Jeffrey T. Csank

Glenn Research Center, Cleveland, Ohio

Ryan D. May

ASRC Aerospace, Inc., Cleveland, Ohio

Jonathan S. Litt and Ten-Huei Guo

Glenn Research Center, Cleveland, Ohio

A Sensitivity Study of Commercial Aircraft

Engine Response for Emergency Situations

NASA/TM—2011-217004

April 2011
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Jeffrey T. Csank

Glenn Research Center, Cleveland, Ohio

Ryan D. May

ASRC Aerospace, Inc., Cleveland, Ohio

Jonathan S. Litt and Ten-Huei Guo

Glenn Research Center, Cleveland, Ohio

A Sensitivity Study of Commercial Aircraft

Engine Response for Emergency Situations

NASA/TM—2011-217004

April 2011

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-217004 1 

A Sensitivity Study of Commercial Aircraft Engine Response for 

Emergency Situations 

Jeffrey T. Csank 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Ryan D. May 

ASRC Aerospace, Inc. 

Cleveland, Ohio 44135 

Jonathan S. Litt and Ten-Huei Guo 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This paper contains the details of a sensitivity study in which the variation in a commercial aircraft 

engine’s outputs is observed for perturbations in its operating condition inputs or control parameters. This 

study seeks to determine the extent to which various controller limits can be modified to improve engine 

performance, while capturing the increased risk that results from the changes. In an emergency, the 

engine may be required to produce additional thrust, respond faster, or both, to improve the survivability 

of the aircraft. The objective of this paper is to propose changes to the engine controller and determine the 

costs and benefits of the additional capabilities produced by the engine. This study indicates that the 

aircraft engine is capable of producing additional thrust, but at the cost of an increased risk of an engine 

failure due to higher turbine temperatures and rotor speeds. The engine can also respond more quickly to 

transient commands, but this action reduces the remaining stall margin to possibly dangerous levels. To 

improve transient response in landing scenarios, a control mode known as High Speed Idle is proposed 

that increases the responsiveness of the engine and conserves stall margin. 

Nomenclature 

Accel Acceleration Schedule 

alt Altitude (ft) 

BL Baseline Limits 

BW Bandwidth 

C-MAPSS40k Commercial Modular Aero-Propulsion System Simulation 40k 

dTamb Degrees Rankine from standard day temperature 

EGT Exhaust gas temperature (degrees Rankine) 

EP Extended Power 

EPR Engine Pressure Ratio 

FAA Federal Aviation Administration 

FastER Fast Engine Response 

Fnt Net Thrust (lbf) 

GM Gain Margin 
NASA/TM—2011-217004 2 

HPC High Pressure Compressor 

HSI High Speed Idle control mode 

Ki Integral Gain 

Kp Proportional Gain 

LPC Low Pressure Compressor 

MAS Modified Acceleration Schedule 

MAX Maximum function 

MIN Minimum function 

MN Mach number 

MP Maximum Power 

Nc Core Speed (revolutions per minute) 

Ncdot Core Acceleration (revolutions per minute per second) 

Nf Fan Speed (revolutions per minute) 

NL No Limits 

PI Proportional plus Integral 

PM Phase Margin 

Ps3 Compressor Discharge Static Pressure (psia) 

Pxx Total pressure at station XX (psia) 

RU Ratio Unit (Wf/Ps3) 

rpm Revolutions per minute 

SLS Sea Level Static: environmental condition defined as an altitude of 0 ft and 

Mach number 0.0 

Tc Time constant (s) 

Td Dead Time (s) 

Tde Time Delay (s) 

Tr Rise Time (s) 

Ts Settling Time (s) 

Txx Temperature at station XX (degrees Rankine) 

VSV Variable Stator Vane 

Wbleed Mass flow rate of air through the Variable Bleed Valve (lbm/s) 

Wcust Mass flow rate of air through the customer bleed valve (lbm/s) 

Wf Fuel flow rate (lbm/s) 

Wxx Mass flow rate through the engine at station XX (lbm/s) 

I. Introduction 

Research is on-going to investigate the use of an aircraft’s engines to stabilize and control a distressed 

aircraft. For the engine to be used as a flight control effector, the engine may be required to perform 

beyond its current limitations: producing thrust greater than maximum rated power or responding faster to 

the pilot’s throttle command. A sensitivity study is conducted using the Commercial Modular AeroPropulsion System Simulation 40k (C-MAPSS40k) (Ref. 1) to determine the effect of extending or 

removing the engine controller’s safety limits to enable the engine to operate beyond its nominal design 

range. Note that this study avoids major changes to the engine controller; all modifications will preserve 

the structure and form of the existing baseline controller (Ref. 2) and work within the constraints of the 

existing engine actuation. 
NASA/TM—2011-217004 3 

It has been shown that in some emergencies the engines can serve as flight control actuators to 

improve the capabilities of the aircraft (Refs. 3 and 4). However, conservative engine controller design 

severely restricts maximum engine performance capability. The controller design is necessarily 

conservative because it is designed to provide safe operation throughout the flight envelope, regardless of 

the age of the engine. However, in emergency situations, this need to preserve safety margins can be 

traded for increased performance to improve the immediate survivability of the aircraft. This sensitivity 

study seeks to quantify how to make these trade-offs for the aircraft engine. 

Previously, a sensitivity study was performed using the Commercial Modular Aero-Propulsion 

System Simulation (C-MAPSS), that considered both increasing the responsiveness of the engine and 

overspeed operation (Ref. 5). The engine modeled in C-MAPSS (Ref. 6) is a 90,000 lbf thrust class 

engine that is similar in design and architecture to the 40,000 lbf thrust class engine of C-MAPSS40k. 

The engine controller architecture is similar in that each imposes limits on various engine parameters such 

as shaft speeds and combustor pressure. The one difference of note is that the C-MAPSS controller 

actively limits the turbine temperature (T48) while C-MAPSS40k does not contain such a limit. In 

commercial aviation, the management of engine temperature is typically left to the pilots; however 

reporting is required when temperatures exceed a maintenance threshold. 

According to Reference 5, during overspeed operation, also referred to as overthrust, the C-MAPSS 

engine outputs, such as the rotor speeds, engine pressure ratio (EPR), turbine temperature, static 

combustor pressure, and thrust, varied linearly with the throttle input. It was also noted that the engine 

limits, mainly rotor speeds, might need to be increased in order for C-MAPSS to reach the desired thrust 

level, but that increasing the rotor speed limits will have a negative impact on engine life. 

While the overthrust case was briefly explored, the majority of Reference 5 focused on increasing the 

responsiveness of the engine to both small and large throttle transients. For small throttle transients of 5, 

it was shown that simply increasing the controller bandwidth did not increase the performance of the 

engine due to the engine controller limits. However, doubling the controller bandwidth and disabling all 

of the controller limits allowed a significant reduction in the settling time. When reviewing the resulting 

data it was found that although all the limits were disabled, the nominal T48 threshold was violated only 

for a brief period. These brief excursions in turbine temperature generally do not cause catastrophic 

failure; rather they deteriorate the engine at an accelerated rate. Thus, momentarily exceeding the nominal 

T48 limit may be acceptable in an emergency scenario if the amount of risk to the engine is acceptable. 

The large throttle transients required the same type of changes made for the small throttle transients with 

an additional modification to the acceleration schedule. 

