NASA/TM--1999-206966

Structures and Acoustics Division

1996 Annual Report

Cynthia S. Acquaviva

Glenn Research Center, Cleveland, Ohio

National Aeronautics and

Space Administration

Glenn Research Center

November 1999

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Introduction

The 1996 Annual Report of the Structures Division

reflects the work areas performed by the Division

staff during the 1996 calendar year. 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.

The Structures and Acoustics Division comprises a

staff of approximately 82 engineers and scientists

plus administrative and support personnel. The

work areas comprise composite mechanics, fatigue,

fracture, and dynamics with emphasis on life and

reliability, as well as mechanical system

technologies and aeroacoustics. 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.

Chief, Structures and Acoustics Division

NASA/TM-- 1999-206966 i
Contents

Structural Mechanics

Design Sensitivity Approximations Improve Performance of Optimization Algorithms ............. 1

Performance of Optimization Methods Improved Through Selected Algorithms .................. 1

Finite-Element Formulation for Smart Structures in Thermal Environments ................... 2

Micromechanics for Particle-Reinforced Composites ...................................... 3

Micromechanics for Two-Dimensional Woven Composites .................................. 4

Fatigue and Fracture

Experimental Techniques Verified for Determination of Yield and Flow Surfaces ................. 6

Creep-Fatigue Behavior of Cast IN 939 Superalloy ....................................... 7

Fatigue Crack Growth of Single-Crystal Superalloys in Hydrogen ............................ 8

Reliability Modeling of Brittle, Anisotropic Solids ........................................ 9

Fatigue Behavior and Life Prediction of a T650-35/PMR-15 (0/90) Weave Under

Isothermal Conditions ......................................................... 10

A General Representation Capturing Reversible and Irrevesible Hereditary Behavior of

a Titanium Alloy ............................................................. 12

Damage Mechanisms and Mechanical Property Degradation in Titanium Matrix Composites

Subjected to TMF Loadings ..................................................... 14

Structural Integrity

CARESLife Ceramics Durability Evaluation Software for Finite-Element Analysis Programs ...... 16

Creep Life Prediction of Ceramic Components Subjected to Transient Tensile and Compressive

Stress States ................................................................ 18

Constitutive Theory Developed for Monolithic Ceramic Materials ............................ 19

Single Transducer Ultrasonic Imaging Method That Eliminates the Effect of Thickness in

Imaging of High-Temperature Structural Materials ................................... 20

Noncontact Measurement of Acousto-Ultrasonics for Ceramic Matrix Composites ............... 21

Structural Dynamics

Dynamic Spin Rig Magnetic Suspension .............................................. 23

High-Temperature Magnetic Bearings for Gas Turbine Engines ............................. 25

Neural Network Control of a Magnetically Suspended Rotor System ......................... 26

Damping Experiment of Spinning Component Plates With Embedded Viscoelastic Material ........ 28

Improved Dual Clearance Squeeze Film Damper for High Loads ............................ 29

User's Guide for MSAP2D, an Euler Flutter and Forced-Response Analysis Code for Multistage

Turbomachinery ............................................................. 30

User's Manual for ASTROP2, Version 2.0. a Program for Aeroelastic Stability Analysis

of Multiblade Structures ....................................................... 31

Using an Euler Aeroelastic Solver to Analyze Gust and Structural Response of a

Two-Dimensional Cascade ...................................................... 31

ii NASA/TM-- 1999-206966
FlutterAnalysisof DuctedRotors................................................... 33

TURBO-AE:An AeroelasticCodeforPropulsionApplications .............................. 35

Usingan IntegratedFiber-OpticLightProbeto MeasureStaticDeflectionsin

RotatingTurbomachinery ...................................................... 36

Feasibilityof UsingNeuralNetworkModelsforAcceleratedTestingof MechanicalSystems........ 36

Acoustics

Scale-Model Acoustic Liner and Low-Tip-Speed Turbofan Testing ...........................

Acoustic Barrier Facilitates Inlet Noise Measurements for Aft-Dominated Fans .................

Jet Noise Prediction for Internal Exhaust Mixer Nozzles ..................................

38

39

41

Mechanical Components

Gear Crack Propagation Studies .................................................... 43

Validated Method to Optimize Load Sharing of Split-Torque Transmissions ................... 44

Evaluation of Face Gears for Aerospace Drive System Applications .......................... 45

Using Hob Offset to Balance Strength of Spur Gears ..................................... 46

Brush Seals Tested for Joint Technology Advanced Gas Generator II Engine ................... 48

Diagnostic Parameters Detect Gear-Tooth Fatigue Crack ................................. 49

Gas Journal Bearing Stability Versus Amplitude Ratio ................................... 51

NASA Selects 1996 Government Invention of the Year .................................... 52

Bibliography .......................................................... 54

NASAfrM--1999-206966 iii

Structural Mechanics

Design Sensitivity Approximations

Improve Performance of Optimization

Algorithms

Computational efficiency in design optimization has

been enhanced through sensitivity approximations.

The sensitivities of behavior constraints have been

obtained with a small number of numerical

calculations. Four algorithms were used to quantify

the performance of the approximations: the

sequence of unconstrained minimization technique

(SUMT); the method of feasible directions (mFD);

the sequence of quadratic programming (SQP); and

the sequence of linear programming (SLP). All four

optimization algorithms, through approximations,

converged to correct optimum solutions. The use of

approximations reduced by one-third the CPU time

that would otherwise be required to solve the

problem with explicit closed-form sensitivities.

The convergence characteristics for a 25-variable,

180-behavior constraints problem determined with

explicit and approximate gradients by SUMT and SQP

optimizers are shown in figures 1 and 2, respectively.

Both algorithms converged to an optimum weight of

308.7 lb. The rates of convergence with and without

approximations were in agreement. The problem,

however, required somewhat fewer reanalysis cycles

when approximate sensitivities were used.

Glenn contacts: Dale Hopkins, (216) 433-3260; and Surya

Patnaik, (216) 433-5916

450 --

360 _

308.73 Ib)

270

03

180

9O

Explicit gradients (73 cycles)

Approximate gradients (70 cycles)

0 25 50 75

Number of reanalysis cycles

Figure l.--Convergence of weight for ring as determined by using

SUMT.

450 --

360 --

__o 270 (Final weight,

308.729 Ib)

180 --

Explicit gradients (63 cycles)

Approximate gradients (50 cycles)

90 -

I I

0 25 50 75

Number of reanalysis cycles

Figure 2.--Convergence of weig/tt for ring as determined by using

IMSL-SQP.

Performance of Optimization Methods

Improved Through Selected Algorithms

Nonlinear programming algorithms play an

important role in structural design optimization. The

performance of optimization methods depends on

the choice of algorithms. Reliable and efficient

algorithms can be found by comparing the

performance of different algorithms for the same

set of design problems. A comparative study--for a

set of about 50 structural design problems solved

by using a dozen different algorithms---could not

identify a single winner that can be called the most

reliable and efficient. Overall, three algorithms

scored high marks for large design problems: two

sequential programming algorithms (SQP1 and

SQP2) and a sequential unconstrained minimization

technique (SUMT). The CPU times required to solve

large problems were comparable for the three

algorithms. Most algorithms, however, could solve

at least one-third of the 50 examples.

Figure 1 shows the performance of 6 optimization

algorithms in solving a set of 10 large problems.

Three algorithms compare well, although with some

variation.

Glenn contacts: Dale Hopkins, (216) 433-3260; and Surya

Patnalk, (216) 433-5916

NASA/TM--1999-206966 1
Problem set

P8a

[] P8b

[_] P8c

[] P8d

[] P24

[] P25

P26

[] P27

[] P28

P29

SUMT SQP2 SQP1 NPSOL RG OC

Figure l.--Performance of different optimizers in solving large problems.

Finite-Element Formulation for Smart

Structures in Thermal Environments

Aeropropulsion systems with piezoelectric

composite components could offer great potential

for improving the performance of aircraft engines

through various vibration suppression, noise

control, and thermal stability applications. To

facilitate the characterization of such smart

structures, we have extended finite-element

methods to take into account the thermal effects

so that the coupled mechanical, electrical, and

thermal behavior of these materials can be analyzed.

Thermal effects typically affect the response of smart

piezoelectric laminates by three distinct physical

mechanisms: (1) thermal strains induced by a

coefficient-of-thermal-expansion mismatch between

the various composite and piezoelectric layers, (2)

pyroelectric effects on the electric displacement of

the piezoelectric material, and (3) temperaturedependent changes in both the composite and

piezoelectric material properties.

We have incorporated the first two thermal effects

mechanisms into a layerwlse laminate theory to

more accurately model both the active and sensory

responses of smart structures. Corresponding finiteelement formulations were derived and implemented

for both beam and plate elements (Lee and

Saravanos, 1996a, b, and c).

The capabilities of the finite-element formulation

were demonstrated on a simply supported [0/+45]s

graphite/epoxy plate with piezoceramlc patches

uniformly attached to both the top and bottom

surfaces. Figure 1 illustrates the geometry and

finite-element mesh of the plate. A linear thermal

gradient (50 °C on the top surface and -50 °C on

the bottom surface) was applied through the

thickness of the plate. With no voltage applied, the

bending deflection to both surfaces of the

piezoceramic patches was as shown in figure 2(a).

By increasing the active voltages applied to the

piezoceramic patches, we gradually eliminated the

thermally induced deflection of the plate.

Figure 2(b) shows the reduction in the deflection

achieved by applying an active voltage of 70 V.

Glenn contacts: Ho-Jun Lee, (216) 433-3316; Dlmltris

Saravanos, (216) 962-3211; and Dale Hopkins, (216)

433-3260

2 NASA/TM--1999-206966
d ........

4 37.2 cm

(a)

Piezoceramic --_

\

_:_ :_;_ °_F_II  ̧

.................. _q_.:i_q:

I

2c . 

r//-- 0.13 mm 16¢ml _\\ 0.75 mm -_ L

..............

.... '3-

(b) _ Graphite epoxy

Figure l.--Model of graphite/epoxy plate with at/ached

piezoceramic patc/tes. (a) Top view. (b) Front vioa"

1.80

1.40 "*"

E 1.00 "_

#,

0.60*

0.20 _ "_'t_

-0.20 ._:_-.<

0.20 0.40 _"_

(a)

1.60

1.20

0.80

"_ 0.40

0.00

0.60

x/a

"A

i I

1,,,/

0.80

1.0

0.00

.20

0

ylb

/- 0.80

" 1.00

-0.40

(b)

v -.

0.20 0.40 0.60 _.

x/a 0.80

1.00

0"00

.20

0

y/b

" 1.00

Figure 2.--Thermal shape control of a [0/_+45]splate subjected

to a 100 °C gradient. (Deflection of plate, W: current position

along the width axis, y; width of plate, b; current position along

length axis, x; length of plate, a.) (a) No vohage applied. (b) 70 V

applied on upper and lower piezoceramic patches.

Micromechanics for Particle-Reinforced

Composites

Recently, there has been a growing interest in the

use of particle-reinforced composites in many widely

varying applications. Concrete can be thought of

as one of the oldest material in this category.

Because of the high costs and technical difficulties

associated with the fabrication of continuous-fiberreinforced composites, their use in certain

applications is somewhat limited. For aerospace

applications, particle-reinforced composites can be

thought of as a viable alternative, especially where

shock and impact properties are important. In fact,

particle-reinforced composites (typically an

aluminum matrix reinforced with SiC or TiC

particles) have already shown great potential for

many automotive applications, such as disk-brake

rotors, connecting rods, cylinder liners, and other

high-temperature applications. In situ reinforced

ceramics, which have higher fracture toughness

than monoliths and are easier to process in nearnet shapes, can also be modeled as particlereinforced composites.

For over two decades, NASA Glenn Research Center

has been developing computer codes based on

simplified micromechanics equations to simulate

the behavior of continuous-fiber-reinforced

composites, primarily for aerospace use. As a part

of this research, several computer codes to simulate

the behavior of polymeric, metallic/intermetallic,

or ceramic matrix composites have been developed.

By following a similar approach, we have developed

a set of micromechanics equations for the analysis

of particulate-reinforced composites. These

equations were programmed as a separate module

in an inhouse-developed code. They can predict the

overall composite properties (mechanical and

thermal) on the basis of volume fraction and the

constituent's thermal and mechanical properties.

The microstress equations were also developed to

decompose the applied stress on the overall

composite in terms of the constituent stresses. A

limited verification study compared the predicted

results with finite- and boundary-element methods

and with some bounding techniques found in the

literature for the analysis of particle-reinforced

composites. The agreement was excellent, and the

overall results demonstrate that the micromechanics equations for particle-reinforced

composites can indeed model their behavior. Figure 1 compares the experimentally measured

normal modulus for a particle-reinforced metal

NASA/rM-- 1999-206966 3
F

/

Micromechanical technique /

• Expermental data //

200 - - Hashin-Shtrikman bounds / / ,./

- Third-order bounds / ..

O

/

/ ./_ _...-J_'_-_ ..._._-

o ! I I I I

0.00 0.10 0.20 0.30 0.40 0.50

Particle volume fraction

Figure 1.--Modulus of AL/SiCp composite. (The subscript p refers to tile reinJbrcing particle.)

I

0.60

matrix composite with predictions made by various

methods.

Glenn contacts: 6ubodh Mltal, (216) 433-3261;

Robert Goldberg, (216) 433-3330; and Pappu Murthy, (216)

433-3332

Micromechanics for Two-Dimensional

Woven Composites

Woven fabric composites are constructed by weaving

two or more fiber tows into each other to form a

layer. These layers are then impregnated with a

coating and some kind of resin or matrix material,

stacked in a desired orientation, and cured to obtain

a composite laminate of specified thickness. The

interlacing of fiber bundles has several advantages,

such as increasing the intralaminar and

interlaminar strength, thereby providing greater

damage tolerance and making possible the

production of near-net-shape structural components. However, these advantages come at the

expense of some loss in the inplane stiffness and

strength, which depend on the weave architecture.

There is certainly a need for sound engineering data

as well as verified and efficient analytical and design

methodologies to evaluate different design

parameters. To accurately model such composites,

one must use design methodologies that account

for processing parameters and microstructural and

geometrical features

We have developed a micromechanics-based

technique that can be used to analyze any kind of

two-dimensional weave architecture. It works in

conjunction with the inhouse-developed codes ICAN

(Integrated Composite Analyzer) and CEMCAN

(Ceramic Matrix Composites Analyzer). The intent

of this work was to develop a technique that can

use the capabilities of already existing codes for

laminated composite behavior, yet account for fiber

waviness, fiber end distributions through the

thickness, and constituent morphology that

accurately represents the micromechanical behavior

of woven composites. Micromechanics techniques

have certain inherent advantages in terms of

computational efficiency over other techniques that

rely on numerical analyses.

Two material systems, a generic graphite/epoxy

plain-weave composite and a SiC/SiC plain-weave

ceramic matrix composite were analyzed with this

technique. Results were compared with some

available experimental data from three-dimensional

finite-element analyses. There was a fair agreement

between the two. This same technique Is currently

being used to develop a computer code for the

4 NASA/TM-- 1999-206966
analysis of plain-weave composites. Figures 1 and

2 show the structure of a plain-weave fabric and

inplane modulus of a SiC/SiC composite as a

function of porosity.

Glenn contacts: Subodh Mital, (216) 433-3261; and

Pappu Murthy, (216) 433-3332

-- Unit structure

-- Cross section

of plain wave

Figure 1.-- Structure of a plain-weave fabric.

400

300

n

_200

"0

o

100

I _ Micromechanics technique

A Experimental data

o I I I I I

0 5 10 15 20 25

Porosity, percent

Figure Z--In-plane modulus as a function of porosit3_ for SiC/SiC

composite with 0.4fiber volume ratio.

NASA/TM--1999-206966 5
Fati ue and Fracture

Experimental Techniques Verified for

Determination of Yield and Flow Surfaces

Many structural components are subjected to

multlaxlal loads when in service. For such

components, a threshold surface is customarily

used to distinguish irreversible response from

reversible response. For elastoplastic materials, a

yield surface can be used to delimit the elastic region

in a given stress space. The concept of a yield surface

is central to the mathematical formulation of a

classical plasticity theory. However, at elevated

temperatures, material response can be highly time

dependent; thus, viscoplastic theories have been

developed to account for the time dependency. Since

many of these theories require experimental

validation, our objective (Lissenden et al., 1996 and

1997) was to verify that current laboratory

techniques and equipment are sufficient to

determine flow surfaces at elevated temperatures.

