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

Подождите немного. Документ загружается.

Appendix M

Example Material Properties

Our purpose here is to provide the basis for representative structural properties

of materials typically found in aircraft engines for use in component design cal-

culations of Sec. 8.2.3. The amount of information that might be displayed can

truly be called bewildering for a number of reasons. To begin with, the material

for each separate engine part is carefully selected, taking into account its usage,

environment, and required durability. Furthermore, for each part there are many

competing materials that have similar capabilities, but different compositions,

processing methods, and brand names. Finally, many properties of a candidate

material must be known before a confident design can be contemplated, such as

allowable tensile and compressive strengths, cyclic strength, creep rates, oxidation,

erosion, and corrosion rates, fracture toughness, crack propagation rate, ductility,

allowable deformation, hardness, modulus of elasticity, thermal conductivity, ther-

mal growth coefficient, and Poisson's ratio (all as a function of time, temperature,

and heat treatment). The permutations are endless.

To cope with this in a textbook, drastic reductions must be made. They are

achieved by only including materials for critical parts (see Sec. 8.2.3), giving

only one example for each generic application, and, most importantly, concen-

trating mainly upon allowable creep strength as required by the considerations of

Sec. 8.2.3.

Although not used directly in the design examples, special attention is drawn

here to the so-called fatigue or cyclic strength of these typical engine materi-

als. This has been done because the life of engine parts is often consumed by

fatigue--or at least by a combination of steady-state and cyclic stress. Exactly

how this combination is accounted for in practice is usually a proprietary

matter. The general nature of the jet engine is such that the major source of cyclic

stress at the "cold end" (i.e., fan and compressor) is vibratory excitation, it and

therefore occurs at about 1000 cycle/s, whereas at the "hot end" (i.e., burner and

turbines) it is a result of pressure and temperature transients and therefore occurs

at about 1 cycle/h. Vibratory and pressure or thermal fatigue are therefore often

referred to as high-cycle and low-cycle fatigue, respectively. The example mate-

rial properties gathered here contain fatigue strength data appropriate to the usual

application of each material, while illustrating the perplexing behavior of this

characteristic.

The information assembled here has been abstracted from the Battelle Memo-

rial Institute's

Aerospace Structural Metals Handbook. 1 The

seven examples are

presented more or less as in their usual order of appearance from the inlet to the

exit of the engine.

623

624

AIRCRAFT ENGINE DESIGN

Aluminum 2124 Alloy (p = 5.29 slug/ft 3)

This is one of a series of premium aluminum alloys that were developed as a

result of studies into the micromechanisms of fracture in high-strength aluminum

compositions. It is normally produced in plate form in a thickness range up to

6 in. and has been thoroughly characterized.

Figure M.1 contains traditional yield and ultimate tensile strength data for

A1 2124 as a function of test temperature and exposure time. Using yield strength

as a design criterion, one would conclude that a part having a required life of 1000

hours while exposed to a temperature of 400°F could withstand tensile stresses up

to 25,000 psi. This, however, would not be a conservative design practice because

Fig. M.1

i 1

2124-1L851 4- to 5-inch Plate

. 80 r

~ 60

2O ~

I,-

--0

O0 200 400 600 800 I000 1200

Temperature, F

Effect of temperature and exposure time on tensile properties of 2124-T851.

APPENDIX M: EXAMPLE MATERIAL PROPERTIES 625

I I I

2124-T851 1.5- to 5-inch Plate

,oo

- I I t

80 75 F

60--

40 ~ -- -- 212 F

Rupture

I 0.2 Percent Creep

lO-I i0 o I01 102 103 104 i0 {i

Time,

Fig. M.2 Creep and creep-rupture curves at temperatures from 75 to 600°F for

2124-T851 plate.

the part will undergo permanent deformation during the exposure time, which

could cause unacceptable distortion and/or interference.

Consequently, it is preferable to base designs on creep and creep-rupture data,

such as those shown in Fig. M.2. The common practice is to allow less than 1%

creep during the life of the part, which, for the preceding example, would allow a

tensile stress of about 15,000 psi. This is generally the more restrictive criterion,

and, therefore, no further yield or ultimate strength data will be presented for any

material.

Figure M.2 reveals that the useful tensile strength of A12124 diminishes rapidly

above a temperature of 200°F; thus, it would seldom find engine application above

400°F. It must also be remembered that the value obtained from Fig. M.2 rep-

resents an upper limit that cannot be directly used as a design criterion. Reduc-

tions must be made to account for cyclic stress, uncertainties in load calculations,

property variations, and safety margins. Experience indicates that a reasonable

estimate of the "allowable working stress" is 80% of the 0.2% creep value ob-

tained from Fig. M.2.

Some idea of the ability of A1 2124 to tolerate cyclic stress can be found on

Figs. M.3 and M.4. A turbine engine part will experience several thousand LCF

cycles caused by throttle motions, but literally billions of HCF cycles caused by

vibrations. Hence, the so-called "run-out" stress, which the material can apparently

withstand forever, becomes an important factor.

