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

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DESIGN: INLETS AND EXHAUST NOZZLES 497

6600

6590

6580

(~f)

6570

6560

6550

1.30 1.35 1.40 1.45 1.50 1.55 1.60 1.65 1.70

Ag[As

Fig. 10.E9 Mission fuel used vs nozzle area ratio.

the required nozzle area ratio

(A9/As)

is plotted versus flight Mach number (M0)

for all mission legs, as was done in Fig. 10.El0, a definite pattern results. The

variable nozzle area ratio shown by the solid line in Fig. 10.El0 is selected for

the AAF and input into the nozzle area schedule in the Mission window of the

AEDsys program (does not include improved inlet estimates). Results of the mis-

sion calculations for this variable nozzle area ratio are listed in Table 10.E7. Note

that this nozzle area ratio schedule gives a mission fuel used that is only 25 lbf

greater that the reference values from Chapter 6 and 226 lbf less than the overall

goal of 6690 lbf from Chapter 3. Two nozzle actuators are required to obtain the

schedule of Fig. 10.E10---one to set A8 and the other to set

A9/A8.

The variable

nozzle area schedule shown in Fig. 10.El0 is selected for the AAF.

10.4.2.2 AAF nozzle geometry.

The preliminary design of the exhaust

nozzle geometry begins with selection of the maximum values for both the primary

nozzle half-angle (0) and the secondary nozzle half-angle (c0 (see Figs. 10.71b,

10.72, and 10.74 in Sec. 10.3.3). The value of the primary nozzle half-angle di-

rectly affects the nozzle discharge coefficient

(CD)

and thus the nozzle throat

area (A8,). An increase in 0 reduces the length and weight of the primary noz-

zle, but may increase the overall weight of the nozzle due to the decrease in

that increases the secondary nozzle inlet and exit areas. The secondary nozzle

498 AIRCRAFT ENGINE DESIGN

A9/A 8

2.2

2.0

1.8

1.6

1.4

1.2

,,,~I

' ' l ' ' ' ' I ' ' '

• • •~/'~"~"~ Selected • AAF Schedule

1.0 , J , ~ I , ~ J , I , , ,

0,0 0,5 1,0 1,5 2,0

Fig. IO.EIO

Required nozzle area ratio vs Mach number.

half-angle (or) affects both the velocity coefficient (Cv) and the angularity coef-

ficient (CA). For a fixed nozzle area ratio (A9/As), an increase in ot increases

the velocity coefficient, decreases the angularity coefficient, decreases the nozzle

length and weight, and changes the gross thrust coefficient. Hence, the selection

of the nozzle half-angles (0 and or) is a complex design problem by itself when

nozzle weight is included.

A maximum primary nozzle half-angle (0) of 30 deg is selected, which corre-

sponds to the nozzle throat at its military power setting. Likewise, a maximum

secondary nozzle half-angle (tx) of 12 deg is selected that corresponds to the

1.6M/5g turn with supersonic area ratio. The resulting nozzles at military and

maximum power for area ratios of 1.2 and 2.0 are shown to scale in Fig. 10.El 1

and Fig. 10.El2, respectively, using the NOZZLE program.

10.4.2.3 AAF nozzle performance. TheperformanceoftheAAFexhaust

nozzle of Figs. 10.El 1 and 10.El2 is determined at each flight condition based

on the general thrust performance method of Sec. 10.3.3, neglecting losses due

to leakage and cooling (ACfu = 0). Values for CD, Cv, and CA were obtained

from Figs. 10.71b, 10.72, and 10.74, respectively. Equation (10.33) was used to

calculate the gross thrust coefficient (Cfg) where the gross thrust (Fg) is given

by Eq. (10.27). The uninstalled thrust (F) was calculated by subtracting the

DESIGN: INLETS AND EXHAUST NOZZLES 499

Table 10.E7 AAF fuel burn for variable exhaust nozzle area schedule

Mission fuel used--

WF,

lbf

Mission phases Mo/Alt, Variable

A 9

Power

Fref,

and segments kft Ref area ratio A8 setting lbf

1-2 A--Warm-up a 0.0/2 214 214 1.200 Mil 8,846

1-2 B--Takeoff acceleration a 0.10/2 105 105 1.200 Max 13,563

1-2 C Takeoff rotation a 0.182/2 40 40 1.200 Max 13,317

2-3 D--Horizontal acceleration a 0.441/2 129 129 1.200 Mil 7,493

2-3 E--Climb/acceleration 0.875/23 428 437 1.200 Mil 4,873

3-4 Subsonic cruise climb 0.900/42 524 522 1.200

= 0 1,403

5-6 Combat air patrol 0.700/30 683 690 1.200 Ps = 0 1,294

6-7 F--Acceleration 1.090/30 391 393 1.453 Max 9,443

6-7 G--Supersonic penetration 1.500/30 1183 1183 2.000 P~ = 0 5,730

7-8 I--1.6M/5g turn 1.600/30 470 470 2.000

= 0 9,612

7-8 J--0.9M/5g turns 0.900/30 498 509 1.200 Ps = 0 7,376

7-8 K--Acceleration 1.195/30 346 347 1.593 Max 10,596

8-9 Escape dash 1.500/30 322 322 2.000 Ps = 0 5,683

9-10 Zoom climb 1.326/39 44 45 1.769 Mil 3,948

10-11 Subsonic cruise climb 0.900/48 491 491 1.200

= 0 1,054

12-13 Loiter 0.397/10 569 568 1.200 P, = 0 923

Total 6439 6464

1.80/40 Maximum Mach number 2.000 Max 10,559

al00°E

momentum of the entering air

(rhoVo/gc)

