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

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416 AIRCRAFT ENGINE DESIGN

References

1Hawthorne, W. R., "The Early History of the Aircraft Gas Turbine in Britain," Deutsche

Gesellschaft ftir Luft- und Raumfarht Symposium, Munich, 1989.

2Lefebvre, A. H.,

Gas Turbine Combustion,

2nd ed., Taylor and Francis, Philadelphia,

1999.

3Zukoski, E. E., "Afterburners,"

Aerothermodynamics of Aircraft Engine Components,

edited by G. C. Oates, AIAA Education Series, AIAA, Washington, DC, 1985.

4pratt, D. T., "Calculation of Chemically Reacting Flows with Complex Chemistry,"

Studies in Convention,

Vol. II, edited by B. E. Launder, Academic Press, 1977.

5Kollrack, R., "Model Calculations of the Combustion Product Distributions in the

Primary Zone of a Gas Turbine Combustor," American Society of Mechanical Engineers,

Paper 76-WA/GT-7, 1976.

6Radhakrishnan, K., and Pratt, D. T., "Fast Algorithm for Calculating Chemical Kinetics

in Turbulent Reacting Flow,"

Calculations of Turbulent Reactive Flows,

AMD-Vol. 81,

edited by R. M. C. So, J. H. Whitelaw, and H. C. Mongia, American Society of Mechanical

Engineers, New York, 1986.

7Spalding, D. B.,

Combustion and Mass Transfer,

Pergamon, Oxford, England, U.K.,

1979.

8Levenspiel, O.,

Chemical Reaction Engineering,

Wiley, New York, 1962.

9pratt, D. T., "Gas Phase Chemical Kinetics,"

Pulverized-Coal Combustion and Gasifi-

cation,

edited by L. D. Smoot and D. T. Pratt, Plenum, New York, 1979.

l°Bragg, S. L., and Holliday, J. B., "The Influence of Altitude Operating Conditions

on Combustion Chamber Design,"

Selected Combustion Problems,

Vol. II, Butterworths,

London, 1956.

l J Kays, W. M., and Crawford, M. E.,

Convective Heat and Mass Transfer,

2nd ed.,

McGraw-Hill, New York, 1980.

12Dittrich, R. T., and Graves, C. C., "Discharge Coefficients for Combustor-Liner Air-

Entry Holes, I-Circular Holes with Parallel Flow," NACA TN 3663, 1956.

13Brodkey, R. S. "Mixing in Turbulent Fields,"

Turbulence in Mixing Operations,

edited

by R. S. Brodkey, Academic Press, New York, 1975.

laSimpson, L. L., "Industrial Turbulent Mixing,"

Turbulence in Mixing Operations,

edited

by R. S. Brodkey, Academic Press, New York, 1975.

~SEroglu, A., "An Experimental Investigation of Entrainment and Mixing in Pulsed and

Exponential Transverse Jets," Ph.D. Dissertation, Univ. of Washington, Seattle, 1991.

16pratt, D. T., "An Analytical Investigation of the Pressure Loss in a Cylindrical Gas

Turbine Combustor," M.S. Thesis, Univ. of California, Berkeley, 1962.

17Beer, J. M., and Chigier, N. A.,

CombustionAerodynamics,

Krieger, Malabar, FL, 1983.

18Dimotakis, P. E., "Turbulent Free Shear Layer Mixing and Combustion,"

High-Speed

Flight Propulsion Systems,

edited by S. N. B. Murthy and E. T. Curran, Vol. 137, Progress

in Astronautics and Aeronautics Series, AIAA, Washington, DC, 1991.

19Schlichting, H.,

Boundary Layer Theory,

6th ed., McGraw-Hill, New York, 1968.

2°Wilson, D. G.,

The Design of High-Efficiency Turbomachinery and Gas Turbines,

MIT

Press, Cambridge, MA, 1984.

21White, E M.,

Fluid Mechanics,

2nd ed., McGraw-Hill, New York, 1986.

