[American Institute of Aeronautics and Astronautics AIAA Infotech@Aerospace Conference - Seattle,...

American Institute of Aeronautics and Astronautics

1

Engine Controls for Off-Nominal Aircraft Operation

Dr. Walter Merrill1, George Mink

2, Hoang Tran Van

3 and Dr. Link Jaw

4

Scientific Monitoring, Inc, Scottsdale, Az, 85258

Transport aircraft engines are designed to meet full-life nominal-usage aircraft

specifications. For situations where the aircraft faces adverse conditions (control surface

failures, runway incursions, etc.), the engine offers the potential to augment aircraft

controllability, by providing differential thrust for steering or additional thrust to increase

climb rate thereby improving flight safety and survivability of an aircraft under abnormal or

emergency conditions. In these conditions, the fast-response controller needs to boost (or

recover) the engine capability by letting it go beyond normal physical and operational limits

for a relatively short period of time until the aircraft lands safely. The enhanced engine

capability is primarily higher and faster thrust; however, it must be produced by balancing

the operability and remaining life of critical engine components. Hence, the goal is a

controller that enables higher and faster thrust with acceptable operability and remaining

life margins. This paper will identify the ability of various existing and new actuators to

effect control beyond current limits, to assess the risk of operating outside of traditional

limits and to design controls that can effectively balance the need for emergency control

effectiveness against operational risks.

Nomenclature

A4 = Turbne Exit Area

A45 = Intermediate Turbine Exit Area

ABV = active bleed valve

ACC = active clearance control

APU = Auxiliary Power Unit

ARMD = Aeronautics Research Mission Directorate

ASC = active stall/surge control

BOM = Bill of Material

BV = bleed valve

DWfP3max = Difference between Wf/P3max and Wf/P3

DWfP3min = Difference between Wf/P3 and Wf/P3min

EGT = Engine Exhaust Temperature

EPR = Engine Pressure Ratio

Fan Bleed = Fan Bleed Flow

FBV = fan bleed valve

HPT = High Pressure Turbine

I/CGV = inlet/compressor guide vane

IRAC = Integrated Resilitent Aircraft Control

LPC = Low Pressure Compressor

MFF = main burner fuel flow

N1 = Low Spool Speed

N2 = High Spool Speed

NASA = National Aeronautics and Space Administration

P2 = Inlet Pressure

1 VP & GM, 8777 E Via de Ventura, Suite 120, Sr. Member

2 VP Engineering, 8777 E Via de Ventura, Suite 120

3 Sr. Engineer, 8777 E Via de Ventura, Suite 120

4 President, 8777 E Via de Ventura, Suite 120

AIAA Infotech@Aerospace Conference <br>and <br>AIAA Unmanned...Unlimited Conference 6 - 9 April 2009, Seattle, Washington

AIAA 2009-1875

Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes.All other rights are reserved by the copyright owner.


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P25 = LPC Exit Pressure

P30 = HPC Exit Pressure

PC = Power Command

PF = Pattern Factor

RCVV = rear compressor variable vane

T/R = thrust reverser

T2 = Inlet Temperature

T25 = LPC Exit Temperature

T3 = HPC Exit Temperature

T4 = Burner Exit Temperature

Thpt_blade = HPT Blade Temperature

VBC = variable bypass control

VSV = variable stator vane

W/GI = water/gas injection

Wf = Fuel Flow

Wf/P3 = Fuel-Pressure Schedule

Wfan = Fan Airflow

I. Introduction

he goal of the NASA IRAC program is to arrive at a set of validated multidisciplinary integrated aircraft control

design tools and techniques for enabling safe flight in the presence of adverse conditions (for example, faults,

damage and/or upsets). The specific objective of this reported research is to develop control algorithms for fast

engine response during emergency operations.

One of the critical engine characteristics that will have a major impact on emergency aircraft recovery from adverse

conditions is the response time of the engine. This effort will focus on research into engine responsiveness. Engine

responsiveness has been shown to be a critical issue for maneuvering an airplane that has lost its hydraulics and thus

its flight control system. More responsive engines will allow a pilot to land an airplane more easily under emergency

conditions. However, responsiveness comes at a price, both in terms of engine component life consumption and

engine operability (stalling). Thus the research shall: 1) study the effects of possible engine control options including

the engine control limits (such as EGT, fan and core speeds, acceleration and deceleration limits, and stall margins)

and available actuators (such as variable geometries and different bleeds); 2) study the impact of the over thrust

operation on the engine component life, and 3) develop the control algorithms that will be able to extend the engine

operation beyond the nominal operation conditions.

This paper addresses the development of advanced control algorithms for a commercial-type high-bypass engine

that include control actions to address fast engine response. Additionally, the work shall incorporate an analysis of

the risk involved in undertaking these control actions.