By conducting multiple probes in axial-torsional

stress space (fig. 1), we were able to define the yield

and flow surfaces. A small offset definition of yield

( 10 _) was used to delineate the boundary between

reversible and irreversible behavior so that the

material state remained essentially unchanged and

multiple probes could be made on the same

specimen. The strain was measured with an offthe-shelf multiaxial extensometer that can measure

the axial and torsional strains over a wide range of

temperatures. The accuracy and resolution of the

extensometer were verified by comparing the

extensometer data with strain gauge data at room

temperature. This showed the extensometer had

sufficient resolution for these experiments. In

addition, the amount of crosstalk (i.e., the

accumulation of apparent strain in one direction

when strain in the other direction is applied) was

found to be negligible.

To determine flow surfaces at elevated

temperatures, the tubular specimens were heated

by induction heating, but the heating system

induced a significant amount of noise in the data.

However, by reducing thermal fluctuations and

using appropriate data-averaging schemes, the

noise was rendered inconsequential. Thus, accurate

and reproducible flow surfaces (fig. 2) were obtained.

It is now possible to validate multiaxial, viscoplastic

theories. Future work will be aimed at examining

multiaxial effects in composites.

Glenn contacts: Brad Lerch, (216) 433-5522; and Rod Ellis,

(216) 433-3340

Torsion

3 9

,3 i 5

i 15

7 angle

1

1..I Tension

4

10 :

Probe

number

1

3

5

7

9

11

13

15

Probe

angle,

deg

12

102

57

147

79

170

125

35

Probe

number

2

4

6

8

10

12

14

16

Probe

angle,

deg

192

282

237

327

260

350

305

215

Figure l.--Probes for determination of yield and flow surfaces.

6 NASA/TM--1999-206966
Shear stress,

°'12, MPa Run

°60 - F- o 1

-- .-. _ 2

.-"" _ ..... -.. o 3

/" -- "'.., [] 4

".. "'., v 5

_-4.-2) -- ._ Axial stress,

(-4.-2)- -- MPa

I t I, ! ,, , , I/i I

""""_v -20 I _r"overlaid :" circle,

_r_-' radius = 54 MPa

___ _ -- Best fit

-60

Figure 2.--Flow surfaces for 316 stainless steel at 650 °C. (Radius

offlow su_. ace, r; stress, Gij.)

Creep-Fatigue Behavior of Cast IN 939

Superalloy

Cast superalloys are being considered as hightemperature structural components for the exhaust

system of the High Speed Civil Transport. These

components are projected to experience service

cycles that include hold periods under tensile

stresses. Low cycle fatigue tests were used to assess

the influence of dwell period on life.

Cast IN 939 (Inconel) specimens were subjected to

load-controlled, elevated-temperature, tensiontension fatigue tests for hold times of 0, 1, 10, and

120 rain at maximum load. A triangular waveform

was employed. A computer-controlled, servohydraulic test machine was used for all the tests.

The specimens were heated with a susceptor, which,

in turn, was heated by an induction coil. All the

tests were conducted at a homologous temperature

of 0.7 (70 percent of the melting temperature).

The fatigue life progressively decreased with an

increase in hold time. The average fatigue life at a

hold time of 120 min was a factor of 300 shorter

than the average life with no hold time. Specimens

were examined after failure, to characterize the

damage mechanisms. Figure 1 is a polished

metallographic section of a specimen tested under

continuous cycling conditions, and figure 2 shows

a specimen tested by using a 120-rain dwell. The

loading direction was vertical. Under continuous

cycling conditions, a fiat, transgranular fracture

path was observed, whereas the 120-min-hold-time

test produced a tortuous, intergranular fracture

surface. Hold times of 1 and 10 min resulted in

both transgranular and intergranular regions on

fracture surfaces. Thus, the change in the mode of

fracture was the cause of the substantial reduction

in the fatigue life as tensile hold time increased.

Glenn contacts: Michael Verrllll, (216)433-3337; and

Sreeramesh Kalluri, (216) 433-6727

Figure l. Meta/lographic section of an IN 939 specimen tested under continuous fatigue

cycling conditions.

NASA/TM--1999-206966 7
(b)

Figure 2. Metallographic section of an IN 939 specimen tested using a 120-rain dwell at

marimum cyclic load. ".

Fatigue Crack Growth of Single-Crystal

Superalloys in Hydrogen

A study was conducted to examine the effect of

hydrogen on the fatigue-crack-growth behavior of

a PWA 1480 (Pratt & Whitney Aircraft) single-crystal

superalloy. The results showed a complex

interrelationship between the fatigue-crack-growth

behavior, hydrogen content, and the active failure

modes. In the low stress intensity range (AK} regime,

the same fatigue failure mode led to similar fatiguecrack-growth rates, regardless of hydrogen content.

Although hydrogen did not significantly influence

the fatigue-crack-growth behavior in the near

threshold regime, as the AK increased it had a major

influence in that it hindered a transition from a more

damaging failure mode to a less damaging one. From

this, we postulated that there is a distinct fatiguecrack-growth curve for each failure mode observed.

Moreover, we suspect that the plateau regions (fig. 1)

for which crack growth rates appear to be

independent of AK are the result of transition from

one fatigue-crack-growth curve to another because

of the gradual change in failure modes. For more

information, see Telesman et al. (1996).

Glenn contact: Jack Telesman, (216) 433-3310

¢

>_

O

O

2

O_

O

L_

10-4

10-5

104

10-7 --

10-8

10-9 __-

10-1o

10o

In 34-MPa H2 /

enviroment --_ /. "

_\ _ " 7 Brittle octahedral

fA// r A_ failure mode

Failure mode / A_"Af

1 rH2 {00 }-_ _ .

\/_Ht_argej/\ _-- Ductile octahedral

__ failure mode

i,' / I I

t' nc"ar0o0i I

| I f t II LKII I I I i I ilia

101 102

Stress intensity range, 4K, MPa-ml/2

Figure 1.--Fatigue-crack-growth cun'es for different failure modes showing the onset of

plateaus due to failure mode transition.

8 NASA/TM-- 1999-206966
Reliability Modeling of Brittle, Anisotropic Solids

A general reliability model was developed for use in

the design of structural components made from

brittle, homogeneous anisotropic materials such as

single crystals or materials with an anisotropic flaw

distribution. This model incorporates a variable

strength distribution and any equivalent stress

criterion such as the coplanar energy release rate

or the minimum strain energy density.

By considering the probability associated with a flaw

existing in a component and the probabilities that

the flaw has a specific orientation and a critical size,

we derived the following reliability function for

anisotropic materials:

PF=I_EXP/____I _ 1 InI_leq(O'x)lrndOdA

where (_leq(0, x) is the mode-I equivalent stress and

60(0) is a function describing strength as a function

of orientation. Strength as a function of orientation

is shown in figure 1 for single crystal NiA1. This

model is similar to the model frequently used for

isotropic, polycrystalline ceramics, with two

exceptions: (1) the strength in this case is a function

of planar angle, and (2) an equivalent stress

formulation defining coplanar extension for an

anisotropic body is used.

The model is applicable to materials that exhibit

anisotropy in both strength distribution and

elasticity or just strength distribution. The strength

anisotropy may result from either the distribution

of flaws or the variation in fracture toughness with

orientation.

In order to verify the model, uniaxial and biaxial

strength data were generated from a single-crystal

NiA1 alloy developed by General Electric Aircraft

Engine Co. The effect of test specimen volume is

shown in figure 2. In the case of uniaxial loading,

the model can predict the effects of test specimen

volume on measured strength as a function of

crystal orientation. Coding of the model for

verification of the multiaxial predictive capability,

via the biaxial test data, is ongoing. For more

information, see Darolia et al. (1996) and Salem

et al. (1998).

Glenn contacts: Jonathan Salem, (216) 433-3313; and

Ronald Noebe, (216) 433-2093

Curve fit

90 ° O Mean strength

1600 _ • Standard deviation

/ Angle

_ / / _. deg

0- 1200

z:"

800

U)

,';" 400

0 _ I I : ;O'--"--4O o

0 400 800 1200 1600

Flexural strength, MPa

Figure /.--Polar plot of flexural strength as a function of crystal

orientation fi_r single crystal NiAI

99

9O

3o

"6

.__,10

-Q 5

2

n

• <100> 3-point

• <100> 4-point

0 <110> 3-point

13 <110> 4-point

o _o ..'o

0

11 •

I I I l I

4 6 8 12 16 18x103

Stress, MPa

Figure 2,--Measured strength distribution for a single co'stal NiAI

alloy

NASA/TM--1999-206966 9
Fatigue Behavior and Life Prediction of

a T650-35/PMR-15 (0/90) Weave Under

Isothermal Conditions

Currently, the gas turbine community is working

toward advanced aircraft that will provide greater

payload deliveries with corresponding increases in

fuel efficiency. Of vital importance in the

development of propulsion systems for these aircraft

is the utilization of cost-effective materials. Among

such materials are lightweight, high-temperature

polymer matrix composites (PMC's). The loadbearing PMC structures are expected to perform at

temperatures of 300 °C or higher, but few reports

are available regarding the mechanical behavior and

life prediction of woven layups at elevated

temperatures (ET). Because of the new expectations,

there is a need for data about the tensile,

compressive, and fatigue characteristics of woven

PMC's subjected to ET. The current investigation

focused on the characterization of the monotonic

tensile, compressive, and fully reversed fatigue

(R_ = -1) behavior exhibited by a T650-35/

PMR-15, 8-satin harness (0/90) weave at room

temperature (RT) and ET. The study documented

the microscopic cracking events and their influence

on macroscopic properties such as stiffness and

residual compressive strength. The isothermal static

and fatigue experiments were conducted at 22 °C

and 316 °C. An analytical model, based on modulus

degradation as a measurable damage variable, was

implemented for damage and life prediction (Ye,

1989). The damage variable was defined as D = 1 -

E/E o where E and E o represent the residual and

initial compressive moduli of the specimen,

respectively. A power law relationship was utilized

to describe the accumulation of damage:

c

ndD A °max

dN BD B-I

{I)

where A, B, and C are material constants. Damage

evolution and life prediction were determined by

integrating equation (i). For life prediction, a definition

of critical damage (Dcnucal) was required; it was

assumed to be a function of load and temperature.

The temperature effects on the moduli and static

tensile strength were minimal. At RT the averages

for the longitudinal moduli were Etensfle = 75 GPa

and Ecomp = 71 GPa; at ET they were reduced by

only 2 percent. Similarly, the 22 °C static tensile

strength and strain were 855 MPa and 1.15 percent,

respectively, whereas at ET, tensile strength was

reduced by 3 percent and compressive strength by

7 percent. The results for the ultimate compressive

stress and strain at RT were -675 MPa and

-1.12-percent strain. Fatigue lives, on the other

hand, were negatively influenced by ET, as can be

seen in figure I. At both temperatures, failure

occurred during compression because of the

residual compressive strength's sensitivity to the

environment and damage. Figure 2 displays the

typical three-stage compressive modulus

degradation behavior and its repeatability as

evidenced by the three RT specimens. For modeling

purposes, failure was defined as the final modulus

value Dcriticat before stage Ill behavior was initiated

(see fig. 2). At this point each specimen had

exhausted 98 percent of its life. For a given stress,

both the RT and ET specimens followed the same

degradation path, although the ET specimens failed

before suffering much damage (i.e., modulus

degradation). At RT, Dcritical values of the specimens

typically exhibited a 16- to 22-percent modulus

reduction, whereas for ET specimens the reductions

in modulus before failure ranged from 5 to 9 percent.

The characterization of equation (1) was achieved

by using only the RT modulus data; this equation

was observed to be dependent only on the applied

stress. Recall that the modulus degradation rates

were equivalent at the two temperatures; thus

temperature dependence was introduced in the

Dcritical function. We observed that Dcritical was a

function of temperature and maximum applied

stress. Figure I compares the model with the

experimental data. The results indicate good

agreement between the theory and the data at the

temperatures used to characterize the model and

at an intermediate temperature. The plateau of the

curves is the point where no damage growth (i.e.,

modulus reduction) is detected before failure; these

results also agree with the experimental

observations. In summary, the temperature caused

a degradation in properties that was dependent on

the resin (e.g., monotonic compression and Rv = -1

fatigue). Also, the residual modulus was shown to

be a good indicator of damage and was successfully

employed as a tool for life prediction.

Glenn contact: Andrew L. Gyekenyesi, (216) 433-8155

10 NASA/TM--1999-206966
120

800 --

110

700 --

a. "_ 100

600 -- _ 90

80

E

E

= 500 -- E

E ._ 70

8 so

400 --

50

300 --

200 --

900 -- 130 m

C_

_ Tensile Temperature,

°C °F

22 72

---£3-- 204 400

-- _ --z_-- 316 600

-- _ _ Solid symbols indicate data used

A

Compressive_ to characterize model

_ "-,_ Lines indicate modeled data

3o ! _J_t,id I_,l,i,i I-,l_hl,I

10 -1 100 101 102 103 104 105 106 107

Cycles to failure, N

Figure .--Stress-based isothermal fatigue life comparison fi)r T650-35/PMR-15 (0/90) weaves.

1.1

1,0 I

0.9

"0

0

E 0.8

10

.N

E O.7

O

z

0.6

0.5

Stage 1

A C'/

_ Stage II A ©Q Dcritical

\

0 A A A A \'-

0 0@

Specimen Cycles to

failure,

Nf Stage II I

• PC2_30 3552

O PC3_17 3721 •

A PC2_24 8428 O

I I I I I

0.0 0.2 0.4 0.6 0.8 1.0

Normalized life, NINf

Figure 2.--Normalized compressive modulus degradation behavior for multiple specbnens

(Oma x = 379 MPa and temperature = 22 °C).

NASA/TM-- 1999-206966 11
A General Representation Capturing

Reversible and Irreversible Hereditary

Behavior of a Titanium Alloy

A number of titanium matrix composite (TMC)

systems are currently being researched and evaluated

for high-temperature airframe and propulsion system

applicaUons. As a result, numerous computaUonal

methodologies to predict both deformation and life

for this class of materials are under development.

Integral to these methodologies is an accurate and

computationally efficient constitutive model for the

metal matrix constituent. In addiUon, because of the

elevated temperatures at which these systems are

designed to operate, such a model must account for

both time-dependent and time-independent

deformations.

To accomplish this, we employed a recently

developed, complete potential-based framework

(Arnold and Saleeb, 1992) that uses internal state

variables and that was put forth for the derivation

of reversible and irreversible constitutive equations.

These reversible and irreversible domains are

posited to be delineated by a temperaturedependent threshold surface, as illustrated in

figure I for low-, mid- and high-temperature

regimes. An example titanium alloy that has

experimentally exhibited both a time- and

temperature-dependent reversible (linear

viscoelastic) and irreversible (viscoplastic) domain)

(Arnold et al., Castelli, 1997) isTIMETAL 21S. Our

previous work (Arnold et al., 1994 and 1995(b))

focused on the specification and characterization

of the irreversible domain for this material. The

'_ij - "i1 -_

I

I

(a) (b)

cry/-

7

_ Time-independent reversible region

_ Time-dependent reversible region

IR Irreversible

(quasUinear) region

(c)

Figure l.--Reversible and irreversible threshold surfaces (a) High temperature. (b) Mid

temperature, (c ) Low temperature.

12 NASA/TM--1999-206966
Temperature,

oC

300

1.0 _. ......e .........::":_"=_"__:.,_'---=y--- o^ Oo

["4A9...... ;_,_%_" "" T o . o_,,,k...,,--..,.-'r--- o oo _

B-v3 " a A_ A-.'_'_ A A AOA A_¢,.'_ - v O °o

_s& A _ ;vA ° o o o

o°;< °o ° oo °°

°.,[,24°°: o/-

k.OTo

"_ 0.6 650

._ 13.4

[ Symbols denote experiments

0.2 _ Lines d=eno_tecorrelations

(a) to = 650 min

0.0

0.0 0.2 0.4 0.6 0.8 1.0

Relative time, t/t o

1.2

1.0

E 0.8

t3

•£ 0.6

U_

¢0

0.4

Temperature,

°C

30O

""_t_,_.,,,_,_,,_,_,_._.,_,,,,_,_,_.., * 482 . ,,

-- _ 565

650

0.2

(b)

0.0 ......... t ......... t ......... I ......... I .........

