

E

_0.4

_e

0.3

0.2

0.1

Crack length at fracture

/

Undetected ,,

crack growth ---_,""

If no cracks found,

return to service

NDE detection limit and assumed initial crack size

0,0 , , _ J , , , , I , _ , _ I , , , L I J , ,

0 1000 2000 3000 4000 5000

Operating hours

Figure 1.mFatigue crack growth under retirement for cause (RFC) inspection regimen. Inspection

interval, 1000 hr.

stages of crack formation can retard the crack

growth rate whereas in the later stages of crack

growth, these same overloads can lead to

catastrophic failure. A population of components

subjected to variable-amplitude loading will exhibit

a distribution in crack growth lives (in excess of

that observed in constant-amplitude loading). To

accurately assess the component reliability, this

variability must be characterized. A NASA/Army

Research Laboratory (ARL) team is in the process

of providing analytical support to Aviation and Troop

Command (ATCOM) in the development of an RFC

program for the T700-700 engine used in the

Blackhawk helicopter. The FASTRAN-II fracture

mechanics analysis code developed at NASA is being

employed to estimate crack growth lives.

Lewis contacts: Pete Bonacuse. (216) 433-3309, Anthony

Calomino, (216) 433-3311, Louis Ghosn, (216) 433-3249,

Pete Kantzos, (216) 433-5202, Jane Mandersheid,

(216) 433-3314. John Shannon, (216) 433-3321, and

Jack Telesman, {216) 433-3310

Thermomechanical Deformation and

Strain Rate Sensitivity of the Dynamic

Strain Aging Alloy Haynes 188

Over the past several years, the Structural Fatigue

Branch at the NASA Lewis Research Center has

had an ongoing effort to develop a comprehensive

high-temperature fatigue data base on the cobaltbase superalloy Haynes 188 as it has been

consistently down-selected for use in many hightemperature structural aeronautics and space

applications. These data are being used for both

characterizing the alloy and for guiding the

development and verification of deformation and

damage models for this class of materials. Initial

research of this alloy revealed that the hightemperature deformation behavior is extremely

complex and strongly influenced by a number of

interacting time- and temperature-dependent

phenomena, such as inelastic flow, dynamic strain

aging (DSA), and dynamic precipitation. Thus, to

gain a thorough mechanistic understanding,

examination of the material deformation behavior

was pursued at the macroscopic (phenomenological)

and microscopic (substructural} levels. To this end,

NASA TM-107300 29
550 ..... I ..... , ..... r

¢0

n

==

¢0

eE

.E_

_E

500

450

Temperature,

T,

°C

n 300

A 550

© 650

<> 700

400

-- _ = 10-4s -1 (solid)

350

3OO

10 o

_ = 10-3s -1 (hollow)

101 102 103 104

Cycles

Figure L---Cyclic stress response during instantaneous strain-rate sensitivity tests with A_ = + 0.4 percent and _ = I0-3s -1 to 10-4s -1.

a comprehensive effort was undertaken to evaluate

the isothermal low-cycle fatigue (LCF) and

thermomechanical fatigue (TMF) behavior of Haynes

188 over the temperature range 25 to I000 °C at

multiple strain rates.

To obtain insight into the macroscopic aspects of

DSA during LCF loadings, the instantaneous swain

rate sensitivity (ISRS} of the cyclic stress was

examined by conducting instantaneous strain-ratechange tests at several temperatures. A negative

ISRS is a well-documented phenomenon that is

generally indicative of DSA. These tests were

accomplished by periodically changing the strain

rate Eto 10-4s -I for one cycle and then returning it

to the higher _ of 10-3s -I. Representative results

from these tests (fig. 1) indicated that the ISRS was

positive below 300 °C, was strongly negative between

400 and 650 °C, was slightly negative at 700 and

750 °C, and was positive again above 800 °C. The

ISRS became substantially negative only in the

temperature range where serrated yielding was also

observed at both _ values (i.e., 400 to 650 °C), thus

possibly indicating that this is the true range over

which DSA mechanisms are operative.

The evolution of tensile and compressive stresses

under in-phase (IP) TMF conditions with various

temperature intervals is given in figure 2 with each

TMF test shown by two curves representing the two

stress amplitudes that occurred at the maximum

and minimum temperatures. In all the TMF tests,

the maximum stresses achieved Just prior to the

onset of failure at the maximum temperatures of

the cycles were comparable to those attained in

corresponding isothermal LCF tests; however, the

number of cycles needed to attain the maximum in

each of the TMF tests was less. In contrast, the

30 NASA TM-107300
750

7OO

650

600

13.

-_ 550

E 500

O3

O3

=¢ 450

E

== 400

350

Temperature

interval,

°C

350 to 550

© 400 to 650

500 to 750

600 to 850

_ Indicates

onset of

failure 600 °C

400°C

500 °C

650 °C

750 °C

350 °C

550 °C

300 _ A_A--A_,A.-A-......_

" A _--'--A_A_ - -- _Ak_

250 850 °C

200 ,I ........ I ........ 1 ....... I ........

100 101 102 103 104

Cycle

Figure 2.---Cyclic maximum (solid symbols) and minimum (absolute value) stress amplitude for inphase thermomechanical fatigue (TMF) tests with _k_= + 0.4 percent and _ = 10-4s -1.

maximum stresses attained at the minimum

temperatures in the TMF tests were not well

represented by the corresponding isothermal LCF

tests. In particular, the tests with AT = 400 to

650 °C, which is contained entirely within the DSA

domain, experienced stress hardening at the

400 °C amplitudes, which was 65 percent greater

than that experienced in the 400 °C LCF test. Even

more noteworthy, this extreme TMF hardening

exceeded the greatest hardening level achieved in

the entire isothermal LCF data base (at 650 °C) by

35 percent. Thus, Haynes 188 clearly exhibits

thermomechanical path-dependent behavior,

indicating the need for such data during the process

of developing and characterizing deformation

theories intended to predict material behavior for

this class of materials.

Lewis contacts: Michael G. Castelli, (216) 433-8464 and

Rod Ellis, (216) 433-3340

A Study of Elevated-Temperature

Testing Techniques for the Fatigue

Behavior of PMC's

The development of polyimides and other advanced

thermoset polymers has expanded the role of

polymer matrix composites (PMC's) to elevated

temperatures, such as airframe and propulsion

system components, where severe operating

environments are encountered. The use of PMC's

in these harsh conditions has raised concerns

relating to the degradation of the physical properties

under simultaneous mechanical and environmental

loads. Few studies have been published dealing with

the high-temperature quasi-static and uniaxial

fatigue behavior of polymer composites, and fewer

still when thermochemical fatigue (TMF) testing is

involved. Because of the limited number of studies

and the wide array of objectives, a focused effort is

required to develop experimental procedures for

elevated-temperature (ET) uniaxial testing of PMC's.

NASA TM-107300 31
1.2

1.0

_ o.8',

•g 0.6

0.4

0.2

PMC block

PMC

specimen -- Mounting clip

Cut-a-way view

-- Thermocouple

wires -_

j -Internal

12345 thermocouple (ITC)

Correlation of block thermocouples

with intemal temperatures

7O

6O

o

__50

8

-a 40

I-

_ 30

E

_ 2O

c_

1- 10

0 I I I I

50 100 150 200 250 300

T block, °C

Figure 1.--Thermocouple block (used as controller) and correlation curve. The graph relates

block to calibration specimen that contains internal thermocouples.

1.4 I I I 70

0

Maximum strain _ _..<_

/ _.,0._ ..,0 __ __ ..0,._ __ _ -''0"-

0.0 , , , I , , , I , , , J 40

0 4xl 05 8xl 05 1.2xl 06

Number of cycles, N

65

_. 60

_ 5O

45

I I ' ' I t

Modulus

• Compressive -I

(b)

, , I , , , I _ , , I /

0 4x105 8xl 05 1.2x106

Number of cycles, N

Figure 2.--Typical behavior of 255 °C fatigue specimen A IO; emax = 0.7(%1_"Re = O.1; Nf = 1189090.

(a) Strain retcheting. (b) Static moduli degradation.

This technology is critical to the structural and longterm durability evaluations of PMC's and is

currently undergoing transition to the Advanced

Subsonic Technology Program (AST) in support of

the PMC's static engine structures element. This

investigation was conducted to evaluate experimental techniques for ET uniaxial mechanical

testing of PMC's. To this end, the tensile and cyclic

fatigue behavior of a carbon/polyimide eightharness satln-weave system, T650-835/AMB21,

was investigated. Isothermal fatigue tests (I_ = 0. I}

were conducted at 22 and 255 °C below T G, glass

transition temperature in an effort to resolve issues

such as optimal specimen design, strain

measurement techniques, and temperature control

and measurement methods for quartz lamp heating

32 NASA TM-107300
(employedforTMFcapability).Finally,the subject

of a measurable fatigue damage metric was

addressed.Testprocedureswereimplementedto

measure the stiffness degradation and strain

ratchetlngthat occurredduringthe fatiguetests.

Suchdata are consideredfundamentalinput for

future life-predictionmodelingefforts.Giventhe

primary focusonET experimentaltechniquesfor

PMC's,onlygeneraltrendsin mechanicalbehaviors

were examined.The results are summarizedas

follows:(1)anuntabbeddog-bonespecimendesign

wasachievedwith consistentgagefailures;(2)an

accuratetemperature control and measurement

technique was developed(figs. I); and (3) postprocessing of stress-strain data allowed for

monitoring of fatigue damagethrough various

moduli degradation curves and strain

accumulationdata(fig. 2). Thestatic tensileand

fatiguebehaviorof the material showedthat the

polyimidecomposite maintained its properties at

255 °C. The static ultimate tensile strength ¢3ui t

displayed a reduction of only 4 percent at the

elevated temperature (800 versus 772 MPa). Also,

the fatigue behavior for these relatively short-term

tests showed no significant differences between the

two temperatures in cycles to failure for the load

levels of 0.7, 0.8, and 0.9 ¢_ult"

Contacts: Andrew L. Gyekenyesi, (216) 433-8155 and Michael

M, CastelU, (216) 433-8464

Model Determined for Predicting Fatigue

Lives of Metal Matrix Composites Under

Mean Stresses

Aircraft engine components invariably are subjected

to mean stresses over and above the cyclic loads.

In monolithic materials, it has been observed that

tensile mean stresses are detrimental and

compressive mean stresses are beneficial to fatigue

life as compared with a base of zero mean stress.

Several mean stress models exist for monolithic

metals, each differing quantitatively in the extent

"o

E

m

m

¢0

o

Z

1.0

0.8

0.6

0.4

0.2

0.0

0.0

! ! ! i

•n

%

12

%

8

•0

0 "

Load ratio,

Re

v -1

A -.6

• -.3

O 0

O .3

• .5

+ .7

t1•

÷

%

l I I I I •

0.2 0.4 0.6 0.8 1.0

Normalized mean stress, Sm/Sy

Figure L--Prediction of mean stress effects on fatigue lives of SiC/Ti-15-3 using Soderberg approach.