This study was conducted using C-MAPSS40k to further the results obtained previously using 

C-MAPSS. C-MAPSS40k is a 40,000 lbf thrust class, two spool, physics-based, component level, high 

bypass turbofan engine simulation and closed loop controller modeled in the MATLAB/Simulink 

(The MathWorks, Inc.) environment (Ref. 1). The C-MAPSS40k open loop engine schematic is shown in 

Figure 1. Figure 1 shows the rotor speeds (Nf and Nc), the Variable Stator Vanes (VSV), the pressures, 

temperatures, and flows at each of the engine’s stations, as well as the flow through the Variable Bleed 

Valve (Wbleed), turbine cooling bleeds (W28, W29, W31, W32), and customer bleed flow (Wcust). 

C-MAPSS40k has the ability to control the engine thrust using either fan speed (Nf) or Engine Pressure 

Ratio (EPR), which is the exit pressure of the low pressure turbine (P50) divided by the pressure at the 

inlet (P2). 

The C-MAPSS40k simulation package includes a representative generic commercial jet engine 

controller (Ref. 2). The aircraft engine controller can be separated into two functions: a power 

management controller, which is responsible for providing thrust and good transient performance, and a 

protection logic controller responsible for protecting the engine from exceeding any of its physical limits. 

Figure 2 shows a diagram of the C-MAPSS40k control system, and highlights the power management and 

protection logic controllers. As shown in Figure 2, the power management controller converts the throttle 

input to a setpoint (EPR or Nf), calculates the error (difference between setpoint and feedback), and 

determines the fuel flow rate required to drive the engine to its setpoint using a PI controller with 

integrator wind-up protection. The protection logic controller protects the engine from physical damage 
NASA/TM—2011-217004 4 

Figure 1.—C-MAPSS40k engine schematic. 

Figure 2.—C-MAPSS40k control system. 
NASA/TM—2011-217004 5 

and maintains operation within acceptable bounds (e.g., it avoids compressor surge and stall and prevents 

combustor flameout) through the use of individual limit controllers, also referred to as limiters. The 

aircraft engine controller contains the following limiters: maximum fan speed (Nf max), maximum core 

speed (Nc max), maximum combustor static pressure (Ps3 max), maximum core shaft acceleration based 

on an acceleration schedule (Accel), minimum combustor static pressure (Ps3 min), and minimum ratio 

unit (RU min) which is the fuel flow divided by the static combustor pressure. A MIN-MAX scheme is 

employed to select the fuel flow rate command from the limit controller closest to its limit, or the power 

management command in cases where no limits are in danger of being violated. This selection scheme 

also serves to provide a smooth transition between the protection logic and power management 

controllers. 

This paper reports the results of the sensitivity study for C-MAPSS40k designed to determine the 

effect of operating the engine at conditions beyond its current limitations. This paper extends the type of 

overthrust analysis performed in Reference 5 to include the risk of engine failure due to producing 

additional thrust. In addition, this paper proposes a new control algorithm to increase the responsiveness 

of the aircraft engine when operating at low power. Section II contains a description of the scenarios 

considered in the sensitivity study. The details on the overthrust control mode, including the engine limits 

of interest and how much additional thrust can be obtained, are found in Section III. Section IV has details 

on increasing the responsiveness of the engine, known as fast engine response. The work is summarized 

in Section V. 

II. Sensitivity Study Scenarios 

The purpose of this sensitivity study is to help determine how far the engine control limits can be 

extended during an emergency that imposes high risk to the aircraft. In such situations, it is posited that 

the increased risk taken by the engine actually decreases the aircraft risk, resulting in the minimal system 

risk, however such system analysis is beyond the scope of this paper. Specifically, there are a few 

scenarios in which engine control modification is being considered to provide additional performance. 

The two scenarios of interest for this study are: 1) a take-off incursion, where the distance for lift-off is 

suddenly decreased, and 2) a rudder/tail failure, where the effectiveness of the rudder control surface or 

vertical stabilizer is reduced. These two scenarios are depicted in Figure 3 taken from Reference 5. 

In the take-off incursion scenario, the remaining runway distance an aircraft has to reach the take-off 

speed is suddenly decreased due to an object entering the runway; alternately, the pilot notices that the 

aircraft is attempting to take off on a runway that is too short. To correct this situation, it has been 

proposed to increase the thrust beyond the engine’s maximum nominal thrust level (Ref. 5) which is 

depicted on the left side of Figure 3. The additional thrust the engine produces would accelerate the 

aircraft at a higher rate, enabling the aircraft to reach its takeoff speed in a shorter distance. The concern 

with producing the additional thrust is that it might require an increase in the turbine temperatures and 

rotor speeds beyond their current limits. 

In the rudder/tail failure scenario, the aircraft has experienced either airframe or internal damage 

(loss of hydraulics), which reduces the effectiveness of the flight control surfaces. This is also illustrated 

in Figure 3. To help control the yaw angle of the aircraft in the case of rudder/tail failure, it has been 

proposed to increase the responsiveness of the engine (Ref. 5). Increasing the responsiveness of the 

engine corresponds to higher accelerations which may lead to a decrease in high pressure compressor 

(HPC) surge margin. 

The engine response for each of the emergency scenarios, overthrust and fast engine response, will be 

investigated separately. Each of these scenarios will be simulated at different environmental conditions. 

Three different altitudes have been chosen at which to perform the study: one near sea level, John F. 

Kennedy International Airport (JFK) at an altitude of 13 ft; the international airport with the highest 

takeoff altitude in the continental United States, Denver International Airport (DEN), at 5,431 ft; and one 

in between, McCarran International Airport (LAS), at an altitude of 2,181 ft. Each test will be conducted 

for both a standard day and a hot day, as well as for a 50 hr engine and an end of life engine. 
NASA/TM—2011-217004 6 

Figure 3.—The two emergency scenarios and corresponding engine responses. Figure from 

Reference 5. 

III. Overthrust (Runway Incursion) 

For the overthrust situation, the engine is commanded to produce additional thrust, which correlates 

to a requirement for the engine to produce a higher EPR or fan speed. Therefore, the engine setpoints are 

extended to demand the additional power, but this also implies that the engine will be operating at speeds, 

temperatures, and pressures greater than the maximum ratings designed to ensure long engine life. 

As discussed earlier, the C-MAPSS40k simulation has several protection logic controllers to ensure 

safe operation of the aircraft engine, shown in Figure 2. For the overthrust operation, the limits of interest 

are the maximum fan speed, maximum core speed, and maximum combustor pressure. The maximum fan 

and core speed limiters protect against disk burst or blade failure, while the maximum combustor pressure 

Requiring Additional 

Thrust

Requiring Faster 

Response

 Takeoff Incursion  Rudder/Tail Failure

Nominal

Thrust

Required

Thrust

Time (sec)

Thrust

Event

Nominal

Response

Required

Response

Time (sec)

Thrust

Event

Air Flow

Compressor

Pressure 

Ratio

Stall / Surge Margins

Limit

Time (sec)

Max Temp/Speed Limits Turbine TemperatureRotor Speed

Requirements via Scenario Analysis

Using Aircraft Flight Simulator

Requiring Additional 

Thrust

Requiring Faster 

Response

 Takeoff Incursion  Rudder/Tail Failure

Nominal

Thrust

Required

Thrust

Time (sec)

Thrust

Event

Nominal

Response

Required

Response

Time (sec)

Thrust

Event

Air Flow

Compressor

Pressure 

Ratio

Stall / Surge Margins

Limit

Time (sec)