22Dixon, S. L.,

Fluid Mechanics, Thermodynamics of Turbomachinery,

3rd ed.,

Pergamon, Oxford, England, 1978.

DESIGN: COMBUSTION SYSTEMS 417

23Evangelista, J. J., Shinnar, R., and Katz, S., "The Effect of Imperfect Mixing on Stirred

Combustion Reactors," Twelfth Symposium (International) on Combustion, Combustion

Inst., Pittsburgh, PA, 1969.

24Swithenbank, J., "Flame Stabilization in High Velocity Flow," Combustion Technology:

Some Modern Developments, edited by H. B. Palmer and J. M. Beer, Academic Press,

New York, 1974.

25 Cornell, W. G., "The Flow in a Vee-Gutter Cascade," Transactions of the ASME, Vol. 78,

Pt. 1, 1956, pp. 573-580.

26Mattingly, J. D., Elements of Gas Turbine Propulsion, McGraw-Hill, New York, 1996.

27Bahr, D. W., "Technology for the Design of High Temperature Rise Combustors,"

AIAA Paper 85-1292, 1985.

28Aronstein, D., and Picirillo, A., Have Blue and the F-117A: Evolution of the Stealth

Fighter, AIAA, Reston, VA, 1997.

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

nology, 19 March 2001.

3°Chigier, N. A., Energy, Combustion, and Environment, McGraw-Hill, New York, 1981.

Engine Component Design:

Inlets and Exhaust Nozzles

10.1 Concept

The inlet and exhaust nozzles are the two engine components that directly in-

terface both with the internal airflow and the flow about the aircraft. In fact, the

integration of the engine and the airframe is one of the most complex and has a

major impact on the performance of the aircraft system. Many technical books,

reports, articles, etc., are available in open literature that concentrate on only small

parts of this major design challenge. This chapter (see also Chapter 6) identifies

the major design considerations of inlets and exhaust nozzles and presents basic

design tools for their preliminary sizing and design.

The results of the preceding engine performance analysis provide a wealth

of information about the required performance of both the inlet and the exhaust

nozzle. For example, the inlet area vs both Mach number and altitude can be ob-

tained from the engine performance (off-design) cycle analysis program of AEDsys

(see Fig. 6.14). Likewise, the engine airflow at specific partial throttle condi-

tions (corresponding to cruise, loiter, etc.) and the assumed inlet total pressure

ratio vs Mach number can be obtained. This design information defines the re-

quirements of the inlet in terms of total pressure ratio and mass flow rate and its

preliminary design starts with this information.

The simplest and most powerful design tool available for preliminary design

of these components is one-dimensional compressible flow. Both the inlet and

the exhaust nozzle can be modeled as simple one-dimensional adiabatic flows or

a series of these flows. The following sections of this chapter present the basic

principles of operation for each component, the major design considerations, and

the basic design tools. Starting at the front of the engine, the inlet is considered

first.

10.2 Inlets

10.2.1 Introduction

The inlet interchanges the organized kinetic and random thermal energies of

the gas in an essentially adiabatic process. The perfect (no loss) inlet would thus

correspond to an isentropic process. The inlet works in concert with the compressor

to give the overall pressure ratio of the engine cycle

Pt3 Pro Pt2 Pt3

__ __

7grTtJd7"(c

PO PO

P.,o

P,2

where ,r, Trd represents the pressure ratio of the inlet (~i = ~rZrd) and ~rc is the

pressure ratio of the compressor. The inlet may be thought of as a compressor

419

420 AIRCRAFT ENGINE DESIGN

with no moving parts. Thus, by analogy with Eq. (4.9a) or (4.9b), the adiabatic

efficiency of the inlet (t/i) can be written as

7r~ Y-1)/Y --

1 rirr~a ×-1)/× - 1

rli -- -- (10.1)

zi - 1 ~i - 1

where "E r

"C i . This expression can be recast in terms of an inlet polytropic

efficiency

(ei)

by using ri

7r{ Y-1)/(yei).