Among the conclusions to be drawn from Figs. M.3 and M.4 are that a notch

reduces the alternating stress capability of A12124 dramatically and that it is less

tolerant of bending than longitudinal cyclic stress. The latter point could be critical

for cantilevered airfoil applications.

626 AIRCRAFT ENGINE DESIGN

100 n

"; 80--

J¢

E 60

I 2124-T851 Plate

i [

Notched Smooth

60 ° | . 1/2 R

K,- 3 .o,o~ .a.

r - 0.013

Im Nominal

.~ Thickness Smooth Notched-

= • \ 2 in. • O

._E I ~ 4 i,,. • z~

x ~tl ] "~. 4.5 in. • r-I

20 I '-."v~ ~Notch.ed

Axial Stress

R-0 I

0103 10 4 10 5 106 10 7

Cycle= to Failure

Fig. M.3 Axial-load fatigue curves for longitudinal smooth and notched specimens

from 2124-T851 plate.

2124-T851 4-inch Plate

R ~ 10 inch

¢,3

~= so

Diameter = 0.3 inch

L&T

0104 t0 5 I0 6 l0 T 108 109

Cycles to Failure

Fig. M.4

Rotational bending

fatigue curves

for smooth specimens from 4 in

2124-

T851 plate.

APPENDIX M: EXAMPLE MATERIAL PROPERTIES 627

Table M.1 Variations in creep deformation with different creep exposures

Time to 0.1% Time to 0.2% Total plastic

Creep exposure creep, h creep, h deformation, %

Temp, °F Stress, ksi Time, h ee-fl forged at 1650°F

750 90 1228 270 840 0.251

750 89 150 0.101

800 75 504 105 420 0.217

800 85 241 40 210 0.234

800 85 150 0.180

1000 20 51 6 27 0.246

1000 30 72 2 8 0.760

1000 30 150 0.820

Form: disc forging, 3 in. thick x 11 in. diameter.

Titanium 6246 Alloy (p = 9.08 slug/ft 3)

This alloy was designed to combine long-time, elevated-temperature strength

characteristics with markedly improved short-time strength properties at both room

and elevated temperatures. It is intended for use in forgings for intermediate-

temperature-range sections of gas turbine engines, particularly in disk and rotor

airfoil components of the fan and compressor.

Ti-6246 is not thoroughly characterized, but is recommended for gas turbine

applications up to 750°E The data of Table M.1 suggest that the allowable stress

based on 1% creep for 1000 hours use at 750°F is about 90,000 psi, whereas at

800°F it is only about 75,000 psi and at 1000°F it is below 20,000 psi.

The data of Fig. M.5, although taken at room temperature, show that Ti-6246

does well with cyclic stress. However, Ti-6246 has poor fracture toughness

200

180

160

140

TI-6246

_ FORGINGS --

• STA COND: 1600 F, IHl~ WQ+II00,8HR,AC

• DA COND: 1600F, I HR, AC+I000F, 8HR, AC

R=0.1

A=0.82

K t =1

103 104 105

CYCLES TO FAILURE

Fig. M.5 Axial fatigue properties of c~-/~ forged materials in two heat-treated

conditions.

628 AIRCRAFT ENGINE DESIGN

100

INCONE L' 601 I 1

SOLUTION

TREATED AT

2100F --

1000F

~1200F

~1300F

1400F

1600F

1800F

~ 2000F

0.1

0. 0001 0. 001 0.01 0.1 1.

MINIMUM CREEP RATE - PERCENT PER HOUR

Fig. M.6 Minimum creep rate at various temperatures and stresses.

and is very susceptible to minor damage. Therefore, either a damage-tolerant

processed version of Ti-6246 or a slightly different alloy is required for practical

applications.

Inconel 601 (p = 15.6 slug/ft 3)

This is a general-purpose alloy for applications requiring resistance to both heat

and corrosion at temperatures up to about 2100°E It has been used in burner liners,

diffuser assemblies, containment rings, exhaust liners, and burner igniters.

The data of Fig. M.6 may be used to show, for example, that Inconel 601 will

creep 1% in 1000 hours at 1600°F if the stress is about 3000 psi. For the higher-

temperature applications it must, therefore, be kept unloaded. In stark contrast with

Ti-6246, however, it retains much of its room temperature strength well beyond

1000°E

Because Inconel 601 is not commonly employed where vibrations matter,

Fig. M.7 shows that it has more than adequate fatigue strength.

Hastelloy X (p = 16.0 slug/ft 3)

This is a nickel-base superalloy with good oxidation resistance at temperatures

up to 2200°F and moderately good strength properties up to 1600°E It has been

used in jet engine exhaust nozzles, afterburner components, and structural parts in

the burner and turbine components.

Referring to Fig. M.8 and using the 1% creep at 1000 hours criterion, one would

conclude that the upper limit to tensile stress at 1200°F is 22,000 psi and at 1500°F

is 7000 psi for Hastelloy X.