Max Temp/Speed Limits Turbine TemperatureRotor Speed

Requirements via Scenario Analysis

Using Aircraft Flight Simulator
NASA/TM—2011-217004 7 

protects from exceeding the combustor casing pressure limit (Ref. 2). For the sensitivity study, the 

maximum combustor pressure will not be extended as even momentary excursions outside of the 

combustor pressure range can cause sudden catastrophic damage. Another limit to monitor is the turbine 

temperature limit; when exceeded this leads to turbine blade erosion. In C-MAPSS40k there is no turbine 

temperature limit, rather an exhaust gas temperature (T50) maintenance guideline exists, but there is no 

automatic limit protection; violating the guideline is at the pilot’s discretion. The measurable exhaust gas 

temperature is typically used to approximate the temperature inside the turbines. The temperature 

maintenance guideline is usually violated in cases when the engine is over-fueled, (Ref. 7) which is 

required to produce additional thrust. In this study, the guideline will not be relaxed but the T50

temperature will be monitored to determine the effect the various control limit modifications have on the 

exhaust gas temperature. The relevant C-MAPSS40k engine’s limits and maintenance guidelines for the 

overthrust scenario are listed in Table 1. 

TABLE 1.—C-MAPSS40k ENGINE’S NOMINAL 

LIMITS/MAINTENANCE GUIDELINES 

Fan speed (Nf) 4,200 rpm 

Core speed (Nc) 12,200 rpm 

Combustor static pressure (Ps3) 433 psi 

Exhaust gas temperature (T50) 1,500 R 

The first step in allowing the engine to produce additional thrust (by design), is to extend the control 

system’s setpoints. The thrust profile indicates the amount of thrust desired as a function of the throttle 

position at a particular flight condition. From the thrust profile, the setpoint profile (EPR or fan speed) 

used to control the engine is created. The thrust profile, and corresponding setpoints, will be extended to 

provide additional power above maximum at each flight condition. This study will use C-MAPSS40k 

with EPR as the controlled variable. A 20 percent extension is arbitrarily chosen as a target. The 

C-MAPSS40k throttle range is from 40 to 80, therefore the extended power range will be from a 

maximum throttle position of 80, corresponding to 100 percent power, to 90 corresponding to 

120 percent of maximum nominal power. Note that since the EPR profile is extended, the engine may not 

produce exactly the requested 120 percent maximum thrust. 

There are two steps taken to determine which variables may limit the amount of additional thrust the 

engine can produce and the life cost of producing additional thrust. The first step is to maintain the 

nominal engine limits and command the additional 20 percent thrust. Ideally, this scenario would indicate 

the power the engine can produce without an increase in risk of failure. The second step is to remove the 

limits, except for the combustor static pressure limit, and command the additional 20 percent thrust. This 

will indicate by how much, if any, each limit must be increased to reach the desired thrust, and the 

amount of risk associated with this new operating point. 

First, consider the engine response for a 50 hr engine at McCarran airport on a standard day (2,181 ft, 

0.17 Mach, 0.0 R from standard day temperature). The responses for maximum power (MP), extended 

power with the baseline limits (EP BL), and extended power with no limits (EP NL) are shown in 

Figure 4. At this flight condition, the 120 percent of maximum thrust can be reached without extending 

the current engine limits. In Figure 4, there is a slight difference between the EP BL and EP NL response 

from 100 percent max power to 120 percent max power. For the EP BL, since the limits remain 

unchanged, the limiters are free to become active and affect the thrust response, whereas for the EP NL, 

only the EPR setpoint controller can affect the thrust response; therefore there may be small differences in 

the transient to the new thrust level. 
NASA/TM—2011-217004 8 

Figure 4.—Overthrust scenario comparing engine outputs at McCarran (2,181 ft, 0.17 

Mach, +0 R from standard day temperature, 50 hr engine). 

Now consider the response for a 50 hr engine at Denver on a hot day (5,431 ft, 0.17 Mach, +48 °R from 

standard day temperature), which is shown in Figure 5. Figure 5 shows that with the baseline engine limits, 

the core speed limit restricts the amount of additional thrust the engine can produce. Additionally, even with 

the core speed limit in effect, the exhaust gas temperature maintenance guideline (1500 °R) is exceeded. In 

order for the engine to produce the desired 120 percent maximum power thrust, both the fan speed and core 

speed limits would have to be extended and a higher turbine temperature would have to be accepted. 

The thrust, fan speed, core speed, and low pressure turbine temperature outputs for the extended power 

with baseline limits and extended power with the limits removed are shown in Table 2. Note that the two 

flight conditions printed in bold in Table 2 (13ft, 0.17 Mach, +0 °R from standard day temperature for both 

a 50 hr engine and end of life engine) are actually limited by the combustor pressure limit, which is not 

under consideration for modification. In Table 2, any individual variable above its limit for the flight 

condition is highlighted, and any variable that is limited is italicized. To reach 120 percent maximum power 

for all conditions, the fan speed limit would have to increase from 4,200 to 4,365 rpm (103.9 percent Nf 

max), the core speed limit would have to increase from 12,200 to 12,530 rpm (102.7 percent Nc max), and 

the T50 limit from 1,500 to 1,724 R (114.9 percent T50 max). Requesting additional thrust while 

maintaining the baseline engine limits results in a maximum exhaust gas temperature of 1,655 R. Note that 

in the EP BL case the thrust may not be able to reach 120 percent due to the new temperature limiter. It is 

also worth restating the fact that the engine is controlled based on EPR, not thrust; thus, the thrust produced 

may exceed the desired 120 percent maximum thrust for the EP NL case. 

5 10 15 20 25

2.8

3

3.2

3.4

3.6

x 104

Fnt, lbf

5 10 15 20 25

3800

3900

4000

4100

4200

Nf, RPM

5 10 15 20 25

1.14

1.16

1.18

1.2

1.22

x 104

Nc, RPM

time, s

5 10 15 20 25

1350

1400

1450

1500

T50,  R

time, s

MP

EP BL

EP NL

Limit
NASA/TM—2011-217004 9 

Figure 5.—Overthrust scenario showing engine outputs at Denver (5,431 ft, 0.17 Mach, 

+48 R from standard day temperature, 50 hr engine). 

TABLE 2.—CRITICAL ENGINE OUTPUTS FOR OVER THRUST OPERATION 

Engine 

life 

Flight condition, 

alt, MN, dTamb 

Extended power base line limits Extended power no limits 

Thrust, 

percent 

Nf, 

rpm 

Nc, 

rpm 

T50, 

R 

Thrust, 

percent 

Nf, 

rpm 

Nc, 

rpm 

T50, 

R 

50 hr 

13 ft / 0.17 / 0 °R°R 119 4019 11911 1460 119 4019 11911 1460 

2181 ft / 0.17 / 0 °R 120 4124 12004 1490 120 4124 12004 1490 

5431 ft / 0.17 / 0 °R 117 4200 12026 1504 121 4260 12104 1531 

13 ft / 0.17 / 48 °RR 116 4073 12200 1548 121 4128 12298 1571 

2181 ft / 0.17 / 48 °R 112 4090 12200 1544 122 4212 12402 1595 

5431 ft / 0.17 / 48 °R 109 4116 12200 1540 125 4365 12530 1642 

End of life 

13 ft / 0.17 / 0 °°R 119 4014 11823 1540 119 4014 11823 1540 

2181 ft / 0.17 / 0 °R 120 4113 11925 1565 120 4113 11925 1565 

5431 ft / 0.17 / 0 °R 118 4200 11961 1586 121 4246 12018 1607 

13 ft / 0.17 / 48 °°R 120 4114 12200 1655 121 4119 12208 1657 

2181 ft / 0.17 / 48 °R 117 4137 12200 1652 122 4201 12305 1679 

5431 ft / 0.17 / 48 °R 113 4169 12200 1648 125 4351 12444 1724 

 