Thus

In {1 + (ri

- 1)/~/}

In{ riZrd (×-l)/y } y -- 1 In (rrd)

= --

-- 1 + -- (10.2)

In(z/) In(r/) y In (rr)

where

7"(" d = ~dmaxFIRspec

(4.12a)

For supersonic flow below Mach five, military specification MIL-E-5008B gives

rlRspec

= 1.0 - 0.075(M0 - 1) 1'35 (4.12c)

Using the above equations, the variation of inlet adiabatic efficiency and equivalent

polytropic efficiency versus flight Mach number are given in Fig. 10.1. The high

inlet efficiencies show that the inlet is superior to the mechanical compressor in

e i

r/i

0.95

0.90

0.85

1.0 1.5 2.0 2.5 3.0 3.5

Fig. 10.1 Inlet efficiencies

('7 --

1.4).

DESIGN: INLETS AND EXHAUST NOZZLES 421

increasing pressure up to Mach 3.5 (see Table 4.4)--a surprising result, and an

extremely useful one for supersonic applications.

The primary purpose of the inlet is to bring the air required by the engine

from freestream conditions to the conditions required at the entrance of the fan or

compressor with minimum total pressure loss and flow distortion. The fan or com-

pressor work best with a uniform flow of air at a Mach number of about 0.5. Also,

since the installed engine performance depends on the inlet's installation losses

(additive drag, external forebody or cowl drag, bypass air, boundary layer bleed

air, etc.), the design of the inlet should minimize these losses. The performance of

an inlet is related to the following characteristics: high total pressure ratio (red),

controllable flow matching, good uniformity of flow, low installation drag, good

starting and stability, low signatures (acoustic, radar, and infrared), and minimum

weight and cost while meeting life and reliability goals. An inlet's overall perfor-

mance must be determined by simultaneously evaluating all of these characteristics

since an improvement in one is often achieved at the expense of another.

The design and operation of subsonic and supersonic inlets differ considerably

due to the characteristics of the flow. For the subsonic inlets, near-isentropic internal

diffusion can easily be achieved and the inlet flow rate adjusts to the demand. The

internal aerodynamic performance of a supersonic inlet is a major design challenge

since efficient and stable supersonic diffusion over a wide range of Mach numbers

is very difficult to achieve. In addition, the supersonic inlet must be able to capture

its required mass flow rate, which may require variable geometry to minimize

inlet loss and drag and provide stable operation. Because of these differences,

the following sections will consider the subsonic and supersonic inlets separately,

beginning with the subsonic inlet.

10.2.2 Design ToolswSubsonic Inlets

Most subsonic aircraft have their engines placed in nacelles and thus this section

will not deal with the inlet alone, but will include the nacelle. The cross section of

a typical subsonic inlet and its geometric parameters are shown in Fig. 10.2. The

Highlight (hl),,xX~ ~

A1 = RD21/4 L '

Dfan Dma x

Fig. 10.2 Subsonic inlet nomenclature.

422 AIRCRAFT ENGINE DESIGN

inlet area,

A1,

is based on the flow cross section at the most upstream opening

of the inlet, known as the inlet highlight. Because the subsonic inlet can draw

in airflow whose freestream area (A0) is larger than the inlet area (A1), variable

inlet geometry is not required (except that blow-in doors are often used to reduce

installation drag during takeoff). The material in this section on subsonic inlets is

based on a fixed geometry inlet.

A list of the major design variables for the inlet/nacelle would include the

following:

1) Inlet total pressure ratio and drag at cruise.

2) Engine location on wing or fuselage (avoidance of foreign object damage,

inlet flow upwash and downwash, exhaust gas reingestion, ground clearance).

3) Aircraft attitude envelope (angle of attack, yaw angle, crosswind takeoff).

4) Inlet total pressure ratio and distortion levels required for engine operation.

5) Engine-out windmilling airflow and drag (nacelle and engine).

6) Integration of diffuser and fan flow path contour.

7) Integration of external nacelle contour with thrust reverser and accessories.