Because Hastelloy X is generally used in situations where only low-cycle fa-

tigue caused by pressure or thermal cycles matters, its fatigue characteristics are

APPENDIX M: EXAMPLE MATERIAL PROPERTIES 629

' 60

m 50

REVERSE BEND

R=-I

INCONE'L 601

COLD-ROLLED

SHEET

ANNEALED

1900F

(FTU

KSB

105 106 107

CYCLES TO FAILURE

Fig. M.7

108

Fatigue properties of annealed sheet.

presented in terms of strain range rather than stress range, as in Fig. M.9. This

must obviously be obtained from a thermal stress calculation involving at least

a knowledge of the coefficient of thermal expansion and modulus of elasticity,

given in Figs. M.10 and M.11, respectively. Note that in the 1200-1600°F range

the modulus of elasticity is approximately 20 million psi. Consequently, the stress

corresponding to a 1% strain range equates approximately to 200,000 psi. This

makes it evident that such materials can tolerate only a small fraction of a percent

strain.

40 "\

Hastelloy )< Plate an¢t Bar, Solution'Treated

at 2150 F, Rapid Cool

' I

Creep Strain

5 Percent

....

I Percent

I0 2 10 3 10 4.

Time,

hours

Fig. M.8 Creep-deformation curves for plate and bar at temperatures of 1200--

1800°F.

630 AIRCRAFT ENGINE DESIGN

11.0

I-

Environment'

Helium Air

•

75F

•

1000 F

[3 Ill 1600 F

0"1102 10 3

i i i i i i

Ha~telloy X 0.5-inch Plate,

Solution Treated at

1150 F, Rapid Cool

10 4 10 5

Cycles to Failure

Fig. M.9 Fatigue life of plate at various temperatures in air and impure helium at

atmospheric pressure.

Mar-M 509 (p = 17.2 slug/ft 3)

This is a high-chromium, carbide-strengthened, cobalt-base superalloy. It was

designed for use in the gas turbine industry, primarily as an investment-cast stator

airfoil alloy to operate at temperatures of 1800°F. Mar-M 509 has high inherent

oxidation and sulfidation resistance and is, therefore, sometimes used without a

protective coating. It also has superior thermal shock characteristics.

The most relevant design tensile stress data for Mar-M 509 are given in Fig. M. 12.

A logical criterion would be 0.5% creep deformation in 500 hours, which would

lead to a stress of approximately 9000 psi at 1800°F. No data were given for

alternating stress.

¢=

~,, I0

'7~ 9

i i

Hastalloy

Between

79 F

and Temperature

Indicated

I I I 1

400 800 1200 1600

Temperature, F

0 2000

Fig. M.10 Effect of elevated temperature on thermal expansion coefficient.

APPENDIX M: EXAMPLE MATERIAL PROPERTIES 631

Haltelloy

Solution

Treated at

2150

30 -- Rapid Cool

-,,

• Sheet, Dynamic Ill lib

• Sheet, Static (47)

• Bar, Static (38)

I L

80 400 800 1200

Temperature, F

Fig. M.11

1600 2000

Effect of elevated temperature on modulus of elasticity.

6O t [ l l t I

Co-24Cr-10Ni-7W-3.5Ta + Ti + Zr As Cast

4c \ I I I I I I

~ 1-1/2 percent Creep in 500 hr

4300

1400 1500 1600 1700 1800 1900 2000

Temperature, F

Fig. M.12 Mar M 509 design curves for typical stresses required to produce creep

strains of 0.5, 1, and 1.5% in 500 hours at temperatures of 1300-2000°F.

632 AIRCRAFT ENGINE DESIGN

I00

Ren~

80 F~ull Heat

"l:reatment'

"Class A: 2225 F, 2 hr, HeQ

+ 2000 F, 4 hr, HeQ

+ 1925

F, 4

hr, FC to 1200

F in 20 min_

1550 F, 16 hr, AC

Stress Rupture

..... 1.0 I

.... 0.5 Percent Creep

------0.2

Specimens ~--~ "'~.

"'-

Cast

to Size ~.

"-'~."''- "''''--

+ Machined 1600 F

Grain Size---__

l~"~. -..] ~"~ ~'-

1/64 to 1/32

inch 18800 F ~" "-~":" ~

) ~--~_ ~. -~.~

010-1 i0 0 I01 102 IO s 104

Time. hr

Fig. M.13 Creep strain and creep-rupture curves at 1400, 1600, and 1800°F for fully

treated cast alloy.

"10

200 m

I0 z

Fig. M.14

! I

Ren6 80

Cylindrical Specimens Cast

to Size, Then Machined Tested

Uncoated, Longitudinal Strain Control

Mode: Axial-Axial, A~ = 0% f = 20 cpm, __

K t = 1.0

Full Heat Treatment: 2225 F, 2 hr,

HeQ + 2000 F, 4 hr, HeQ + 1925 F,

4 hr, FC to 1200 F in 20 min, AC

+ 1550 F, 16 hr, AC

Pseudo Stress = Elastic Modulus x c

~--..~- 1200

1800 F 1700

10 3 10 4 IO s

Cycles to Failure

Axial low cycle fatigue behavior at 1200-1800°F.