5 10 15 20 25

2.5

3

3.5

x 104

Fnt, lbf

5 10 15 20 25

4000

4100

4200

4300

4400

Nf, RPM

5 10 15 20 25

1.18

1.2

1.22

1.24

1.26

x 104

Nc, RPM

time, s

5 10 15 20 25

1400

1500

1600

1700

T50,  R

time, s

MP

EP BL

EP NL

Limit
NASA/TM—2011-217004 10 

To determine if these new operating conditions are acceptable, an example risk function is 

implemented similar to the one proposed in Reference 8. This risk function, which is a probabilistic 

model of engine life based on rotor speeds and turbine temperatures, will determine the probability of 

engine failure for a specified operating time at the given condition. Although a runway incursion may 

only require the additional thrust for a few minutes, the overthrust function may be utilized for long 

periods during flight in order to overcome drag due to aircraft damage. For this study, the risk function 

assumes that the engine will have to operate in the overthrust condition for 90 min. The 90 min time is a 

worst case estimate for a distressed plane flying over the continental United States to find an appropriate 

airport at which to land. Unfortunately the data necessary to determine true engine risk values are not 

readily available, thus the risk function implemented is approximate and will only allow trends to be 

discovered. Actual values for specific limits would necessitate detailed engine study and testing prior to 

implementation and control law development. The reader is referred to Reference 8 for more information 

regarding the risk function implementation and its limitations. For the purposes of this sensitivity study, 

we will proceed as if the risk model is accurate, however this approximation must be noted. 

The thrust and risk values for maximum power, extended power with baseline limits, and extended 

power with no limits are shown in Table 3. In Table 3 any condition that produced a risk greater than 

10 percent is highlighted. From Table 3, the extended power with baseline engine limits risk reaches up to 

18.83 percent risk of failure for an end of life engine on a hot day, even though none of the active limits 

are exceeded. To prevent this excessive risk we can observe, from Table 2, that in this case T50 is beyond 

its maintenance guideline by more than 150 °R. In order to allow us to take advantage of the additional 

thrust capability and ensure that the risk of operation does not exceed the risk level deemed acceptable 

(Ref. 9) a T50 limiter can be added to the controller. The addition of the T50 limiter will prevent the 

engine from reaching 120 percent full power in some cases, but does increase the safety by limiting the 

risk of engine failure to a level deemed acceptable by the pilot/flight controller. 

Table 4 compares the thrust, T50, and risk for the extended power using the baseline limits but with 

the addition of a T50 limiter. By setting the T50 limit to 1,600 R, the risk can be reduced from 

approximately 15 percent to less than 5 percent, as highlighted in Table 4. However, the reduction in risk 

also corresponds to a reduction in the thrust produced by approximately 2,000 to 3,000 lbf. 

TABLE 3.—THRUST AND RISK FOR MAX POWER AND OVERTHRUST OPERATION 

Engine 

life 

Flight condition, 

alt, MN, dTamb 

Max power Extended power 

current limits 

Extended power 

no limits 

Thrust, 

percent 

Risk, 

percent 

Thrust, 

percent 

Risk, 

percent 

Thrust, 

percent 

Risk, 

percent 

50 hr 

13 ft / 0.17 / 0 °R 100 0.001 119 0.568 119 0.568 

2181 ft / 0.17 / 0 °R 100 0.001 120 2.256 120 2.256 

5431 ft / 0.17 / 0 °R 100 0.002 117 3.862 121 8.061 

13 ft / 0.17 / 48 °R 100 0.017 116 3.993 121 9.087 

2181 ft / 0.17 / 48 °R 100 0.160 112 4.177 122 28.831 

5431 ft / 0.17 / 48 °R 100 1.348 109 4.503 125 99.956 

End of life 

13 ft / 0.17 / 0 °R 100 0.002 119 1.528 119 1.528 

2181 ft / 0.17 / 0 °R 100 0.003 120 3.913 120 3.913 

5431 ft / 0.17 / 0 °R 100 0.014 118 9.096 121 16.714 

13 ft / 0.17 / 48 °R 100 0.804 120 14.827 121 15.953 

2181 ft / 0.17 / 48 °R 100 1.603 117 16.262 122 45.324 

5431 ft / 0.17 / 48 °R 100 2.225 113 18.830 125 100.000 

 
NASA/TM—2011-217004 11 

TABLE 4.—THRUST, EGT, AND RISK FOR OVERTHRUST OPERATION 

Engine 

life 

Flight condition, 

alt, MN, dTamb 

Extended power 

current limit 

Extended power current limit and 

T50 limited 

Thrust, 

percent 

T50 

R 

Risk 

percent 

Thrust 

percent 

T50 

R 

Risk 

percent 

50 hr 

13 ft / 0.17 / 0 °R 119 1460 0.568 119 1460 0.568 

2181 ft / 0.17 / 0 °R 120 1490 2.256 120 1490 2.256 

5431 ft / 0.17 / 0 °R 117 1504 3.862 117 1504 3.862 

13 ft / 0.17 / 48 °R 116 1548 3.993 116 1548 3.993 

2181 ft / 0.17 / 48 °R 112 1544 4.177 112 1544 4.177 

5431 ft / 0.17 / 48 °R 109 1540 4.503 109 1540 4.503 

End of life 

13 ft / 0.17 / 0 °R 119 1540 1.528 119 1540 1.528 

2181 ft / 0.17 / 0 °R 120 1565 3.913 120 1565 3.913 

5431 ft / 0.17 / 0 °R 118 1586 9.096 118 1586 9.096 

13 ft / 0.17 / 48 °R 120 1655 14.827 108 1600 2.654 

2181 ft / 0.17 / 48 °R 117 1652 16.262 106 1600 3.105 

5431 ft / 0.17 / 48 °R 113 1648 18.830 105 1600 4.086 

TABLE 5.—RISK AS A FUNCTION OF DESIRED OPERATING TIME 

Engine 

life 

Flight condition, 

alt/MN/dTamb 

Extended power no limits 

15 min 30 min 45 min 60 min 90 min 

50 Hr 

13 ft / 0.17 / 0 °R 0.001 0.003 0.016 0.069 0.568 

2181 ft / 0.17 / 0 °R 0.001 0.008 0.062 0.274 2.256 

5431 ft / 0.17 / 0 °R 0.002 0.028 0.224 1.002 8.061 

13 ft / 0.17 / 48 °R 0.002 0.031 0.254 1.135 9.087 

2181 ft / 0.17 / 48 °R 0.004 0.109 0.900 3.991 28.831 

5431 ft / 0.17 / 48 °R 0.066 2.426 18.542 60.331 99.956 

End of Life 

13 ft / 0.17 / 0 °R 0.001 0.006 0.042 0.185 1.528 

2181 ft / 0.17 / 0 °R 0.001 0.014 0.107 0.478 3.913 

5431 ft / 0.17 / 0 °R 0.003 0.059 0.485 2.167 16.714 

13 ft / 0.17 / 48 °R 0.002 0.056 0.461 2.060 15.953 

2181 ft / 0.17 / 48 °R 0.006 0.193 1.591 6.974 45.324 

5431 ft / 0.17 / 48 °R 0.198 7.163 46.246 93.912 100.000 

The desired operation time is an additional factor for the overthrust scenario. Table 5 contains the risk 

for the extended setpoint and no limit case for different desired operating times, to show the effect the 

operating time has on risk. As the desired operating time increases, the risk of failure also increases. For 

instance, at 5,431 ft, 0.17 Mach, on a hot day with a 50 hr engine, if 120 percent of maximum power is 

requested for 15 min, there is a 0.066 percent chance of a failure, however if the time is increased to 

90 min there is a 99.956 percent chance of a failure using the risk function described earlier (Ref. 8). In 

addition, the ambient temperature also plays a big role; notice that at the same altitude and Mach number, 

the risk increases for the hotter day. The reason for this is that the engine has to work harder to produce 

the required thrust on hotter days, especially at higher altitudes. 