The overall objective of the proposed research is to develop a generic, enhanced engine controller to enable effective

thrust management with acceptable risk under abnormal conditions. The enhanced engine control will

1. be able to extend engine operations beyond normal operating conditions and limits

2. include unconventional control modes and control laws to enable fast thrust response in engines

3. facilitate the study of the effect of over thrust conditions on engine component life

4. facilitate design and development of new adaptive control strategies and higher-level decision or

optimization concepts by other researchers and engineers

This paper will describe a candidate control architecture as well as some enhanced control results for two adverse

scenarios. The result will demonstrate potential for substantially increased thrust response. Also included in the

paper will be a discussion of possible actuation and emergency control modes which could be used to address engine

operation in adverse situations.

T


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II. Candidate Control Architecture

SMI has investigated two candidate control architectures. These are shown in Figure 1 and Figure 2. Note that

both architectures have several elements in common. Both architectures have separate vehicle and engine risk

assessment and health management capabilities. How information is exchanged between engine and airframe

control represents the key differences in the two architectures.

Adaptive Flight

Control

Engine

Control

Control Law Actuators

Sensors

Engine or

Simulator

Health Assessment

EHM Sensors

Eng. Risk

Assessment

Control Mode

Perf. Target

EHM

Vehicle Health

Mgt.

-

+

Vehicle Risk Mgt.

Limit Offsets

Adaptive Flight

Control

Engine

Control


Sensors

Engine or

Simulator

Health Assessment

EHM Sensors

Eng. Risk

Assessment

Control Mode

Perf. Target

EHM

Vehicle Health

Mgt.

-

+

Vehicle Risk Mgt.

Limit Offsets

Figure 1. Centralized engine control architecture.

The architecture of Figure 1 assumes that the overall assessment of vehicle condition and operational risk will be

performed by the flight control. The flight control will then provide specific limits, control mode definitions and

performance targets to the engine control. Since the critical decision making is handled by the flight control, this is

referred to as a centralized architecture.

An alternative viewpoint is the architecture of Figure 2. Here, the flight control continues to provide an assessment

of vehicle condition but will provide a level of “risk tolerance” to the engine control based upon that assessment.

From this situational risk tolerance the engine control will decide how best to adapt or modify its control modes,

which limits to modify and which performance targets to achieve. Because important decision making about engine

operation and control is contained within the engine control this is referred to as a distributed architecture. Implied

in this architecture is a hierarchical decision making structure.

There are valid reasons for either approach. The centralized approach would provide the more ‘optimal’

performance approach. However, computational demands might result in slow response in the centralized approach.

This points to the advantage of the decentralized approach which would enable parallel resources to be applied to

provide more immediate response to changing conditions. At this point in the study there is insufficient design

experience to conclusively select one architecture. However experience with engine control systems suggests that

the decentralized approach is the more likely implementation approach and, thus, will be the architecture assumed in

this project.


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Adaptive Flight

Control

Engine

Control


Sensors

Engine or

Simulator

Health Assessment

EHM Sensors

Eng. Risk

Assessment

Thrust Target

EHM

Vehicle Health

Mgt.

-

+

Vehicle Risk Mgt.

Situational Risk

Tolerance

Control Adapter

Selected Mode Selected Actuators

Adaptive Flight

Control

Engine

Control


Sensors

Engine or

Simulator

Health Assessment

EHM Sensors

Eng. Risk

Assessment

Thrust Target

EHM

Vehicle Health

Mgt.

-

+

Vehicle Risk Mgt.

Situational Risk

Tolerance

Control Adapter

Selected Mode Selected Actuators

Figure 2. Distributed engine control architecture.

III. Candidate Fast Response Thrust Control Actuation Options

There are several possible actuation options for fast engine response and these are given in Table 1. The

actuator options are defined for all of the major engine components and are given for three classes of actuation, 1)

existing, 2) higher response (but contained within existing size and footprint constraints) and 3) new or advanced

actuation options.

Table 1 Fast thrust response actuation options

Existing Commercial

Engine Actuation

Higher Resp Actuation

(in Existing Package)

New or Advanced

Actuation

Fan IGV, VSV, IGV, VSV, W/GI ASC, VBC, FBV

Compressor CGV, RCVV, BV, ACC ABV, CGV, RCVV, ABV,

ACC

ACC, ASC, Aspirated Tip,

water injector, gas injector

Combustor MFF MFF PF, Active Comb Ctrl

Turbine ACC ACC ABV, ACC, A4, A45

Mixer VBC

Nozzle T/R T/R T/R

Engine

System

Air turbine/Starter Generator

Motor, APU

Air turbine/Starter

Generator Motor, APU

Although we first believed that modified or advanced actuation would be necessary to achieve desired engine

operation, our initial simulation results indicate that project objectives will be achieved with existing fuel actuation

capabilities and variable compressor geometry actuation. However, either higher response or new capabilities will

be required for improved bleed response actuation (particularly in the high bypass fan) to achieve some of the

proposed candidate emergency control modes. Finally, there may be some utility to studying the improvements

possible with active compressor control. Because the major operational risk element in all of the rapid thrust

scenarios is compressor (or fan) stall/surge avoidance, developing an assessment of the value of an active


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compressor control mode to alleviate stall or surge during emergency maneuvers would enable an informed future

programmatic decision on the necessity for new or advanced active compressor stall control actuation.

