Mattingly J.D., Heiser W.H., Pratt D.T. Aircraft Engine Design

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APPENDIX M: EXAMPLE MATERIAL PROPERTIES 633

J¢

40 ~ " 1200 F

"-.2

4 ,, 1795 F

~,~

0.1

percent total plastic

..... 0.5

percent total plastic strain

~-~

Rupture

I ] I l I I

tO0

tO00 I0,000

Time, hours

'Nimoni¢ 10S Extrmdled Bar'

2100 F, 4 hrs, AC + 1920 F, 16 hrs.

AC + 1560 F, 16 hi'=, AC

"~ ,¢~.,,~. "'-" ".~... "" "-,,

, ~"~\

, ~::~ 1500 F

-,,\

~'~'~ 1600 F

100,000

Fig. M.15 Creep and rupture curves for Nimonic 105 extruded bar.

Rene' 80 (p = 15.9 slug/ft 3)

This is a cast, precipitation-hardenable, nickel-base superalloy. It has excellent

creep-rupture strength up to 1900°F, combined with good elevated temperature

ductility and superior hot corrosion resistance. The main use of Rene' 80 is in

investment-vacuum-cast turbine stator and rotor airfoils, which are coated for any

jet engine applications.

Using the 1% creep deformation at the 1000 hour criterion, Fig. M.13 reveals

that Rene' 80 has the remarkably high tensile stress capability of about 32,000 psi

at 1600°F and 13,000 psi at 1800°E

Alternating stress considerations are often important in cooled turbine rotor

airfoils, which are usually cantilevered and subject to the aerodynamic buffeting

(as a result of both random turbulence and organized wake passage) that produces

634 AIRCRAFT ENGINE DESIGN

vibratory loads. It is especially important to avoid resonance between a regular

upstream disturbance and a natural frequency of the rotor airfoils.

Thermal stresses can cause low-cycle fatigue failure if the limits given in

Fig. M.14 are reached. This means that cooled airfoils must be designed with the

greatest care so that no local areas of large thermal stress are created at any time

during the normal engine operation. In particular, it can be readily seen that a thin

airfoil "skin" with a large temperature difference between coolant and mainstream

will not last long.

Nimonic 105 (p = 15.5 slug/ft 3)

This is a wrought nickel-cobalt-chromium-base superalloy having creep and

oxidation resistances that make it suitable for service at temperatures up to 1750°E

It is used for turbine rotor airfoils, disks, forgings, ring sections, bolts, and fasteners.

The data of Fig. M.15 show that Nimonic 105 falls between Rene' 80 and

Hastelloy X in tensile strength. Under the same ground rules, one finds a stress of

about 17,000 psi at 1600°F and perhaps 3000 psi at 1800°E No data were given

for the low-cycle fatigue behavior of Nimonic 105.

Reference

IAerospace Structural Metals Handbook, Battelle Memorial Inst., Columbus Lab.,

Columbus, OH, 1984.

Appendix N

Turbine Engine Life Management

N.1 Introduction to the Engine Lifing Process

While performing their intended aerodynamic tasks, the rotating compressor and

turbine parts must also be almost certain to last the design lifetime of the engine. The

turbine engine lifing (pronounced "life"-ing) process is evolutionary and has been

unfolding over the past 50 years. In various degrees it is affected today by many

of the factors depicted in Fig. N. 1. These factors have all evolved through lessons

learned and state-of-the-art changes. Some of the most notable advances that have

been made are in the areas indicated by the shaded blocks in Fig. N. 1. Some

of these include: 1) finite element modeling techniques used in stress analysis;

2) improved materials (strength, temperature capability, life, fracture toughness);

3) engine controls; 4) testing techniques; 5) manufacturing and inspection; 6) sur-

face treatments that instill beneficial residual stresses such as laser shock peening;

7) data recording systems for usage monitoring; and 8) usage tracking systems.

When taken together, all of these advances in the "state of art" have had a very

positive effect on the ability to predict engine component life and on safety.

N.2 Engine Life Management

N.2.1 Lessons Learned

Through many development programs (commercial and military) and customer

engine problems the following summary of generic lessons learned has been

collected:

1) It is dangerous to assume defect-free structures in safety of flight components.

2) Critical parts and potential failure modes must be identified early in devel-

opment and appropriate control measures implemented.

3) Thermal and vibratory environments must be identified early in engine

development.

4) For complex components analytic stresses must be verified by testing.

5) Materials and processes must be adequately characterized (particularly, the

fracture properties).

6) Design stress spectra, component test spectra, and full scale engine test

spectra must be based on the anticipated service usage of the engine.

7) Potential engine/airframe structural interactions must be defined and taken

into account.

8) Closed-loop management procedures must be defined and enforced through

the application of realistic inspection and maintenance requirements, individual

engine tracking procedures, deficiency reporting, and updates in procedures based

on actual usage.

635

636 AIRCRAFT ENGINE DESIGN

Fig. N.I Turbine engine lifing factors.