In order for the engine to generate additional thrust at a given condition, the risk of engine failure will 

rise, mainly due to increased speeds and temperatures. The addition of the T50 limiter and an estimated 

operation time can ensure the engine does not exceed the risk level determined by the flight controller, 

minimizing the overall aircraft risk while maximizing the engine performance. 
NASA/TM—2011-217004 12 

IV. Fast Engine Response (FastER) 

Fast Engine Response (FastER) is concerned with how quickly the engine can follow the throttle 

input. There are two different throttle transients of interest, small and large. In the rudder/tail failure case 

discussed earlier, the throttles are used collectively for longitudinal control and differentially for 

lateral/directional control. In an emergency situation, small throttle transients (generally less than 5° of 

throttle movement) have been shown to dampen out the phugoid mode, (Ref. 10) and stabilize the aircraft. 

The small throttle transient response is dominated by the tuning of the PI setpoint controller. To increase 

the performance of the closed loop system, the controller bandwidth is increased by increasing the 

integral term of the PI controller. The concern with increasing the bandwidth of the controller is the effect 

on the stability of the system, measured in terms of gain margin and phase margin. Increasing the 

bandwidth leads to a reduction in the gain margin and phase margin and could even degrade the 

performance if the margins are low enough. This is usually seen as an excessive amount of overshoot, 

which increases the settling time, and possibly results in oscillations about the setpoint. 

The second type of throttle command of interest is a large change, from flight idle to full power. This 

type of throttle command would most likely be seen during an aborted approach or landing. The 

acceleration schedule is designed to protect the high pressure compressor from stalling during these large 

throttle transients. To increase the responsiveness of the engine for a large throttle transient, the 

acceleration schedule can be modified at the cost of stall margin reduction. 

For either range of throttle transient, the performance is measured in terms of the following metrics 

and is shown in Figure 6: 

 Time Constant (Tc): Time in seconds for the engine to reach 63.2 percent of the difference 

between the final thrust value and the initial thrust value, measured from when the engine 

response to a throttle step change begins. 

 Delay Time (Tde): Time in seconds from when the throttle transient is initiated to 50 percent of 

the difference between the final thrust and initial thrust values. 

 Dead Time (Td): Time in seconds from when the throttle transient is initiated to when the engine 

initially responds to the command. 

 Settling Time (Ts): Time in seconds from when the throttle transient is initiated to the time when 

the output is within 2 percent of the difference between the final thrust and initial thrust values. 

 Rise Time (Tr): Time in seconds it takes to transition from 10 to 90 percent of the difference 

between the final thrust and initial thrust values. 

The following three sections, respectively will discuss the results of the sensitivity study for small 

transients, large transients, and finally for a new control mode known as High Speed Idle (HSI). 

A. Linear Controller (Small Throttle Commands) 

The bandwidth of the linear setpoint controller is modified to increase the responsiveness of the 

engine for small throttle transients. An increase in the linear controller’s bandwidth should lead to an 

increase in the closed loop performance (decrease settling time, rise time, etc.) but at the cost of gain 

margin and phase margin. 

The performance and stability margins of the closed loop system for changes in the controller 

bandwidth are shown in Table 6. Table 6 compares the rise time, settling time, gain margin (GM) in 

decibels, and phase margin (PM) in degrees, for a 50 hr and an end of life engine at three approach 

conditions for the nominal PI controller and various controllers where the bandwidth is increased by 

modifying the integral term. While there is minimal change in the gain margin, modifications that cause 

the phase margin to decrease below the 45 design point are highlighted. 

 
NASA/TM—2011-217004 13 

The results shown in Table 6 indicate that the lower the gain margin and phase margin are pushed, the 

better the performance. Increasing the integrator term by a factor of 1.6 results in phase margins below 

40°, but does not have a significant improvement in the performance over increasing the integrator by a 

factor of 1.4, which results in a phase margin above 40°. Figure 7 shows an example of an engine with the 

integrator increased by a factor of 2.4, decreasing the phase margin to 24.82°. The response has a large 

overshoot, resulting in the same settling time as the nominal response; thus there is no advantage to 

increasing the integrator term to this extreme, especially considering the loss in stability margin. 

Figure 6.—Illustration of the performance metrics used for FastER. 

 

TABLE 6.—CLOSED LOOP COMPARISON FOR CHANGES MADE 

TO THE PI CONTROLLER FOR A 5 THROTTLE TRANSIENT 

Flight condition, 

alt/MN/dTamb 

50 hr engine End of life engine 

Tr, 

s 

Ts, 

s 

GM dB PM, 

° 

Tr, 

s 

Ts, 

s 

GM, 

dB 

PM, 

° 

Nominal performance 

2013 ft / 0.26 / 48 °R 1.62 2.52 16.17 47.54 1.70 2.64 16.46 48.63 

4181 ft / 0.26 / 48 °R 1.64 2.56 15.73 46.47 1.74 2.68 16.28 50.00 

7431 ft / 0.26 / 48 °R 1.71 2.59 16.25 49.49 1.77 2.71 16.43 50.49 

Integral term increased by factor of 1.2 

2013 ft / 0.26 / 48 °R 1.43 2.23 15.86 43.81 1.50 2.32 16.15 44.83 

4181 ft / 0.26 / 48 °R 1.46 2.28 15.43 43.01 1.53 2.37 15.97 46.20 

7431 ft / 0.26 / 48 °R 1.52 2.32 15.95 45.80 1.56 2.40 16.15 47.02 

Integral term increased by factor of 1.4 

2013 ft / 0.26 / 48 °R 1.31 2.05 15.52 40.29 1.37 2.13 15.82 41.24 

4181 ft / 0.26 / 48 °R 1.32 2.08 15.11 39.70 1.40 2.17 15.65 42.62 

7431 ft / 0.26 / 48 °R 1.40 2.14 15.63 42.30 1.44 2.20 15.88 43.71 

Integral term increased by factor of 1.6 

2013 ft / 0.26 / 48 °R 1.20 1.92 15.18 36.97 1.26 1.99 15.49 37.88 

4181 ft / 0.26 / 48 °R 1.23 1.96 14.79 36.57 1.28 2.02 15.32 39.25 

7431 ft / 0.26 / 48 °R 1.31 2.02 15.29 38.99 1.32 2.07 15.59 40.58 
NASA/TM—2011-217004 14 

Figure 7.—Plot showing the effect of increasing the bandwidth to a value that decreases 

the phase margin below 40. The integrator gain corresponding to the wider bandwidth 

system was increased by a factor of 2.4 over the baseline, which reduced the phase 

margin from 48.24 in the nominal case to 24.82. 