IV. Actuator Effectiveness

In this section the effectiveness of the actuators for fuel flow and fan bleed is determined through simulation. To

simulate a propulsion controlled landing scenario, sea level, static standard day conditions were assumed. Also PLA

was nominally set at 60 degrees. The engine control was overridden and two tests were conducted. First was a

direct perturbation of the fuel flow. Second, was a direct perturbation of fan bleed flow. During these tests, all

inputs other than the perturbed input were held constant to isolate the effect to either fan bleed or fuel flow.

In each case the results presented are obtained by performing the fuel flow or fan bleed change as input and

recording all major engine performance changes in steady-state as outputs. The engine parameter names and

descriptions used in this effectiveness study are defined in the nomenclature section.

At the end of each group, there are three (3) calculated parameters used as performance metrics, and they are listed

as follows.

a. Delta ……...Parameter change due to input change

b. Delta (%)….Parameter change due to input change in %

c. Ratio………%Delta Parameter divided by %Delta Input

The data for the fan bleed case are given in Table 2 and for the fuel flow case in Table 3

.

Table 2 Fan bleed perturbation case

Wfan Wf EPR N1 N2 T25 T3 T4 Thpt_blade

Unit (lb/s) (pph) (none) (rpm) (rpm) ( R ) ( R ) ( R ) ( R )

SS Value 1025 2.584 1.299 3793.0 11010.0 692.0 1353.0 2788.0 2252.0

Perturbed Value 1028 2.584 1.300 3808.0 10832.0 694.1 1355.0 2789.0 2253.0

Delta 3 0.000 0.001 15.00 -178.00 2.10 2.00 1.00 1.00

Delta(%) 0.29 0.00 0.077 0.40 -1.62 0.30 0.15 0.04 0.04

Ratio 0.10 0.00 0.03 0.13 -0.54 0.10 0.05 0.01 0.01

EGT P25 P30 Wf/P3 DWfP3max DWfP3min VSV Fan Bleed Net Thrust

Unit ( R ) (psia) (psia) (pph/psia) (pph/psia) (pph/psia) (deg.) (lb/s) (lbf)

SS Value 1389.0 36.0 316.0 29.40 15.59 14.21 9.47 205 23410.0

Perturbed Value 1390.0 36.3 316.1 29.42 14.93 15.71 9.47 211 23276.0

Delta 1.00 0.33 0.10 0.02 -0.66 1.50 0.00 6 -134.0

Delta(%) 0.07 0.92 0.03 0.07 -4.23 10.56 0.00 3.02 -0.57

Ratio 0.02 0.30 0.01 0.02 -1.40 3.49 0.00 1.00 -0.19


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Table 3 Fuel flow perturbation case


Unit (lb/s) (pph) (none) (rpm) (rpm) ( R ) ( R ) ( R ) ( R )

SS Value 1025 2.584 1.299 3793.0 11010.0 692.0 1353.0 2788.0 2252.0

Perturbed Value 1023 2.571 1.296 3785.0 10858.0 691.4 1352.0 2789.0 2254.0

Delta -2 -0.013 -0.003 -8.00 -152.00 -0.60 -1.00 1.00 2.00

Delta(%) -0.20 -0.50 -0.231 -0.21 -1.38 -0.09 -0.07 0.04 0.09

Ratio 0.39 1.00 0.46 0.42 2.74 0.17 0.15 -0.07 -0.18


Unit ( R ) (psia) (psia) (pph/psia) (pph/psia) (pph/psia) (deg.) (lb/s) (lbf)

SS Value 1389.0 36.0 316.0 29.40 15.59 14.21 9.47 205 23410.0

Perturbed Value 1390.0 35.8 314.2 29.46 14.99 15.54 9.47 205 23278.0

Delta 1.00 -0.13 -1.80 0.06 -0.60 1.33 0.00 0 -132.0

Delta(%) 0.07 -0.36 -0.57 0.20 -3.85 9.36 0.00 0.00 -0.56

Ratio -0.14 0.72 1.13 -0.41 7.65 -18.60 0.00 0.00 1.12

The normalized effectiveness (Ratio values in the tables) for the various engine parameters is plotted in Figure 3 and

Figure 4. Here you can clearly see the effect that each of the inputs has on various engine parameters and their

relative strengths. As expected, fuel flow is a more effective control input. However, fan bleed has considerable

control authority over key variables. This, along with the ability to rapidly modulate fan bleed, makes fan bleed an

important control variable.

-1.00

-0.50

0.00

0.50

1.00

1.50

2.00

2.50

3.00


No

rma

lize

d e

ffec

tiv

en

es

s

Fan Bleed Case

Fuel Flow Case

Figure 3. Normalized effectiveness - part 1.


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-20.00

-15.00

-10.00

-5.00

0.00

5.00

10.00


No

rmalized

efe

cti

ven

ess

Fan Bleed Case

Fuel Flow Case

Figure 4. Normalized effectiveness - part 2.