NASA TM-107300 33
to which detrimental or beneficial effects are

ascribed.Therehave been limited attempts to apply

these models to metal matrix composites (MMC's).

However, because most of the fatigue data has been

limited to tension-tension loading, there has been

a limited range of mean stresses over which models

could be assessed. In this work, a unidirectional

SiC/Ti-15-3 composite was tested with both

tension and compressive stresses, thus extending

the range of imposed mean stresses. The results

showed that tensile mean stresses were detrimental

and compressive mean stresses were beneficial to

the fatigue lives.

Several mean stress models were examined for

applicability to this class of composite materials.

The models were Smith-Watson-Topper, Walker,

Normalized Goodman, and Soderberg. The

Soderberg approach, which normalizes the mean

stress to the 0.02-percent yield strength,

represented the effect of mean stresses best over

the range covered (fig. I). The other models had

significant variation in their predictability and often

failed to predict the composite behavior at very high

tensile mean stresses. This work is the first to

systematically demonstrate the influence of mean

stresses on MMC's and to model their effects.

Attention was also given to fatigue-cracking

mechanisms in the Ti-15-3 matrix and to

micromechanics analyses of mean stress effects.

Lewis contacts: Brad Lerch, (216) 433-6622 and

Gary Halford, (216) 433-3266

Oxidation Embrittlement Observed in

SiC/SiC Composites

As part of a comprehensive materials characterization program, tensile creep-rupture tests were

performed on a SiC fiber-reinforced SiC matrix

composite. The results of these tests and

subsequent analysis revealed an oxidation

embritflement phenomenon that occurs readily at

a discreet temperature range below the maximum

use temperature. Figure 1 shows rupture lives for

a creep stress of 83 MPa as a function of

temperature. Note that the rupture time is constant

at an intermediate temperature range of 700 to

982 °C. Also shown in the figure is the failure

location, as measured from the center of the

specimen. Whereas failure occurred in the specimen

gage section for temperatures of 500 to 700 °C, at

10 3 45 "_

p- 102-

2

o

E

_-. 101 -

10 0

• Rupture time

/X Failure location

/

, I

/ Gage section

/ 1

I 1 I

600 800 1000

t-

--35 °

e-

¢D

E

- 30"_

W

-- 25E

-- 20_

- 15_

- 10_

o

o

- 5¢

0_

400 1200

Gage section temperature, °C

Figure 1.--Stress rupture data for [0190] Dupont enhanced CV! SiClNicaion SiC

composite as function of gage section temperature at creep stress of 83 MPa.

Also shown is failure location as function of gage section temperature.

34 NASA TM-107300
Damaged fiber tows Grip

end

982 °C

Symmetry location of fracture surface

Figure 2.--Creep damage distribution of two [0/90] Dupont enhanced CVI SiC/Nicalon SiC specimens tested at stress of 83 MPa. Damaged fiber tows were

found throughout gage section of 700 °C specimen. However, for the 982 °C

specimen, damage was concentrated at point outside of gage section where

temperature was 700 to 800 °C.

higher temperatures, the failure location migrated

toward the cooled grip ends. Although the results

initially suggested that the test procedure was

influencing the measured creep-rupture lives and

was also driving the failure location out of the gage

section, subsequent additional experiments and a

thermal stress analysis verified the robustness of

the test method employed.

Metallurgical examination of failed specimens

revealed an environmentally assisted material

degradation operating in the temperature range of

700 to 800 °C. Figure 2 shows the damage

distribution of two specimens tested at the same

stress but at different gage section temperatures.

The views shown in figure 2 are sections polished

in the thickness direction, showing the broad

specimen face. The loading direction is horizontal.

In the 700 °C specimen, creep damage is distributed

throughout the uniformly heated test section;

however, in the 982 °C specimen, damage is

concentrated only at the point outside the gage

section where the temperature was 700 to 800 °C

and less damage occurred within the hotter gage

section. Both specimens failed because of an

aggressive environmental attachment of the SiC

fibers at a location where the temperature was in

the 700 to 800 °C range.

SiC/SiC composites are candidates for a combustor

liner application in the engine of the High Speed

Civil Transport (HSCT). During its service cycle, the

combustor liner will experience a thermal gradient,

being cooled near attachment regions and exposed

to combustor gases on the inner wall. The existence

of a minimum in creep-rupture behavior over a

discreet temperature range indicates that the

kinetics of this process are unconventional. Thus,

material properties must be well characterized over

the temperature range of expected operation. Also,

the design and lifing of components must account

for this behavior.

Lewis contacts: Michael J. Verrili, (216) 433-3337,

Anthony M. Calomino, (216} 433-3311, and David N, Brewer.

(216) 433-3304

NASA TM-107300 35
Structural Dynamics

Rotordynamics on the PC: Transient

Analysis With ARDS

Personal computers can now do many Jobs that

formerly required a large mainframe computer. For

example, the program ARDS (Analysis of

RotorDynamic Systems) uses the component mode

synthesis method to analyze the dynamic motion

of up to five rotating shafts. As originally written in

the early 1980's, this program was considered large

for the mainframe computers of the time.

Written in Fortran 77, ARDS has been successfully

ported to a 486 personal computer. Plots appear on

the computer monitor via calls programmed for the

original CALCOMP plotter; plots can also be output

on a standard laser printer. The executable code

uses the full array sizes of the mainframe version

and easily fits on a high-density floppy disk. The

program runs under DOS with an extended memory

manager.

In addition to transient analysis of blade loss, step

turns, and base acceleration, with simulation of

squeeze film dampers and rubs, ARDS calculates

natural frequencies and response to imbalance.

Some examples of ARDS-PC capability are seen in

its use to analyze a magnetic-bearing-supported

rotor as it experiences a sudden increase in

imbalance or drops onto backup bearings. ARDS

drew an outline of the configuration (fig. 1) of the

small rotordynamics demonstrator rotor that was

modeled with 9 elements, resulting in 10 rotor

stations. Concentrated masses were attached to the

shaft at five of the stations. An electromagnetic

bearing was at station 3 and a bronze bushing

supported the shaft at station 8. Magnetic bearings

are customarily used with backup bearings that can

support the rotor in the event of magnetic bearing

failure. For the present example, a backup bearing

in the form of a loose bushing was modeled in

addition to the magnetic bearing at station 3. No

contact occurs in the backup bearing during normal

(magnetically suspended) operation; the backup

bearing is therefore nonlinear in that the stiffness

is zero until the radial clearance is taken up. It was

then assumed to have a constant stiffness in the

radial direction; the tangential force was calculated

as the radial force times a friction coefficient. This

bearing model was built into ARDS. Each computer

run used 100 time steps per revolution; on the 50-

MHZ 486 computer, 4000 time steps took slightly

less than 2 min of calculation time.

A blade loss in a turbine engine introduces a sudden

imbalance that can be many times the normal

operating imbalance. Under this condition, the

magnetic bearing supporting the rotor can become

overloaded such that the backup bearing comes into

0 <) <) <>

1'

0.00 3.31 6.63

i D i i i

9.94 13.25 16.56 19.88 23.19

Configuration scale

Oq/l 4/_5

14:t6

Beatings

Linear

Stn 1 Stn 2

3 0

9 0

Nonlinear

Stnl Stn2

@ 3 o

26.50

Figure 1.--ARDS drawing of example rotor.

36 NASA TM- 107300
DISPLACEMENTCROSS PLOT

STATION NUMBER 3

8_IoCk_ loss tronsN_|

TRAN. Z-DIR. VS TRAN, Y-DIR

OqlO61q*,

10:06

Figure 2._Orbital motion of magnetic and backup bearing for sudden imbalance.

operation. The situation as the bearing makes

contact is similar to that of a turbine wheel

contacting its outer shroud. Figure 2 shows 10

revolutions of a blade-loss transient. The imbalance

is applied to the rotor at station 5. A friction

coefficient of 0.4 was assumed for the backup

bearing; the linear magnetic bearing is also active.

Figure 2 is an orbital plot of the rotor at station 3,

the magnetic bearing location, and shows that the

rotor flies out, hits the backup bearing, bounces

off, and repeats this behavior for the entire time

period plotted.

Figure 3 plots the dynamic behavior of the rotor

when the magnetic support fails and the rotor drops

onto the backup bearing. The rotor walks up the

side of the bearing, although the more sensitive scale

for the horizontal axis in figure 3 exaggerates the

motion. The vibratory motion eventually dies down.

In the analysis, excitation forces can be combined;

for example, a turn in combination with blade loss

and base acceleration.

Lewis contact: David P. Fleming. (216} 433-6013

NASA TM-107300 37
o

O

O

i i

-,o.oo -8. o -6.bo -,.oo -2. o

"10-3 "o

t

0

u_

t

2.00

DISPLACEMENT CROSS PLOT

STATION NUMBER 3 TRAN. Z-DIR.VS TRAN. Y-DIR

t5:32

Figure 3.---Rotor drop onto backup bearing.

A Magnetically Sealed Bearings System

The goal of this project is to develop a retainerless,

noncontacting magnetically sealed bearing that can

be directly substituted for current satellite bearings.

Bearings are becoming llfe-limiting components in

long-duration satellite missions. Typically bearings

are poorly sealed, resulting in premature failure by

lubricant loss and contamination of instrument

sensors by lubricant vapor. Current bearings also

can cause mechanical noise when the ball retainer

rubs or Jams against the raceways. This project

focuses on the development of a bearing for use in

satellite scanning mechanisms.

The magnetically sealed bearing will use permanent

magnets and a magnetic fluid to form a complete

seal on both external faces of the bearing. The ball

retainer will be eliminated and the ball pocket will

be flooded with liquid lubricant. Ceramic bails will

be utilized with metal raceways. Both the magnetic

fluid and the liquid lubricant will have low vapor

pressure characteristics.

Three major milestones for the first year were the

development of a magnetic fluid, the design of firstgeneration bearing prototypes, and the tribological

testing of the magnetic fluid. The development of

the magnetic fluid was conducted in conjunction

with the largest U.S. magnetic fluid production

company. A special hydrocarbon-based fluid was

designed and synthesized. The hydrocarbon base

was NYE--2001, a newly developed, extremely lowvapor-pressure space lubricant. Two magnetic

strength levels of fluid were developed: 100- and

400-G-saturation magnetization, designated N100

38 NASA TM-107300
and N400. No existing magnetic fluids met the

requirements for this system. The successful

completionofthis milestoneprovidedtheenabling

technologyfor this project.