B. Acceleration Schedule (Large Throttle Commands) 

The acceleration schedule is designed to protect an end of life engine from stalling the high 

pressure compressor during a quick throttle transient from flight idle to full power. The acceleration 

schedule used in C-MAPSS40k was designed based on core acceleration versus core speed (Ncdot

versus Nc). Figure 8 shows the baseline acceleration schedule for C-MAPSS40k along with a modified 

acceleration schedule (MAS). The modified acceleration schedule is the nominal acceleration schedule 

shifted by some value. In Figure 8, an example offset of 300 rpm per second is used, but any suitable 

value can be chosen. By modifying the acceleration schedule, the performance can be increased at the 

cost of reducing the stall margin. 

In this study, the stall margin value is an indirect measurement of risk. One of the uncertainties that 

the acceleration schedule must account for is normal engine-to-engine variation. Since a below average 

engine must be able to accelerate safely, the schedule must be designed to accommodate a 3  variation in 

stall margin to essentially guarantee that no engine will stall. If the standard deviation install margin due 

to engine-to-engine variation is 1 percent, then in the worst case 99.86 percent of all engines will 

accelerate safely for a 3 percent designed minimum stall margin, as shown in Figure 9. Any adjustment of 

the acceleration schedule for faster response means that the below average engines are more likely to 

stall. This increased acceleration can be related to risk through a probability distribution function, where 

the risk of stalling is the portion of the distribution below 0 percent stall margin. The C-MAPSS40k 

simulation stalls when the surge margin reaches 0 percent. The percent of engines that stall increases as 

the designed stall margin value decreases. If there is overwhelming confidence that a specific engine is 

above average, then using a more aggressive acceleration schedule is acceptable. 

15 20 25 30

1

1.2

1.4

1.6

x 104

Fnt

15 20 25 30

1.02

1.04

1.06

1.08

x 104

Nc

15 20 25 30

1150

1200

1250

1300

1350

T50

15 20 25 30

15

20

25

30

HPC SM

Nominal

Increased BW
NASA/TM—2011-217004 15 

Figure 8.—Nominal and an example modified acceleration schedule value of 

300 rpm/s (MAS(300)). 

Figure 9.—Plot of the expected engine-to-engine variation in stall margin 

in an active fleet, indicating the need to retain a required minimum stall 

margin. 

7000 7500 8000 8500 9000 9500 10000 10500 11000 11500 12000

0

200

400

600

800

1000

1200

1400

1600

1800

Core Acceleration, RPM/s

Core Speed, RPM

Nominal

MAS(300)

-2 0 2 4 6 8 10

0

0.005

0.01

0.015

0.02

0.025

0.03

0.035

0.04

Percent of Engines

HPC Stall Margin, %

Data

Stall

Designed SM
NASA/TM—2011-217004 16 

Figure 10.—Current response and response with the modified acceleration schedule 

(7,431 ft, 0.26 Mach, standard day temperature, 50 hr engine). 

The engine response using the proposed modified acceleration schedule compared to the baseline 

response is shown in Figure 10. The quicker thrust response is associated with an increase in T50, Nc, and 

a reduced HPC stall margin. 

Table 7 compares the dead time, rise time, time constant, settling time, and minimum stall margin for 

the current acceleration schedule and the proposed MAS (300). Comparing the settling times and stall 

margins for the nominal and MAS response in Table 7, the performance of the engine increases 

(decreased settling time) with a reduction in the stall margin with the end of life engine, in some cases the 

stall margin is below zero meaning the engine would stall. Further investigation into Table 7 indicates 

that at some flight conditions the performance could still be increased by increasing the offset to the 

nominal schedule, while at other flight conditions the stall margin is less than 1.0 percent and a more 

conservative offset should be utilized. A lookup table for the offset value added to the nominal 

acceleration schedule can be created based on the flight condition, engine degradation level, and the 

acceptable minimum stall margin. This type of function would allow for more performance increase while 

still ensuring safe operation since the table would be designed to provide an average engine with the 

desired stall margin. For the case of an end of life engine where the minimum stall margin with the 

baseline acceleration schedule is less than 1.0 percent, no offset would be added to the engine since 

stalling the engine will likely lead to failure of the aircraft to land successfully. 

One of the shortcomings with the modified acceleration schedule is that the increase in performance 

correlates with a reduction in stall margin. The nominal acceleration schedule shown in Figure 8 is 

designed such that the core acceleration is severely limited at lower core speeds, such as at flight idle 

conditions. If the core speed could be increased at flight idle conditions, then with no modification of the 

acceleration schedule the allowed core acceleration would be increased, allowing an increase in the 

performance of the engine without a reduction in the minimum stall margin. There would be no loss in the 

minimum stall margin since the performance is still bounded by the nominal acceleration schedule; the 

difference is the initial condition of the transient. This idea of increasing the idle power level is known as 

high speed idle, and is the subject of the next section. 

15 20 25 30 35

0

1

2

3

x 104

Fnt, lbf

15 20 25 30 35

0.9

1

1.1

1.2

1.3

x 104

Nc, RPM

15 20 25 30 35

800

1000

1200

1400

1600

T50,  R

time, s

15 20 25 30 35

0

10

20

30

40

HPC SM, %

time, s

Nominal

MAS

Limit
NASA/TM—2011-217004 17 

TABLE 7.—C-MAPSS40K PERFORMANCE WITH MODIFIED ACCELERATION SCHEDULE FOR 

A THROTTLE TRANSIENT FROM FLIGHT IDLE TO FULL POWER 

Flight condition, 

alt, MN, dTamb 

Nominal performance MAS (300) 

Td, 

s 

Tr, 

s 

Tc, 

s 

Ts, 

s 

SM, 

percent 

Td, 

s 

Tr, 

s 

Tc, 

s 

Ts, 

s 

SM, 

percent 

50 hr engine 

2013 ft / 0.26 / 0 °R 0.385 2.010 2.175 4.150 11.926 0.325 1.740 1.960 3.505 7.705 

4181 ft / 0.26 / 0 °R 0.385 2.085 2.280 4.405 10.480 0.325 1.800 2.020 3.700 6.272 

7431 ft / 0.26 / 0 °R 0.370 2.235 2.400 4.690 8.147 0.325 1.950 2.065 3.940 3.851 

2013 ft / 0.26 / 48 °R 0.520 3.105 2.985 6.205 9.117 0.400 2.925 2.635 5.365 3.787 

4181 ft / 0.26 / 48 °R 0.520 3.240 3.000 6.460 7.577 0.385 3.060 2.635 5.605 2.313 

7431 ft / 0.26 / 48 °R 0.505 3.285 3.000 6.655 5.329 0.385 3.060 2.605 5.770 0.300 

End of life engine 

2013 ft / 0.26 / 0 °R 0.385 2.055 2.055 4.180 7.360 0.340 1.785 2.005 3.520 3.237 

4181 ft / 0.26 / 0 °R 0.385 2.130 2.145 4.465 5.955 0.340 1.845 2.050 3.715 1.746 

7431 ft / 0.26 / 0 °R 0.385 2.295 2.205 4.840 3.631 0.325 1.995 2.125 4.015 –0.61 

2013 ft / 0.26 / 48 °R 0.520 2.985 2.850 6.220 4.394 0.400 2.775 2.695 5.290 0.679 

4181 ft / 0.26 / 48 °R 0.595 3.135 3.165 6.805 2.817 0.415 2.895 2.860 5.665 –1.20 

7431 ft / 0.26 / 48 °R 0.550 3.150 3.225 6.955 0.591 0.400 2.910 2.845 5.770 –3.61 

C. High Speed Idle 

As proposed in the previous section, an increase in the idle core speed could lead to faster transients 

for large throttle movements without a reduction in the nominal minimum stall margin, due to the higher 

starting point. The challenge associated with this approach is that increasing the core speed, or flight idle 

condition, results in a higher fan speed that corresponds to more thrust being produced. During approach, 

when the throttle is at flight idle, producing additional thrust will likely increase the speed of the aircraft 

as well as the distance required to land. Note that the aircraft’s control surfaces may be able to spoil some 

additional thrust, but not a significant amount. 