0.0 0.2 0.4 0.6 0.8 1.0

Relative time, t/t o

Figure 2.--Nonisothennal viscoelastic correlations with experiments. (a) Creep. (b)

Relaxation.

primary objective of our present work is to construct,

within the context of a complete potential structure,

a multiaxial, nonisothermal viscoelastic model to

describe the reversible strain component of the total

strain decomposition. To construct a potentialbased viscoelastoplastic mathematical framework,

we had to specifically partition both the strain

(reversible and irreversible) and stress (equilibrium

and nonequilibrium) variables. By making the

reversible portion of the model specific and

assuming all mechanical elements (i.e., springs and

dashpots) have the same Poisson ratio effect, we

significantly reduced the required number of

material parameters and characterization effort for

the model. Thus, a successful nonisothermal

characterization of this reversible viscoelastic model

has been completed for TIMETAL 2 IS. The resulting

normalized creep and relaxation response histories

are shown, respectively, in parts (a) and (b) of

figure 2, for four temperatures (300, 482,565, and

650 °C).

Glenn contact: Steven M. Arnold, (216) 433-3334

NASA/TM-- 1999-206966 13
Damage Mechanisms and Mechanical

Property Degradation in Titanium Matrix

Composites Subjected to TMF Loadings

Over the past decade, continuously reinforced

titanium matrix composites have been the focus of

many research programs aimed at evaluating these

materials for high-temperature aeronautics and

aerospace applications. Specifically, the use of

theses materials in future supersonic and

hypersonic flight vehicles and advanced propulsion

systems is viewed as a critical enabling technology.

Since the vast majority of such applications involve

combined cyclic stress and temperature conditions,

it is essential to characterize and evaluate material

behavior under thermomechanical fatigue {TMF}

loadings if the ultimate goal is to develop robust

deformation and damage/life models applicable to

in-service conditions.

Detailed examinations of TMF deformation, damage

progression, and life behaviors have been completed

and summarized for multiple laminates of

continuous SiC-fiber-reinforced titanium matrix

composites (Castelli, 1996). Fundamental behaviors

resulting from in-phase (IP) and out-of-phase (OP)

TMF loadings were addressed with particular

emphasis on the micromechanical damage

mechanisms leading to macroscopic failure. This

was accomplished by using both fractography and

metallography to examine microstructural details.

In addition, macroscopic damage progression was

determined on the basis of mechanical property

degradation from data collected with an advanced

TMF test technique (Castelli, 1994) that allowed

explicit measurements of the macroscopic

isothermal static moduli and the coefficient of

thermal expansion (CTE) as functions of the TMF

cycles.

Zero-tension TMF life results indicate analogous

trends for both [0] and [0/90] titanium matrix

composite laminates (see fig. 1). High stress IP TMF

and mid- to low-stress OP TMF loadings are more

life-limiting than maximum temperature isothermal

conditions. The dominant damage mechanisms

changed with cycle type. Damage resulting from IP

200

160

]

120

,=

._ 80

1400

1200

_. 1000

8oo

_ 600

_E 400

Orientation

V [0°]

0 [0o/90o1

[90o]

Open symbols indicate in phase

Solid symbols indicate out of phase

_7

A \ tD 40

.......

200-

101 102 103 104 105

Cycles to failure, Nf

Figure l.--Thermomechanical fatigue life behaviors of SiC-reinforced titanium matrix composite

laminates.

14 NASA/TM--1999-206966
TMFconditionsproducedmeasurabledecreasesin

static moduli, but only minimal changesin CTE

occurred. Microstructural damageis generally

associatedwith [0] fiber cracking,sparsematrix

damage,and no surface-initiatedcracking.Under

OPTMFconditions,notablestaticmoduliandCTE

degradationswere experienced.Here, conditions

promotedenvironment-enhancedsurface-initiated

cracking.The [0/90] titanium matrix composites

alsoexhibitedmatrixcrackinginitiatedatdebonded

[90]fiber/matrix interfaces.Bothlaminatesshowed

little to no [0]fiber crackingunder OPTMF.Zerotension TMF of a [90] titanium matrix composite

revealed significant changesin both the static

moduli and CTE under IP and OP conditions.

Mechanicalpropertydegradationpatternswerevery

similar to those exhibited under isothermal

loadings.Further, damagemechanismspromoted

by isothermal,IPTMF,and OPTMFloadingswere

essentiallyidentical,consistingof debonding at the

fiber/matrix interface followed by matrix cracking

that initiated at the debond and propagated

transverse to the applied load. For more

information, see Castelli (1996b), Castelli and Rao

(1996), Verrilli and Castelli (1996a and b), Lei

et al. (1996), and Arnold et al. (1994 and 1995a).

Glenn contact: Michael G. Castelli, (216) 433-8464

NASA/TM--1999-206966 15
Structural Inte rit

CARES/L_fe Ceramics Durability Evaluation Software for Finite-Element Analysis Programs

The CARESLife program is a state-of-the-art

software design tool that predicts the timedependent reliability of monolithic ceramic

components. The program has many features and

options for materials evaluation and component

design. Its capability, flexibility, and uniqueness has

attracted many users representing a broad range

of interests and has resulted in numerous awards

for technological achievements and technology

transfer.

CARESLife is an integrated package that predicts

the probability of a monolithic ceramic component's

failure as a function of time in service. It couples

commercial finite-element programs---which resolve

Dental Crown Turbocharger Wheel

ZnSe Vacuum Chamber Window Television Picture Tube

Evacuated

tube stresses

Figure/.--Examples of components designed with CARESLife.

16 NASA/TM--1999-206966
acomponent'stemperatureandstressdistribution--

to reliability evaluation and fracture mechanics

routinesformodelingstrength-limitingdefects.These

routines are basedon calculationsof the brittle

material'sprobabilisticstrength.

Thekeyto maintaininginterestin CARESLifehas

been active maintenance and promotion of the

programaswell asenhancementof the program's

easeofuse.Towardsthis endtheCARES/Graphics,

and CARESInputmoduleswerecreated.CARES/

Graphics enables the plotting of specimen rupture

data and the corresponding regression lines that

characterize the data. CARESInput is an interactive

input preparation program to CARESLife. It helps

to eliminate the bewildering array of input files and

keywords confronting the user, performs error

checking, and includes online explanations as

appropriate. Incorporating the many suggestions

from users and reviewers has significantly improved

the CARESInput program in 1996.

CARESLife has continued to be in high demand

worldwide, although present technology-transfer

efforts are primarily focused on U.S.-based

organizations. Success stories can be cited in

numerous industrial sectors such as aerospace,

automotive, biomedical, electronic, glass, and

nuclear- and conventional-power generation

industries. For more information, see Nemeth et al.

(1996a and b), Salem et al. (1996), and Sokolowski

et al. (1996).

Glenn contacts: Noel N. Nemeth, (216) 433-3215;

Lynn M. Powers, (216) 433-8374; and Lesley A_ Janosik,

(216) 433-5160

NASA/TM--1999-206966 17
Creep Life Prediction of Ceramic Components Subjected to Transient Tensile

and Compressive Stress States

Since ceramics have desirable properties at high

temperatures, there is interest in using them in

structural applications such as advanced turbine

systems. However, the design lives for such systems

can exceed 10 000 hr. With such long life

requirements, the stresses on the components must

be kept relatively low. The combination of high

temperatures and low stresses typically places failure

mode for monolithic ceramics in the creep regime.

A design methodology for predicting the lifetimes of

structural components subjected to multlaxial creep

loading has been developed. This methodology

utilizes commercially available finite-element

packages and takes into account the time-varying

creep stress distributions {stress relaxation). In this

methodology, the creep life of a component is divided

into short time steps, during which the stress and

strain distributions are assumed constant. The

damage D is then calculated for each time step on

the basis of a modified Monkman-Grant creeprupture criterion. For components subjected to

predominantly tensile loading, failure is assumed

to occur when the normalized accumulated damage

at any point in the component is greater than or

equal to unity.

Some ceramic components, such as vanes and

rotors, are subjected to concurrent tensile and

compressive stress fields. For such components,

failure generally starts at or near the most highly

stressed point and subsequently propagates across

the section. The creep-rupture life for members

subjected to concurrent tensile and compressive

loading is divided into two stages. The first is called

the stage of latent failure (damage initiation). During

this stage, damage accumulates until it reaches

unity at some point within the component and

failure begins. (Hence. this portion of life represents

the entire predicted life for predominantly tensile

components.) Damage due to compressive stresses

is assumed to be negligible, although in this

methodology, it can be accounted for very easily if

it is determined to be of any significance.

Subsequently, the second stage, named the damage

propagation stage, takes place. During this stage a

damage front defined by the condition D = 1 will

travel through the body or surface of the component.

Component failure occurs at the end of this stage

when its total load-carrying capacity is expended.

In CARESCreep (Ceramics Analysis and Reliability

Evaluation of StructuresCreep), this means that

failure is assumed to occur when D = l at the

periphery of the expanded critical damage zone. The

corresponding time will be the creep-rupture life

for that component. This size of the critical damage

zone corresponding to creep-rupture failure varies

depending on the load conditions and component

configuration. One way to estimate the duration of

the propagation stage is to assume it is equal to

the time required for the damage zone to penetrate

the initial tensile stressed portion of the structure.

An example of this technique applied to a flexure

specimen, is shown in table I.

This creep-rupture life methodology has several

advantages. First, it yields a cumulative damage map

for the component, showing the critical locations

where failure would originate; this can be very helpful

in the redesign of such components. In creep-type

loading applications, it is not a trivial task to predict

the location of failure since the multiaxial stress

components are redistributed as time elapses. Thus,

failure does not necessarily occur at the location

where stresses are highest at the beginning of loading

or at the time of failure; it can take place elsewhere.

A second advantage is that this design methodology

could incorporate the primary creep strain effect into

the analysis (thereby influencing the stress state),

which could then predict shorter lives (conservative

predictions) than can be predicted when only the

secondary creep stain effect is used. The third

advantage is that any equivalent stress criterion can

be used to predict the component's life. Thus, as

multiaxial creep data emerge in the future and we

understand better how ceramics fail under such

applications, the CARESCreep code can be modified

accordingly. Fourth, any creep-rupture criterion

(continuum damage mechanics, Larson-Miller,

minimum commitment method, etc.) can be utilized

to compute damage and predict life. For more

information, see Jadaan et al. (1996).

Glenn contacts: Lynn M. Powers, (216) 433-8374; and

Lesley A. Janosik, (216) 433-5160

Table I.--FLEXURE SPECIMEN RESULTS

Initial maximum

stress,

MPa

20O

250

30O

350

Experimental Damage Predicted

failure time, initiation, lifetime,

hr hr hr

830 200 830

55 14.5 80

45 12 8.6

1.3 .12 1.3

18 NASA/TM--1999-206966
Constitutive Theory Developed for

Monolithic Ceramic Materials

With increasing use of advanced ceramic materials

in high-temperature structural applications such

as advanced heat engine components, the need

arises to accurately predict thermomechanical

behavior that is inherently time-dependent and

hereditary--in the sense that current behavior

depends not only on current conditions but also on

thermomechanical history. Most current analytical

life prediction methods for both subcritical crack

growth and creep models utilize elastic stress fields

to predict the time-dependent reliability response

of components subjected to elevated service

temperatures. The inelastic response of these

material systems at high temperature has been well

documented in the materials science literature, but

this issue has been ignored by the engineering

design community. However, from a design

engineer's perspective, it is imperative to emphasize

that accurate predictions of time-dependent

reliability demand accurate stress field information.

Ceramic materials exhibit different time-dependent

behavior in tension and in compression. Thus,

inelastic deformation models for ceramics must be

constructed in a fashion that admits both sensitivity

to hydrostatic stress and differing behavior in

tension and in compression. A number of

constitutive theories for materials that exhibit

sensitivity to the hydrostatic component of stress

have been proposed; they characterize deformation

by using time-independent classical plasticity as a

foundation. However, none of these theories allow

a different behavior in tension and in compression.

Furthermore, these theories are somewhat lacking

in that they are unable to capture the creep,

relaxation, and rate-sensitive phenomena exhibited

by ceramic materials at high temperature.

The objective of our effort has been to formulate a

macroscopic continuum theory that captures these

time-dependent phenomena. Specifically, we have

focused o_ the inelastic deformation behavior

associated with these service conditions by developing

a multiaxial viscoplastic constitutive model that

accounts for time-dependent hereditary material

deformation (e.g., creep, stress relaxation, etc.) in

monolithic structural ceramics. Using continuum

principles of engineering mechanics, we have derived

the complete viscoplastic theory from a scalar

dissipative potential function. Constitutive equations

for the flow law (strain rate) and evolutionary law

have been formulated on the basis of a threshold

function, identified here as F (see fig. 1), that is

sensitive to hydrostatic stress and allows different

behavior in tension and compression. For illustration,

a set of threshold flow stress values has been adopted

that roughly corresponds to values anticipated for

isotropic monolithic ceramics. Specifically, the

compressive uniaxial threshold stress value is (_c =

2.00 MPa, and the tensile uniaxial threshold stress

value is (_t = 0.20 MPa. In addition, Inelastic

deformation is treated as inherently time-dependent,

and a rate of inelastic strain is associated with every

state of stress. Thus, creep, stress relaxation, and

rate sensitivity are phenomena resulting from applied

boundary conditions and are not treated separately

in an ad hoc fashion.

Complete details of the model and its attending

geometrical implications are presented in Janosik

and Dufflj (1998). A quantitative assessment has

yet to be conducted since the material constants

have not been suitably characterized for a specific

material. Incorporating this model into a nonlinear

finite-element code would provide industry a means

to numerically simulate the inherently timedependent and hereditary phenomena exhibited by

these materials in service. This approach could

make life prediction more accurate for structural

ceramics used in high-temperature applications.

Glenn contacts: Lesley A. Janosik, (216) 433-5160; and

Stephen F. Duffy, (216) 433-5626

,=o'22,

0 5 I MPa

- Ml-'a

Figure ].-- Threshold fimction F,projected onto the GIt -- c22 stress

plane.

NASA/TM--1999-206966 19
Single Transducer Ultrasonic Imaging

Method That Eliminates the Effect

of Thickness in Imaging of HighTemperature Structural Materials

NASA Glenn Research Center Life Prediction Branch,

in partnership with Sonix, Inc., and Cleveland State

University, recently advanced the development,

refinement, and commercialization of an advanced

nondestructive evaluation inspection method entitled

the Single Transducer Thickness-lndependent

Ultrasonic Imaging Method (Roth 1996; and Roth

et al. 1996a and b). Selected by R&D Magazine as

one of the I00 most technologically significant new

products of 1996, the method uses a single

transducer to eliminate in ultrasonic images of

materials the effects of superimposing thickness

variation (fig. I). As a result, any variation seen in

the ultrasonic image is solely due to microstructural

variation. This method precisely and accurately

characterizes material gradients (pore fraction,

density, chemical) that affect the uniformity of

physical performance (mechanical, thermal,

electrical) for the material. Among the advantages of

this method over conventional ultrasonic imaging are

(1) elimination of machining costs (for precision

thickness control) during quality control stages of

material processing and development and (2)

elimination of labor costs and the subjectivity inherent

to further image processing and image interpretation.

At NASA, this method has primarily been used for

accurate inspection of high-temperature structural

materials such as monolithic ceramics, metal matrix

composites, and polymer matrix composites. This

year, data were published for plate-like samples,

and the current research is focused on applying

the method to tubular components.

Initial publicity about the development of this

method has generated 150 requests for further

information. These requests came from a wide

variety of institutions and individuals, including the

Federal Bureau of Investigation, Lockheed Martin

Co., Rockwell International, Hewlett Packard Co.,

and Procter & Gamble. In addition, NASA has been

solicited by the 3M Co. and Allison Abrasives to

investigate the use of this method on composite

materials manufactured by these companies.