V. Candidate emergency control modes

Eight candidate emergency control modes are considered and shown in Table 4. The table

provides a brief description of the mode, an indication of the major technology challenge

required to implement the mode, an assessment of the risk level and the overall thrust objectives

of the mode. For example mode 1 is stall margin estimation. This is essentially the notion of

using improved information about the remaining stall margin to allow closer operation to stall

and thus to enable faster acceleration or deceleration. However, there are different methods to

implement margin estimation and this control mode. Other modes exhibit similar generality.

An initial, limited assessment of risk is provided for each mode. This assessment exists at only

the most superficial level and much additional work will be required to appropriately evaluate

these risks in a realistic context. However, it is believed that the assessment represents a good

first step in opening a discussion about risks and benefits. The chart is color coded to reflect

risk and the key to the risk assessment is

Color Code Risk Level

Red High

Yellow Medium

Green Low


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Table 4 Candidate Emergency Control Modes Emergency

Control Mode Mode # Control Mode Description Technology Challenge Risk Assessment

Thrust

Objective

Stall Margin

Estimator

1 Use Delta P/P or HPC Mach Est. to estimate

actual Stall Margin and use to provide improved

Acceleration Schedule by operating closer to

stall. May also use measurements to reduce

stall margin by eliminating portions of margin

"stack up" during emergency operations and

allowing accelerations closer to stall

Reliable stall margin

estimation

Operating closer to stall/surge would

increase the likelihood of such an

event. Encountering a surge during

emergency operation would probably

result in a Major accident

Increase

acceleration

and

deceleration

response

Gas Injection 2 Use high pressure gas injection at compressor to

accelerate rotor

The impact of high

pressure injection is

not well understood

Potential stall/surge initiation site. See

above.

Increase

acceleration

and

deceleration

response

Variable Thrust

Reverser

3 Use continuously variable thrust reverser to

provide 'negative' thrust capability for high

response, high differential thrust two engine

operation

Reliable and low

weight actuation

TRs have been deployed in-flight

accidentally and short term effects are

known. Long term operability effects

unknown. Hot Gas reingestion, e.g.

Increase

acceleration

and

deceleration

response and

delta thrust

Reduced

Temperature

Margin

4 Allow the turbine to operate hotter to increase

total thrust

Improved turbine

engine life estimation

Encountering blade deformation and/or

minor melting could be tolerated during

an emergency operation. However,

disk failure would cause substantial

engine damage and possibly result in a

Major accident

increase thrust

Mass Injection 5 Use either water or gas injection to lower cycle

temperatures

The weight penalty

incurred

Water injection has been used in

engines with success. Water injection

would have limited lifetime utility

increase thrust

Rotor Torque

Augmentation

6 Use either an air turbine or electric motor

(possibly an integral starter/generator) to add

acceleration or deceleration torque to the engine

rotors. Power for this may come from APU or

from a low speed spool generator.

The actuator and

power source would be

heavy and complex.

Integral starter/generators exist.

Interaction with aerothermodynamics

of engine unknown

Increase

acceleration

and

deceleration

response

High Speed

Flight Idle

7 Enable higher rotor speed operation at flight idle

(or other glide condition) by "dumping"

power/thrust (bleeds or TR) to enable faster

accelerations

Localized overheating

or distortion due to

dumping large

amounts of airflow

Similar to long term TR operation Increase

acceleration

and

deceleration

response

Improved BOM

Modes

8 Use higher response, existing actuation to

provide more responsive performance and

tighter limit protection

Heavier and more

costly higher response

versions of existing

actuation

Well understood risks/benefits Increase

acceleration/d

eceleration

and thrust

Mode 1. Stall Margin Estimation

In this mode either measurements or estimates of ∆P/P, HPC Mach number or airflow are used to estimate actual

Stall Margin and this information is used to provide improved acceleration rates by operating the engine closer to

the stall line. Control feedback about the “distance” of the compressor from stall or indicators and precursors of

stall, may be incorporated into the control modes to enable, safe, reduced stall margin operation. Another, related

approach is to use these measurements/estimates to reduce stall margin by eliminating portions of margin “stack up”

during emergency operations. The margin stack includes elements associated with engine deterioration, engine inlet

distortion, manufacturing variability and other effects. These element margins are normally added together to

achieve a conservative total margin. It may be feasible to incur some additional operational risk during an

emergency situation in the following ways. First, if an aircraft encounters an emergency situation and has relatively

new engines, the margin element associated with deterioration may be eliminated. Second, if aircraft maneuvers

will be restricted during an emergency situation, it may be feasible to reduce the distortion element portion of the

margin.

The risk associated with this mode was judged to be high due to the difficulty in generating a reliable and accurate

estimate of surge and stall margin for both the fan and compressor.

Nevertheless, because of the benefit to improved operability both in emergency and conventional control modes, the

mode was one of those down-selected for additional study.

Mode 2. Gas Injection


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In this mode high pressure gas injection into the low pressure compressor is used to increase local pressure, freeing

up some of the turbine rotor torque to accelerate the fan rotor. One possible approach is to use a storage container to

accumulate high pressure air from the turbine for use in a burst or transient mode when required. Of course the

repetitive usage of such an approach would depend on accumulator size and acceleration requirements. This mode

was assessed to be high risk because injection pressures on the scale required to impact rotor acceleration would

have a serious, negative impact on compressor stability. Gas impingement on a compressor blade may make it

rotate backwards. Injection may also cause stages before the gas injection to go toward surge. This combined with

the practical issue of the availability of a useful, high pressure air source makes this mode one with a low probability

of success.