One way to increase the core speed and not produce a large amount of additional thrust is to change 

the variable stator vane (VSV) input. The VSV is at the inlet of the high pressure compressor shown in 

Figure 1 and is nominally scheduled as a function of the corrected core speed. Using off-nominal VSV 

commands results in a change in the high pressure compressor operating point. By moving to a less 

efficient point on the map we can force the power management controller to increase the fuel flow rate, 

and therefore the core speed, in order to maintain the desired EPR or Nf setpoint. Since most of the thrust 

is produced by the fan, this adjustment does not produce a large increase in thrust. For an altitude of 

4,181 ft, 0.26 Mach, standard day temperature, a 50 hr engine, and a fixed idle fuel flow rate (open loop), 

the baseline engine outputs are compared to the outputs with a 5 decrease in the VSV, shown in 

Figure 11. Figure 11 indicates that modifying the VSV causes a decrease in the thrust, fan speed, 

combustor static pressure, and low pressure turbine exit pressure (P50), and an increase in both the core 

speed and exhaust gas temperature. 

For closed loop operation using EPR as the setpoint, modifying the VSV angle decreases P50, 

decreasing EPR, which causes an error between the desired and actual EPR. The fuel flow rate increases 

to drive the EPR to the setpoint and the thrust increases to approximately the same thrust level as with 

nominal control. With the modified VSV schedule, the core speed is higher and therefore the transition 

time from flight idle to full power will be shorter due to the higher idle core speed. 

For some flight idle conditions, the ratio unit (RU) limiter will be active instead of the EPR controller. 

The RU value multiplied by the current combustor static pressure is the fuel flow rate determined by the 

RU control limiter. Modifying the VSV decreases the combustor static pressure, which will decrease the 

fuel flow rate. Therefore, the RU min value has to be increased to offset the pressure loss due to 

modifying the variable stator vane angle and to create the high speed flight idle condition when the RU

limiter is active. 
NASA/TM—2011-217004 18 

Figure 11.—Open loop engine outputs comparing the nominal engine outputs and 

modified VSV (Mod VSV) for a constant fuel flow. 

In addition to the VSV and RU min modifications, the modified acceleration schedule is included in 

the high speed idle control mode to provide the best results. Figure 12 compares the response for the 

nominal, MAS with a function that determines the offset values based on the operating condition, and 

high speed idle controllers at 2,013 ft, 0.26 Mach, 48 °R above standard day temperature, for a 50 hr 

engine. Table 8 contains the dead time, rise time, delay time, time constant, setting time, and minimum 

stall margin for three flight conditions for the nominal response, MAS, and high speed idle. 

Table 8 indicates that high speed idle produces a better response than MAS but with a slightly lower 

stall margin remaining. The stall margin for high speed idle could be increased by reducing the 

modification made to the MAS acceleration schedule. 

Another way to increase the performance of the high speed idle control mode is to further decrease 

the efficiency of the engine by bleeding some of the core flow during idle conditions. Increasing the 

customer bleed flow should increase the core speed further, resulting in a higher idle condition. The 

customer bleed valve is shown in Figure 1, and reduces the flow through the core (W36). The Nf, Nc, net 

thrust, Ps3, P50, and T50 engine outputs with a constant fuel flow (open loop) for both the nominal 

engine configuration and with the bleed valve active are shown in Figure 13. Figure 13 indicates that the 

core speed increases while the net thrust decreases with the customer bleed valve active. 

The results of the high speed idle with MAS and bleed valve are compared against both the nominal 

engine response and high speed idle with MAS for a large throttle transient in Figure 14. Figure 14 shows 

the increase in core speed with HSI-Bleed and approximately the same thrust at idle. The high speed idle 

with bleed valve clearly has the fastest thrust response at the flight condition shown in Figure 14. 

Table 9 contains the performance metrics and stall margins for the nominal, FastER, high speed idle, 

and high speed idle with customer bleed valve engine responses. Table 9 indicates that in terms of settling 

time and stall margin, high speed idle with customer bleed valve provides the best response. 

20 30 40 50

1940

1960

1980

2000

Nf, RPM

20 30 40 50

9250

9300

9350

9400

Nc, RPM

20 30 40 50

3000

3100

3200

3300

Fnt, lbf

20 30 40 50

98

100

102

104

Ps3, psi

time, s

20 30 40 50

13.48

13.5

13.52

13.54

13.56

P50, psi

time, s

20 30 40 50

1020

1030

1040

1050

1060

T50,  R

time, s

Nominal

Mod VSV
NASA/TM—2011-217004 19 

Figure 12.—Comparison of Nominal, Modified Acceleration Schedule (MAS), and High Speed Idle (HSI) 

responses for flight idle to full power throttle transient (2,013 ft, 0.26 Mach, 48 above standard day 

temperature, 50 hr engine). 

TABLE 8.—COMPARISON OF NOMINAL, MODIFIED ACCELERATION SCHEDULE, 

AND HIGH SPEED IDLE DATA FOR A THROTTLE TRANSIENT FROM 

FLIGHT IDLE TO FULL POWER 

Operating condition, 

alt/MN/dTamb 

Td, 

s 

Tr, 

s 

Tde, 

s 

Tc, 

s 

Ts, 

s 

SM, 

percent 

Nominal 

2013 ft / 0.26 / 48 °R 0.520 3.105 3.205 2.985 6.205 9.117 

4181 ft / 0.26 / 48 °R 0.520 3.240 3.220 3.000 6.460 7.577 

7431 ft / 0.26 / 48 °R 0.505 3.285 3.220 3.000 6.655 5.329 

MAS 

2013 ft / 0.26 / 48 °R 0.390 2.910 2.280 2.175 5.175 4.147 

4181 ft / 0.26 / 48 °R 0.405 3.090 2.445 2.310 5.550 5.777 

7431 ft / 0.26 / 48 °R 0.450 3.210 2.820 2.640 6.105 5.160 

High speed idle with MAS 

2013 ft / 0.26 / 48 °R 0.360 2.865 1.905 1.815 4.815 3.792 

4181 ft / 0.26 / 48 °R 0.360 3.060 2.010 1.935 5.130 4.146 

7431 ft / 0.26 / 48 R 0.375 3.165 2.235 2.130 5.505 4.907 

10 15 20

0

2

4

x 104

Fnt, lbf

10 15 20

0.5

1

1.5

x 104

Nc, RPM

10 15 20

0

2000

4000

Nf, RPM

10 15 20

0

2

4

Wf, lb/s

10 15 20

1000

1500

2000

T50,  R

time, s

10 15 20

0

20

40

HPC SM, %

time, s

Nominal

MAS

HSI
NASA/TM—2011-217004 20 

Figure 13.—Engine outputs showing the effect of bleeding off 10 lb/s from the core flow 

(4,181 ft, 0.26 Mach, standard day temperature, 50 hr engine). 