The designof first-generationbearing prototypes

resultedin four different concepts.A 40-mm-bore

bearingsizewas chosenfor the first designs. This

bearing is larger than the scanner bearing, enabling

easier testing at this stage. Second-generation

prototypes will be based on the MPB 17-mm-bore

scanner bearing. Three of the four bearing designs

utilize the fully flooded concept whereas the final

bearing is only lubricated on the wear track.

Bearings using magnets with both axial and radial

magnetization were designed. One prototype was

designed to reuse the magnetic rings after the

bearing failed. All the bearings have been fabricated

and the assembling and preliminary testing have

begun on two of the designs.

A pin-on-disk tribometer was used for the

tribological testing that was conducted to determine

the effect of magnetic fluid migration into the wear

region. Eight fluids were tested: four existing

magnetic fluids, the two fluids developed for this

program, the NYE-2001 base oil used in the

magnetic fluid, and the current liquid lubricant used

in space. The results showed that the base oil

performed best; however, the N 100 magnetic fluid

outperformed all other fluids, including the current

space lubricant.

We have successfully met all the first-year

milestones. Three goals for the second year are (I)

two test rigs must be constructed, one for seal

pressure testing and one for bearing tests; (2) the

existing bearing prototypes must be tested, and the

second-generation prototypes need to be designed

and fabricated; (3) a technology transfer agreement

with ITT Aerospace needs to be formulated.

Lewis contact: Ralph Jansen, (216) 433-6038

An Active Magnetic Bearing With

High-Temperature Superconductor Coils

and Ferromagnetic Cores

A proof-of-feasibility demonstration showed that

high-temperature superconductor (FITS) coils can

be used in a high-load active magnetic bearing in

liquid nitrogen (figs. 1 and 2). A homopolar radial

bearing with commercially wound HTS (Bi 2223)

bias and control coils produced a radial load

capacity of over 890 N (200 Ib) (measured nonrotating) and supported a shaft to 14 000 rpm. The

Figure 1.---Components of high-temperature superconducting magnetic bearing.

NASA TM-107300 39
Figure 2.---Stator assembly of high-temperature superconducting magnetic bearing.

goal was to show that HTS coils can give stable

operation with ferromagnetic cores in a feedbackcontrolled system at a current density similar to

that for copper in liquid nitrogen.

The bias coil, wound with a nontwisted

multifllament I-rrs conductor, dissipated negligible

power for its direct current. The control coils, wound

with a monofilament HTS sheathed in silver,

dissipated negligible power for direct current.

Altemating-current losses increased rapidly with

frequency and quadratically with ac amplitude.

Above about 2 Hz, the effecUve resistance of the

control coils exceeds that of the silver which is in

electrical parallel with the oxide superconductor.

These results show that a twisted multifilament

conductor is not needed for stable levitation but

may be desired to reduce control power for sizable

dynamic loads.

This type of bearing could have application in

cryogenic rocket turbopumps and possibly in future

cryogenically fueled (e.g., liquifled natural gas or

slush hydrogen} aircraft turbine engines. Large HTS

electric motors being developed for industrial

applications could have HTS magnetic bearings

because the cryogenic cooling would already be

present.

High-temperature superconductor technology has

already improved since the coils for this bearing

were purchased and much higher current density

is already possible; hence, still more compact coils

are now possible, allowing cryogenic magnetic

bearings to be either more compact or stronger.

Lewis contacts: Gerald Brown, (216} 433-6047,

FJ/seo DiRuleo and Andy Pxovenza, (216| 433-6025

40 NASA TM-107300
Application of Magnetic Bearing to

Dynamic Spin Rig

The Lewis Research Center's Dynamic Spin Rig,

located in building 5, test cell CW-18, is used to

test turbomachinery blades and components by

rotating them in a vacuum chamber. The rotor is

supported by two mechanical bearings, and during

rotation it is vibrated by voice-coil-type linear

electromagnetic shakers to obtain natural

frequencies of blades. As an alternative exciter, a

magnetic bearing was incorporated in the rig. The

operation of the magnetic bearing as an exciter was

successful and compared favorably with the

electromagnetic shakers.

The rig was modified to allow the magnetic bearing

to operate not only as an exciter but as a fully

functioning magnetic bearing. The magnetic bearing

provided support for the rotor at one end that had

previously been supported by a mechanical bearing.

Additionally, the magnetic bearing provided active

control of stiffness and damping at the point of

action. Testing yielded successful magnetic bearing

operation for rotational speeds up to 9000 rpm in a

vacuum, with control gains producing stiffness up

to 60 000 Ib/in. (67 818 N/mm). Plans are to replace

the second mechanical bearing with a magnetic

bearing.

Lewis contacts: Dexter Johnson, (216) 433-6046, Oral

Mehmed, (216) 433-6036, and Gerald Brown, (216) 433-6047

Measurement of Gust Response on a

Turbine Cascade

Benchmark experimental data on the gust response

of an annular turbine cascade are presented. The

experiment was particularly designed to provide

data for comparison with the results of a typical

linearized gust response analysis. Reduced

frequency, Mach number, and incidence were varied

independently. Except for the lowest reduced

frequency, the gust velocity distribution was nearly

sinusoidal. For the high-inlet-velocity series of tests,

the cascade was near choking. The mean flow was

documented by measuring blade surface pressures

and the cascade exit flow. High-response pressure

transducers were used to measure the unsteady

pressure distribution. Inlet-velocity components and

turbulence parameters were measured using a hot

wire. In addition to the synchronous time-average

pressure spectra, typical power spectra are included

for several representative conditions.

The gusts were generated by a rotor consisting of

pins having either a 3.17- or a 4.7-mm diameter.

The number of pins on the rotor was either 6, 12,

or 24, which resulted in gust reduced frequencies

of 2.5, 5, and 10. The annular turbine cascade had

24 blades and was positioned 3.9 axial chord

lengths behind the rotor. Figure I illustrates the

instrumentation port locations, and the

representative root-mean-square (rms) spectra are

illustrated in figure 2. The synchronous peaks in

this figure could also have been obtained using

linear phase-lock averaging; however, the rmsaveraged spectra also include the nonsynchronous

origin and the random pressure fluctuations. The

frequency units are engine orders; thus, the

synchronous peaks appear at the frequency equal

to the number of rotor pins used to generate wakes.

The position-axis units correspond to blade port

numbers in figure 1; thus, position 0 corresponds

to the port nearest to the leading edge on the suction

• Dynamic

O Steady

1 2 3

10

Figure 1._Instrumentation ports.

0.100

0.075

Cp 0.050

0.025

0"0001

Position 54 3 60

12 Engine order

Figure 2.--Root-mean-square pressure spectra. (a) Reduced frequency, ¢o,10; inlet Mach number, M 1, 0.27;

positive incidence, J32> O.

NASA TM-107300 41
surfaceside,and 11correspondsto theportnearest

to the leadingedgeon the pressureside.

The gust amplitude varied somewhatwith the

reducedfrequency;however,it did not appear to

have a dominant effect. Unsteady, synchronousresponse blade pressures are strongly dependent

on reduced frequency and incidence. Mach number

dependence is weak for negative incidence and

significant for positive incidence at lower reduced

frequencies. The mean blade pressure distribution

is to some extent dependent on the reduced

frequency, particularly for the negative incidence

and the higher inlet Mach number. At a reduced

frequency of c0 = 10, an inlet Mach number of

M I = 0.27, and a positive incidence 132 > 0, a

magnification of the turbulent pressure fluctuations

on the suction side of the aft portion of the blade

resulted in a significant excitation concentrated at

an integral engine order much higher than the

synchronous excitation frequency.

lewis contacts: Anatole Kurkov, (216) 433-5695 and

Barbara Lucci, (216) 433-5902

Blade

number

0.0020 _ 1

*--* 2

*---" 3

0.0012

.-0.0012

-0.0020 ' ' ' '

0.0 2.4 4.8 7.2 9.6 12.0

Time

Figure 1.--Response of rotor blades for tip gap = oo

(unducted rotor). Free-stream Mach number, Mo_

0.5; advance ratio, J, 3.55; setting angle, 13,61.2°.

0.0004

"0

_ -0.0004

Flutter Analysis of Ducted Rotors

An aeroelastic analysis procedure (DuctE3D) was

developed to investigate the flutter of turbomachine

configurations. In the analysis, the unsteady air

loads on vibrating blades are obtained by solving

three-dimensional unsteady Euler equations, and

the blade structural properties are obtained from a

three-dimensional finite element model. The duct

is assumed to be infinitely long and structurally

rigid. Any number of structural modes of blade

vibration can be included in the analysis. The

aeroelastic equations are formulated in normal

mode and are solved for flutter in the time domain,

allowing the analysis of all possible interblade phase

angles. For verification, the analysis is applied to a

ducted fan configuration.

The three-dimensional Euler equations are solved

using an implicit-expllcit hybrid scheme. The

scheme results in a large saving of CPU time as

well as of memory requirements because only two

matrix inversions are required as opposed to three

for fully implicit schemes. The structural analysis

is carried out using the normal mode approach,

assuming the hub and duct to be rigid. The flutter

analysis is carried out by perturbing the blades

from steady state and solving the combined

aeroelastic equation, aerodynamic and structural,

in a sequential fashion while marching in time. The

response of the blades to initial perturbation is

calculated. Increasing the amplitude of the

response implies instability. Any number of normal

modes can be included in the analysis. Further,

no knowledge of interblade phase angle or flutter

frequency is required a priori. A single run is

sufficient to provide information on the blade

aeroelastic characteristics for the particular flight

condition. The response of the blades may be

Fourier analyzed to provide flutter frequency and

interblade phase angle.

The analysis was applied to an eight-bladed ducted

fan obtained by enclosing the SR3C-X2 propfan in

a rigid cylindrical duct. The SR3C-X2 propfan, in

the unducted configuration, showed flutter in the

the NASA Lewis Research Center wind tunnel tests

at a free-stream Mach number M_ of 0.6 for the

eight-bladed configuration operating at an advance

ratio J of 3.55 and a setting angle J3of 61.2 °. The

analysis was conducted at M= = 0.5 using the first

three normal modes for two different tip gaps (I

and 0.4 percent of the rotor radius) and the

unducted configuration. The analysis showed the

42 NASA TM-107300
Blade

0.0020 number

1

•-_ 2

0.0012 ,,--- 3

•--" 4

"0

Q.

E -41.0004

-0.0012

-0.0020

0.0 4.4 8.8 13.2 17.6 22.0

Time

Figure 2._esponse of rotor blades for tip gap of I percent. Free-stream Mach number, M_, 0.5; advance

ratio, J, 3.55; setting angle, [J,61.2 °.

0.0080

0.0048

0.0016

"ID

.i_

O.