Due to the low probability of success, this mode was not selected for additional study.

Mode 3. Variable Thrust Reverser

In this mode continuously variable thrust reversers are used to provide “negative” thrust capability for high

response, high differential thrust, two engine operation. Additionally, if the engine were to be operated at an over

thrust condition with the reverser “spilling” some of the thrust, both positive and negative, high response, variable

thrust could be achieved. Regarding the thrust reverser, although the engine does not accelerate or decelerate

faster, this mode has the effect of creating differential moments about the yaw and partially the pitch and roll axes,

depending upon the aircraft. The relative position of the engines versus the center of gravity of the airplane will

determine how large the differential moments can be. If we can use the thrust reverser and modulate it successfully,

this means that we can control the airplane attitude with or without varying the engine operating conditions. Usually,

it takes a couple of seconds to deploy the thrust reverser (Nozzle deployment) and the reversed thrust can be

modulated by varying the mass flow via a series of internal bleed valves. If the nozzle is deployed and ready to be

used (this will create an additional drag and eliminate the delay time), we can rapidly modulate the backward thrust

by modulating the internal bleed valves.

This approach was judged to be a medium risk one since there is substantial experience with thrust reversers in

current commercial operations. Additionally, there is anecdotal experience as thrust reversers have been deployed

in-flight accidentally. The short term effects of thrust reversing are known, however the long term operability

effects (as during a sustained emergency scenario) on propulsion/airframe integration of redirecting thrust are

unknown. Potentially, there may be hot gas re-ingestion or structural heating due to long duration thrust reversal.

A variation on this approach would be to exhaust gas from the engine directly via bleeds. Using fan bleed, in

particular, will be investigated as part of the High Speed Flight Idle mode.

The variable thrust reverser mode will not be considered for additional study.

Mode 4. Reduced Temperature Margin

In this mode the engine’s turbine is allowed to operate at higher temperatures principally to enable the engine to

achieve a higher total thrust but also to accelerate the engine and provide faster thrust response as well. The key

technology development required to implement this mode will be an improved understanding of engine life

estimation. Substantial work is ongoing to develop improved turbine life modeling and this project will take

advantage of those efforts. Operation risks to the engine include encountering blade deformation and/or minor

melting of the blades and eventually total disk failure. It is assumed that minor deformation and melting could be

tolerated during an emergency operation. However, disk failure would not be acceptable in any situation. This

mode was judged to be medium risk.

Due to the importance of achieving higher thrust, in particular in runway incursion or “go around” maneuvers, this

mode was selected for additional study.

Mode 5. Mass Injection

In this mode either water or gas injection is used to lower cycle temperatures to increase engine maximum thrust.

Water injection has been used in engines before with success hence this approach is judged to be of low risk.

However, water injection would have limited lifetime utility due to the requirement to store large amounts of water

on each flight.

This mode was not selected for additional study.


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Mode 6. Rotor Torque Augmentation

In this mode either an air turbine or electric motor (possibly an integral starter/generator) would be used to add

acceleration or deceleration torque to the engine rotors. Power for the air turbine or motor may come from the

auxiliary power unit or from an electric generator attached or integrated with the low speed rotor. This approach

was judged to be of low risk because integral starter/generators systems exist. However, their operation is usually

restricted to very low speed regimes. The interaction of such a system, when operated during operational speeds,

with the aerothermodynamics of the engine is unknown. Additionally, the power source for such a system would

heavy and complex.

This mode was not selected for additional study.

Mode 7. High Speed Flight Idle

In this mode we can enable higher rotor speed operation at flight idle (or any other glide condition) by “dumping”

engine airflow by engine bleeds or thrust reversers to enable faster acceleration and deceleration of the engine. As

in the Thrust Reverser Mode, both positive and negative thrust control is possible. This mode was judged to be low

risk because thrust reversers and engine bleed actuators exist today. However, localized overheating or engine inlet

distortion due to dumping large amounts of airflow is a concern. Another way to get lower flight idle thrust is to

lower the idle N2 speed. This approach risks combustor lean blowout. One solution to lean blowout is to shut off

some atomizers keeping only atomizers upstream of the igniters on.

Due to the likelihood of success, this mode was selected for additional study.

Mode 8. Improved Bill of Material Mode

In this mode we will use higher response, existing actuation to provide more responsive performance and tighter

limit protection for the engine. This approach was judged to be of low risk the technology risks and benefits are

well understood. The higher response, continuously variable bleed actuation which likely will be required will be

heavier and more costly than existing actuators. However, since the expected bandwidth and flow requirements are

well within today’s state of the art, no substantial impediments to developing these actuators are expected.

This mode was selected for additional study.