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

50-MHz peak

amplitude c-scan

4O

Percent of full-scale height

_-88

40

20-MHz time-of-flight

('rOF) c-scan

20 20

0 O Z0 mn 8 _

• Neither microstructure Thickness

nor thickness variations • Thickness variation masks

(insensitive) microstructure variations

Scan

duration,

_sec

20-MHz thicknessindependent image

4O

20-

0

Scanning

speed,

mm/Hsec

-il ,oa

-18._

-18 .'tO

8 ZB _t

• Microstructure variations

(low, high, intermediate

pore fraction)

Figure /.--Conventional ultrasonic imaging versus thickness-independent ultrasonic imaging of silicon nitride monolithic

ceramic with pore fraction variations and l O-percent thickness variation.

20 NASA/TM--1999-206966
Noncontact Measurement of AcoustoUltrasonics for Ceramic Matrix

Composites

Advancements in the processing and fabrication of

ceramic matrix composites (CMC's) for use in engine

components require appropriate mechanical and

nondestructive testing methods. These methods are

needed to characterize the properties, assess the

integrity, and predict the life of engine components.

Acousto-ultrasonics (AU) has demonstrated the

capability to detect fatigue degradation of CMC's in

experiments that caused matrix cracks to be formed

and propagated in a manner typical of that caused

by the mechanical stresses experienced during

service (Periera et al., 1994; and Brewer et al., 1995).

The AU measurements' performed for the cited work

required contact coupling of the ultrasonic

transducers to the surface of the specimens used.

However, being able to do AU measurements without

contacting the specimens would be more desirable.

This would allow AU experiments to be done at high

temperatures without the limitations imposed by the

ultrasonic equipment. Beyond this, it would allow

AU to be done on structures while they are in service.

In this latter application, probes could be designed

to monitor the health of aircraft engine parts.

For a number of years (Huber and Green, 1990)

lasers have been under investigation as remote

ultrasonic input sources and ultrasonic detectors.

The use of ultrasonic transducers, coupled through

an air gap, has also been studied for some time

(Safaeinili, et al., 1995). At NASA Glenn we have

been more successful in using the laser as an

ultrasonic input medium than as an output device.

On the other hand, we have been more successful

using an air-coupled transducer as an output device

than as an input device. For this reason, we have

studied the combination of laser input with aircoupled transducer output.

With this laser-in/air-out noncontact combination,

the first antisymmetric plate wave velocity in SiC/

SiC specimens was successfully measured (Kautz,

1996). Figure 1 shows the noncontact measurement

system. Tests of 18 specimens of 3 architectures

measured by both this noncontact and a contactcoupling method gave the same results (see fig. 2).

Now the noncontact technique need to be extended

to other AU measurements, such as ultrasonic

decay, and to other material systems of interest in

aeronautics, such as metal- and polymer-matrix

composites. This future work will utilize higher

intensity laser beams and more sensitive aircoupled transducers for better imaging of the

ultrasonic signals that must be analyzed.

This AU technique, once calibrated to a composite

system and component geometry, may be useful in

predicting the life of aerospace components.

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

Waveform

array out

Waveform

array in

Trigger in

delay Time

Delay

trigger out

[ Delay

trigger in

__] Waveform Si_.gnal

[ digitizer ]-in

-- computer [pre=

[ li!

• NdYAG

laser

90o Prism

_,,,_ _Signal

ut -_- Air-coupled, Foot switch 1-7

_----_,_ 0.5-MHz

_l_'_ transducer with

'eo"'_ focal length of 5.08 crr

Signal in

Figure l.--Experimental arrangement for collecting laser-in, air-out acousto-u/trasonic waveforms.

NASAfrM--1999-206966 21
0.40

0.39 --

0.38-

o 0.37

,-_ 0.36

¢3

_o

a)

>

._ 0.35

_ 0.34

0.33

0.32

°_/o

Q

8o/,_ o

- o Jo %o

i I

0.32 0.34 0.36 0.38

Contact velocity, VS, cm/lasee

Figure 2.--Laser-air versus contact plate veloci_. ; VS.

0.40

22 NASA/TM--1999-206966
Structural D namics

Dynamic Spin Rig Magnetic Suspension

The Glenn Research Center's Dynamic Spin Rig,

located in Building 5, test cell CW-18 (see fig. 1), is

used to test turbomachinery blades and components

by rotating them in a vacuum chamber. A team from

the Machine Dynamics Branch successfully

integrated a magnetic bearing and control system into

the Dynamic Spin Rig (see fig. 2). This magnetic

bearing worked very well to both support and shake

the shaft. We demonstrated that the magnetic

bearing can transmit more vibrational energy into

the shaft and excite some blade modes to larger

amplitudes than can the existing electromagnetic

shakers. In cooperation with the University of

California at San Diego, experiments conducted on

damping of composite plates successfully

demonstrated the system's robustness for long-term

iiii_iiiiiii!iiii

Figure 1.--NASA Lewis Dynarnic Spin Rig Facility:

NASAYrM-- 1999-206966 23
Figure 2.--Radial magnetic bearing in NASA Lewis Dynamic Spin Rig Facili_.'

along with viscoelastic damped composite plates attached to rotor.

testing. Existing magnetic bearing system hardware

and software were modified to give enhanced

capabilities. New features were added, for example,

a user-friendly interface, and flexibility in

orientation and phasing while applying the

excitation force. We discovered that the bearing can

also provide blade damping by using feedback from

the blade strain gauges. This Is an additional benefit

since insufficient blade damping Is a critical problem

in advanced turbomachinery blades. The success

of this initial work will lead to the design and

implementation of a full magnetic suspension

system for the Dynamic Spin Rig by September

1997. The facility can now provide either a

mechanical or magnetic support system for rotors.

The magnetic support will enable longer run times

at higher speeds and larger vibration amplitudes

for rotating blades.

Glenn contacts: Dexter Johnson, (2161 433-6046;

Oral Mehmed, (216) 433-6036; and Gerald V. Brown, (216)

433-6047

24 NASA/TM--1999-206966
High-Temperature Magnetic Bearings for

Gas Turbine Engines

Magnetic bearings are the subject of a new NASA

Glenn Research Center and U.S. Army thrust, with

significant industry participation and cooperation

with other Government agencies. The NASA-Army

emphasis is on high-temperature applications for

future gas turbine engines, so magnetic bearings

could increase the reliability and reduce the weight

of such engines by eliminating the lubrication system.

They could also increase the DN limit (i.e., the

diameter of the bearing times the revolutions per

minute) on engine speed and allow active vibration

cancellation systems to be used, thereby yielding a

more efficient, "more electric" engine. In addition, the

Integrated High Performance Turbine Engine

Technology (IHPTET) Program, a joint Department of

Defense-industry program, identified a need for a

high-temperature (1200 °F) magnetic bearing that

could be demonstrated in a phase-III engine.

The magnetic bearing being developed is similar to

an electric motor. It has a laminated rotor and stator

made of cobalt steel. Wound around the stator are

a series of electrical wire coils that form a series of

electric magnets around the circumference. The

magnets exert a force on the rotor. A probe senses

Figure 1.--NASA lO00 °F Magnetic Bearing Test Rig.

NASA/TM--1999-206966 25
the positionof the rotor, and a feedback controller

keeps it centered in the cavity. The engine rotor,

bearings, and casing form a flexible structure that

accommodates a large number of modes. The

bearing feedback controller, which could cause some

of these modes to become unstable, can be adapted

to varying flight conditions to minimize the seal

clearances and monitor the health of the system.

Cobalt steel has a curie point greater than 1700 °F,

and copper wire has a melting point beyond that.

Therefore, practical limitations associated with the

maximum magnetic field strength in cobalt steel and

the stress in the rotating components limit the

temperature to about 1200 °F. The objective of this

investigation was to determine the speed and

temperature limits of the magnetic bearing operating

in an engine. Our approach was to make use of the

Glenn expertise in magnets, mechanical components,

high-temperature materials, and surface lubrication

to build and test a magnetic bearing in both a rig

and an engine. Testing would be done at Glenn or

through cooperative programs in industrial facilities.

In 1995, we made significant progress. We initiated

with Allison Engine a cooperative program to work

on a high-temperature magnetic thrust bearing.

During this work, we uncovered a problem with the

conventional design of the magnetic thrust bearing:

Because the thrust bearing was not laminated, it

caused eddy currents that severely reduced the

bandwidth. We also worked at Allison to bring their

high-temperature magnetic bearing rig to full speed.

In addition, we predicted the stability limits for both

the inhouse and Allison magnetic bearings and

tested a high-temperature displacement probe

Currently, our flexible casing rig is being converted

to a high-temperature magnetic bearing rig (see

fig. 1). Testing should start in the third quarter of

1997. Our plan is to develop a high-temperature,

compact wire insulation and to fiber-reinforce the

core lamination to operate at higher temperatures

and DN values. We plan to modify our stability

analysis and controller theory by including a

nonlinear" magnetic bearing model. We are

developing an expert system that adapts to changing

flight conditions and that diagnoses the health of

the system. Then, we will demonstrate the bearing

in our rotor dynamics rig and, finally, in an engine.

Glenn contacts: Albert Kascak, (216) 433-6024; and

Gerald Brown, {216) 433-6047

Neural Network Control of a Magnetically

Suspended Rotor System

Magnetic bearings offer significant advantages over

conventional bearings because they do not contact

other parts during operation, which can reduce

maintenance. Furthermore, they are far superior

because they offer higher speeds, no friction, no

lubrication, weight reduction, precise position

control, and active damping. However, there are

technical barriers, that limit the application of

magnetic bearings in industry. One of them is the

need for a nonlinear controller that can overcome

the system nonlinearity and uncertainty inherent

in magnetic bearings. For our work, a neural

network was selected as a nonlinear controller

because it generates a neural model without any

detailed information regarding the internal working

of the magnetic bearing system; it can even be used

for systems that are thought to be too complex or

where an accurate system model cannot be derived.

A feed-forward architecture with a back-propagation

learning algorithm was selected because of its

proven performance, accuracy, and relatively easy

implementation.

The neural network plant emulator was first trained

to emulate a theoretical model of the nonlinear

plant. A discrete theoretical model of the plant

dynamics in state-space notation was used to

choose both the present states of the plant {rotor

displacement and velocity} and the plant input

{control current) as input to the emulator. The next

states, the rotor displacement and velocity after one

sample time, were chosen as output from the plant

emulator. During the learning procedure, the errors

between the actual network output and the desired

values were minimized by upgrading the weights.

After training, the neural emulator perfectly

predicted the next states (delayed by one sample

time} of the magnetic bearing system for the current

states and control force, which were not in the

training sample data.

In the second step, the trained neural emulator was

used to train a neural net controller to make the

whole system meet conventional performance

specifications such as bandwidth, settling time,

overshoot, and such. We wanted the controller to

take the current magnetic bearing states x(t) and

demand r as input parameters, and to deliver a

control force u(t) as output to the magnetic bearing

26 NASA/TM-- 1999-206966
system.Thesecurrent state valuesshould make

the magneticbearing'snextstatevectorxIt + 1) be

identical to that defined by the desired linear

reference system, which satisfies the performance

specifications in either the frequency or the time

domain.

Figure 1(a) shows the Bode plot of a simple secondorder linear reference model derived from frequency

domain specifications; figure 1(b) shows the closedloop magnetic bearing system after training. They

are almost identical, even after 200 training epochs.

Another neural controller based on time domain

specifications was trained and tested by simulating

its response for the initial condition and by

comparing the results to the actual magnetic

bearing response (see fig. 2). The neural network

controller was so accurate that it perfectly

overlapped the magnetic bearing response

(+ symbols).

In summary, a neural network controller that

circumvents the magnetic bearing's nonlinearity

was developed and successfully demonstrated on a

small Bentley-Nevada magnetic bearing rig. The

neural plant emulator and neural controller were

so accurate that the neural network controller did

a nearly perfect job of making the nonlinear

magnetic bearing system act like the linear reference

model. This novel approach demonstrated the

feasibility of using it for advanced turbomachinery

systems with large-scale magnetic bearings of

unknown dynamics.

Glenn contacts: Benjamin Choi, (2161 433--6040; Gerald

Brown, (216) 433--6047; and Dexter Johnson, (216) 433-6046

1.2_ ._== o.8 ....

"O 0.4_-- --_ ......

0.01 1 I Lllll=[ I J lilllJ["_-,LlllilB[

101 102 103 104

Frequency, Hz/sec

0.8 ............

g_ 0.4

:_ 0.0 / I t lllllll I t llllll] "N.-_IJlJl

101 102 103 104

Frequency, Hz/sec

o 0; "_ -100-- "o -100

-2oo - ¢n -200

e- e-

-.300 I I t lilt1] I I I llll_l I _ I_lJl_[ o. -300-- I L I_lJl_l I l=lJl

101 102 103 104 101 102 103 104

Frequency, Hz/sec Frequency, Hz/sec

Figure 1.--Bode plots of desired and neural network models in frequency domain showing averaged transfer function. ( a ) Simple,

second-order linear reference model (rap = 1.1 and w c = I000 Hz). (b) Trained closed-loop system.

0.02

- 0.01 tO

8 o.oo

12.

-0.01

0.00

-- Neural net controller prediction

+ Magnetic bearing response

4

0

-4

.=

o

o

-8

, , i t , -12 ' J ' ' '

0.02 0.04 0.06 0.08 0.10 0.00 0.02 0.04 0.06 0.08 0.10

Time, sec Time, sec

Figure 2.--Desired and actual trained, closed-loop system response for simple, second-order linear reference model in time

domain. Overshoot, P_e 4.3 percent; settling time, Ts, 0.0001 sec; initial rotor position, x o, 0.011 in.: initial rotor velocio; Xce 6

in./sec; and reference position of rotor, rf, -0.005 in.

NASA/TM-- 1999-206966 27
Damping Experiment of Spinning

Component Plates With Embedded

Viscoelastic Material

One way to increase reliability and durability of gas

turbine engine blades is to reduce blade vibration. It

is well known that vibration can be reduced by adding

damping to metal and composite blade-disk systems.

Toward this end, NASA Glenn and the University of

California at San Diego jointly investigated the use

of Integral viscoelastic damping to reduce the

vibration of rotating composite fan blades. The

objectives of the experiment were to verify the

structural integrity of composite plates with

viscoelastic material patches embedded between

composite layers while the plates were under large

steady forces from spinning and to measure the

damping and natural frequency variation with

rotational speed.

Experimental data were obtained from in-vacuum

vibration spin experiments of flat and twisted graphite

composite plates damped by embedding 3M ISD 113

viscoelastic material patches between the composite

layers. Figure 1 shows the rotor installation In the

spin facility. The rotor has a tip diameter of 792 mm;

the plates have an aspect ratio of 3 and a chord of 76

mm. Damping was calculated from measured transfer

functions of blade base acceleration to blade strain.

Damping was repeatable, and there were no failures

or delaminations of the plates. This is significant since

3M ISD 113 has a low creep modulus at room

temperature and the plates had centrifugal loads of

up to 28 000 g at the tip. There was a great deal of

centrifugal stiffening of the plates, which caused a

significant drop in the damping ratios; however, the

viscoelastic material damping remained nearly

constant. Even though the damping ratios decreased,

they were always greater than 2 times the damping

ratios of undamped control plates. Real fan blades,

Figure 1.--Viscoelastic damped composite plates in the NASA Dynamic Spin FaciliO:

28 NASA/TM--1999-206966
though,havesmallerincreasesinnaturalfrequencies

withrotationalspeed, and therefore, the decrease in

fan blade damping ratio should be smaller than that

measured in this experiment.

From the results, we concluded that the presence

of centrifugal forces, which are well known to

increase blade bending stiffness and corresponding

natural frequencies, will decrease the damping

ratios. This occurs because as the blade stiffens

the corresponding modal strain levels and strain

energy in the damping material decrease, thus

decreasing the modal damping ratios. To further

improve damping, new designs should consider

ways to increase the strain energy level in the

viscoelastic material; for example, use a stiffer

viscoelastic damping material than was used here.

This study not only reveals the potential of integral

viscoelastic material damping in composite fan

blades, but it also illustrates that there are technical

challenges that still must be overcome before such

damping can be effectively used as a design option.

Glenn contact: Oral Mehmed, (216) 433-6036

Improved Dual Clearance Squeeze Film

Damper for High Loads

Squeeze film dampers are widely used to control

vibration in high-speed rotating machinery.