E -0.0016

-0.0048

Blade

number

1

2

,--- 3

Z

V

P

i

\/

t

-0.0080 I I I , ,

0.0 3.2 6.4 9.6 12.8 16.0

Time

Figure 3._esponse of rotor blades for tip gap of 4 percenL Free-stream Mach number, M_ 0.5; advance

ratio, J, 3.55; setting angle, _, 61.2 °.

rotor in the unducted configuration to be stable at

Moo = 0.5, correlating with the experimental

observation. The initial perturbations died out

rapidly with the decaying response of all the blades.

The response is shown in figure I for only the first

normal mode response of the four blades. A similar

response was observed for other modes and the

other four blades. The analysis for the tip gap of

1 percent indicated the rotor to be marginally

unstable. The first mode response of four of the

blades is shown in figure 2. For the configuration

with a 0.4-percent tip gap, the initial perturbation

of one blade results in a rapidly increasing response

in all the blades, indicating a flutter instability. The

response for this case in shown in figure 3. These

calculations show that the presence of the duct

changed the aeroelastic characteristics of the rotor.

Lewis contacts: R. 6rivastava, (216) 433-6045 and

T.S.R. Reddy, (216} 433-6083

Fan Blade Deflection Measurement and

Analyses Correlation

Steady deflection measurements were taken of a

NASA/Pratt & Whitney fan blade design while it

was rotating in a vacuum in the Lewis Dynamic

Spin Facility. The 1/5-scale fan blades had a tip

diameter of 22 in. (55.88 cm), a pin-root retention,

and a spar-sheU construction. The purpose of the

test was to measure the radial deflection of the blade

tip and the blade angle change at selected radial

stations along the blade span. The fan blades were

unducted for the test. The procedure for radial

deflection measurement had no precedent and was

newly developed for this test.

The data were compared with analytical predictions

made at NASA Lewis. Radial deflection measurements were used to assure that adequate tip

clearance existed between the fan blades and the

duct for a follow-on wind tunnel test. Also, blade

angle deflection measurements were desired before

pitch-setting parts for the wind tunnel test were

finish machined.

Imser beams were aimed across the blade path at

photodiodes, the signals of which were used to

determine blade angle change or tip radial

deflection. These laser beams were set parallel to

the spin axis at selected radial stations.

NASA TM-107300 43
Results

Calculated Measured

Cruise © Cruise

Reverse D Reverse

ooo

E 0.04 l °r\

o.o2

= o.oo- o.ool-

¢ _0.40 I ] I I l 1

0 20.0 40.0 60.0 80.0 100.0

Percent chord

Figure 1.--Measured and calculated blade tip radial displacements at 10 000 rpm.

"0

t1.0

0.5

0.0

-0.5

-1.0

Radius,

in.

Calculated Measured

7.00 _ 5.25

8.00 © 7.94

9.00 [] 10.73

10.00

10.80

2 4 6 8 10

Rotational velocity, rpm

Figure 2.--Measured and calculated blade twist at cruise.

I

12x103

Some results from the test are shown in figure I,

which compares measured and calculated values

for tip radial deflection at 10 000 rpm for the cruise

and reverse staggers, and in figure 2, which shows

the measured and calculated change in blade angle

versus the rotational speed for the cruise stagger

measured at three radial stations.

Lewis contacts: Oral ]dehmed, (216) 433-6036,

Dave Janetake, (216) 433-6041, and Anatole Kurkov,

(216) 4ss-se95

Space Mechanisms Lessons Learned

Study

A Space Mechanisms Lessons Learned Study was

conducted (1) to determine which mechanisms have

worked in the past and which have failed and (2) to

insure against a U.S. corporate memory loss with

regard to building space mechanisms (mechanical

moving components) for long life and reliability. A

large number of satellite failures and anomalies

have occurred recently (e.g., Galileo and Hubble).

In addition, because of more demanding operating

requirements, failures or anomalies have occurred

during the qualification testing of future satellite

and space platform mechanisms, even before they

were launched (GOES-NEXT, CERES, Space Station

Beta Joint Gimbal, etc.). For these reasons, it was

imperative to determine what worked and what

failed so that the best selection of mechanical

components could be made and to make timely

decisions on initiating research to develop any

needed technology. Thus, the purpose of this study

was to capture and retrieve information relating to

the performance of mechanical moving equipment

operating in space to determine which components

have operated successfully and which have

produced anomalies such as those listed in table I.

Data were obtained from various sources: (1) an

extensive literature review that included

government contractor reports and technical

journals; (2) communication with and visits to

(when necessary) the various NASA and DOD

centers and their designated contractors, including

contact with project managers of current and prior

NASA satellite programs as well as wlth their

industry counterparts. Unpublished information

was then requested from NASA and industry. (4) A

mail survey was designed to establish specific

mechanism experience and also to solicit opinions

of material to be included in a future Space

Mechanisms Design Guidelines Handbook.

The majority of the work was done at Mechanical

Technologies Inc. (MTI) under contract NAS3-

27086.

Lewis contact: Robert L. Fusaro. (216) 433-6080

44 NASA TM-107300
TABLE I.--MECHANICAL ANOMALIES IN SPACECRAFT

System

Momentum wheel spin

bearing

Sensor support bearing

Sensor launch clamp

Harmonic drive ATP

control

Sliprings, brush

contacts

Potentiometer for ATP

control

Control moment

gyroscope

Momentum wheel

Propellent pump

gearbox

Sliprings, brush

contacts

Gear mechanism

Synchronous motor

assembly

Momentum wheel

spin bearings

Inertial guidance

synchronous motor

bearing

Harmonic drive

Momentum wheel active

lubrication system

Solar Array Deployment

Mechanism (SADM)

Gimbal bearings on telescope

Spin bearing

Gas bearing, gyroscope

Foil bearings for

turbomachinery

Conditions

Grease-packed bearings; 3600 rpm,

room temperature to 100 *F

Preloaded ball bearings; oscillatory

motion

Located inside craft; thermal blanketed

Fluorocarbon lubricant; very low

speed; temperature <1 to 50 *F;

boundary condition

MoS2/Ag/C brushes on Ag rings;

numerous recurrences

Low-temperature, light-load fluid

lubricant

Oil injection on beating land

Very high torque for slewing

Grease-lubricated

High speed

MoS2/Ag/C brushes on Ag rings

Fluorocarbon grease; high loads;

boundary conditions

Mineral-oil-grease-packed bearings

Mineral oil grease; high speed

Mineral oil grease; high speed

Fluorocarbon grease; lowtemperature operation

High-speed, long-life requirement

Large launch loads on MoSr

lubricated beatings

Dry MoS 2 lubricant; low

temperature

Large-diameter, thin-cross-section

bearing

Alumina surfaces, stearate lubricant

High-strength alloy; CFx-polyamide

lubricant; temperature extremes

Problem

Torque/temperature anomalies

Failure in test

Seizure on launch pad

Excessive wear; lubricant failure in test

Excessive electrical noise due to

moisture and corrosion

Electrical noise and open circuit due to

lubricant thickening

Beating failure; lubrication design wrong

Bearing failure

Torque and temperature anomalies

Contractor switching lubricants

Excessive noise due to oxidation of

MoS 2

Lubricant degradation

Motor failure due to increased bearing

drag

Possible lubricant degradation in testing

Possible chemical reaction between

grease and iron surface during storage

Excessive torque caused by lowtemperature viscosity of grease

Inability to deliver adequate lubricant

quantity

Testing of static loads

Friction increase due to testing in air

Humidity-induced dimensional instability of cotton-phenolic retainer

Erratic friction on startup; uneven

lubrication during testing

High-friction startup after standing

Impact

Single-point mission failure;

possible indication of failure

>$500K for additional testing

Single-point failure; launch

prohibited or mission failed

Degraded mission or possible

mission failure; changed

lubricant

Inability to point communications

antennas; reduced mission

objective

Mission reduced due to loss of

pointing; ~$500K for additional

testing

Premature mission failure

Loss of mission; >$1M for

additional testing and annealing

Possible mission failure

Possible launch failure with new

lubricant

Reworking of brushes and rings;

delivery delay

System failure

Degraded mission due to failure

Single-point mission failure

Guidance failure

Degraded mission due to failure

Lifetime requirement not met by

system

Possible single-point failure;

passed test

Modified specification to do inert

gas test; passed

Possible target acquisition

failure; changed to metal ball

separator

Reliability problem for flight

units; major rework if failure

Potential system failure; inability

to start turbine

NASA TM-107300 45
Alternative Operating Modes for

Magnetic Bearing Control

Most magnetic bearing control schemes use a bias

current with a superimposed control current to

linearize the relationship between the control

current and the force it delivers. For most operating

conditions, the existence of the bias current

requires more power than alternative methods that

do not use conventional bias. With the existence

of the bias current, even in no-load conditions,

there is always some power consumption. In

earthbound applications of magnetic bearings, this

constant power loss may not be of critical

importance but in aerospace applications, it

becomes an important concern. This research

examines alternative methods that diminish or

eliminate bias current.

The alternative methods presented here

consequently reduce electrical power loss but do

not necessarily provide equal or better

performance. Comparisons of the performance of

these methods are made with the industry standard

bias method. Some of the characteristics that are

examined are slew rate, force output, response to

saturation effects, control amplitude and frequency

dependency, and steady state operation. When

several methods are evaluated, an optimization

strategy is used to choose control parameters that

result in the best tradeoff between power

consumption and control performance.

Subsequently, all methods are compared and

ranked according to their power loss and

performance tradeoff results.

This analysis is performed using computer

simulation, after which the proposed methods are

used in a physical laboratory experiment. The

computer simulation uses Matlab/Simulink tm

simulation software on a microcomputer. The

experiment involves a magnetic bearing apparatus

and supporting electronics. The magnetic bearing

apparatus includes a flexible shaft, two magnetic

bearings, various disks that are used for rotational

inertias, and a motor. Only one magnetic bearing

is used for control.

Lewis contact: Dexter Johnson, (216) 433-6046

A New Antiwear Additive for Bearings

Used in Spacecraft

Bearings currently used in spacecraft for rotating

systems are not completely sealed. If they were used

on satellites, the lubricants used for ground-based

applications would vaporize, form a "cloud" around

the spacecraft, and then condense on the sensors

and solar cells, impairing their sensitivity and ability

to operate. For this reason, low-vapor pressure

lubricants called PFPE's are used on spacecraft.

Unfortunately, the PFPE lubricants used are so

stable that commonly used antiwear additives are

not soluble in them. In addition, they tend to

degrade under boundary lubricating conditions,

limiting the llfe of bearings. The purpose of this

study was to determine if a sflane added to a PFPE

as an emulsion or applied as a coating to the

counterfaces could reduce the wear of bearings

and/or the degradation of a PFPE lubricant. To

evaluate the effect, long- and short-sliding-duration

tests were performed using a pin-on-disk

tribometer.

