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

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674 AIRCRAFT ENGINE DESIGN

engine start and augmentor light-off. The power conditioning unit converts the

alternator alternating current output to direct current and steps the voltage up or

down to meet the needs of the other electrical loads on the engine. Because digital

controls require continuous power, modern engines typically bring in an aircraft

power bus to serve as backup in case of alternator failure.

A standard engine sensor set includes devices to measure temperatures, pres-

sures, rotor speeds, and actuator positions. Specialized diagnostic sensors (e.g.,

vibration, lubrication system chip detection) may also be included. The specific

temperatures and pressures measured may vary slightly from engine to engine,

depending on the engine manufacturer and the control strategies each employs.

A typical sensor set for a two-spool turbofan will measure engine inlet temperature

and pressure

(Tt2

and

Pt2),

compressor inlet temperature (T/2.5), bypass duct pres-

sure (Pt16), combustor inlet pressure (Pt3), tailpipe temperature (Ttr), and position

feedbacks from the compression system variable stator actuators, fuel metering

valve, and exhaust nozzle actuators. Because of the harsh environment, hot section

sensing between the high-pressure turbine inlet and low-pressure turbine exit is

generally avoided or limited to a non-intrusive method such as optical pyrometry.

Optical pyrometers are not accurate enough to use for control purposes, but are

useful as limiters to avoid excessive temperatures in the turbine system.

The heart of the electrical subsystem, indeed of the control system, is the digital

electronic controller. Depending on the manufacturer and engine, it may be re-

ferred to as a DEC (digital electronic controller), DEEC (digital electronic engine

control), DECU (digital electronic control unit), or FADEC (full authority digital

electronic control). A typical FADEC (Fig. 0.8) consists of a computer running

a program that implements the control algorithms described previously, analog

sensor signal conditioning circuitry with analog-to-digital converters to feed in-

formation into these control algorithms, and output signal conditioning circuitry

interfaced to digital-to-analog converters that turn the computed commands into

the current and voltage levels necessary to drive valves and actuators to the required

positions. A modern FADEC will take readings, compute commands, and adjust

actuator and valves settings up to 80 times per second. Because of the critical role

the FADEC plays in the operation of the engine, digital control is typically imple-

mented in a redundant architecture. Redundancy is usually limited to the electrical

subsystem, since the mechanical portion of the control system is highly reliable.

Millions of hours of field experience have demonstrated that dual, or duplex, redun-

dancy raises electrical subsystem reliability to a level commensurate with that of

the mechanical components. More recent military engines also have some type of

Fig. 0.8 Full authority digital electronic control (FADEC).

APPENDIX O: ENGINE CONTROLS 675

health or maintenance diagnostic unit interfaced with the FADEC. Future engines

will employ predictive or prognostic sensor and data processing to increase safety

and improve logistic support. The level of sophistication ranges from simple data

logging (limit exceedances, occasional performance snapshots, cycle counting,

time at temperature) to complex health tracking with Fast Fourier Transform

(FFT) based vibration analysis, automated failure diagnosis, and continuous per-

formance trending. Although FADECs and diagnostic units are similar in many

respects, there is one critical difference. FADECs are flight-critical (sometimes

called safety-critical) devices, i.e. loss of FADEC function directly impacts safety

of flight. Failure of a diagnostic unit will cause the loss of important maintenance

information, but has no bearing on flight safety. Therefore, redundancy require-

ments are not imposed on the diagnostic system in order to save weight and cost.

0.2.2 Fuel Defivery Subsystem

The fuel delivery subsystem consists of one or more pumps, filters, heat ex-

changers, valves, and plumbing. Its primary functions are to deliver fuel to the

main combustor (and angmentor, for afterbuming engines) at the appropriate mass

flow rate and to cool, either directly or indirectly, engine or aircraft components.