However, when imbalance occurs at values

appreciably above the design value (e.g., due to

turbine blade loss), a conventional squeeze film will

be overloaded and, thus, no longer effective in

controlling vibration amplitudes and bearing forces.

The dual clearance damper (fig. 1), characterized

by two oil films separated by a sleeve, overcomes

this problem. The sleeve is fixed in place for normal

operation, and the thin inner damper film (typically

about 0.1 mm) provides all of the damping. In the

event of rotor blade loss or some other occurrence

that increases the imbalance significantly, the sleeve

is released by shearing of the pins holding it (see

fig. 1) and the outer damper film becomes active.

The two films then operate in series; that is, the

bearing load is transmitted first through the inner

film and then through the sleeve and outer film to

the machine structure. The outer film has a larger

clearance than the inner Film to accommodate the

larger vibration amplitude necessarily accompanying

the higher imbalance. Sleeve mass, which was not

accounted for in previous analyses, was found to be

detrimental to damper performance. Inclusion of a

spring support for the sleeve mitigated the effect of

sleeve mass and markedly improved performance of

the dual clearance damper.

A dual clearance damper was designed for a fivemass flexible rotor system. The rotor, 62 cm long

with a mass of 5.7 kg, was assumed to have an

imbalance of 290 g-cm distributed over the rotor

length. For all cases in which sleeve mass was

accounted for, the mass was taken as 360 g--about

what it would be for aluminum.

For this rotor system, vibration amplitude is largest

at midspan. In figure 2, midspan amplitude is

plotted over a range of speeds for several conditions.

The solid line represents the previously analyzed

case, neglecting sleeve mass; its vibration is very

close to that obtained with an ideal damper whose

stiffness and damping are independent of amplitude

and speed. When the sleeve mass was included,

but not the spring support, vibration was somewhat

lower at low speeds, but rose considerably higher

at the rigid-support critical speed of 8280 rpm. Next,

three spring stiffnesses were investigated. A stiffness

of 0.18 MN/m (the dot-dash line) produced the

lowest amplitudes overall; the peak vibration was

not much higher than for the massless sleeve case.

Higher spring stiffnesses resulted in progressively

higher peak amplitudes.

The dual clearance damper, with proper spring

support of the sleeve, holds the promise of providing

good rotordynamic control of flexible shaft systems

at both normal and extraordinarily large imbalances.

Glenn contact: David Fleming, (216) 433-6013

Sleeve

_-- Shear pin

Ball bearing

Ball bearing

housing and

damper joumal

Inner damper film

Outer damper film

Figure l.---Dual clearance squeeze fibn damper

NASA/TM--1999-206966 29
E

E

G

E

101

100

10-1

10-2

2

5--

_..',- ,,-2' f ...... ,- ,/'

_ ,-Y

/... '.\

S Sleeve Spring

mass, stiffness,

m, k,

g MN/m

0 0

360 0

360 0.18

360 0.35

360 0.53

i I I E I I

4 6 8 10 12 14xl 0 3

Shaft speed, rpm

Figure 2.--Rotor midspan amplitude.

User's Guide for MSAP2D, an Euler Flutter and Forced-Response Analysis Code

for Multistage Turbomachinery

Most of the current aeroelastic analyses for

turbomachines analyze only one blade row (i.e., an

isolated blade row). The multistage effects, those due

to the presence of other blade rows, have been

neglected or included in an approximate manner.

Prediction of flutter and forced-response

characteristics of turbomachinery blades requires

that the multibladed-row turbomachine be accurately

analyzed as a single unit. Towards this goal, an

aeroelastic program, MSAP2D, was developed. The

program can be used to analyze unsteady

aerodynamic and aeroelastic characteristics of

multiblade rows, such as compressor or turbine

stages. In addition, it can be used for an isolated

blade row as well.

In this program, the unsteady aerodynamic forces

are obtained by using a finite volume method with

a combination of flux vector splitting and flux

difference splitting to solve two-dimensional Euler

equations. 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

passage-centered H-grids. The interface boundary

between the blade rows is handled as an additional

fluid boundary; it is updated explicitly after the

interior of the computational domain is updated.

The structural model is a typical section with

bending and torsional degrees of freedom for each

blade in the blade row and for all blade rows. The

aeroelastic equations are solved in the time domain

by using Newmark's method. This program has been

calibrated for several examples.

We have written a user's guide to help the user

prepare the input data file required by the MSAP2D

code. The user's guide completely describes the

input data, and gives the input and output

information for three examples from published

papers. References directing the user to detailed

explanations of the aerodynamic analysis, the

numerical algorithms, and the aeroelastic analysis

are provided in the manual. A job control file for

executing the program on the Cray-YMP computers

is also included.

Glenn contacts: Tondapu S.R. Reddy, (216) 433-6083;

Rakesh Srivastava, (216) 433-6045; and Oral Mehmed, (216)

433-6036

30 NASA/TM--1999-206966
User's Manual for ASTROP2, Version 2.0,

a Program for Aeroelastic Stability

Analysis of Multiblade Structures

Flutter is an aeroelastic instability that must be

considered in designing a safe aircraft engine.

Flutter in fan, compressor, or turbine blading can

cause delays in the development of new engines,

and a consequent increase in developmental costs.

ASTROP2 (aeroelastic stability and _response of

12ropulsion systems--two-dimensional analysis) is

an aeroelastic code that can perform flutter analysis

of rotating propfan, compressor, and turbine blades.

It was developed as part of a project aimed at

developing advanced propellers for aircraft capable

of cruising at 0.8 the speed of sound.

To formulate the aeroelastic eigenvalue equations,

the code uses strip theory to combine the unsteady

aerodynamic forces due to vibrating blades with the

blade structural characteristics. The stability

information (flutter frequency and damping) is

inferred from the eigenvalues obtained from these

equations. In ASTROP2, the unsteady aerodynamic

forces are obtained from two-dimensional cascade

theories, and the structural characteristics are

obtained from a three-dimensional structural

model. An interface with NASTRAN is included in

the program; however, the interface can be modified

to suit any structural program. The program

requires minimal computational resources, and it

can be used in a desktop workstation setting.

In version 2.0, flutter analysis is performed by two

separate programs: (1) 2DSTRIP, a program that

calculates the structural dynamic information at

selected stations, called strips; and (2) 2DASTROP,

a program that calculates unsteady aerodynamic

force coefficients and the aeroelastic stability. This

allows users to check their structural model and

aerodynamic model separately. In Version 1.0 of

ASTROP2, these two steps were performed in a

single program. Other improvements to version 2.0

include an option to account for counterrotation,

improved numerical integration, accommodation for

nonuniform inflow distribution, and an interative

flutter calculation for frequency convergence. The

ASTROP2 program is written in FORTRAN 77; the

current version (2.0) has been run on a Cray

computer and a Silicon Graphics Indigo2

workstation. It is now available through COSMIC.

The user's manual for version 2 of ASTROP2

completely describes the input data required and

gives the input and output for a propfan that

fluttered during wind tunnel testing. References

directing the user to detailed explanations of the

aerodynamic analysis, the numerical algorithms,

and the aeroelastic analysis are provided in the

manual. A job control file for executing the program

on the Cray-YMP is also given.

Glenn contacts: Tondapu S.R. Reddy, (216) 433-6083; and

John M. Lucero, (216) 433-2684

Using an Euler Aeroelastic Solver to Analyze Gust and Structural Response of a

Two-Dimensional Cascade

The forced response from a nonuniform, unsteady

gust caused by upstream blade rows, and

aeroelastic instability are two areas that must be

considered in the safe design of turbomachines. To

our knowledge, there is no code in the open

literature that includes gust response calculation,

rotor-stator interaction, and structural (aeroelastic)

response calculation in a single program. To fill this

void, the multistage aeroelastic analysis program

MSAP2D was modified to include gust response

analysis. This was done by prescribing the unsteady

inflow conditions at the inlet and using onedimensional nonreflecting boundary conditions.

Then these modifications were validated by

comparing the predicted unsteady pressure and

phase angles with measured data.

The unsteady aerodynamic forces are obtained by

using a finite volume method with a combination

of flux vector splitting and flux difference splitting

to solve two-dimensional Euler equations. 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 degrees of

freedom for each blade in the cascade. The

aeroelastic equations are solved in the time domain

by using Newmark's method.

Figure 1 plots the amplitude and phase of the

unsteady pressure coefficient on the lower surface

of the turbine airfoil as a function of the chord

fraction. The Mach number is 0.27, and the flow

enters axially and turns about 40 ° . The steady inflow

angle is 2.98 ° , and the reduced frequency of

oscillation based on the chord is 10.0. The

NASA/TM--1999-206966 31
amplitudes of the unsteady pressurescompare

reasonablywell with measureddata. The phase

comparisons are fair in the regions where

amplitudesarehigher,but arepoorin the areaof

lower amplitudes. However,predictions for the

uppersurfaceandforaMachnumberof0.2showed

poor to fair correlation;thus, this needsfurther

investigation.

Glenn contacts: Tondapu S.R. Reddy, (216) 433-6083;

Rakesh Srivastava. (216) 433-6045; and Oral Mchmed, (216)

433-6036

.. 0.20

._.

o

o

o 0.15

,_ O.lO

c

"6 o.o5

0.00

0.0

F . Experimental data

. ! i I

0.2 0.4 0.6 0.8 1.0

Fraction of chord

200

150

100

i

e_

i o

m

• Experimental data

Euler results

t,- -50

_ -lOO

e-

-150

0.0 0.2

I I I

0.4 0.6 0.8

Fraction of chord

Figure 1.--Characteristics of unsteady pressure coefficient on lower surface as a function of

chord fraction. (a) Amplitude. (b) Phase angle, (Mach number = 0.27, reduced frequency

based on chord = lO.O: angle of attack = 2.98 °. gap-to-chord ratio = O. 9744, phase angle =

0.0%

32 NASA/TM--1999-206966
Flutter Analysis of Ducted Rotors

An aeroelastic analysis procedure (DuctE3D) has

been developed to investigate the flutter in

turbomachines. DuctE3D determines the unsteady

airloads on vibrating blades by solving threedimensional unsteady Euler equations and obtains

the blade structural properties from a threedimensional finite-element model. The duct is

assumed to be infinitely long and structurally rigid.

Any number of structural modes of vibration of the

blades can be included in the analysis. The

aeroelastic equations are formulated in normal

mode and are then solved for flutter by using a

frequency domain analysis technique for all possible

interblade phase angles. For verification purpose,

the analysis was applied to a ducted fan

configuration of a propfan enclosed within an

infinitely long and structurally rigid duct.

The three-dimensional Euler equations are solved

with an implicit-explicit hybrid scheme. Since only

two matrix inversions are required, as opposed to

three for fully implicit schemes, this scheme saves

large amounts of CPU time and requires much less

memory. The structural analysis is carried out by

using the normal mode approach and assuming the

hub and duct to be rigid. The flutter analysis is carried

out by moving one of the blades from steady state in

a prescribed pulse motion and obtaining the unsteady

aerodynamic forces on all the blades. These forces

1.0xl 0-.4

are then superposed, and the roots of the aeroelastic

equation are obtained by solving an eigenvalue

problem. The roots provide a measure of aerodynamic

damping at the combined aeroelastic frequency.

Negative aerodynamic damping would imply that the

blade will extract energy from the surrounding air

and, thus, indicates instability. Any number of normal

modes can be included in the analysis. A single run

is sufficient to provide information about the blade

aeroelastic characteristics for the particular flight

condition for each mode of interest.

The analysis was applied to an eight-bladed, ducted

fan, the SR3C-X2 propfan enclosed in a rigid

cylindrical duct. This propfan, in the unducted

configuration, showed flutter in the wind tunnel

tests at NASA Glenn at a freestream Mach number

M = 0.6 for the eight-bladed configuration operating

at an advance ratio J = 3.55 and setting angle

fl = 61.2 °. The analysis was carried out at M = 0.5

by using the first three normal modes for a tip gap

of 0.4 percent of rotor radius. First, the analysis

was used to obtain the unsteady generalized forces

on all the blades, and then a post-processing

program was used to solve the eigenvalue problem

that provided the aeroelastic characteristics.

Figures 1 and 2 show the variation of generalized

forces F 11 with time. Observing the generalized

forces, we found that their amplitudes fall off rapidly

with increasing distance from the reference blade

(the blade being moved). Thus, the argument was

"O

E

0.0

-1.0

-2.0

-3.0

--4.0

• --_: ......

I

I

t

II

I

I

I I

..... Reference blade

+1

........... +2

-1

. | | | ! ! i

0 1 2 3 4 5 6 7

Nondimensional time

Figure/.--Time histo O' of generalized forces for t/re eight-bladed fan operating at M = 0.5, J = 3.55, and

fl = 61.2 °.

NASA/TM--1999-206966 33
o

.N

¢-

o

4.0xl 0-5

+1

3.0 ..... +3

2.0

| I | | ! ! J

1.0

0.0

-1.0

0 1 2 3 4 5 6 7

Nondimensional time

Figure !--Plus signs refer to blade location clockwise from reference blade. Minus signs refer to blade

location counterclockwise from reference blade, (while standing infront of rotor and looking downstream).

(a) Blades closest to reference blade. (b) Blades furthest from reference blade.

325

320

_. 315

==31o

"e 305

E 300

Interblade

phase angle,

- deg

// \_90 ---e.-- 8 blade passages

0 ///_0_., --_-- 4 blade passages

1

,,'_ _ 315

270 IL_

&

225

295 l I I I I

80 1O0 120 140 160 180

Real eigenvalue (damping), Hz

Figure 3.--Eigenvalue plots for analysis of eight- and four-bladed rotors in the eight-bladed fan

operating at M = 0.5, J = 3.55, and fl = 61.2 °.

34 NASA/TM--1999-206966
madethat if only the blades in close proximity to

the reference blades are included in the analysis,

say four blades, the characteristics of an eightbladed rotor should be able to be calculated within

reasonable accuracy. This could potentially reduce

the computational costs by at least 50 percent. In

order to test this, the generalized forces were

calculated for the eight-bladed rotor by using only

four blade passages and applying the periodicity

across the four blade passages. The resultant

eigenvalue plots for both the eight-bladed and fourbladed rotor analyses are shown in figure 3. At this

stage the results are not very encouraging. Even

though the analysis of the four-bladed rotor

accurately predicts the interblade phase angles for

which it is valid, the other interblade phase angles

are poorly predicted. We feel that this method

requires more investigation.

Glenn contacts: Rakesh Srivastava, (216) 433-6045; and

Tondapu S.R. Reddy, (216) 433-6083

element analysis code, such as NASTRAN, is used

to calculate the mode shapes and frequencies.

To determine aeroelastic instability (flutter), a workper-cycle approach is used. That is, the motion of

the blade is prescribed to be a harmonic vibration in

a specified in-vacuum normal mode. The

aerodynamic forces acting on the vibrating blade and

the work done by these forces on the vibrating blade

during a cycle of vibration are calculated. If positive

work is being done on the blade by the aerodynamic

forces, the blade is dynamically unstable, since energy

will be extracted from the flow, and an increase in

the amplitude of oscillation of the blade will result.

Initial calculations have been done for a

configuration representative of the Energy Efficient

Engine fan rotor. Figure 1 shows the work-per-cycle

after each cycle of vibration, which can be seen to

vary only slightly after the fourth cycle of blade

vibration. The negative sign of the converged workper-cycle shows that the fan blade is dynamically

stable and will not flutter.

TURBO-AE: An Aeroelastic Code for

Propulsion Applications

NASA's Advanced Subsonic Technology (AST)

Program seeks to develop new technologies to

increase the fuel efficiency of commercial aircraft

engines, improve the safety of engine operation, and

reduce emissions and engine noise. With the

development of new designs of ducted fans,

compressors, and turbines to achieve these goals,

a basic aeroelastic requirement is that there should

be no flutter or high resonant blade stresses in the

operating regime. This, in turn, makes necessary

an accurate prediction/analysis code to verify the

aeroelastic soundness of the design. Such a threedimensional viscous propulsion aeroelastic code,

named TURBO-AE, is being developed at the NASA

Glenn Research Center.

TURBO-AE is based on a three-dimensional

unsteady aerodynamic Euler/Navier-Stokes

turbomachinery code TURBO, which was developed

under a grant from NASA Glenn. TURBO can model

viscous flow effects that play an important role in

certain aeroelastic problems, such as flutter with

flow separation, or stall flutter, and flutter in the

presence of shock and boundary-layer interaction.