N.2.2 Life Management Scope

Life management is important because its goal is to ensure the safe operation of

engines subjected to complex usage. This is accomplished through the calculation

of the lives of high-energy rotating and pressurized components, which would not

be contained if they were to fail in an engine. The goal of life management is to

safely maximize the use of critical parts to minimize ownership costs. The cost of

some of these parts, because of the use of high-temperature exotic materials, com-

plex design and manufacture, and rigid quality control, is considerable. Examples

of some of these high-energy components would be rotating disks, spools, airfoils,

spacers, shafts, and seals and high-pressure cases.

Life management requires the selection of one of the lifing concepts that will

be discussed in Sec. N.4, Lifing Concepts, or the lifing concepts may be dictated

by the customer and stated in the Request for Proposal (RFP).

For example, a typical commercial engine flight usage shown in Fig. N.2 would

lean heavily on the commercial safe life management concept. The stress response

of a typical turbine rim from this usage is shown in Fig. N.3. In simple terms, the

maximum stress range at the critical location is used to enter the low cycle fatigue

(LCF) minimum life material curve at the corresponding operating temperature to

give the component minimum life.

In contrast, Fig. N.4 shows engine usage that would be common to a military

fighter aircraft. In this case, to life manage would require combining the safe life

concept with damage tolerance (required by the military customer in the RFP) and

involve a more rigorous thermal, stress, and fatigue and facture analysis to establish

minimum life and safe inspection intervals.

In addition to the effects of usage on engine life, the effects of mission mix and

the effects of an ambient temperature mix and warm and cold starts would also

require an evaluation.

N.2.3 Definitions

N.2.3.1 Damage tolerance approach.

The damage tolerance approach is

a structural philosophy that recognizes that undetected metallurgical defects

APPENDIX N: TURBINE ENGINE LIFE MANAGEMENT 637

16,000--

~14,000--

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.o 8,000-

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6,000-

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GROUND ID E

85 MINS

START UP

SECS

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20 40 60 80 100 120

Time - Minutes

Fig. N.2 Typical commercial mission.

100--

90--

80--

70--

60--

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20-

10-

O-

-10-

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Maximum

Stress

i S

Minimum

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20 40 60 80 100 120

Minutes

Stress

Range

140

Fig. N.3 Typical turbine rim transient stress analysis.

638 AIRCRAFT ENGINE DESIGN

Throttle

Altitude

Math

i,,_" ...... ~

Number

..........

60 80 100

Time - Minutes

Fig. N.4 Typical F-16 flight recorder data.

and/or manufacturing flaws or damage (knicks, voids, tears, laps, burrs, etc.) can

progress to failure. It assumes that damage, in the form of a crack of minimum

detectable size, is present at all locations and should not grow to critical size in

two inspection intervals for the intended usage.

N.2.3.2 Fracture critical components.

Fracture critical components are

those whose failure would result in loss of the aircraft, and for single-engine

aircraft, prevention of sustained flight. This includes all major rotating structures

such as disks, shafts, and rotating seals. For single-engine applications this also

includes blades, bearings, and critical external accessories. Damage tolerance re-

quirements are applied only to fracture critical components and not, in general, to

durability critical components.

N.2.&3 Durability critical components. These are components whose

cracking or failure could cause significant engine structural damage or perfor-

mance/reliability/maintainability problems.

N.2.&4 Hot parts. Hot parts are those parts in the hot gas stream. Usually

the life requirement is one-half that of cold parts life.

N.2.3.5 Coldparts. Cold parts are all parts not in the hot gas stream. This

includes turbine disks, spacers, etc. Usually the life requirement is the same as that

required for the aircraft system in which the engine is to be used.

N.2.3.6 Initial flaw size. Initial flaws are assumed to exist in fracture crit-

ical components. Experience has shown that premature cracking occurs at high

stressed areas, and initial flaw conditions have included both material and man-

ufacturing related quality variations. Damage tolerance uses a sharp crack initial

flaw assumption to characterize these abnormal initial conditions. Assumed initial

embedded flaw sizes are based on the intrinsic material defect distribution or the

nondestructive evaluation (NDE) methods to be used during manufacture. An in-

spection reliability of 90% probability of detection (POD) at the lower bound 95%

APPENDIX N: TURBINE ENGINE LIFE MANAGEMENT 639

confidence level (CL) is required for the assumed initial flaw sizes. The basis for

these flaw sizes is the data on the various nondestructive evaluation methods. 1

N.2.3.7 Residual strength.

Residual strength is defined as the load car-

rying capability of a component at any time during service exposure period con-

sidering the presence of damage and accounting for the growth of damage as a

function of exposure time. The requirement is to provide limit load or maximum

expected service load residual strength capability throughout the service life of the

component. One example of this is normal or expected overspeed caused by control

system tolerance.

N.2.3.8 Inspection interval

The inspection interval for damage tolerance

considerations should be compatible with the overall engine maintenance plan.

It is highly desirable that the inspection interval be equal to the hot part design

service life as this is the expected minimum maintenance interval for the engine

or module. The damage tolerance inspection interval should be contained in the

contract specification or RFP.