The results indicated that sflane coatings and/or

emulsions, when used in conjunction with PFPE

otis, reduced the wear of 440C balls and disks as

compared with similar tests on an untreated PFPE

oil. Figure 1 and table I compare the wear rate of

the balls for short and long sliding durations. For

the short-term tests, the silane decreased the pin

wear rate by an average of about 50 percent, and

there did not seem to be any clear advantage in

using coatings instead of emulsions or in combining

coatings with emulsions. However, it should be

emphasized that neither the coatings nor the

emulsions were optimized in this study. In general,

ball wear rates were an order of magnitude lower

for the long-term tests than for the short-term tests,

which could be attributed to run-in effects.

Optical and microFourier transform infrared

(_ FTIR) microscopy observations performed on the

sliding specimen surfaces after the tests were

completed indicated that the sflane disk coatings

and/or emulsions could form thin layerlike transfer

films on ball wear surfaces. Strong evidence was

also found that the sflane mitigated the degradation

of this particular PFPE oil.

Lewis contact: Robert L. Fusaro, (216) 433-6080

46 NASA TM-107300
E

E

0

E

E

0

5xl 0 -16

4 1

3--

2--

1 --

(a)

0

None

• Mean wear rate

7- Standard deviation

O_

¢-

v == .d_

E.C_

I I I I I

Treatment

10x10 -17

8--

6 --

4 --

2--

(b)

None

• Mean wear rate

m Standard deviation

r

.o _"

E o_

Q) e3

O

o

_o _, .__ .c_

•_ _ o

.o E .c ._o.o

"ID _._ W'ID

.=__ 8o ==

" '°i

og =-_

I I I I I I I

Treatment

Figure 1.---Comparison of mean ball wear rates and standard deviations for (a) shortterm and (b) long-term test intervals.

NASA TM-107300 47
TABLE L--COMPARISON OF BALL WEAR RATES FOR SHORT- AND LONG-DURATION TESTS

[Experimental conditions: temperature, 25 °C; atmosphere, dry air (100 ppm moisture content);

load, 2 kg; speed, 200 rpm; 440C stainless steel specimens.]

Disk or ball

treatment

with oil

None

Oil emulsion

Coating

(disk)

Coating (disk

and ball)

Oil emulsion/

coating (disk)

Oil emulsion/

coating (disk

and ball)

Short-term tests (0 to 30 kc) Long-term tests (50 to 4500 kc)

Number Pin wear rate,

of tests m3/m

Maximum Standard

deviation

Minimum Median

131x 10 -18 240x 10- Is

21x lffla 81x10-ts

34>< 10 -is 84x 10 -18

62x 10.18 91x 10 -Is

51x 10 -Is 65× 10- Is

93x lif Is 107× 10 -18

Number

of tests

Pin wear rate,

m3/m

Maximum Minimum Median Standard

deviation

455x 10 -18 120>< 10- TM 7 40x 10 -Is 23x 10 -Is 48x 10 -Is 31x 10 -Is

129><10.18 36><10 -18 4 45× 10 -18 21x lO -Is 31x 10-18 13x lt)- 18

139x 10 qs 34x 10- TM 6 49><10 -Is 15×10-18 27x 10- Is 12x 10- Is

134x 10 -18 38× 10-18 2 18x 10 -Is 9>< 10- Is 14><10-18 6x 10 -Is

79><10.18 12x 10-18 4 29>< 10 -18 lOx 10- ts 21x 10- Is 8x 10-18

120x 10 -18 19x lO -18 2 16><10 -18 9><10 -Is 12x 10 -Is 5× 10 -18

Users Guide for ECAP2D: An Euler

Aeroelastlc Stability Analysis Code for

Oscillating Cascades

An accurate analysis of the blade row aeroelastic

stability is essential for predicting the flutter

characteristics of turbomachinery blades. Toward

this goal, the aeroelastic program ECAP2D was

developed. The program can be used for unsteady

aerodynamic and aeroelastic analyses of isolated

blade rows.

The unsteady aerodynamic forces are obtained from

solving two-dimensional Euler equations using a

finite volume method with a combination of flux

vector splitting and flux differencing splitting. The

Euler solution scheme is third-order accurate in

space and second-order accurate in time. The flow

equations are solved on one or more passagecentered H-grids. The structural model is a typical

section with bending- and torsional-motlon degrees

of freedom for each blade in the cascade. The

solution methods include the harmonic oscillation,

the influence coefficient, the pulse response, and

time integration using Newmark's method. The

program has been calibrated for several examples.

A guide was written to help the user prepare the

input data file required by the ECAP2D code. Also

provided in the guide are a complete description of

the input data; input and output for four examples

from published papers; references for detailed

explanations of the aerodynamic analysis, the

numerical algorithms, the aeroelastic analysis;

and a job control file for executing the program

on Cray YMP computers.

Lewis contacts: T.8.R. Reddy, (216) 433-6083 and

G.L. 8te£ko, (216) 433-3920

Users Guide for FPCAS3D: A Three

Dimensional Full-Potential Aeroelastic Solver

The FPCAS3D computer code was developed

for the aeroelastlc stability analysis of bladed

disks such as those in fans, compressors,

turbines, propellers, and propfans. A guide was

written to help the user prepare the input data

fries required by the code.

The users guide contains a complete description

of the input data and inputs and outputs for

six examples from published papers. Detailed

explanations of the aerodynamic analysis, the

numerical algorithms, and the aeroelastic

analysis are not given. Instead, the reader is

directed to specific references that deal with

these items.

48 NASA TM-107300
TheFPCAS3Dcodewasdevelopedin theStructural

DynamicsBranch.Theaerodynamicanalysisused

in this code is based on the unsteady threedimensionalfull-potentialequation,whichissolved

for a completebladerowor for aspecifiednumber

of bladepassagesin a bladerow.A finite volume

approachisadoptedandaNewtoniterationmethod

is usedto solvethe nonlinearproblemasa series

oflinearproblemsat eachtimestep.Thestructural

analysisisbasedonafinite elementmodelforeach

blade.Eitherafrequency-domainoratime-domain

flutter analysisis possiblewith this code.

Lewis contact: Milind A. Bakhle, (216) 433-6037

Accelerated Testing of Spacecraft

Mechanisms Study

The goal of building longer life unmanned satellites

and space probes has created a demand for

meaningful accelerated test methods to simulate

long-term service in space. This is particularly true

for tribological components such as bearings, seals,

and gears that are used in space. There is an urgent

need for lightweight, low-torque, durable

mechanisms that can operate efficiently in a hard

vacuum environment.

In response to this need, a study was conducted to

determine for space operation the significant

mechanical components that would benefit from

accelerated test techniques. Table I identifies those

mechanisms that could benefit and table II lists the

components used in them. The study also looked at

the current types of accelerated testing techniques,

their shortfalls, and the need to develop new

techniques.

An accelerated testing technology roadmap was

developed to assess the life and reliability of

spacecraft mechanical systems and consisted of

integrating system components testing, analytical

modeling, computer codes, computer smart systems,

and so forth into a methodology that could be used

to predict or verify the life and reliability of a

mechanical system.

Based on the results of the accelerated testing

technology roadmap study, a space mechanism

TABLE I.--MECHANISMS IDENTIFIED

FOR ACCELERATED TESTING

Momentum/reaction wheels

Gyroscopes

Scanning devices for Earth sensors

Solar array devices

De-spin mechanisms

Motors

Storage life, deployment mechanisms

Speed reducers, harmonic drives, and other gear systems

Slider devices (linear actuators)

Tape recorders

Power/signal transmission devices

Robotic joints

Connectors

TABLE II._COMPONENTS NEEDED FOR RELIABLE

OPERATION OF SPACE MECHANISMS

Bearings (rolling element, journal, slider)

Gears

Rings (slip, roll, O-)

Commutators

Gears

Flex harnesses and cables

Low-cycle devices, (springs, switches, cams, etc.)

Seals

Lubricants

mechanical system was suggested for testing to

demonstrate that the methods developed would

adequately predict the life and/or performance of

a mechanism. The roadmap developed includes

experimental equipment, test procedures, time

guidelines, and a cost analysis.

Lewis contact: Robert L. Fusaro, (216) 433-6080

NASA TM-107300 49
GE Uses NASA Seals to Meet JTAGG

Engine Goals

The NASA Lewis Research Center developed and

delivered to General Electric (GE) high-temperature

(> 1500 °F) braided rope seals to provide a combined

seal/compliant mount to minimize thermal stresses

in high-temperature nickel-aluminide (NiAl) turbine

vanes (fig. I). The new seal/compliant-mount

approach has allowed GE to overcome the serious

NiAI vane-cracking problem (observed while using

the conventional brazing approach) that was

threatening the future use of the NiAl alloy system.

The NASA-provided seals consist of a ceramic fiber

core overbraided with fine superalloy wires,

providing a high-temperature (>1500 °F),

compliant, flow-resistant structure well suited for

GE's application.

GE successfully tested the vane/rope seal system

in a full-scale JTAGG (Joint Technology Advanced

Gas Generator) engine, meeting phase 1+

temperature goals (4th quarter of 1995 calendar

year). The NiAI vane/seal system was tested for

hours and no thermal distress or cracking

problems were revealed, thus enabling this system

to be considered for future IHPTET (Integrated High

Performance Turbine Engine Technology) program

elements.

Lewis contact: Bruce 6teinetz, (216) 433-3302

//"'-_ NASAcompliant

Lightweight

high-temperature

• r" turbine vane

/

Figure 1 .DHigh-temperature breded rope seal.

50 NASA TM-107300
Structural Inte rity

Acousto-Ultrasonics and Tensile

Cycle Degradation in Ceramic Matrix

Composites

Advancements in the processing and fabrication of

ceramic matrix composites (CMC's) require

appropriate mechanical and nondestructive testing

methods to characterize properties, assess integrity,

and predict the life of engine components. A serious

form of degradation in CMC's is fatigue from the

formation and propagation of matrix cracks due to

mechanical stresses experienced during service.

In the work described here, mechanical stress was

introduced by way of tensile cycling on SiC/SiC

specimens having four architectures: twodimensional balanced weave, [0/90], [+45], and

[0,+45,90,-45] s. At selected intervals during the

cycling, acousto-ultrasonic (AU) measurements

were made under specific conditions so as to

determine the exponential time rate of the ultrasonic

decay (UD) of the signal as it passed from the point

of introduction into the specimen to the point of

detection. The UD is a measure of the signal

attenuation. These measurements were performed

at room temperature and employed 2.25-MHZcenter-frequency broadband transducers. The UD

rates were taken from the power spectrum of the

detected AU signal. The 1.5- to 2.0-MHZ portion of

the spectrum was used.