In-vacuum free-vibration modes are used to model

the structural dynamics of the blade. Then, a finiteThe TURBO-AE code will provide a useful aeroelastic

prediction/analysis capability for engine

manufacturers. It will reduce design cycle times by

allowing new blade designs to be verified for

aeroelastic soundness before the blades are built

and tested. This prediction capability will make it

possible to build thinner, lighter, and faster rotating

blades without encountering aeroelastic problems

like stall flutter and high-cycle fatigue due to forced

vibrations.

Glenn contacts: Milind A. Bakhle, (216) 433-6037; and

George Stefko, (216) 433-3920

_¢

o

o

o

10--

0

-10 --

-20

Unstable

Stable

I I t 1 I

0 2 4 6 8 10

Vibration cycle

Figure l.--Aeroelastic work-per-cycle for blade vibrating infirst mode.

NASA/TM--1999-206966 35
Using an Integrated Fiber-Optic Light

Probe to Measure Static Deflections in

Rotating Turbomachinery

An integrated fiber-optic light probe system for

monitoring blade-tip deflections, vibrations, and to

some extent, changes in blade-tip clearances in a

turbomachinery fan or a compressor rotor has been

designed, fabricated, and tested. The system is

composed of a set of integrated fiber-optic light

probes that are positioned to detect the passing

blade tip at the leading and the trailing edges. Each

probe consists of an integrated fiber-optic

transmitting channel and numerical aperture

receiving fibers, all mounted in the same cylindrical

housing. This configuration permits both

nonsynchronous blade vibrations and steady-state

blade deflections to be measured by using the timing

information provided by each light probe. With

integrated fiber-optic technology, a spatial

resolution of 50 _m is possible while keeping the

outer diameter below 2.5 mm. In addition, one fiber

sensor can monitor changes in blade-tip clearance

on the order of 10 _m. These probes were evaluated

by making measurements in a single-stage

compressor facility and an advanced fan rig in the

NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel.

Glenn contact: Anatole Kurkov, (216) 433-5695

Feasibility of Using Neural Network

Models for Accelerated Testing of

Mechanical Systems

Verification testing is an important aspect of the

design process for mechanical systems. Full-scale,

full-length life testing is typically used to qualify

any new component for use in space. However, as

the required life specification is increased, fulllength life tests become more costly and also

lengthen development time. In fact, this type of

testing becomes prohibitive if the mission life

exceeds 5 years, primarily because of the high cost

of and slow turnaround time for new technology.

As a result, accelerated testing techniques are

needed to reduce the time required for testing

mechanical components.

In current accelerated testing methods, typically the

speeds, loads, or temperatures are increased in

order to simulate a high cycle life in a short period

of time. However, two significant drawbacks exist

with this technique. The first is that it is often not

clear what the accelerated factor is when the

operating conditions are modified. Second, if the

conditions are significantly changed (on the scale

of an order of magnitude or more), the component

Is forced to operate out of its design regime. When

the system is operated under such conditions, the

material and mechanical systems' parameters are

often exceeded, and the test is rendered

meaningless.

We believe that neural network systems may be able

to model the operation of a mechanical component.

If so, these neural network models could simulate

long-term mechanical testing by using data from a

short-term test. This combination of computer

modeling and short-term mechanical testing could

then be used to verify the reliability of mechanical

systems, thereby eliminating the costs associated

with long-term testing. Neural network models could

also enable designers to predict the performance of

mechanical systems at the conceptual design stage

by entering the critical parameters as input when

running the model.

The purpose of this study was to assess the potential

of using neural networks for predicting the

performance and life of mechanical systems. To do

this, we generated a neural network system to model

the wear as obtained from three accelerated testing

devices: (1) a pin-on-disk tribometer, (2) a linecontact rub-shoe tribometer, and (3) a four-ball

tribometer. Critical parameters such as load, speed,

oil viscosity, temperature, sliding distance, friction

coefficient, wear, and material properties were used

to produce models for each tribometer.

The study showed that neural networks were able

to model these simple tribological systems, thereby

illustrating the feasibility of using neural networks

to perform accelerated life testing on more

complicated mechanical systems (e. g., bearings,

etc.). Figure 1 compares actual wear data generated

on a rub-shoe tribometer to data that were

generated from a neural network. As the figure

illustrates, the correlation is very good.

Neural networks can also be used to predict input

variables for conditions that have not been run

experimentally. Figure 2 is a neural-networkgenerated three-dimensional plot of wear rate (pinon-disk tribometer) as a function of sliding distance

and sliding speed. The figure depicts wear rate

36 NASA/TM--1999-206966
valuesthat would occurat differentdistanceand

speedvalues.

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

20xl 0 -3

18 --

16 --

6--

14

¢O

E

-_ 12

E

_= 10

o

>

8

O Actual

O Network

4_ !.

0 5 10 15 20

Data point number

I

25

Figure 1.--Comparison of previously unknown rub-shoe wear

volumes (actual data) to that of neural network model

approximation.

%_. 4o_ _ _,1_ o's°

3O

•,o_, o'_7oo.oso._o _,_,

o_ • SP

Figure 2.--Neural network-generated three-dimensional plots

illustrating the relationship between speed, load, and pin-ondisk wear rate.

NASA/TM--1999-206966 37
Acoustics

Scale-Model Acoustic Liner and Low-TipSpeed Turbofan Testing

In support of the Advanced Subsonic Technology

{AST) project, a cooperative NASA Glenn/Pratt &

Whitney test was initiated in October of 1996 to

evaluate advanced fan-noise-reduction technologies.

The test was conducted in NASA Glenn's 9- by 15-Foot

I_w-Speed Wind Tunnel (LSWT), which is configured

for acoustic testing. The existing Glenn drive rig was

used to drive 22-in. fan models for various types of

testing. In addition to steady-state aerodynamic data,

unsteady aerodynamic data, and structural data, two

types of acoustic test data were taken. Far-field

acoustic data (fig. 1) and internal duct rotating-rake

microphone data were obtained. The rotating-rake

data were used to measure acoustic modes present

in the duct, especially those due to the interaction

of the fan wakes with the stators. The test was

divided into two :primary segments. The first

evaluated the acoustic effectiveness of advanced

acoustic liners installed in the inner and outer fan

duct. Figure 2 shows the location of these liners in

the 22-in. model nacelle duct. Several advanced

liners were tested and compared to a baseline that

was set (by the AST Project), to 1992 liner

technology. During the second test segment, a

Figure 1.--NASA/Pratt & Whitn©' Fan 1 in acoustic configuration.

38 NASA/TM--1999-206966
id-acoustic liner

I i=_le acoustic liners

____ _'_ _ _, ,v - -

' ...... ' ....L i]i

Figure Z--Location of advanced liners as tested b7the 22-in. mode/nacelle.

highly loaded, low-tip-speed fan (called P&W Fan 2),

was tested for aerodynamic performance, operability

(which also relates to airplane safety), and acoustic

performance.

(1) Concepts validate to improve nacelle duct

treatment effectiveness by 25 percent relative to

1992 technology. (Target completion date,

January 1, 1997)

(2) Validate technology to reduce engine noise

by 6 dB relative to 1992 technology.

Acoustic and aerodynamic performance testing was

completed at the end of May 1997 for the advanced

liners and also for P&W Fan 2. The data has been

shared with the AST community and is being

evaluated to determine the noise-reduction

effectiveness advanced acoustic liners and the lower

tip speed, P&W Fan 2 relative to P&W Fan 1.

Glenn contacts: Brian Fire, (216) 433-3892; and David

Elliott, (216) 433-8711

Acoustic Barrier Facilitates Inlet Noise

Measurements for Aft-Dominated Fans

Noise from modern high-bypass-ratio subsonic

turbofans tends to be aft-dominated. That is, the

highest levels of flyover noise radiate from the fan

exit. Therefore, to measure the fan inlet sound

radiation without contamination from the aft

radiation, we had to selectively suppress the aft

noise. To do this, we used an acoustic barrier in

the NASA Glenn 9- by 15-Foot Low Speed Wind

Tunnel (9xl 5 LSWT), which effectively isolated the

inlet noise field on a model of an advanced turbofan.

This proof-of-concept test was performed on a model

turbofan manufactured for NASA Glenn by the

Allison Engine Company as part of the Advanced

Subsonic Technology (AST) Program. The barrier

was of wood frame construction, 8 cm thick with,

typically, 0.64-cm tempered fiberboard skins. It was

constructed in sections that were joined upon

installation, and it was mounted on tracks on the

tunnel floor and ceiling at a sideline distance of

15 cm from the fan nacelle. The barrier sections

extended from floor to ceiling and had an axial

length of 61 cm. An elliptical leading edge was faired

into the upstream barrier section, where nominally

full-height by 46-cm-axial-length acoustic treatment

was installed on the fan side of the barrier just

downstream of the leading edge. This acoustic

treatment consisted of a bulk absorber with a

perforated metal skin. Tests were made with the

barrier leading edge at the fan inlet highlight plane

and at a distance 15 cm further aft. The barrier

extended downstream virtually to the end of the

treated tunnel test section.

Figure 1, a photograph of the Allison fan in the 9x 15

I_SWT, shows the acoustic barrier in the upstream

position. The translating microphone, which was

used to collect the acoustic data, is on the 224-cm

sideline. Figure 2 is a directivity plot showing the

aft fan noise suppression due to the barrier along

the 224-cm sideline. These data are for one-third

octave band levels at 6300 Hz, which corresponds

to the second harmonic of the fan rotor-stator

interaction tone. Results are shown for the fan

without the barrier and for the fan with the barrier

leading edge at the two different axial test locations.

Aerodynamic data indicate that the barrier does not

influence the performance of the fan.

Glenn contacts: Richard P. Woodward, (216) 433-3923; and

Christopher E. Hughes, (216) 433-3924

NASA/TM--1999-206966 39
Figure l._Allison fan installed in Lewis" 9x15 LSWT with acoustic barrier in place. Translating microphone probe used to acquire

sideline noise data is visible on the sideline.

Acoustic barrier location _e. _

1_ = None te" "l"

o At inlet highlight . ./¢ -

i I I I , I I I I I , I i I , I , I , I i I i I

20 40 60 80 100 120 140

Figure 2.---Sound pressure level direetivities for fan 2 blade passage frequency tone showing acoustic shielding effect of the barrier wall.

40 NASA/TM--1999-206966
Jet Noise Prediction for Internal Exhaust

Mixer Nozzles

At NASA Glenn, advanced exhaust systems designed

to reduce jet noise from low-bypass-ratio turbofan

engines are being studied as part of NASA's Advanced

Subsonic Technology Program. Our efforts are

directed at developing engine components that will

reduce acoustic emissions and not compromise

engine performance. Thus, experiments were

conducted to map aeroacoustic features of internal

exhaust gas mixers. Test configurations included 12-

and 20-lobe mixers, as well as a reference splitter

nozzle at one-seventh scale. At the same time,

computational tools were developed to help designers

with the parametric study and development of new

generation mixers. Aeroacoustic measurements of

these nozzles will serve as a data base for developing

improvements to existing aerodynamic and acoustic

prediction capabilities.

A computational-based acoustic code that predicts

turbulence-generated noise on the basis of a

Reynolds-averaged Navier-Stokes solution was

developed. Flow predictions were made with the

NPARC code and an appropriate two-equation k-¢

turbulence model.

In acoustic calculations, we assumed noise is

dominated by small-scale turbulence, and that noise

sources are those described by Lilley's equation,

namely self and shear noise; these were modeled to

account for anisotropy of turbulence. Figure 1

shows a 12-lobe mixer operating at a primary nozzle

pressure ratio of 2.77. Sample acoustic predictions

only for noise sources external to the mixer exit

plane are shown in figures 2 and 3. Figure 2 portrays

far-field sound pressure level directivity on an arc

of 15.25 m, and figure 3 shows the noise spectra at

0 = 150 ° from the nozzle inlet.

Glenn contact: Abbas Khavaran, (216) 977-1120

_- Hot.core w

Mixed nozzle

exit flow

Figure l.--Geometry of _.'pica/ lobed forced mixer e.rhausr

O_

¢)

-o

e. ...I

o

110

105

100

95

90

85

80

2O

O Experimental data O

Self noise ......... _3

.... Shear noise .... '!.. "-- ,-,,O'Q

.......... Self + shear noise O Q'(5"°_1_"_-_-_

o._-'_....-C_ "

oo0O_-'"

..................a.a /

I '

f

t

40 60 80 100 120 140 160 180

Angle from inlet, deg

Figure Z--Sound pressure level (SPL) directiviO'.

NASA/TM--1999-206966 41
110

m 100

"0

>

--_ 9O

I/)

80

"0

ru_ 70

60 L

101

0 Self noise

Shear noise

Experimental data

.......... Self + shear noise

_,.__.___.o

_.f'f 0

// \\.o

,, /

!" / " _'\

,.,', .... iv/ , , , ..... , , ....... , , _I,\",.

102 103 104

Frequency, Hz

Figure 3.--Spectra at 0 = 130°from nozzle inlet.

42 NASA/TM--1999-206966
Gear Crack Propagation Studies

Designers of gears for helicopter or turboprop power

transmissions have a common goal: reduced weight.

To help meet this goal, some design gears with thin

rims. Rims that are too thin, however, may lead to

bending fatigue problems and cracks. The most

common methods of gear design and analysis are

based on standards published by the American Gear

Manufacturers Association. Among these standards

are rating formulas for gear tooth bending to prevent

crack initiation. These standards take into account

the effect of rim thickness on tooth bending fatigue;

however, they do not indicate the crack propagation

path or the remaining gear life once a crack has

started.

Recently, analytical and experimental studies were

performed to investigate the effect of rim thickness

on gear tooth crack propagation. The goal was to

determine, for various rim thicknesses, whether

cracks grow through gear teeth or through gear

rims. A finite-element-based computer program

FRANC (Fracture Analysis Code) simulated gear

tooth crack propagation by using principles of linear

elastic fracture mechanics. Quarter-point,

triangular elements were used to represent the

stress singularity at the crack tip. The program had

an automated crack propagation option that used

an automated remeshing scheme to numerically

simulate crack growth. To determine crack

propagation direction, crack tip stress intensity

factors were estimated. Crack propagation life was

predicted on the basis of the calculated stress

intensity factors by using various fatigue crack

growth models. In addition, gear bending fatigue

tests were performed in a spur gear fatigue rig.

Gears with various backup ratios (rim thickness

divided by tooth height) were tested to validate crack

path predictions and crack propagation life.

For backup ratios of 3.3 and 1.0, the analysis

predicted that cracks would propagate through the

teeth and not the rims. This was validated by the

experiments (fig. l(a)). For a backup ratio of 0.3,

the analysis predicted cracks that would propagate

through the rim, which was also validated by

experiments (fig. l(b)). For a backup ratio of 0.5,

/

II[]]

Figure 1. Experimentally de/ermined crack propagation paths. (a)

Thick-rimmed tooth produced tooth fracture. (b) Thin-rimmed tooth

produced rim fracture.

the experiments produced rim fractures whereas

the analysis showed instability (tooth or rim

fracture), depending on various initial conditions.

The analytical results at the point of transition from

tooth to rim failure should be used cautiously to

produce a safe gear design. A good correlation

between predicted and measured gear crack growth

was achieved when the predictions were based on

the Paris crack growth equation and the concept of

fatigue crack closure. For thin rims, a decrease in

rim thickness caused an increase in both the stress

intensity factor and the compressive cyclic stress

in the gear tooth fillet region. The increase in stress

intensity factor promoted crack growth, whereas the

increase in cyclic compressive stress tended to

NASA/TM-- 1999- 206966 43
!E 25

._15

.==

.'_ 10

g 5

Backup ratio,

mB

0.3 1.0

3.3

0.5

"o

o '_a)

0 = , i I J

0 1 2 3 4 5

Crack length, mm

6x106 Backup ratio,

4 0.3

o 1.0

2

t( 3.3

1 b

0 1 2 3 4 5

Crack length, mm

Figure 2.--Effect of backup ratio on crack propagation. (a)

Mode-I stress intensit)'factors. (b) Crack propagation cycles.

retard crack growth and increase the number of

propagation cycles to failure {see fig. 2). For more

information, see Lewicki and Ballarini (1996a and

b} and Zakrajsek and Lewicki (1996).