These tasks are complicated by the temperature limit of the fuel, around 350°F for

the kerosene-based Jet-A, JP-5, and JP-8 fuels used in commercial and military

engines. By the time fuel reaches the engine's main fuel pump from an aircraft

tank, it has already cooled various vehicle subsystems (Fig. 0.9). Although current

design practice is to keep fuel temperature below 200°F at the engine interface,

higher temperatures are not uncommon. Between the pump discharge and com-

bustor fuel nozzles, fuel is routed through a heat exchanger to cool engine oil.

When the maximum operating temperature of approximately 350°F is exceeded,

various compounds precipitate from the fuel. These solids deposit in fuel lines

and nozzles, leading to performance deterioration and ultimately engine failure.

At certain flight conditions maintaining fuel below this critical temperature at

the main combustor nozzles (and angmentor spray bars in afterburning turbo-

fan engines) is extremely difficult. This challenge has become more severe on

newer military fighter aircraft, leading to the need for a disciplined systems engi-

neering approach to the design of the aircraft fuel delivery system to ensure that

Fuel Tanks

T_< 160~F

Rectrculafing.Fuel,

T = 200 ° F

I t

Fuel

T~ 200°F

Fig. 0.9 Typical aircraft fuel delivery system.

676 AIRCRAFT ENGINE DESIGN

fuel temperature is kept under control at all operating conditions. This has led

to the search for fuels that can absorb more heat before destabilizing, as well as

endothermic fuel that can absorb additional heat by cracking. Also fuel may be re-

circulated to the aircraft tanks during periods of high heat rejection (see Fig. 0.9).

For such systems pump thermal performance, that is, the change in fuel temperature

between pump inlet and discharge, is critical. Variable geometry fuel pumps that

minimize fuel heating over the entire flow range have been developed to address

this critical thermal management issue.

An engine fuel delivery system is required to supply the required mass flow of

fuel at sufficient pressure to ensure proper mixing with compressor discharge air

in the combustor. The main fuel pump is the critical element in this system. It is

responsible for taking low pressure (less than 100 lbf/in. 2) fuel from the aircraft fuel

delivery system and driving it through the engine fuel system such that it enters the

combustor fuel injectors at much higher pressure, often in excess of 1000 lbf/in. 2

Today's engines use various types of fixed geometry pumps to deliver main fuel

flow to their combustors. Gear-type main fuel pumps are the most commonly

used because of their simplicity and ruggedness. Centrifugal pumps have also

begun to appear in newer engines. A fuel metering valve is located downstream

of the pump and is used to vary flow to a manifold that supplies the combustor's

individual fuel nozzles. The flow range of the main fuel pump and metering valve

is related to the speed range of the pump drive system, contained in the engine

accessory gearbox. Because of the relatively small variation in engine core speed,

and therefore gearbox drive speed, compared to the huge difference in fuel flow

between idle and full power, it is impossible to match delivered fuel flow to required

fuel flow at all power settings. Thus, all engine fuel systems include a recirculation

loop that takes flow in excess of requirement back either to the pump inlet or to

the aircraft tanks. In addition, afterburning engines include a separate loop in their

fuel systems to deliver fuel to their augmentors. This loop includes its own pump

and plumbing, terminating at a manifold that feeds the augmentor spray bars.

0.2.3 Actuation Subsystem

The actuation system consists of valves and actuators required to manage fuel

flows, secondary and bleed airflows, and variable stator positions. Depending on

the type of engine, turbine system clearance, nozzle throat area, and thrust vec-

tor angle are other potential control parameters of interest. Actuation power is

supplied by readily available sources on the engine, usually hydraulic (from the

high pressure fuel supply) and occasionally pneumatic (compressor interstage or

discharge bleed).

Linear actuators are used to position compression system variable stators and

nozzle flaps (for engines with variable area exhaust). This device consists of

a servo-valve controlled by a torque motor mounted on a cylinder and piston

(Fig. O. 10). In certain applications the servovalve may drive more than one cylin-

der. The servovalve modulates flow of high pressure hydraulic fluid (fuel) to either

side of the piston head to control displacement from a predefined null position.