8) Flowfield interaction between nacelle, pylon, and wing.

9) Community and/or stealth noise suppression requirements.

10) Stealth electromagnetic signal suppression requirements {low Radar Cross-

Section (RCS)}.

Basic design tools for many of these items will be identified in this section. The

reader is encouraged to research open literature for a more in-depth analysis of

any single item. Special attention is drawn to Ref. 1, which is a recently updated

textbook on the aerodynamics of inlets. The recent developments in stealth have

impacted inlet design as seen in the application of this technology to new subsonic

military aircraft (e.g., F-117A Nighthawk stealth fighter and B-2 Spirit stealth

bomber).

10.2.2.1 Inlet total pressure ratio (rrd).

In the cycle analysis used in the

preceding chapters, the inlet total pressure ratio (Zrd) was assumed to be constant

for subsonic inlets and equal to ~dmax

(the

total pressure ratio due to friction).

Due to the complexity of the flow, a method will not be presented for calculating

the inlet total pressure ratio. However, Fig. 10.3 presents attainable Zrd and its

variation with flight Mach number and corrected engine mass flow. If desired, the

impact of the varying zrd on engine performance can be estimated using this figure

in concert with the off-design cycle deck and mission profile.

10.2.2.2 Inlet sizing--throat diameter ( Dth ).

Thediameteratthethroatof

the subsonic inlet is sized such that the Mach number at this location (based on one-

dimensional flow) does not exceed 0.8. This will provide some margin for growth

or error since the one-dimensional Mach number at the throat corresponding to

actual inlet choke is about 0.9. The maximum corrected engine mass flow that

the throat must pass will then correspond to this limiting Mach number and the

diameter is easily calculated using

k Pto ]max

MFP@M

= 0.8

DESIGN: INLETS AND EXHAUST NOZZLES 423

7~d

0.985

Desi r pt. X M 0 : 0

100

Corrected engine airflow (percent)

Fig. 10.3 Typical

subsonic inlet total pressure

ratio. 2

which can be expressed in terms of the corrected mass flow as

Dth=~ 4 ~

t~tco max

2116

MFP@M

= 0.8

and reduces to

Dth=0.1636~ or

Dth=l.15OvFA~max

(10.3)

where///c0max

has units of Ibm/s,

Dth

has units of ft, and

A~max

has units of

ft 2.

Figure 10.4 is a representative output from the engine performance (off-design)

portion of the AEDsys computer program showing the corrected mass flow for a

high-bypass ratio turbofan engine with the engine's cruise envelope superimposed.

A figure like this can be used for selecting the point of maximum corrected mass

flow that determines the throat diameter

(Dzh).

100

Fig. 10.4

10kft 20 kfi 30 kfi

Typical subsonic engine airflow requirements.

424

AIRCRAFT ENGINE DESIGN

20 ° -

~A/C

10 o-

0 o

0.2

,'~ VTO =

1.2

VSTAL L

,j 2 ° margin

1.6 V Maximum Operational

~ Requirement

~, ~ Normal Operation

I I I I I I I

0.3 0.4 0.5 0.6 0.7 0.8 0.9

Fig. 10.5 Variation in angle of attack. 2

10.2.2.3 Effect of aircraft attitude and crosswind. During an aircraft's

flight, the angle of attack of the aircraft can vary substantially, which will change

the angle of the airflow with respect to the centerline of the engine. The largest

variations normally occur during the takeoff rotation and landing phases of flight.

Figure 10.5 shows the variation in angle of attack for a typical subsonic passenger

aircraft. Note that, during normal operation, the major changes in angle of attack

occur between 1.2

VSTALL and 1.6 VSrALL.

The design of the inlet can compensate

for this variation in aircraft angle of attack by inclining the face of the nacelle and

changing the shape of the inlet as shown in Fig. 10.6. Note the difference in shape

of the inside inlet contour of top and bottom that is required for high angle of attack

and cruise operation. The incidence of the nacelle face can be estimated 2 using

Olna c -~

0.5 deg + 1.4 deg

OtA/c(IOw

speed aircraft) (10.4a)

Otnac = 0.5 deg + 1.1 deg

OlA/C

(high subsonic speed aircraft) (10.4b)

where

OlA/C

is the cruise angle of attack for the aircraft.