N.2.3.9 Flight data recorder.

An in-flight recorder records usage time his-

tory of aircraft and engine, is used to sample usage, and is not required on every

aircraft.

N.2.3.10 Engine life monitor.

This monitor is an in-flight counter that

tracks certain engine parameters for life purposes and is required on every engine.

N.2.3. 11 Creep.

Creep is permanent deformation that occurs as the result

of the prolonged application of stress. The applied stress is below a level that would

be expected to cause plastic deformation. Creep occurs at temperatures that are

high relative to the melting point of the metal.

N.2.3.12 Low cycle fatigue (LCF).

LCF is damage that is related to stress

cycles that lead to component failure in less than 10,000 cycles.

N.2.3.13 High cycle fatigue (HCF).

HCFisdamagethatisrelatedtostress

cycles that lead to component failure in more than one million cycles, caused

primarily by vibration.

N.2.3.14 Stress rupture.

Rupture is failure when a component has been

statically loaded at elevated temperature for a long period.

N.2.3.15 Laser shock peening (LSP).

LSP is a new process for using

compressive residual stress to desensitize airfoil leading edges to foreign object

damage (FOD). This process is available and is in production on some military

engines. The process is so effective that retired blades have had leading edges laser

shock peened and severely damaged, simulating FOD in a critical area, and when

tested have shown remarkable high cycle fatigue strength and life that was even

higher than new undamaged blades. The process uses a high-energy pulsed split

laser beam to create a high amplitude shock wave on both sides of the blade or

vane simultaneously, as shown in Fig. N.5.

640 AIRCRAFT ENGINE DESIGN

OSCILLATOR

PRE-AMPLIFIER

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Fig. N.5 Pulsed laser beam used in peening.

Figure N.6 shows the airfoil with ablative material on the area to be laser shock

peened with a water curtain flowing over the area to be processed. The laser is

focused on the blade surface and when fired passes through the water and quickly

vaporizes the ablative material. The rapid vaporization of the albative material is

contained by the water curtain and causes a large spike in pressure, which drives

a shock wave into the host material. Yielding of the material occurs on both sides

of the airfoil, and deep residual compressive stresses are formed in the spot by the

reactive response of the elastic material that surrounds the laser spot.

N.2.3.16 Crack initiation life.

Crack initiation life is used in fatigue anal-

ysis to establish a component life based on the attainment of a 0.030-in. crack

length used in safe life analysis.

N.2.3.17 Crack propagation fife.

This is the cyclic life of a component

where a crack from an initial flaw size (usually determined by NDE) grows to a

critical size and failure occurs. 2

"~ -~- "~ ~ LASER BEAM

VAPOR PRESSURE

~M)WATER CURTAIN

FOCUSED ll ~

LASERBEAM "

TRAVELING SHOCK WAVES

Fig. N.6 Airfoil with ablative material in laser shock peening.

APPENDIX N: TURBINE ENGINE LIFE MANAGEMENT 641

N.3 Engine Failure Modes

Each component of the engine can react differently, from a life standpoint,

to the duty cycle to which the engine is subjected. For example, turbine blades

and vanes can be critical in creep or stress rupture if the usage requires long

periods of time at maximum power. On the other hand, a large number of throttle

transients in this duty cycle can alter the failure mode of the same turbine blade

and vanes to that of LCF. Because of this sensitivity, engine usage must be moni-

tored for significant changes relative to that assumed for design. Table N. 1 shows

the failure modes for various components in the engine for typical usage.

N.4 Lifing Concepts

Lifing concepts have evolved and improved since the early period of the gas

turbine engine. Early on, passing a 150-h test was the main "pass or fail" criteria

used to qualify a new engine design to enter production and service. Engine life

development and life prediction techniques evolved mostly in response to dura-

bility problems, customer demands, and/or regulatory involvement. The safe life

(see Fig. N.7) and damage tolerance concepts have been the two most widely

used design methods for producing components to meet life requirements. More

recently, however, the combination of both life concepts, as shown in Fig. N.8, is

the preferred lifing method for both commercial and military and is required by

the military in their engine structural development programs.

N.4.1 Safe Life

Criterion

The safe life fatigue design approach is shown in Fig. N.7. Components are

designed so that the LCF limit exceeds the required usage interval. The life limit

is usually based on minimum or minus 3-sigma value of the material life property

curve for crack initiation. The main concern with this safe life concept is that there

is no recognition or provision for what effect initial defects can have on component

life such as machining errors, burrs, nicks, etc.

CRACK

SIZE

DISTRIBUTION OF CRACK

LCF INiTIATiON TIMES

ENGINE USAGE (CYCLES, EFH, ETC.)

CRITERIA: LOW CYCLE FATIGUE (LCF) LIMIT BASED ON LOWER BOUND

(-3G OR 1/1000) DISTRIBUTION OF CRACK INITIATION TIME

ACTION: 100% PART REPLACEMENT AT LCF LIMIT

Fig. N.7 Safe life fatigue design.

642 AIRCRAFT ENGINE DESIGN

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