VI. Emergency Control Mode Preliminary Results

In this section we use modeling and simulation of a commercial-type high-bypass engine, and the development of

advanced control algorithms for the engine that address fast engine response. In particular we develop several

different control mode options to determine if the example engine has additional thrust slew rate capability. In

addition to demonstrating additional thrust slew rate, we begin to expose some of the operational issues and risk

associated with obtaining increased engine thrust response. A schematic drawing of the example engine used in this

study, a twin spool high bypass ratio engine (the “Pax200” engine), is shown in Figure 5. This engine is typical of

commercial engines in the 40,000 lbf thrust class.

In this preliminary simulation study, a generic controller (developed by Scientific Monitoring, Inc. under contract to

NASA Glenn Research Center, March 2007) was used to control the engine model and serves as the baseline

control. In order to add realism to the study, the actuator and engine characteristics shown in Table 5 were included

in the simulation model.


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Table 5 Engine actuation and characteristics added to the simulation

Actuators Engine Characteristics

Fuel Valve LPC surge margin

LPC bleed valve HPT blade temperature

Fan bleed valve

Variable Stator Vanes (VSV)

In addition to the baseline actuators defined in Table 5, higher response versions of the fuel valve and VSV actuators

were tested. These higher response actuators had no measurable effect beyond the baseline on engine thrust

response and are not considered further.

Figure 5. PAX 200 Twin Spool High Bypass Ratio Engine – Schematic.

VII. Test Cases

Table 6 shows the different test case scenarios that were simulated in this fast thrust response study. All cases were

run with EPR control and at sea level and zero Mach number conditions (T2=518.67R, P2 = 14.69psia). The table

gives a brief description of the case, the case number which will be used for ease of referral, the engine power

command setting (called PC) and the special conditions which define the unique characteristics of the test.


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Table 6 List of test case scenarios

Case Description Case # PC (deg.) Special Actions

Baseline 1a 80.5-40-80.5 None

Modified Wf/P3

schedule

1b 80.5-40-80.5 Increase the acceleration and decrease the

deceleration limits in the controller

Baseline 2a 80.5-50-80.5 None

VSV Schedule

Modification

2b 80.5-50-80.5 Increase VSV from 12.5 to 24.5 deg. @

PC=50 deg.

Same as 1b 3a 80.5-40-80.5 Increase the acceleration and decrease the


VSV Slam with 1b 3b 80.5-40-80.5 Increase the acceleration and decrease the

deceleration limits in the controller and slam

opened the VSV during acceleration (PC=40

to 80.5)

Baseline 4a 80.5-47.5-80.5 None

Fan Bleed 4b Maintain 80.5 Modulate fan bleed to produce the same

steady state thrust as in case 4a

Same as 1b 5a 80.5-40-80.5 Increase the acceleration and decrease the


LPC Bleed 5b 80.5-40-80.5 Increase the acceleration and decrease the

deceleration limits in the controller. Open

LPC bleed valve during idle & fully close

during acceleration to full throttle

Same as 5b 6a 80.5-40-80.5 Open LPC bleed valve during idle & fully

close during acceleration to full throttle (same

as 5b)

Wf/P3 Schedule,

plus VSV plus LPC

Bleed (1b, 3b, 5b)

6b 80.5-40-80.5 Increase the acceleration and decrease the

deceleration limits in the controller. Slam

opened VSV during acceleration (PC=40 to

80.5) and Open LPC bleed valve during idle

& fully close during acceleration to full

throttle

Case Numbered Transient

Case 1a represents the normal acceleration/deceleration schedule with a generic controller and normal actuator

response rates. Under sea level, static conditions, power command angles PC of 80.5 and 40 degrees correspond to

an EPR demand of 1.458 and 1.021 respectively. In this simulation case, a normal Wf/P3 schedule was used to

prevent engine from surge, over temperature, and flameout. This action has the effect of limiting the engine speeds

and thrust responses. To accelerate to full thrust (Pc=80.5 deg.) from ground idle (Pc=40deg.), the engine needs

about 13.5 seconds.

Case 1b represents the same case as in Case 1a with modified Wf/P3 schedules. This test case corresponds to the

BOM Mode #8 from Table 4. Additionally, modifying the Wf/P3 schedules also allows higher turbine

temperatures. This corresponds to Mode #4, Reduced Temperature Margin. Specifically, the acceleration and

deceleration Wf/P3 schedules have been partially pushed out to give the engine more speed and thrust response. For

this simulation scenario, the engine only needs about 7.5 seconds to produce full thrust, a gain of 6 seconds

compared to case 1a. Note that higher transient T4 and HPT blade temperature rates were experienced and this could

affect turbine component life.

Case 2a represents the deceleration to flight idle (PC=50 deg.) followed by an acceleration to full power (PC=80.5

deg.) with normal schedules. For this case, the engine needs about 5.75 seconds to accelerate to full thrust.


13

Case 2b represents the same case as described in case 2a with modified VSV schedule. This test case is another

variation corresponding to BOM Mode #8. Specifically, during the flight idle (PC=50 deg.) the VSV was increased

from 12.5 to 24.5 degrees in order to raise N2 speed. The rationale of increasing N2 is to raise the speed and thrust

responses of the engine. The EPR response of case 2b has indeed increased over case 2a but the thrust remained

practically the same.