The major effect of crosswind is flow separation inside the inlet with the resulting

flow distortion and possible fan stall. This effect can be reduced by making the

leading edges on the side of the inlet thicker to minimize flow separation with

crosswinds.

The ratio of

Dhl/Dth

for the top, sides, and bottom of the inlet are influenced by

takeoff performance, angle of attack capability, and crosswind capability. Design

values for

Dhl/Dth

range between 1.10 and 1.16 for conventional takeoff and

landing subsonic aircraft. 2

selage Center Line

_ 0- --- __

Fig. 10.6 Inclined nacelle face.

DESIGN: INLETS AND EXHAUST NOZZLES 425

Run -470

Mach No. = 0.30

c~ nacelle = 28.0

Corrected engine flow = 1627 lbm/sec

-10~

Forward looking aft 0

-12 ~20

Steady-state pattern Dynamic pattern

(Ae/P)max = 0.167 (AP/P)max = 0.227

Typical steady-state and dynamic total pressure distortion patterns. 2

Fig. 10.7

10.2.2.4 Inlet flow distortion. Inlets operated at high angles of flow inci-

dence are susceptible to flow separation from the inside contour of the inlet. This

flow separation causes large regions of low total pressure as shown in Fig. 10.7.

The magnitude of this distortion from the desired uniform flow is measured by the

term called "inlet distortion" and calculated by

Inlet Distortion

(Pt

max - et

min)/Pt avg

(10.5)

Both the instantaneous (dynamic) and time-averaged (steady-state) distortion

levels are used to measure the quality of an inlet's flow. When using Eq. (10.5) for

calculation of dynamic inlet distortion, the average total pressure in the denomina-

tor is the spatial average at an instant in time and the maximum and minimum total

pressures are for that same instant in time. Determination of the steady-state inlet

distortion requires time averaging of the total pressures in the inlet. The average

total pressure (Pt avg) of Eq. (10.5) is the spatial average of the time averaged total

pressures.

The magnitude of the inlet distortion is a function of the inlet's geometry, mass

flow rate, flight Mach number, and flow incidence angle. The importance of each

of these variables on inlet distortion is shown in Figs. 10.8 and 10.9. The effect

of high distortion is to shift the fan or compressor surge line to values of lower

pressure ratio and/or higher mass flow rate as shown in Fig. 10.10 (see Sec. 5.3.2).

10.2.2.5 Inlet drag and engine-out drag. Detailed design of the nacelle

geometry is beyond the scope of this textbook and this material is included only

to give a qualitative understanding of several nacelle design issues. The size of

the nacelle forebody (Dhl/Dmax and L/Dmax) is a design compromise between

the requirement of low cruise drag and avoiding catastrophes when one or more

426 AIRCRAFT ENGINE DESIGN

0.2 --

Steady-State

Inlet Distortion

/ Mo=O

(a=O)

/ M o =0.85

Dist.<O.1

/ / (a=2-4 °)

n Point

Corrected Engine Mass Flow Rate

Fig. 10.8 Typical variation in inlet distortion with Mach number. 2

0.20

Steady-State

0.10'

Inlet Distortion

(z,,a, M 0 = 0.3

11~: .,~ Distortion

A 22 °

A o / ~ Pattern of

u/ wF~g'10"6

0 I

I I

1400 1500 1600 1700

Corrected Engine Mass Flow Rate

(Ibm/s)

Typical variation in inlet distortion with angle of attack?

Fig. 10.9

Surge line (clean) /-~j~. 1 t ~"

~jf/ j"

f~J

~ // ~ Surge line

/ (with distortion)

Corrected Mass Flow Rate

Fig. 10.10 Effect of inlet distortion on the fan or compressor surge line.