Figure 1 shows typical results for a tensile fatigue

cycling experiment on a two-dimensional balancedweave SiC/SiC specimen. The UD rate monitors the

fatigue state up to and at failure. This result was

obtained for all specimens for all four architectures

tested. It is concluded that this AU technique, once

calibrated to a composite system and component

geometry, may be used for lifing aerospace

components.

This work was done in conjunction with Pratt and

Whitney and General Electric by virtue of their

participation in the EPM program.

In general, AU has been applied to manufacturing

problems for wire rope and wood fiber hardboard

and work was done in cooperation with the

MacWhyte Company and the Masonite Corporation.

Acousto-ultrasonics has also been applied to the

problem of determining the physical properties of

human bones for the Hadassah Medical Center in

Jerusalem, Israel.

Lewis contact: Harold E. Kautz, (216) 433--6015

I

v-

.06 er

../ failure

o .04 _ /t_After one cycle

"o

.L2

co .02 I

_<_ Before testing

= I I I I

/

0

0 10o 101 102

Tensile cycle history

©

©

I

103

+1

°_i_llllliil

-1

o -il _-2 Jill

+2

*_I_l _i i i i t i t i I

-2

0 250 500

Time, i_s

Figure 1.--Acousto-ultrasonic decay rate monitors tensile cycle fatigue in

SiCSiC specimens.

NASA TM-107300 51
Award-Winning CARES/Life Ceramics

Durability Evaluation Software: Making

Advanced Technology Accessible

Products made from advanced ceramics show great

promise for revolutionizing aerospace and terrestrial

propulsion and power generation. However, ceramic

components are difficult to design because brittle

materials in general have widely varying strength

values. The CARES (Ceramics Analysis and

Reliability Evaluation of Structures)/Life software

eases this task by providing a tool to optimize the

design and manufacture of brittle material

components by using probabilistic reliability

analysis techniques.

CARES/LIfe is an integrated package that predicts

the probabfllty of a monolithic ceramic component's

failure as a function of time in service. It couples

commercial finite element programs, which resolve

a component's temperature and stress distribution,

to reliability evaluation and fracture mechanics

routines for modeling strength-limiting defects.

These routines are based on calculations of the

probabilistic nature of the brittle material's strength.

The program has many features and options for

materials evaluation and component design. The

capability, flexibility, and uniqueness of CARES/

Life has attracted much interest. To maintain this

interest as well as to keep abreast with fastchanging operating systems and applications

software, CARES/Life has been upgraded with

graphic templates for common business

presentation software such as Lotus Freelance

Graphics. Additionally, an interactive input

preparation program has been prepared and guides

the user through various program control options

and the specific data input formats. A grinding

damage module has been added to account for flaws

introduced from finishing (grinding) operations and

specimen rupture data (fig. 1). This grinding damage

module, in conjunction with finite element analysis,

can now be used to characterize the material

fracture behavior.

CARES/Life has been in high demand worldwide

although present technology transfer efforts are

entirely focused on U.S.-based organizations.

Success stories can be cited in numerous industrial

sectors including aerospace, automotive, biomedical

(fig. 2), electronic, glass, nuclear, and conventional

.==

O

el

la

el

2

a.

99

" Estimated m = 16.11 ,'

9O ¢re = 937.4 ,' _...._'-

s• _ "°°

8o .........

-- •

50-

2o--.......

5 __unds based on the 90-percent 0 J ,' confidence limit for Weibull modulus m and

1ti" ' Weibull characteristic strength _re.

j_

1 .J"" _ __

.1 .... _-- Kanofsky-Srinivasan 90-percent

confidence bands

I I I t

750 800 850 900 950 1000 1050

Fracture stress

Figure L--Specimen rupture data depicted using common business presentation graphics

packages that significantly enhance utility of CARESLife for design engineers.

52 NASA TM-107300
Body porcelain

Opaque

porcelain

NiCr alloy

Dentin

Pulp

chamber

(void)

Figure 2.--Tilted view of cross section and top of finite

element model of porcelain-fused-to-metal molar crown

analyzed with CARESLife.

power generation. In 1995 Lewis Research Center

in partnership with Philips Display Components

Company (PDCC) won a presUgious R&D 100 Award

from R&D Magazine for the development of CARES/

Life (by Lewis) and its application to the design and

manufacture of an improved television picture tube

(by PDCC) for the U.S. consumer market. The

exceptional technical accomplishments and efforts

of the CARES/Life development team were also

recognized with the Lewis Research Center's 1995

Steven V. Szabo Engineering Excellence Award.

Lewis contacts: Noel N. Nemeth. (216) 433-3215. Lynn M.

Powers. (216) 433-8374. Lesley A. Janosik, (216} 433-5160

Integrated Design Software Predicts

Creep Life of Monolithic Ceramic

Components

Significant improvements in propulsion and power

generation for the next century will require

revolutionary advances in high temperature

materials and structural design. Advanced ceramics

are candidate materials for these elevated

temperature applications. As design protocols

emerge for these material systems, designers must

be aware of several innate features, including the

degrading ability of ceramics to carry a sustained

load. Generally, time-dependent failure in ceramics

occurs because of two different delayed failure

mechanisms, slow crack growth (SCG) and creep

rupture. Slow crack growth initiates at a preexisting

flaw and continues until a critical crack length is

reached, causing catastrophic failure. Creep

rupture, on the other hand, occurs because of bulk

damage in the material in the form of void nucleation

and coalescence that eventually leads to

macrocracks which then propagate to failure.

Successful application of advanced ceramics

depends on the proper characterization of material

behavior and the use of an appropriate design

methodology. The life of a ceramic component can

be predicted with the CARES (Ceramics Analysis

and Reliability Evaluation of Structures) integrated

design programs. CARESCREEP determines the

expected life of a component under creep conditions

and CARES/LIFE predicts the component life due

to fast fracture and subcritical crack growth.

CARES/LIFE was previously developed and has

been used in numerous industrial and government

applications.

The advent of new techniques in ceramic processing

technology has yielded a new class of ceramics that

are highly resistant to creep at high temperatures.

Such desirable properties have generated interest

in using ceramics for turbine engine component

applications where the design lives for such systems

are on the order of I 0 000 to 30 000 hr. These long

life requirements necessitate subjecting the

components to relatively low stresses. The

combination of high temperatures and low stresses

typically places failure for monolithic ceramics in

the creep and creep rupture region of a timetemperature-failure-mechanism map.

NASA TM-107300 53
mm

m

m

m

m

SPIN

DISK

1_ 137mm _ ...............

o.oo

.11

.22

.34

.45

.56

.68

.79

.90

1.00

Figure 1.---Silicon nitride spin disk (AlliedSignal Engines) and analyzed to determine creep life.

(a) Spin disk. (b) Damage map.

An analytical methodology in the form of an

integrated design program (CARES/CREEP) has

been developed for the life prediction of ceramic

structural components subjected to creep-rupture

conditions. This methodology utilizes commercially

available finite element packages and takes into

account the transient state of stress and creep strain

distributions (stress relaxation). The creep life of a

component is discretized into short time steps,

during which the stress distribution is assumed

constant. The damage is calculated for each time

step based on a modified Monkanan-Grant (MMG)

creep-rupture criterion. In a manner similar to

Miner's rule for cyclic fatigue loading, the cumulative

damage is subsequently calculated as time elapses.

Failure is assumed to occur when the normalized

cumulative damage at any point in the component

reaches unity. The corresponding time will be the

creep-rupture life for that component.

CARES/CREEP has been distributed for beta

testing. The first test site is Solar Turbines for their

work in the development of industrial gas turbines

of the 5.5-MW size. The code will also be

disseminated to other engine companies such as

Allison and AlliedSignal Engines. Benchmark

examples from AlliedSignal have been used to

validate the code as shown in figure I.

Lewis contacts: Lynn M. Powers, (216l 433-8374, Leeley A.

Janosik, (216) 433-5160, Noel N. Nemeth, (216) 433-3215

Thermomechanical Analysis of Ceramic

Matrix Composite Components

Ceramic matrix composite material systems for use

in the High-Speed Civil Transport (I-ISCT) combustor

are currently being developed and evaluated. In

addition, model development is also underway to

allow accurate predictive capabilities for these new

materials. Thermomechanical analyses were

provided to support both the material and model

development issues.

Damage mechanics models being developed under

the Enabeling Propulsion Materials (EPM) program

are being included in the finite element code

CSTEM. Various test specimens are being analyzed

to validate the models and to implement them in

the code (fig. I).

At the NASA Lewis Research Center, finite element

analysis support is being provided for EPM-related

testing, and at times is being used to explain

unexpected material responses observed in the

laboratory. A case in point is a high-temperature

tensile endurance test where measured lifetimes

were orders of magnitude lower than expected. A

finite element analysis (FEA) revealed a pesty

temperature problem. At other times, FEA support

has been used before testing to establish run

parameters necessary to achieve desired stress

levels, an example of which is the work being done

for cylindrically shaped CMC combustor liners.

Analysis issues have been primarily focused on the

complicated thermal boundary conditions: backside

impingement cooling, hotside film cooling, and cold

54 NASA TM-107300
{a)

L., I / I l

L,, t , ,,,,l ,, I +i_: ....... +i I i I ,:

(b)

Figure l.--Thermomechanically induced axial stresses and temperature profile for EPM tensile specimen. (a) Heated

gage section and cold grips. (b) High-stress regions outside gage section.

contact regions. Initial analysis showed that the

stress levels were far below the desired test levels.

By modifying the impingement cooling, it is hoped

that stresses can be raised locally to the desired

levels.

Support is also being provided to NASA's Marshall

Space Flight Center for designing and analyzing a

CMC nozzle being considered for future use on the

space shuttle main engine. Preliminary work has

been done to establish baseline material

requirements.

Lew/s contact: David J. Thomas (216_ 433-5664

Modeling the Effective Elastic Behavior

of Transversely Cracked Laminated

Composite

During loading, composite materials typically

experience damage in the form of transverse matrix

cracking (i.e., through-the-thickness cracks that

run parallel to the fibers and are caused by the

presence of the in-plane transverse stress).This form

of damage may be noncatastrophic; it does not

necessarily result in the immediate failure of the

laminate. However, transverse matrix cracking does

adversely affect the mechanical response of the

material and is often a precursor to additional

modes of failure. As such, a significant body of

research has been devoted to modeling the effects

of progressive transverse cracking. Continuum

damage models as well as many shear-lag

approaches have been proposed for analyzing crossply laminates. Shear-lag models implicitly assume

that transverse matrix cracks occur at regularly

spaced intervals. This assumption allows the

solution of the stress state for the damaged laminate

to be reduced to the solution for a characteristic

volume element bounded by two transverse cracks

and having a length equal to the average crack

spacing. Thus, the effective elastic constants can

be calculated. In this work, the validity of this

assumption was investigated qualitatively by

examining the probability density function for

transverse crack location. The work proceeded to

extend the method of Lee and Daniel for determining

the shear-lag parameters to a general symmetric

multilayer system. The elasticity problem for the

region of the laminate between two parallel matrix

cracks having an arbitrary off-axls orientation was

established from equilibrium considerations in

terms of the average (through-the-thickness)

stresses and was solved using the newly developed

generalized shear-lag (GSL) relation and the

appropriate boundary conditions. Modeling of the

effective elastic properties that result from the

250xl 06

_= 200

Q.