Case 3a represents the same case as described in Case 1b.

Case 3b represents the same case as described in Case 1b with modified VSV schedule. This test case represents a

third variation to the BOM Mode #8. In fact, during the acceleration to maximum power from ground idle, the VSV

was slammed open instead of following a designed opening schedule as in case 3a. This action gives more

responsiveness to the engine speed and thrust with a net gain of 1.5 seconds. However, the HPC surge margin was

sharply reduced momentarily and could cause the engine to surge. An optimum VSV schedule could be defined but

it will be highly dependent upon realistic LPC and HPC surge lines.

Case 4a represents another case of deceleration to flight idle. This time, a PC=47.5 degrees was used for the flight

idle condition (thrust=14,600 lbf) and a PC=80.5 degrees was selected for the full throttle (thrust=36,600 lbf). For

this case, normal acceleration and deceleration Wf/P3 schedules were used that resulted in a thrust response time of

about 6.5 seconds.

Case 4b indicates that a constant PC=80.5 degrees was used but the engine thrust was modulated by operating the

fan bleed. This test case corresponds to the Flight Idle Mode #7. In this case, when the fan bleed was opened (at

time =2 seconds), the N1 speed increased and hit the N1max limit. Consequently, fuel flow and EPR decreased to

prevent N1 over speed with reduced engine thrust (14,600 lbf). As the fan bleed closed at time=20 seconds, the

engine thrust increased to 36,600 lbf but with shorter time compared to the previous case (1.5 seconds compared to

6.5 seconds obtained for the case 4a). The LPC and HPC surge margins did not changed significantly between

simulation cases 4a and 4b.

Case 5a is the same as case 1b.

Case 5b shows the improved thrust response relative to the case 5a, with a net gain of almost 1 second. This faster

response was due to operating the engine with higher N1 and N2 speeds at idle by opening the LPC bleed valve.

This test case represents still another variation on the BOM Mode #8.

Case 6a is the same as case 5b.

Finally, case 6b shows the results of combined maneuvers used in cases 1b, 3b, and 5b. This test case is also another

variation on the BOM Mode #8. In this case, during the acceleration to full throttle from idle, the total thrust

response time is a little more than 5 seconds.

Comparison Transients

The first example in this section, Figure 66, shows a comparison of the LPC bleed for Cases 1a, 6a, and 6b. Note

that Case 1a did not use LPC bleed and that there is no difference between Cases 6a and 6b. This is the amount of

bleed used to help achieve the rapid thrust performance of 6a and 6b.


14

LPC Bleed Flow

0

10

20

30

40

50

60

70

0 5 10 15 20 25 30

Time (seconds)

LP

C B

lee

d F

low

(lb

/s)

LPC Bleed (1a)

LPC Bleed (6a)

LPC Bleed (6b)

Figure 6 LPC Bleed comparison - Cases 1a, 6a, and 6b.

The next set of figures compares responses for Cases 4a and 4b. Thus, they compare engine response for the

baseline control and a rapid thrust control (fan bleed) capability. In Figure 7 the thrust for Cases 4a and 4b are

shown. Here, both thrust acceleration and deceleration are considerably faster than the baseline, about 4 seconds

faster for deceleration and about 5 seconds for acceleration.

The level of fan bleed used to achieve the thrust response of Cases 4a and 4b is shown in Figure 8. Note that this

level of fan bleed is substantial as would be expected to effect a thrust change of over 20,000 lbs. Initially, it was

felt that large amounts of fan bleed may present a gas re-ingestion problem. However this is probably not the case.

Generally, gas re-ingestion can happen only below about 100 knots. We should be going fast enough in the

scenarios considered that reverser flow cannot reach the engine inlet.

In Figure 9 the LPC surge margin is shown. Here a substantial reduction in surge margin is experienced during the

deceleration transient and this level is maintained due to the “off nominal” engine operation cause by the large

amount of fan bleed. An assessment needs to be made of the risk associated with moving to within 2% of the surge

margin boundary. Note that accelerations are inherently more stable than the decelerations in the low pressure

compressor.


15

Thrust Modulation

0

5000

10000

15000

20000

25000

30000

35000

40000

45000

0 5 10 15 20 25 30

Time (seconds)

Th

rus

t (l

bf)

Thrust (4a)

Thrust (4b)

Figure 7 Thrust response for Cases 4a and 4b.

The results of Figure 10 show the surge margin performance of the HPC for the baseline and fan bleed transient

examples. Note that the only time surge margin performance for the fan bleed case is lower than the baseline is

during engine deceleration. Note that this value is not any lower than the value for the baseline during acceleration.

Thus, the fan bleed control mode, in this example, does not generate any more risk to the HPC than does the

baseline control.


16

Fan Bleed

0

100

200

300

400

500

600

700

800

900

1000

0 5 10 15 20 25 30

Time (seconds)

Fan

Ble

ed

(lb

/s)

Fan Bleed (4a)

Fan Bleed (4b)

Figure 8 Fan bleed for Cases 4a and 4b.