from

F = Fg - fnoVo/gc

(10.E5)

The values used in calculating the nozzle performance and the results are pre-

sented in Table 10.E8. The ratio of the uninstalled thrust (F) based on this nozzle

design to the uninstalled thrust from the performance analysis

(Fref)

of Table 10.E7

is very close to unity for most flight conditions. Hence, the 7rn = 0.97 and P0 = P9

used in the engine performance analysis predicts the performance of the AAF ex-

haust nozzle quite well over the flight conditions considered.

10.4.2.4 AAF nozzle installation losses.

The nozzle installation losses

((bnozzte)

of Chapter 6 were based on nozzle area ratios corresponding to ideal

expansion (P9 = P0). Now that the geometry and nozzle area ratios of the AAF

exhaust nozzle are known, a revised estimate of nozzle installation losses will be

made based on the methods of Chapter 6 (see Sec. 6.2.3). The throat flow areas

(As) listed in Table 10.E8 and the dimensions shown in Figs.10.E11 and 10.E12

are based on two-dimensional axisymmetric flow and result in nozzle areas that

are 6.4% larger than the one-dimensional area calculated by the AEDsys cycle

calculations. The minimum throat area was increased from its reference value by

the reciprocal of the discharge coefficient (1/Co) and the nozzle dimensions scaled

accordingly. Based on the dimensions of Figs.10.E11 and 10.E12, the afterbody

500 AIRCRAFT ENGINE DESIGN

[, 12.0 in )1( 19.4 in )[

A9/~

= 2.0

"~ ~,,. 10"59°

8.77 in

A9M 8 = 1.2

Fig. 10.Ell AAF exhaust nozzle---military power.

area (A10) is estimated to be 6.30

ft 2

(rl0 = 17 in.) and the afterbody length (L) is

estimated to be 5 ft [two times the nozzle length of 31.4 in. (12.0 in. + 19.4 in.)].

These values are larger than the afterbody area of 5.153

ft 2

and afterbody length

of 4.611 ft estimated in Chapter 6. The increase in afterbody area is mainly due to

the 31.4 in inside diameter of the afterburner.

The AAF exhaust nozzle schedule was input into the Mission portion of the

AEDsys program and the mission flown. The resulting installation losses for the

AAF exhaust nozzle are listed in Table 10.E9. The ratios of installed thrust

(T/Tref)

are also listed and are based on

T F

(1 --

~inlet -- ~nozzle)

--~-- where

~)inlet:~ginletref

Zref Fref (1 - ~ginlet -- ~nozzle)ref

The results show that nozzle installation losses

(fb,ozzte)

do not appreciably change

from the estimates of Chapter 6 (see Table 10.E9) except for the two acceleration

flight conditions (segments 6-7 F and 7-8 K).

I. ,20in .[. ,94,n .[

'~ A9/A 8 =

2.0

15.7 in [ N

11.3 in

Ag/A 8 = 1.2

Fig. 10.El2 AAF exhaust nozzle---maximum power.

.o ~

502 AIRCRAFT ENGINE DESIGN

..=

.~ "o

O Q

~ ~ r,.J u..,

DESIGN: INLETS AND EXHAUST NOZZLES 503

0 0

ii'il , •

504 AIRCRAFT ENGINE DESIGN

10.4.3 Closure

The present design configuration status is summarized by Figs. 10.E3, 10.E6,

10.El0, 10.Ell, and 10.El2. Although this is a promising start because no real

barriers to a successful design have been encountered, further iteration would focus

on the following: 1) improvement of nozzle losses for Subsonic Cruise Climb

legs; 2) improvement of inlet installation performance for Combat Air Patrol; and

3) impact of inlet and nozzle internal and external performance on the overall

engine installed performance.