Control of the actuated variable is accomplished by changing the position of the

piston, which is mechanically linked to the control effector. For example, vari-

able stator position is controlled by displacement of an actuator piston rod that is

APPENDIX O: ENGINE CONTROLS 677

Fig. 0.10 Inlet guide vane actuator.

mechanically linked to a "synch" ring. The synch ring is mounted to the outer case

of the engine and has individual tabs or levers that are connected to shafts that

extend from each stator. Linear displacement of the actuator piston causes rotation

of the synch ring, which in turn causes rotation of the variable stators.

Turbine system clearance control, used in some high bypass commercial tur-

bofans, uses an airflow modulation approach as opposed to the direct mechanical

positioning method just described. This typically involves using a valve to route

"cool" (relative to the turbine section) compression system bleed air across the

turbine case to cause it to contract. The contraction reduces airfoil tip clearance,

resulting in reduced specific fuel consumption via increased turbine system effi-

ciency. This is done during cruise portions of the mission, where constant throttle

settings are the norm for commercial engines.

0.3 Summary

This discussion addressed the basic and fundamental elements of the gas tur-

bine engine control system. Developing control modes and logic for gas turbine

engines have evolved over many decades of research and experiment from simple

speed govemors to more elaborate, multivariable functions. As the application of

digital electronic control systems to commercial and military gas turbine engines

has become more accepted and widespread, more sophisticated, performance en-

hancing control modes and logic have, likewise, evolved. The ability to optimize

engine performance at various flight conditions is now possible. Adaptive control

modes that can optimize engine performance based on a variety of requirements

and conditions, such as the particular mission being flown or the health of the

engine, are being aggressively pursued.

Reference

1D'Azzo, John J., and Houpis, Constantine H., Feedback Control System, Analysis and

Synthesis, McGraw-Hill, New York, 1966.

Appendix P

Global Range Airlifter (GRA) RFP

This abbreviated Request for Proposal (RFP) is presented as the second design

example of this textbook. It is based on anticipated global airlift needs of the

Department of Defense and commercial carriers, including passengers and cargo.

The solution is carded out on the accompanying CD-ROM.

P,1 Background

The C-5 Galaxy was developed in the 1960s to provide strategic airlift needs

during the Cold War where the majority of material was prepositioned. The first

high-bypass turbofan engine, the General Electric TF39, was developed to power

this aircraft. Without aerial refueling, this aircraft can fly about 2,000 n miles with

its maximum payload of 290,000 lbf and 6,300 n miles empty.

Recent hostilities such as Desert Storm, Bosnia, and Afghanistan demonstrated

that a new Global Range Airlifter (GRA) that could fly without being refueled

anywhere in the world from the continental United States was needed. The follow-

ing section lists specific requirements that will be the basis of the design exercise.

Please note that several of the requirements have been imposed in order to allow the

GRA to meet civilian flight restrictions, reflecting the modem reality that military

aircraft are seldom exempt from peacetime considerations.

P.2 Requirements

1. Maximum gross takeoff weight (GTOW) of 1,000,000 lbf.

2. Minimum payload of 150,000 lbf.

3. Minimum unrefueled range of 10,000 n miles.

4. Maximum fully loaded takeoff distance of 10 kft (no obstacle) at sea level on

a hot day (90°F).

5. Maximum landing distance (50% reverse thrust) of 6 lift at sea level on a

standard day.

6. Maximum takeoff distance (no payload, return fuel load) of 6 kft at sea level

on a hot day (90°F).

7. Maximum one engine out takeoff distance (no payload, 2,000 n miles fuel

load) of 6 lift at sea level on a standard day.

8. Maximum fully loaded distance for climb to 35 kft altitude of 200 n miles on

a Mil Std 210 hot day.

9. Minimum fully loaded takeoff climb gradient with one-engine out of 3 deg

at sea level on a hot day (90°F).

10. Minimum cmise ceiling of 35 lift at 95% GTOW on a standard day.

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