Glenn contact: David G. Lewicki, (216) 433-3970

Validated Method to Optimize Load Sharing of Split-Torque Transmissions

Most helicopters in service have a planetary

transmission design. Studies show that some

helicopters would be lighter and more reliable if a

split-torque transmission design were employed.

Historically, U.S. helicopter manufacturers have

been hesitant to use this design because of

difficulties in ensuring balanced load sharing among

the load paths--an essential design consideration.

Recently, a Sikorsky-Boeing team chose to use this

design in the U.S. Army Comanche helicopter, with

Sikorsky Aircraft designing and manufacturing the

transmission. To ensure that such a helicopter

would be safe, NASA and Army researchers studied

the load-sharing problem. A theory was proposed

that balanced load sharing could be achieved by

proper configuration of the geartrain. From this

theory, scientists at the NASA Glenn Research

Center developed a computer code to determine the

optimal configuration. The code calculates how

elastic deformations in the gear teeth, shaft,

housing, and bearings affect the relative load carried

in each load path. Sikorsky Aircraft applied the

information from the new computer code to the

design of the Comanche gearbox.

A prototype split-torque transmission had been

designed and built prior to the development of the

theory and code. Measurements on the prototype's

four final drive gears showed an unbalanced load

distribution, with one gear beyond its allowable

design limit. The Sikorsky theory and the NASA

Glenn code duplicated the load trends, and inhouse

experiments showed similar results. The code was

then applied to the prototype transmission to

determine its optimum design configuration.

Incremental design changes were made to the

transmission, and subsequent testing showed

nearly equal load sharing on the four final drive

gears--with all of them below the design limit (see

fig. 1). Because of this analysis, a significant amount

of development time and money was saved.

Government and Sikorsky personnel have applied

for a patent as coinventors of the validated theory

and computer code. Glenn and the Army continue

to collaborate with Sikorsky Aircraft by sharing

transmission technology that will bring the

Comanche gearbox design to maturity. For more

information, see Krantz (1996) and Krantz and

Delgado (1996).

Glenn contact: Timothy L. Krantz, (216) 433-3580

44 NASA/TM-- 1999-206966
Left _ -

engine

input

Right

engine

input

Allowable design load

Gear Loads

i

I

L--._

i

1 2 3 4

m

1 2

i

m

3 4

Before new theory After new theory

Figure l.--Measurements of loads on Jbur final drive gears prove that new theory ensures

load sharing.

Evaluation of Face Gears for Aerospace

Drive System Applications

Reduction of system weight is the highest priority

for any helicopter drive system. With this in mind,

face gears have been chosen as a candidate gear

mesh component for advanced helicopter drive train

systems. This gear system can transfer power from

a single input pinion to two output gears, thereby

reducing by 50 percent the level of loading applied

to the teeth. In a helicopter drive system this

capability results in a substantial weight savings

as the gears can be designed to have approximately

half the load capacity of standard spiral bevel gears.

Figure 1 shows face gears in a future helicopter

drive system.

In the past, face gears were limited to low-load and

low-speed arrangements. In helicopters they would

be operating well beyond their known operational

envelope. Furthermore, there are no design criteria

for face gears. And their manufacture presents

another problem: Aerospace gears are produced to

very tight tolerances, with the best gear steels being

used for strength and durability, and Computer

Numerically Controlled (CNC) machines often being,

used to manufacture them; however, the

manufacturing process for face gears is not this

modern. Thus, the Defense Advanced Research

Projects Association (DARPA} is currently funding

development of a procedure to final-grind the face

gear members.

Since 1991, NASA Glenn has been experimentally

evaluating the face gear system (see the test

hardware in fig. 2). During this time period, several

variations in the tooth geometry and heat treatment

have enabled the gears to operate at high rotational

speeds and loads (19 100 rpm, 360 hp). The long

life required for these gears, however, has still not

been realized. The hardware tested to date has been

compromised by less than adequate materials and

dated manufacturing technology. As a result,

current aerospace gear materials and modern final

NASA/TM--1999-206966 45
grindingprocessesarebeingusedto manufacture

newfacegears,which areexpectedto betestedby

October 1997. For more information, see

Handschuhet al. (1997).

Glenn contacts: Robert F. Handschuh, (216) 433-3969; and

David G. Lewicki, (216) 433-3970

Rotor shaft

output --_\ Sun gear7

/

//

//

//

L_ Combining _ / _-_Engine spur

gear Face gears __v pinion input

Figure/.--Helicopter transmission with face gear drives.

Figure 2.--Face gear specimen used in test program.

Using Hob Offset to Balance Strength of

Spur Gears

In a gear set with a speed ratio other than unity,

the smaller gear (the pinion) tends to be weaker

than the larger gear, particularly if the pinion has a

small number of teeth. Also, in some cases, tooth

interference from the pinion may cause undercut.

Undercutting further reduces tooth strength.

Gear designers sometimes specify nonstandard

gears, such as one with the tooth addendum on

the pinion increased slightly and the gear addendum

reduced. This adjustment to the tooth addendum

is made by withdrawing the hob (cutting tool)

slightly as the pinion is cut and by advancing the

hob as the gear is cut.

For nonstandard gears operated at standard center

distance, the cutter offsets for the pinion and gear

are equal and opposite. In some designs, especially

those with large module teeth, long addendum

pinion teeth may cause undercutting on the gear.

In these cases, the gears can be operated at a slightly

extended center distance.

The NASA gear dynamics code Dynamic Analysis of

Spur Gear Transmission {DANST) was used to

investigate the effect of various amounts of hob offset

on the dynamic stress to a set of extended center

distance spur gears operated over a wide speed range.

Results were compared with the optimum value

computed from an older static procedure. A contour

diagram (fig. 1) shows how various hob offset values

over the speed range of 1 000 to 15 000 rpm affect

the dynamic tooth stress. At lower speeds (below

5 000 rpm) the dynamic stress is minimized by hob

offsets that are less than the optimum static value,

which is shown as a dashed line. (See the two areas

marked with letter "A" in fig. 1). Above 7 000 rpm,

the dynamic stress is minimized by a greater amount

of hob offset (letter "B" in fig. 1).

Often the goal in gear design is to create a balance

between the tooth strength of the pinion and the

gear. Maximizing pinion tooth strength may

excessively weaken the teeth of the gear. Another

study was done to find the amount of hob offset

required for balanced tooth strength. Figure 2 shows

these results. The gear dynamics code DANST can

be used to determine an optimally balanced design.

For more information, see Linet al. (1997) and Liou

et al. (1995).

Glenn contact: Fred B. Oswald, (216) 433-3957

46 NASA/TM--1999-206966
0.24

0.20

0.16

C

.

4.e

O_

_=

o

0.12 o

¢,.

.2

C

0.08

0.04

B

0.00

1 3 5 7 g 11 13

Rotating speed, rpm

I I

15x103

Figure 1.--Contour plot showing effect of/rob offset and operating speed on pinion dynanlic stress. (Dashed

line indicates optimal static procedure hob offset. )

0.25

• 0.20

i0.15

t-

_- 0.10

._o

._c

ft.

0.05

-- /-- Maximum allowable

/ value (0.213 in.)

-{. .............

value (0.159 in.)

_- Minimum allowable

/

/ value (0.015 in.)

0.00 7 ---_ 1 1 I

0 4 8 12 16xl 03

Rotating speed, rpm

Figure Z--Hob offset value required to balance dynamic tooth strength of

pinion and gear plotted as a function of operating speed.

NASA/TM--1999-206966 47
Brush Seals Tested for Joint Technology Advanced Gas Generator H Engine

Brush seals for the Joint Technology Advanced Gas

Generator (JTAGG) II engine being developed by

AllledSlgnal Engines were tested at the NASA Glenn

Research Center. This engine uses two identical

brush seals in tandem to separate the hot highpressure compressor air from the cooler buffer air

for the number 2 carbon ring seal. At intermediate

rated power, the engine operating conditions for

the brush seal are estimated to be 50 000 rpm

(899 ft/sec), 175-psid air-to-air-pressure

differential, and 1200 °F inlet air temperature. Both

static and performance tests were done on the brush

seals. In addition, they were tested with a radial

offset of 0.003 in. to simulate maneuver conditions

and with a radial runout of 0.003 in. in the rotor to

simulate transient conditions such as traversing

through critical speeds. A 50-hr endurance test was

also conducted. Although the maximum operating

conditions could not be attained simultaneously in

the seal test rig, we concluded that the brush seal

design should be able to meet the air-leakage flow

factor goal of less than 0.004 at maximum operating

conditions.

In post-test inspection, small amounts of wear were

found in the Haynes 214 bristle tips and the lnconel

718 rotor that was coated with Triboglide (EG&G

CM NASA

2.5

10-25-96

Figure l.--Advanced low-hysteresis brush seal.

48 NASA/TM-- 1999-206966
Sealol, Cranston, RI). Triboglide is a chromium

carbide coating with barium calcium fluoride (a

high-temperature solid lubricant) dispersed in the

coating. Overall, the seal and rotor were in excellent

condition; thus, we predict that the seals should

easily meet the JTAGG II engine test requirement

of 100 hr.

We tested both conventional and advanced low

hysteresis design brush seals. An advanced low

hysteresis design brush seal is shown in figure 1.

The bristles are made of Haynes 214, and the side

plates are made of INCO 625 (Inconel). The lowpressure side plate was relieved where it contacts

the bristles, and a thin deflector plate was inserted

between the high-pressure side plate and the

bristles. Initially, this seal exhibited little hysteresis,

but after 50 hr of testing, substantial hysteresis

occurred. However, the leakage flow factor remained

under the 0.004 goal. For more information, see

Arora and Proctor (1997).

Glenn contact: Margaret P, Proctor, (216) 977-7526

Diagnostic Parameters Detect GearTooth Fatigue Crack

Drive-train diagnostics is one of the most significant

areas of research in rotorcraft propulsion. An

investigation of serious rotorcraft accidents

resulting from fatigue failures showed that

32 percent of the failures occurred in engine and

transmission components. Now, Government

aviation authorities are demanding that in the near

future the safety record of civil helicopters match

that of conventional fixed-wing jet aircraft. Such

an accomplishment would require that helicopter

safety improve 30-fold over fixed-wing aircraft.

Practically, this can be accomplished only with the

aid of a highly reliable, online health and usage

monitoring (HUM) system. For rotorcraft

transmissions, a critical element of a reliable HUM

system is the ability to accurately detect gear-tooth

damage as early as possible. A gear-tooth fatigue

crack, if undetected, can propagate into complete

tooth fracture or gear rim fracture, both of which

can result in catastrophic failure of the transmission

system. To assess the performance of a variety of

gear damage detection techniques, gear crack

propagation tests were conducted at the NASA

Glenn Research Center.

The spur gears shown in figure 1 were part of a

series of crack propagation tests conducted on a

spur gear fatigue test rig at NASA Glenn. To initiate

cracks, a notch was placed in the fillet region (loaded

side) on one tooth of each of the test gears. The

notches were located at a radius of 40.49 mm on

the fillet, which is the position of maximum tensile

stress. The spur gear test rig was then used to cause

the spur gear specimen to fail through bending

fatigue. To verify the crack growth rate, crack

detection gauges were installed on each side of the

notched tooth. As the crack propagated, it broke

through each of the nine crack wires that

circumferentially surrounded the fillet area on each

side. In figure 2 crack length is plotted as a function

NASA/TM--1999-206966 49
=j

e.

¢.)

e"

r_

e.

0

O.

lO0 -

I/

80- I/

71

If

60 - / /

/ /

40 - / /

/ /

o I

0.00 0.04 0.08 0.12 0.16

Run time, hr

Figure Z--Crack length as afunction of run time.

o Front side

D Rear side

x Complete tooth

fracture

Z

120 -

Crack length, percent of total

11 15

3.5 5.8 8.1 13 118 100 iil i

100 -

80-

NB4 doubled at 60-

11 percent of total

40 - crack length -

20-

0 ....... _-t---

0.00 0.04 0.08 0.12 0.16

Run time, hr

Figure 3.--Parameter NB4 as a function <_"run time.

of run time for the test gear shown in figure 1. As

can be seen in figure 2, the crack gauges covered

only the first 18 percent of the total crack length

before failure occurred (complete fracture of the

tooth). Although the crack gauges covered only

18 percent of the total crack length, they covered

75 percent'of the time from the start of the crack to

gear or tooth failure. Two regions can be seen in

figure 2: in the first there is moderate crack growth

rate (first 75 percent of propagation time); and in

the second, accelerated crack growth rate, where

the final 82 percent of total crack length occurs

(last 25 percent of crack propagation time).

A number of gear damage detection techniques were

applied to the vibration signal emitted by an

accelerometer mounted on the housing of the gear

test rig. All of the gear fault detection methods

investigated reacted., to the fatigue crack in the region

of accelerated crack growth, prior to complete

fracture. Recall that this is the region where the

last 82 percent of total crack length is achieved in

only 25 percent of the total crack growth time.

Although in this test all of the methods were able

to detect the crack during propagation in this last

region, more advance warning of the potentially

catastrophic crack is needed. Only three of the

methods investigated were able to show a trend that

was increasing beyond the nominal value during

the first phase of crack propagation. Of these three,

the technique called NB4, which was developed at

NASA Glenn, gave the most robust indication of the

fatigue crack. NB4 is a demodulation technique that

uses normalized kurtosis to produce a dimensionless fault-indication parameter with a value of three

under nominal conditions. Figure 3 shows the NB4

parameter (measure of gear faults) detection as a

function of run time for this test. For reference, the

crack length as a percent of total length is shown

along the top of this figure. The NB4 parameter not

only gave a robust indication of the crack in the

accelerated crack growth region, prior to failure,

but also, more importantly, it reacted to the growing

fatigue crack during the moderate crack growth

region. When the crack had advanced to only

i I percent of its total length to failure, NB4

indicated a value double the nominal value. This is

significant in that the earlier a fatigue crack can be

detected in an aircraft transmission, the more time

and options that are available to the operator. For

more information, see Zakrajsek and Lewicki (1996).

Glenn contact: James J. ZakraJsek, (216) 433-3968

50 NASA/TM--1999-206966
Gas Journal Bearing Stability Versus

Amplitude Ratio

A gas journal bearing with a wavy surface was tested

over a range of speeds up to 18 000 rpm to

determine its stability in an unloaded condition as

a function of the wave amplitude. The bearing was

50 mm in diameter, 58 mm long, and had

0.0165 mm of radial clearance. Three waves were

created on the inner surface by deforming the

bearing sleeve. The ratio of the wave amplitude to

the radial clearance (i.e., wave amplitude ratio WAR)

was varied from 0 to 0.3.

Unlike the plain journal bearing (fig. 1(a)), a wave

journal bearing reduces the susceptibility to

subsyncronous frequency whirl (SSFW) motion. A

wave journal bearing (Dimofte, 1995) is a bearing

with a slight, but precise, variation in the circular

profile of the bearing such that a waved profile is

circumscribed on the inner bearing diameter. The

profile has a wave amplitude that is a fraction of

the bearing clearance. Figure 1(b) shows a threewave bearing. Note that the clearance, the wave,

and the wave's amplitude are greatly exaggerated

in figure 1 so that the concept may be visualized.

Experiments showed that for wave amplitude ratios

of 0.187 or larger the bearing ran in a stable mode,

up to a critical speed, after which it began to

experience SSFW motion with a frequency close to

one half the synchronous frequency. The speed

threshold when the SSFW motion began was found

to be a function of the WAR. For example, the

threshold speed was 9 650 rpm for a WAR of 0.187

and 17 210 rpm for a WAR of 0.30. However, for

WAR's smaller than 0.187 (such as 0.024 and

0.157), the bearing became sensitive to SSFW

motion at very low speeds. Moreover, it developed a

supersynchronous whirling motion with a frequency

close to twice the synchronous frequency. A WAR

equal to 0.187 or greater suppressed both the

subsynchronous and supersynchronous whirl

motion and allowed the bearing to run in a very

stable mode.

A stability bearing analysis code was used to predict

the critical mass of an experimental bearing

operating in an unloaded condition. WAR's from zero

(true circular) to 0.35 were used. The results are

plotted in figure 2, where we can see that a true

circular bearing and a wave bearing with a WAR of

0.15 are unstable for any bearing mass. However,

if the WAR is increased, a portion of the predicted

curves lie above the bearing mass line. In these

regions the bearing can run in a stable mode free of

SSFW motion.