_<150

i1=

_100

==

5o

0 i I I I , i i

0 0.001 0.002 0.003 0.004 0.005 0.006

Strain, (Ex)applied

Figure 1.--Predicted average through-the-thickness

laminate stress erx versus applied axial strain (f-x)applied

for [OI301601-301-60]s laminate using GSL model.

NASA TM-! 07300 55
2O

E

¢J

r-

"O

O

o

15 60°---_

\ _ ....... ,--'°

j..'" \

_30o -'-x

\

0 ........ _ ....

0 0.001 0.002 0.003 0.004 0.005 0.006

Strain, (ex)applied

Figure 2.--Predicted evolution of crack densities in

[0/30/60/-30/-60]s laminate as function of applied

axial strain (cx)applie d.

damaged layer was detailed. The analysis method

was benchmarked against published results for a

cross-ply laminate, and the advanced modeling

capabflitles of the model were demonstrated for a

[0/30/60/-30/-60] s laminate (figs. I and 2).

Lewis contact: David J. Thomas, (216) 433-5664

Capability of Thermographic Imaging

Defined for Detection of Defects in HighTemperature Composite Materials

The ability of a single-sided thermograhic imaging

technique (fig. 1) to detect fiat-bottom-hole defects

of various diameters and depths was evaluated in

four composite systems (two types of ceramic matrix

composites (CMC's), one metal matrix composite

(MMC), and one polymer matrix composite (PMC))

of interest as high-temperature structural materials.

The holes ranged from l to 13 mm in diameter and

from 0.1 to 2.5 mm in depth in samples

approximately 2 to 3 mm thick. The thermographic

imaging system utilized a scanning mirror optical

system and an infrared (IR) focusing lens in

conjunction with a Mercury-Cadmium-Telluride

infrared detector element to obtain high-resolution

infrared images. High-intensity flash lamps located

on the same side as the infrared camera were used

to heat the samples. After heating, up to 30 images

were sequentially acquired at 70- to 150-msec

intervals.

Limits of detectability based on the depth and

diameter of the fiat-bottom holes were observed for

each composite material. For the SiC/CAS CMC

A-- Xenon flash lamps with

i_ /I aluminum parabolic reflectors

/ I //1/-- Mounting plate

/I (to the right)

J t.2/-<.

Power _ Sample

-- supply I

monitor

• Defect boundaries inside material slow down diffusion of thermal front as it propagates,

resulting in different temperatures near defects as compared with good areas.

Figure 1.--F_xperimental therrnography technique description.

56 NASA TM-107300
12 1

o

o

o

O Detected

Barely detected

[] Not detected

O O

O

o o o o

o

o o _ o

Empirical rule detectability

threshold region -_

o o 0 0 0 0 _

oo

0.8

0

0.0 0.4 1.2 1.6 2.0 2.4

Depth, mm

Figure 2.--Thermography detectability for SiC/CAS CMC

results. (Defects _<1.8mm deep with diam _>1.6mm will

be detected.)

samples, defects of depth less than or equal to

1.8 mm having diameters greater than or equal to

1.6 mm probably will be detected with the

thermography methodology used in this study

(fig. 2). For the SiC/SiC CMC samples, defects of

depth less than or equal to 1.8 mm having diameters

greater than or equal to 2.6 mm probably will be

detected. For the SiC/Ti MMC samples, defects of

depth less than or equal to 1.6 mm having diameters

greater than or equal to 3.2 mm probably will be

detected. For the graphite/polyimide PMC samples,

defects approximately 3 to 12 mm in diameter and

less than or equal to 1.8 mm in depth probably will

be detected. Depth appears to be the limiting

variable with regard to detectability in the PMC

system. The thermography imaging results were

consistent with the empirical rule which states that

defects of diameter less than or equal to 0.5 to

1.0 x d _where d Is the depth below the surface)

probably will not be detected.

The thermographic images were compared with

ultrasonic and conventional film radiographic

images. Radiographic images clearly revealed all

fiat-bottom holes and provided the highest quality

images of the three imaging methods. The ultrasonic

imaging results were material dependent. For the

SJC/CAS CMC material, thermographic imaging

revealed defects as clearly as or more clearly than

the ultrasonic imaging did. For the SiC/SiC CMC

material, pulse-echo ultrasonic imaging had

difficulty clearly revealing all defects whereas

through-transmission ultrasonic imaging enabled

the visualization of all defects; thermographic

images revealed the shallowest and intermediatedepth defects but could not reveal the deepest

defects (those at depths > ~ 2.0 mm below the

surface). For the SiC/Ti MMC material, ultrasonic

imaging revealed all defects while thermographic

images did not reveal the smallest, deepest defects

(those I to 3 mm in diam at depths of 1.6 to 1.7 mm

below the surface). For the graphite/polyimide PMC

material, ultrasonic images barely revealed

indications of the deepest defects (>1.8 mm below

the surface, ~ 1 to 12 mm in diameter) whereas

thermographic images did not reveal any of the

deepest defects.

Lewis contact: Don J. Roth (216) 433-6017

Crack Extension Measurement In Brittle

Materials

Ceramics are of great interest as structural

components in engines because of their low density

and high-temperature load-bearing capabilities.

However, ceramics and glasses are brittle and

exhibit time-dependent crack growth. A simple

technique was developed to monitor crack extension

in brittle materials such as ceramics and glasses.

A variety of specimens with sharp cracks have been

used to monitor crack extension in glasses and

ceramics; however, many of them are complicated

or use grooves to guide the cracks. Measurements

using the simple beam specimen with a straightthrough precrack were improved by placing a straln

gage on the backface after precracking, as shown

in figure 1. Finite element analysis was used to

Figure 1.---Simple beam specimen.

NASA TM- 107300 57
1.0

0.8

ee=

_e 0.6

J<

O

° 0.4

_ 0.2

o

Z

0.0

I I I [ I

°_°_o\

0

\

0

I I I I I

-25 -20 -15 -10 -5

Normalized backface strain

\

O

\

o

O

0

Figure 2._Backface strain as function of crack length.

5O I I I I I

40

_ 3O

"D

0

_ 20

10

0 I I J I I

0 100 200 300 400 500 600

Backface microstrain

Figure 4.--Experimental load as function of backface strain.

1.0 I I ! I I

. 0.8

O)

t-

_e 0.6

J¢

O

e0

o 0.4

"O

.N

_0.2

0

z

0.0

--,30

O_

o

\

0

\

o

\

0

I , I I I J ,

-25 -20 -15 -10 -5

Normalized backface strain

0

Figure 2.--Backface strain as function of crack length.

calibrate the backface strain as a function of crack

length (fig. 2). The finite element modeling (FEM)

results were verified by conducting an experimental

calibration in which saw-notched specimens were

used to simulate cracks. The results are shown in

figure 3, and experimental load versus backface

strain is shown in figure 4.

The results allow simple, inexpensive monitoring

of stable crack extension, crack growth resistance,

and fatigue crack growth in brittle materials.

Customers include Oak Ridge National Laboratory,

which funded the research, and members of the

American Society for Testing Materials Committee

C28 fracture task group, such as the National

Institute of Standards and Technology, AUiedSignal,

and Allison Gas Turbine.

Lewis contacts: Jon Salem, (216) 433-3313, Bung Choi,

(216} 433-8366, Lou/s Gho_n, (216) 433-3249, and Ralph

]Pawiik, (216) 433-8563

Experimental Verification of Crack

Growth Models for Ceramics

Because of their low density and high-temperature

load-bearing capabilities, ceramics are of great

interest as structural components in engines.

However, ceramics and glasses are brittle and

exhibit tlme-dependent crack growth.

A variety of models for room- and elevatedtemperature crack growth exist and have been

incorporated in the CARES (Ceramics Analysis and

Reliability Evaluation of Structures)/Lffe design

code. The code allows designers to analyze and

58 NASA TM-107300
Figure 1.--Failed biaxial flexure specimen.

predict the probability of failure of structural

components made from brittle materials over their

lifetime. Experimental verification of the CARES/

Life crack growth code was undertaken by testing

silicon nitride, a structural ceramic of industrial

interest, at 1300 °C in air. Four-point flexure

loading was used to generate simple specimen

(design) data, and biaxlal flexure of plates was used

to simulate a larger test volume with a multiaxial

stress state.

A failed biaxial flexure specimen is shown in figure 1 and the test results are shown in figure 2.

550

450

.c"

e-

.=

O_

350

Four-point loading

Disk

.... 2_,- _ " "-- Pnncipa| of

"_" \\ independent action

\_ _-C = 0.82 Griffith crack

\

_-- C = 0.82 Semicircular crack

250 = = ' =

10-2 10-1 100 101 102 103

Stressing rate, MPa/s

Figure 2.---Silicon nitride test in air at 1300 °C.

Agreement was good for most of the failure criteria

used in the predictions, indicating that the code

can predict failure probabilities of multiaxial stress

states when the governing mechanism is slow crack

growth.

Customers included ORNL, which funded the

research, and other participants in the Ceramic

Technology Program: AlliedSignal, Allison Gas

Turbine, and Solar Turbine.

Lewis contacts: Jon Salem. (216] 433-3313. Sung Cho|,

(216) 433-8366. Ralph Pawlik, (216) 433-8563

NASA TM-107300 59
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Minnetyan, L.; and Gotsis, P.K.: Progressive Fracture in

Adhesively Bonded Concentric Cylinders. Proceedings of the

40th International SAMPE Symposium and Exhibition, vol. 40,

book I, 1995, pp. 849--860.

Mital, S.K.; and Chamis, C,C.: Microfracture in Toughened

Epoxy Composites. Proceedings of the 40th international SAMPE

Symposium and Exhibition, 1995, pp. 1330-40.

Mital, S,K.; Chamis, C.C.; and Gotsis, P.K.: Micro-Fracture

in High-Temperature Metal-Matrix Laminates. Comp. Sci. T.,

vol. 50, 1994, pp. 59-70.

Mital. S.K.; Murthy, P.L.N.; and Chamis, C.C.: Micromechanics

for Ceramic Matrix Composites Via Fiber Substructul-ing. J.

Compos. Mater., vol. 29, no. 5, 1995, pp. 614-633.

Mital, S.K.; Murthy, P.L.N.; and Chamis, C.C.: Micmmechanics

for Woven Composites. HITEMP Review 1995, vol. Ill, NASA

CP-10178, pp. 52-1 to 52-12.