The combined influences of the inlet and nozzle designs on the installed thrust

of the AAF with respect to the results of Chapter 6 (T/Tref)

are

not apparent by

viewing Tables 10.E4 and 10.E9. Since the AEDsys program does not currently

include an improved inlet model, the overall performance of the AAF will be

estimated using the data already gathered. The data now available in Tables 10.E3,

10.E4, 10.E5, 10.E8, and 10.E9 are used to estimate their combined influence on

T~ Tref using

T ..~ Jr,/ F (1

~inlet -- (Pnozzle)

"~ref -- 7gdspec ~eref nozzle

(1 --

4)inlet -- 4)nozzle)ref

where F/Frefl nozzle is obtained from Table 10.E8. The results are presented in

Table 10.El0. Note that the improvements in inlet performance more than com-

pensate for the decrease in nozzle performance for the majority of mission phases

and segments. The improved performance over the two Subsonic Cruise legs,

Loiter leg, and other portions of the AAF mission will more than offset the

lower installed thrust performance during the Combat Air Patrol and several other

mission phase.

References

1Seddon, J., and Goldsmith, E. L., Intake Aerodynamics, 2nd ed., AIAA Education Series,

AIAA, Reston, VA, 1999.

2younghans, J., "Engine Inlet Systems and Integration with Airframe," Lecture Notes for

Aero Propulsion Short Course, Univ. of Tennessee Space Inst., Tullahoma, TN, 1980.

3"Stealth Engine Advances Revealed in JSF Designs," Aviation Week and Space Tech-

nology, 19 March, 2001.

4Hawker Siddeley Aviation Ltd., The Hawker Siddeley Harrier, Bunhill Publications

Ltd., London, 1970 (Reprint from Aircraft Engineering, Dec. 1969-Apr. 1970).

5Fabri, J. (ed), Air Intake Problems in Supersonic Propulsion, Pergamon, New York,

1958.

6Heiser, W., and Pratt, D., Hypersonic Airbreathing Propulsion, AIAA Education Series,

AIAA, Reston, VA, 1994.

7Sedlock, D., and Bowers, D., Inlet~Nozzle Airframe Integration, Lecture Notes for Air-

craft Design and Propulsion Design Courses, U.S. Air Force Academy, Colorado Springs,

CO, 1984.

8Swan, W., "Performance Problems Related to Installation of Future Engines in Both

Subsonic and Supersonic Transport Aircraft," March 1974.

9Oates, G. C.,

Aerothermodynamics of Gas Turbine and Rocket Propulsion (revised and

enlarged), AIAA Education Series, AIAA, Reston, VA, 1988.

DESIGN: INLETS AND EXHAUST NOZZLES 505

l°Surber, L., "Trends in Airframe/Propulsion Integration,"

Lecture Notes for Aircraft

Design and Propulsion Design Courses,

U. S. Air Force Academy, Colorado Springs, CO,

1984.

11 Kitchen, R., and Sedlock, D., "Subsonic Diffuser Development for Advanced Tactical

Aircraft," AIAA Paper 83-168, 1983.

12Aronstein, D., and Piccirillo, A.,

Have Blue and the F-117A: Evolution of the "Stealth

Fighter,"

AIAA, Reston, VA, 1997.

13Hunter, L., and Cawthon, J., "Improved Supersonic Performance Design for the F-16

Inlet Modified for the J-79 Engine," AIAA Paper 84-1272, 1984.

14Stevens, C., Spong, E., and Oliphant, R., "Evaluation of a Statistical Method for De-

termining Peak Inlet Flow Distortion Using F-15 and F-18 Data," AIAA Paper 80-1109,

1980.

15Oates, G. C. (ed),

Aircraft Propulsion Systems Technology and Design,

AIAA

Education Series, AIAA, Reston, VA, 1989.

16Oates, G. C. (ed),

Aerothermodynamics of Aircraft Engine Components,

AIAA Edu-

cation Series, AIAA, Reston, VA, 1985.

17Stevens, H. L., "F-15/Nonaxisymmetric Nozzle System Integration Study Support Pro-

gram," NASA CR-135252, Feb. 1978.

18Summerfield, M., Foster, C. R., and Swan, W. C., "Flow Separation in Overexpanded

Supersonic Exhaust Nozzles,"

Jet Propulsion,

Vol. 24, Sept.-Oct. 1954, pp. 319-321.

19Tindell, R., "Inlet Drag and Stability Considerations for M0 = 2.00 Design," AIAA

Paper 80-1105, 1980.

Epilogue

The Air-to-Air Fighter (AAF) Engine design study has reached a successful

conclusion. Satisfying solutions have been found to all of the important technical

problems, and it is evident that the AAF systems requirements can be met. The

competition may now proceed in earnest.

Further refinements and/or improvements, based on the quantitative results and

physical insights gained during this first iteration, are certainly possible. The most

likely candidates for further exploration have been noted along the way. We rec-

ommend, however, against doing this unless the system requirements change or

superior analytical tools become available.

If no adequate AAF Engine design had been found, the next logical step would

be to revise the engine cycle parameters and try again. It is even possible that

no reasonable design point solution exists, in which case the aircraft parameters

or even the mission requirements must be re-examined. Fortunately, that did not

happen here, and we can happily close the book on the AAF Engine project.

We hope that you have found this journey as challenging, instructive, and re-

warding as we have, and that you find the concepts, analyses, and software useful

for their original purposes, and in ways that we have not envisioned.

507