17

%LPC Surge Margin

0

5

10

15

20

25

30

0 5 10 15 20 25 30

Time (seconds)

LP

C S

urg

e M

arg

in (

%)

%LPC Surge Margin (4a)

%LPC Surge Margin (4b)

Figure 9 LPC surge margin for Cases 4a and 4b.

%HPC Surge Margin

0

5

10

15

20

25

30

35

40

0 5 10 15 20 25 30

Time (seconds)

HP

C S

urg

e M

arg

in (

%)

%HPC Surge Margin (4a)

%HPC Surge Margin (4b)

Figure 10 HPC surge margin for Cases 4a and 4b.

Thrust Response Comparison


18

Table 7 below shows the comparison of the thrust response for several of the test cases. As indicated in Table 7, the

thrust modulation methods used for cases 6b and 4b are the most attractive with the highest thrust slew rate. The

risk of surge is also highest, as the open VSV at lower N2 and higher Wf/P3 reduce HPC surge margin. Note that

case 5b is only slightly less responsive than case 6b (approximately 10% slower). However, this was achieved with

substantially less risk of a compressor stall/surge. An analysis of thrust response required to achieve controlled

flight during emergency scenarios will provide additional insight into which control of the tested control approaches

will best meet the project objectives.

Table 7 Thrust Response Comparison

Case Special Actions PC(deg) Time to full thrust

(secs)

Mean Thrust Slew

Rate (lbf/sec)

1a Baseline 80.5-40-80.5 13.5 2600

1b Increase acceleration and decrease

deceleration limits

80.5-40-80.5 7.5 4666


deceleration limits and Slam open

VSV during acceleration (PC=40 to

80.5)

80.5-40-80.5 6.2 5645


deceleration limits and Open LPC

bleed valve during idle & fully close

during acceleration to full throttle

80.5-40-80.5 6.2 5645


deceleration limits, Slam open VSV

during acceleration (PC=40 to 80.5)

and Open LPC bleed valve during

idle and fully close during

acceleration to full throttle

80.5-40-80.5 5.5 6365

4b Modulate fan bleed to produce the

same steady state thrust as in case

4a

Maintain

80.5

1.5 14666

VIII. Conclusion

The paper has presented some initial results that demonstrate that, for situations where the aircraft faces adverse

conditions (control surface failures, runway incursions, etc.), the engine offers the potential to augment aircraft

controllability. This augmentation is achieved by providing differential thrust for steering or additional thrust to

increase climb rate thereby improving flight safety and survivability of an aircraft under abnormal or emergency

conditions. In these conditions, the fast-response controller needs to boost (or recover) the engine capability by

letting it go beyond normal physical and operational limits for a relatively short period of time until the aircraft lands

safely. The enhanced engine capability is primarily higher and faster thrust; however, it must be produced by

balancing the operability and remaining life of critical engine components. Hence, the goal is a controller that

enables higher and faster thrust with acceptable operability and remaining life margins.

In this paper we identified the ability of various existing and new actuators to affect control beyond current limits.

One important finding is that almost all of the engine response requirements can be achieved using existing

actuation. In addition an engine control architecture that includes engine condition and risk assessment during

emergency conditions is described. Several candidate emergency control modes are described and some preliminary

simulation results of these modes are presented.


19

Specific results of the simulation study show that the thrust slew rate from idle to max thrust may be increased from

2600 lbf/sec to 14666 lbf/sec using specific actuators (including fan bleed, which is a new actuator) and control

logic. Risks have yet to be evaluated, but HPC surge risk is highly dependent upon VSV and thus may be mitigated

with VSV control changes. Increased maximum thrust was achieved with higher turbine temperatures. The impact

of these higher temperatures and the durations for which these higher temperatures exist must be studied to quantify

their effects on turbine life and overall operational risk.

In many engines the capability for implementing fan bleed is already in place. In some instances these are cascades

that are also used for the thrust reverser and they are sized to pass up to 100% of the fan flow. The actuators are also

in place, though they may need to be modified in order to enable the flow to act like a bleed instead of the normal

thrust reversal. For example, the thrust reverser has an inside portion that shuts off flow to the nozzle. We may

want this to leak some of the flow to the nozzle and divert the rest through the bleed slots.

In addition to adding extra sophistication to the simulated results to explore these operational considerations, future

work will focus on the development of design techniques that can successfully perform the trade off between control

effectiveness and operational risks during various emergency scenarios.

Acknowledgments

Scientific Monitoring, Inc. (SMI) received support for this research under NASA Contract NNC08CA54C which

was awarded from the NRA NNH07ZEA001N-IRAC1, to develop a fast response engine controller design. This

effort is being performed under the Integrated Propulsion Control and Dynamics element of the IRAC project as part

of the NASA ARMD Aviation Safety Program. The authors would like to acknowledge the willing and useful

collaboration of the project team partners from Pratt and Whitney, who are led by Mr. Bruce Wood and Boeing ,

who are led by Mr. James Urness. Finally, the authors would like to thank Dr. Ten-Huei Guo and Mr. Jonathan Litt

for their direction and management of the project and Mr. Litt for his editorial review of the paper.

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