Mital, S.K.; Murthy, P.L.N.; and Chamis, C.C.: Simplified

Micromechanics of Two-Dimensional Plain Woven Fabric Composites. ASME International Congress, 1995, pp. 829-840.

Mital, S.K.; Murthy, P.L.N.; and Chamis, C.C.: Simulation of

Ceramic Matrix Composites Behavior Including Progressive Fracture and Load Redistribution. AIAA Paper 95-1215--CP, 1995.

Mittelman, A., et al.: Acousto-Ultrasonic Characterization of

Physical Properties of Human Bones. Acousto-Ultrasonics:

Theory and Application, J.C. Duke, ed., Plenum Press, New York,

NY, 1988, pp. 305-309.

Morales, W., et al.: A New Antiwear Additive/Surface Pretreatment for PFPE Liquid Lubricants. NASATM-107038, 1995.

Murray, S.F.; and Heshmat, H.: Accelerated Testing of Space

Mechanisms, NASA CR-198437, 1995.

Murthy, P.L.N.; and Chamis, C.C: Integrated Composite Analyzer (ICAN}: Users and Programmers Manual. NASA TP2515,1986.

Nemeth, N.N., et al.: Designing Ceramic Components for I)tLrability.

Am. Ceram. Soc. Bull., vol. 72, no. 12, Dec. 1993. pp. 59--66.

Nemeth, N.N., et al.: Durability Evaluation of Ceramic Components Using CARES/LIfe. ASME Paper 94-GT-362, 1994.

Nemeth, N.N., et al.: Lifetime Reliability Evaluation of Structural Ceramic Parts With the CARESLIfe Computer Program.

AIAA paper 93-1497-CP, 1993, pp. 1634-1646.

Nemeth, N.N., et al.: Time-Dependent Reliability Analysis of

Monolithic Ceramic Components Using the CARES/Life Integrated Design Program. ASTM Special Technical Publication,

Proceedings of the Symposium on Life Prediction Methodologies

and Data for Ceramic Materials, _ STP- 120 I, C.IL Brinkman

and S.F. Duffy, eds., American Society for Testing and Materials. Philadelphia, PA, 1994, pp. 390-408.

Noebe, R.D.; Lerch, B.A.; and Rao, K.B.S.: Assessment of

Microalloying Effects on the High Temperature Fatigue Behavior of NiAI. High-Temperature Ordered Intermetalllc Alloys VI

Symposium, part 1, Materials Research Society, Pittsburgh, PA,

1995, pp. 291-296.

Pai, S.S.; and Chamis, C.C.: Probabilistic Assessment of

Combustor Liner Design. ASME International Congress, 1995.

Patnaik, S., et al.: Completed Beltrami-Mlchell Formulation for

Analyzing Mixed Boundary Value Problems in Elasticity. AIAA

Journal, vol. 34, no. 1, Jan. 1996, pp. 143-148.

Patnaik, S.N.; Guptill, J.D.; and Berke, L.: Merits and

Limitations of Optimality Criteria Method for Structural

Optimization. Int. J. Numer. Methods Eng., vol. 38, no. 18, 1995,

pp. 3087-3120.

Pindera, M.-J.; Aboudi, J.; and Arnold, S.M.: Limitations of the

Uncoupled, RVE-Based Micromechanical Approach in the Analysis of Functionally Graded Composites. Mech. Mater., vol. 20,

no. 1, 1995, pp. 77-94.

Portlock, L.E.; Schetky, L. McD; and Steinetz. B.M.: Shape

Memory Alloy Adaptive Control of Gas Turbine Engine Blade

Clearance. AIAA Paper 95-2762, 1995.

Powers, L.M., et al.: Lifetime Reliabfllty Evaluation of Structural

Ceramic Parts With the CARES/Life Computer Program.

Collection of Technical Papers from the 34th Structures, Structural Dynamics and Materials Conference. AIAA, Washington,

DC, part. 3, 1993, pp. 1634-1646.

Powers, L.M.; Jadaan, O.M.; and Gyekenyesi, J.P.: Creep Life of

Ceramic Components Using a Finite Element Based Integrated

Design Program (CARESCREEP). ASME Paper 96-GT-369,

1996.

Rao, K.B.S.; Castelli, M.G.; and EUis, J.R.: On the Low Cycle

Fatigue Deformation of Haynes 188 Superalloy in the Dynamic

Strain Aging Regime. Scrip. Metal. et Materialia, vol. 33, no. 6,

1995, pp. 1005-1012.

NASA TM-107300 63
Rao,K.B.S., etal.:TemperatureDpendent CyclicDeformation

Mechanisms inHaynes188Superalloy. NASATM-107016, 1995.

Rao,K.B.S.;Lereh,B.A.;andNoebe, R.D.:EffectofProcessing

RouteonStrainControlledLowCycle Fatigue Behavior of Polycrystalline NiAI. Proceedings of the Fatigue and Fracture of

Ordered IntermetaUic Materials II Symposium, W.O. SoboyeJo,

T.S. Srivatsan, and R.O. Ritchie. eds., TMS, Warrendale, PA,

1996, pp. 245--271.

Rao, K,B.S., et a1.: Temperature and Strain-Rate Effects on

Low-Cycle Fatigue Behavior of Alloy 80OH. Metall T-A, vol. 27A,

no. 2, 1995, pp. 255--267.

Reddy, T.S.R.: User's Guide for ECAP2D: An Euler Unsteady

Aerodynamic and Aeroelastic Analysis Program for Two Dimensional Oscillating Cascades. NASA CR-189146, 1995.

Roberts, G.D, et al.: Adhesively Bonded Strain Gages for Extended Use on PMR-15 at 316 °C. HITEMP Review 1995, vol. I,

NASA CP-10178, 1995.

Rokhlin, S.I.; Chu, Y.C.; and Baaklini, G.Y.: Assessment of

Damage in Ceramics and Ceramic Matrix Composites Using

Ultrasonic Techniques. J. Eng. Gas Turbines Power, vol. 117,

no. 3, 1995, pp. 417-423.

Roth, D.J.; Bodis, J.R., and Bishop, C.: Thermographic Imaging

for High-Temperature Composite Materials: A Defect Detection

Study. NASA TM-106950, 1995.

Roth, D.J., et al.: An NDE Approach for Characterizing Quality

Problems in Polymer Matrix Composites. NASA TM-I06807,

1995.

Salem, Ja_, et al.: Reliability Analysis of Unlaxially Ground Brittle

Materials. 95-GT-31, Proceedings of ASME International

Congress, 1995.

Saxavanos, D-_.; and Hopkins, D.A.: Local Stresses in Composite Plates with Piezoelectric Actuators and Sensors Using a

Discrete-Layer Theory. Proceedings of ASME International

Congress, 1995.

Shah, A.R.; and Chamis, C.C,: Cyclic Load Effects on Long Term

Behavior of Polymer Matrix Composites. NASATM-107007, 1995.

Shah, A.R.; Murthy, P.L.N.; and Chamis, C.C.: Effect of Cyclic

Thermo-Mechanical Loads on Fatigue Reliability in Polymer

Matrix Composites. NASA TM- 10709 I, 1996.

Shapiro, W., et al.: Space Mechanisms Lessons Learned Study,

Volume I: Summary. NASATM-107046, 1995.

Shapiro, W,, et al.: Space Mechanisms Lessons Learned Study.

Volume II: Literature Review, NASA TM-I07047, 1995.

Shiao, M.C.; Singhal, S.N.; and Chamis, C.C.: Network Based

Parallelism of Probabilistic Composite Structural Analysis. AIAA

Paper 95-1259-CP, 1995, pp. 889--899.

Shiao, M.C.; Singhal, S.N.; and Chamis, C.C.: Reliability, Risk

and Cost Trade-Offs for Composite Designs. Proceedings of the

40th International SAMPE Symposium and Exhibition, 1995,

pp. 881--891.

Srivastava, R.; and Reddy, T.S.R.: Aeroelastic Analysis of Ducted

Rotors. ASME 1995 International Congress, 1995.

Steinetz, 13. M.: Aerospace Seal Requirements. Proceedings of

the Seals Code Development Workshop, 1995, pp. 41-53.

Steinetz. B.M. and Kren, L.A.: Resilient Braided Rope Seal.

Lew 16, 231-1, Invention Disclosure, January 1995.

SutJahjo, E.; and Chamis, C.C.: Multldisciplinary _nite Elements

for Coupled Analysis of Fluid Mechanics, Heat Transfer, and

Solid Mechanics. AIAA Paper 95--1309-CP, 1995, pp, 1318--1327.

Verrilli, M.J., et al,: Standardization Activities in TMF Test

Methodologies. Presented at the Structures and Materials Panel,

AGARD CP-569, 1996.

Weaver, R.L.: Diffuse Field Decay Rates For Material CharacterizaUon. Solid Mechanics Research for Quantitative NonDestructive Evaluation, J.D. Aehenbaeh and Y. RaJapaskte, eds.,

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Zaretsky, E.V.; August, R.; and Coe, H.H.: Effect of Hoop Stress

on Ball Bearing Life Prediction. NASATM-106874, 1995.

Zaretsky, E.V., Poplawski, J.V.; and Peters, S.M.: Comparison

of Life Theories for Roliing-Element Bearings. NASA TM-I06585,

1995.

Zaretsky, E.V.: STLE Life Factors for Rolling Bearings. STLE

SP-34, 1992.

64 NASA TM-107300

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July 1997

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Structures Division 1995 Annual Report

6. AUTHOR(S)

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National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

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National Aeronautics and Space Administration

Washington, D.C. 20546-0001

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13. ABSTRACT (Maximum 200 words)

The NASA Lewis Research Center Structures Division is an international leader and pioneer in developing new

structural analysis, life prediction, and failure analysis related to rotating machinery and more specifically to hot

section components in air-breathing aircraft engines and spacecraft propulsion systems. The research consists of both

deterministic and probabilistic methodology. Studies include, but are not limited to, high-cycle and low-cycle fatigue

as well as material creep. Studies of structural failure are at both the micro- and macrolevels. Nondestructive evaluation methods related to structural reliability are developed, applied, and evaluated. Materials from which structural

components are made, studied, and tested are monolithics and metal-matrix, polymer-matrix, and ceramic-matrix

composites. Aeroelastic models are developed and used to determine the cyclic loading and life of fan and turbine

blades. Life models are developed and tested for bearings, seals, and other mechanical components, such as magnetic

suspensions. Results of these studies are published in NASA technical papers and reference publication as well as in

technical society journal articles. The results of the work of the Structures Division and the bibliography of its

publications for calendar year 1995 are presented.

14. SUBJECT TERMS

Structural mechanics; Structural fatigue; Structural dynamics; Structural integrity

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