The experimentally observed threshold to SSFW

motion for a bearing with a WAR equal to O. 187 or

larger is also plotted in figure 2 as a set of scattered

points located on the horizontal line denoting the

actual bearing mass value of 5.2 kg. Essentially,

this line intersects with the predicted critical mass

curves at the speeds at which a bearing with a mass

of 5.2 kg would begin SSFW motion for the given

WAR values. Figure 2 demonstrates that the

experimental data points are in good agreement with

the predicted values. For more information, see

Dimofte (1995).

Glenn contact: Florin Dimofte, (216) 977-7468

I Load

Speed a 0

eO 1

(a)

Shaft

Bearing _- Bearing

(b)

Figure/.--Journal hearings. (a) True circular (b) Three wave. (Shaft center, O; bearing centel; 01. )

NASA/TM--1999-206966 51
25

2O

15

Predicted critical

mass at WAR of

_7 0.0

[] .15

O .20

O .30

A .35

Observed critical

mass at WAR of

• 0.18

• .22

• .28

• .30

Bearing

mass of

5.2kg

"'-u ....... [_j..-__

0 5 10 15 20 25 30xl 03

Speed, rpm

Figure 2.--Predicted critical mass and experimentally observed threshold of whirling.

NASA Selects 1996 Government Invention of the Year

A patented seal developed for the National Aerospace

Plane project by Dr. Bruce M. Steinetz and Mr. Paul

J. Sirocky (retired) of NASA's Glenn Research

Center, Cleveland, Ohio. won NASA's 1996

Government Invention of the Year award.

The High-Temperature, Flexible, Fiber Preform Seal

has been used for numerous NASA aeronautic

applications since it was patented in 1992. It was

also used by General Electric Aircraft in the joint

NASA/DOD/GE Integrated High Performance

Turbine Engine Technology Program (IHPTET). The

hybrid rope seal was used to seal the perimeter of

advanced nickel-aluminide turbine vane airfoils;

thus the vane could grow relative to the supporting

structure, thereby overcoming the thermal shock

failure experienced with conventional seals. The

high-temperature turbine vane/seals, combined

with several other advanced technologies,

contributed to meeting the program goals of

reducing specific fuel consumption by 20 percent

and increasing engine power-to-weight by

40 percent, in a successful full-scale Joint

Technology Advanced Gas Generator (JTAGG)

engine test.

The invention Is being evaluated by Pratt & Whitney

as a possible replacement for the sealing interfaces

between the large turning vanes and the flowpath

fairing elements of the F 119 engine, which will be

used in the F22 fighter, the country's next

generation premier fighter plane. Because of the

potential for improved durability. Pratt & Whitney

is also evaluating an abrasion-resistant hybrid

version of the seal (ceramic core/superalloy sheath).

And Williams International is evaluating NASA's

fiber preform seal as a means to both seal and

structurally support a high-temperature, uncooled,

engine transition segment for an advanced

52 NASA/TM--1999-206966
demonstratorturbine engine.Temperaturesin this

applicationare in excessof 1000°F--beyondthe

capabilitiesof anyother seal.

In additionGlenn,undera reimbursableSpaceAct

Agreement,is working with Praxair,a major U.S.

producer of industrial gases,to adapt the seal

technologyfor use in the company'sproprietary

high-temperatureindustrial gassystems.

Someof the attributes of the seal are as follows: It

is a flexible seal that can be used at very high

temperatures (2000+ °F) in numerous aerospace

and industrial applications; it can bend around

sharp radii (about equal to the seal's diameter),

conforming to and sealing complex turbine

components; and it exhibits low leakage, retains

resilience after high-temperature cycling, and can

support structural loads. The seal can be braided

of emerging ceramic and/or superalloy fibers to

meet specific application requirements.

Dr. Steinetz, a Senior Research Engineer at Glenn,

is responsible for coordinating NASA's engine seal

development projects. He earned his Ph.D. in

Mechanical and Aerospace Engineering at Case

Western Reserve University, Cleveland, Ohio, in

1991. Currently, Dr. Steinetz leads a research team

that includes Michael Adams and Lawrence Kren,

who helped perform the laboratory tests qualifying

the seal for subsequent engine tests.

Mr. Sirocky was a Senior Project Engineer at Glenn

until 1985, when he retired with 30 years of

Government service. He was working for Sverdrup

Technology, Inc., when the patent for the seal was

issued.

Glenn contact: Bruce M. Steinetz, {216) 433-3302

NASA/TM--1999-206966 53
Arnold, S.M.; and Saleeb, A.F.: On the

ThermodynamicFrameworkofGeneralizedCoupled

Thermoelastic-Viscoplastic-DamageModeling.

NASATM-I05349, 1992.

Arnold, S.M.; Saleeb,A.F.; and Castelli, M.G.:

Specialized Deformation Tests for the

Characterization of a ViscoelastoplasticModel:

Applicationto aTitaniumAlloy.NASATM-106268,

1997.

Arnold, S.M.; Saleeb,A.F.; and Castelli, M.G.:A

Fully Associative, Nonisothermal, Nonlinear

Kinematic,UnifiedViscoplasticModelforTitanium

Alloys.NASATM-106926,1995.

Arnold, S.M.; Saleeb,A.F.; and Castelli, M.G.:A

Fully Associative,Nonlinear Kinematic, Unified

ViscoplasticModelfor Titanium BasedMatrices.

NASATM-106609,1994.

Arora,G.K.;andProctor,M.P.:JTAGGII BrushSeal

Test Results. AIAA Paper 97-2632, (NASA

TM-107448) 1997. (Available online: http://

letrs.lerc.nasa.gov/cgi-bin/LeTRS /

browse, pl? 1997/TM- 107448. html).

Brewer, D., et al.: Measurement of Mechanical

Damage in SiC/SiC Composites Using Vibration,

Acousto-Ultrasonic, and Thermographic NDE.

HITEMP Review, NASA CP-10178, Vol. II, paper

no. 36, 1995, p. 1.

Castelli, M.G.: A Summary of Damage Mechanisms

and Mechanical Property Degradation in Titanium

Matrix Composites Subjected to TMF Loadings.

Thermal-Mechanical Fatigue of Aircraft Engine

Materials, AGARD CP-569, 1996a, pp. 12: 1-12.

Castelli, M.G.: Characterization of Damage

Progression in SCS-6/Timetal 21S [0] 4 Under

Thermomechanical Fatigue Loadings. Life Prediction

Methodology for Titanium Matrix Composites, ASTM

STP-1253, W.S. Johnson, J.M. Larsen, and

B.N. Cox, eds., American Society for Testing and

Materials, Philadelphia, 1996b, pp. 412-431.

Castelli, M.G.; and Bhanu Sankara Rao, K. Cyclic

Deformation of the Superalloy Haynes 188 Under

Isothermal and Thermomechanical Loadings,

Superalloys 1996, The Minerals, Metals & Materials

Society, 1996, pp. 375-382.

Darolia, R., et al.: Design Methodology of

Intermetalllcs. IHPTET F33615-94-C2414, R&D

Status Report no. I 0, 1996.

Dimofte, F.; Wave Journal Bearing With

Compressible Lubricant, Part I: The Wave Bearing

Concept and a Comparison With a Plain Circular

Journal Bearing. STLE Tribology Transactions,

vol. 38, no. 1, 1995, pp. 153-160.

Handschuh, R., et al.: Experimental Evaluation of

Face Gears for Aerospace Drive System Applications.

NASA TM-I07227 (ARI_TR-I 1109), 1997,

Huber, R.D.; and Green, R.E., Jr.: AcoustoUltrasonic Nondestructive Evaluation of Materials

Using Laser Beam Generation and Detection. NASA

CR-185282, 1990.

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

Creep Life of Ceramic Components Using a Finite

Element Based Integrated Design Program (CARES/

Creep). ASME Paper 96-GT-369, 1996.

Janosik, L.A.; Duffy, S.F.: A Viscoplastic

Constitutive Theory for Monolithic Ceramics--I.

ASME J. Eng. Gas Turbines and Power, vol. 120,

no. 1, Jan. 1998, pp. 155-161.

Kautz, H.E.: Non-Contact Determination of

Antisymmetric Plate Wave Velocity in Ceramic

Matrix Composites. NASA TM-107125, 1996.

Krantz, T.L.: A Method to Analyze and Optimize the

Load Sharing of Split Path Transmissions NASA

TM-107201, 1996. (Available online: http://

letrs, lerc. nasa.gov / cgi-bin / LeTRS /

browse.pl? 1996/TM- 107201 .html).

54 NASA/TM--1999-206966
Krantz,T.L.;andDelgado,I.R.:ExperimentalStudy

of Split-Path TransmissionLoad Sharing. NASA

TM-107202,1996.

Lee, H.-J.; and Saravanos, D.A.: Active

Compensationof Thermally Induced Bending and

Twisting in Piezoceramic Composite Plates.

AIAA Paper 96-1277-CP, Apr. 18-19, 1996a,

pp. 120-130.

Lee, H.-J.; and Saravanos, D.A.: Coupled Layerwise

Analysis of Thermopiezoelectric Composite Beams.

AIAAJ., vol. 34, no. 6, June 1996b, pp. 1231-1237.

Lee, H.-J.; and Saravanos, D.A.: Layerwise Finite

Elements for Smart Piezoceramic Composite Plates

in Thermal Environments. AIAA Paper 96-1277

(NASA TM-106990), 1996c.

Lei, J.-F., et al.: Comparison Testings Between a

High Temperature Extensometer and a Resistance

Strain Gage With Multiple Installation Techniques.

Experimental Mechanics, vol. 36, no. 4, Dec. 1996,

pp. 434-439.

Lewicki, D.G.; and Ballarini, R.: Gear-Crack

Propagation Investigation. NASA TM-107147,

1996a.

Lewicki, D.G.; and Ballarini, R.: Effect of Rim

Thickness on Gear Crack Propagation Path. NASA

TM-107229, 1996b. (Available online: http://letrs.

lerc.nasa.gov/cgi-bin/LeTRS/browse.pl? 1996/

TM- 107229. html).

Lin, H.H., et al.: Balancing Dynamic Strength of

Spur Gears Operated at Extended Center Distance,

NASA TM- 107222, 1997. (Available online: http: / /

letrs.lerc, nasa.gov/cgi-bin/LeTRS /

browse.pl? 1997/TM- 107222.html).

Liou, C.-H., et al.: Using Hob Offset to Balance

Dynamic Strengths in Spur Gears. NASA

TM- 106934, 1995.

Lissenden, C.J., et al.: Experimental Determination

of Yield and Flow Surfaces under Axial-Torsional

Loading. Multiaxial Fatigue and Deformation

Testing Techniques, ASTM STP- 1280, S. Kalluri and

P.J. Bonacuse, eds., 1997, pp. 92-112.

Lissenden, C.J., et al.: Verification of Experimental

Techniques for Flow Surface Determination, NASA

TM-107053, 1996. (Available online: http://

letrs.lerc.nasa.gov/cgi-bin/LeTRS/

browse, pl? 1996/TM- 107053. html).

Nemeth, N.N., et al.: Designing Ceramic

Components for Durability. Advanced Ceramic

Matrix Composites--Design Approaches, Testing

and Life Prediction Methods, E.R. Generazio, ed.,

Technomic Publishing Co., Lancaster, PA, 1996a,

pp. 3-16.

Nemeth, N.N., et al.: Durability Evaluation of

Ceramic Components Using CARESLife. J. Eng.

Gas Turb. Power (NASA TM-106475), vol. 118,

no. 1, 1996b, pp. 150-158.

Periera, J.M.; Kautz, H.E.; and Roth, D.J.:

Measurement of Mechanical Damage in Ceramic

Matrix Composites Using Vibration and AcoustoUltrasonic NDE. HITEMP Review. NASA CP- 10146,

Vol. I, 1994, pp. 23-1 to 23-2. (Permission to cite

this material was granted by Carol A. Ginty,

December 1997.)

Roth, D.J.: Single Transducer Ultrasonic Imaging

Method That Eliminates the Effect of Plate

Thickness Variation in the Image. NASA

TM-107184, 1996. (Available from NASA Glenn's

Subsonic Systems Office.)

Roth, D.J., et al.: Commercial Implementation of

NASA-Developed Ultrasonic Imaging Methods via

Technology Transfer, Materials Evaluation, vol. 54,

no. 11, November 1996a, pp. 1305-1309.

Roth, D.J., et al.: Commerical Implementation of

Ultrasonic Velocity Imaging Methods via Cooperative

Agreement Between NASA Glenn Research Center

and Sonix, Inc. NASA TM- 107138, 1996b. (Available

online: http: / /letrs.lerc.nasa.gov / cgi-bin /LeTRS /

browse.pl? 1996/TM- 107138.html).

Salem J.A.; Noebe, R.; and Mandersheid, J.:

Reliability Modeling of Brittle Anistropic Materials.

Proceedings of the 22nd Annual Conference on

Composite Advanced Ceramics Materials and

Structures, Jan. 20-24, Cocoa Beach, FL, 1998.

NASA/TM--1999-206966 55
Salem,J.A.,et al.: ReliabilityAnalysisofUniaxially

Ground Brittle Materials. J. Eng. Gas Turbines

Power(NASATM-I06852), vol. 118,no. 4, 1996,

pp. 863-87I.

Safaeinili,A.; Lobkis,O.I.,andChimenti,D.E.:AirCoupledUltrasonicCharacterization of Composite

Plates. Materials Eval., Oct. 1995, pp. 1186--1197.

Sokolowski, T.M., et al.: Stress and Reliability

Analysis of a Metal-Ceramic Dental Crown. NASA

TM-107178, 1996.

Telesman, J.; Ghosn, L.J.; and DeLuca, D.P.:

Influence of the Failure Mode on Fatigue Crack

Growth Behavior in Single Crystal Superalloys.

Hydrogen Effects in Materials, A.W. Thompson and

N.R. Moody, eds., The Minerals, Metals & Materials

Society, Warrendale, PA, 1996, pp. 943-952.

Verrilli, M.J. and Castelli, M.G., eds.:

thermomechanical Fatigue Behavior of Materials:

Second Volume, ASTM STP-1263. American Society

for Testing and Materials, Philadelphia, 1996.

Ye, L.: On Fatigue Damage Accumulation and

Material Degradation in Composite Materials.

Comp. Sci. Technol., vol. 36, no. 4, 1989,

pp. 339-350.

Zakrajsek, J.J.; and Lewicki, D.G.: Detecting Gear

Tooth Fatigue Cracks in Advance of Complete

Fracture, NASA TM-I07145, 1996. (Available

online: http: / /letrs.lerc.nasa.gov / cgi-bin /LeTRS /

browse, pl? 1996/TM- 107145. html).

56 NASA/TM-- 1999-206966

Form Approved REPORT DOCUMENTATION PAGE OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services. Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway. Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

November 1999

4. TITLE AND SUBTITLE

Structures and Acoustics Division 1996 Annual Report

6. AUTHOR(S)

Cynthia S. Acquaviva

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

3. REPORT TYPE AND DATES COVERED

Technical Memorandum

5. FUNDING NUMBERS

WU-910-31-11--00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-IIII8

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--1999-206966

11. SUPPLEMENTARY NOTES

Responsible person, Cynthia Acquaviva, organization code 5900, (216) 433-3306.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 39 Distribution: Nonstandard

This publication is available from the NASA Center for AeroSpace Information. (301) 621-0390

12b. DIS_HIBUTION CODE

13. ABSTRACT (Maximum 200 words)

The Structures and Acoustics Division of NASA Glenn Research Center is an international leader in rotating structures,

mechanical components, fatigue and fracture, and structural aeroacoustics. Included are disciplines related to life

prediction and reliability, nondestructive evaluation, and mechanical drive systems. Reported are a synopsis of the

work and accomplishments reported by the Division during the 1996 calendar year. A bibliography containing 42

citations is provided.

14. SUBJECT TERMS

Structural mechanics; Structural fatigue; Structural dynamics; Structural integrity;

Mechanical components; Acoustics

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OF REPORT OF THIS PAGE

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OF ABSTRACT

Unclassified

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65

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A04

20, LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-1B

298-102