Figure 14.—Engine response comparing nominal, high speed idle with MAS (HSI), and high 

speed idle with MAS and customer bleed (HSI-bleed) (4,181 ft, 0.26 Mach, standard day 

temperature, end of life engine). 

20 30 40 50

1750

1800

1850

1900

1950

Nf, RPM

20 30 40 50

9150

9200

9250

Nc, RPM

20 30 40 50

2200

2400

2600

2800

3000

Fnt, lbf

20 30 40 50

85

90

95

100

Ps3, psi

time, s

20 30 40 50

13.3

13.35

13.4

13.45

13.5

P50, psi

time, s

20 30 40 50

1000

1050

1100

T50,  R

time, s

Nominal

Bleed

10 15 20

0

2

4

x 104

Fnt, lbf

10 15 20

0.8

1

1.2

x 104

Nc, RPM

10 15 20

0

2000

4000

Nf, RPM

10 15 20

0

2

4

Wf, lb/s

Time, s

10 15 20

1000

1500

2000

T50,  R

Time, s

10 15 20

0

50

HPC SM, %

Time, s

Nominal

HSI

HSI-Bleed
NASA/TM—2011-217004 21 

TABLE 9.—RESPONSE OF AN END OF LIFE ENGINE COMPARING THE NOMINAL, MAS, 

HIGH SPEED IDLE, AND HIGH SPEED IDLE WITH CUSTOMER BLEED VALVE, 

FOR A THROTTLE TRANSIENT FROM FLIGHT IDLE TO FULL POWER 

Operating condition, 

alt, MN, dTamb 

Td, 

s 

Tr, 

s 

Tde, 

s 

Tc, 

s 

Ts, 

s 

SM, 

percent 

Nominal 

2013 ft / 0.26 / 0 R 0.385 2.055 2.440 2.265 4.180 7.360 

4181 ft / 0.26 / 0 R 0.385 2.130 2.530 2.37 4.465 5.955 

7431 ft / 0.26 / 0 R 0.385 2.295 2.590 2.475 4.840 3.631 

MAS 

2013 ft / 0.26 / 0 R 0.34 1.785 1.81 1.665 3.52 3.237 

4181 ft / 0.26 / 0 R 0.34 1.845 1.855 1.71 3.715 1.746 

7431 ft / 0.26 / 0 R 0.325 1.995 1.87 1.9 4.015 –0.605 

High speed idle with MAS 

2013 ft / 0.26 / 0 R 0.345 1.980 2.070 1.92 3.675 7.221 

4181 ft / 0.26 / 0 R 0.360 2.040 2.145 2.01 3.915 5.790 

7431 ft / 0.26 / 0 R 0.360 2.145 2.235 2.115 4.200 3.502 

High speed idle with MAS and bleed valve 

2013 ft / 0.26 / 0 R 0.360 1.830 1.815 1.65 3.450 6.915 

4181 ft / 0.26 / 0 R 0.345 1.890 1.845 1.71 3.615 5.887 

7431 ft / 0.26 / 0 R 0.345 2.010 1.890 1.785 3.870 4.306 

Summary 

This paper presents the results of a sensitivity study using the Commercial Modular Aero-Propulsion 

System Simulation 40k (C-MAPSS40k) to obtain additional performance from a commercial aircraft 

engine in terms of additional thrust production and faster engine response during critical high risk 

situations. The main idea is to accept an increase in the risk of engine failure, but in return minimize the 

overall risk to the vehicle. This paper specifically investigates the risk of failure for either an increase in 

thrust production beyond the rated maximum, or an increase in the responsiveness of the engine to the 

pilot’s throttle command. 

For the overthrust scenario, this study indicates that additional thrust production is possible using the 

current controller architecture with an increase in the risk of failure due to increased turbine temperatures. 

To limit the amount of risk taken, a turbine temperature limit could be added to the control structure for 

overthrust operation only. In addition, the time for which the additional thrust is requested has a large 

impact on the risk of failure. It is worth restating that the risk model considered in the study of engine 

overthrust is only representative of the trends observed in the deterioration of an engine due to excessive 

temperatures and speeds. If and when a more accurate model is developed, these tests should be repeated. 

In the fast engine response scenario, three different methods of increasing the responsiveness are 

proposed. The first is to modify the linear proportional integral controller implemented in C-MAPSS40k, 

which specifically addresses small throttle transients. This method involved modifying the integrator 

term, thus increasing the bandwidth, which causes a decrease in the stability margins (gain and phase). It 

was determined, based on the original design specification, that increasing the integral term by 40 percent 

would be the greatest modification that could be made and still maintain an appropriate amount of phase 

margin. Modification of the proportional integral controller will be used for any size throttle command, 

and is the only effective controller modification for small throttle commands when the protection logic 

controllers are not limiting the engine response. 

The second method of increasing the responsiveness of the engine is to modify the acceleration 

schedule. This method involves adding an offset to the nominal acceleration schedule, which increases the 

performance at the cost of reduced stall margin. This method does work, but if the available stall margin 

is low, this method does not provide much improvement, thus it is only helpful for large throttle 

command at flight conditions where the available minimum stall margin is relatively high. 
NASA/TM—2011-217004 22 

The final method to increase the responsiveness of the engine is to implement a high speed idle 

controller along with the other techniques. The idea behind the high speed idle control mode is to operate 

the engine at a higher core speed for a given amount of thrust during critical periods when increased 

responsiveness is warranted. Two modifications made to increase the core speed while maintaining the 

same thrust are to change the variable stator vane schedule, and to increase the flow rate through the 

customer bleed valve. In addition, the Ratio Unit minimum limiter is increased to ensure a higher idle 

power level, in terms of core speed. The data indicate that at worst there is a small reduction in the stall 

margin, while achieving a performance improvement in terms of decreased settling time, time constant, 

rise time, etc. 

References 

1. May, R.D., Csank, J., Lavelle, T.M., Litt, J.S., and Guo, T.-H., “A High-Fidelity Simulation of a 

Generic Commercial Aircraft Engine and Controller,” AIAA–2010–6630, 46th AIAA/ASME/ 

SAE/ASEE Joint Propulsion Conference & Exhibit, Nashville, TN, July 25–28, 2010. 

2. Csank, J., May, R.D., Litt, J.S., and Guo, T.-H., “Control Design for a Generic Commercial Aircraft 

Engine,” AIAA–2010–6629, 46th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 

Nashville, TN, July 25–28, 2010. 

3. Burcham, F.W., Jr, and Fullerton, C.G., “Controlling Crippled Aircraft—With Throttles,” NASA 

TM-104238, 1991. 

4. Burcham, F.W., Jr., Fullerton, C.G., and Maine, T.A., “Manual Manipulation of Engine Throttles for 

Emergency Flight Control,” NASA/TM—2004-212045, January 2004. 

5. Litt, J.S., Frederick, D.K., and Guo, T.-H., “The Case for Intelligent Propulsion Control for Fast 

Engine Response,” AIAA–2009–1876, AIAA Infotech@Aerospace Conference, Seattle WA, 

June 6–9, 2009. 

6. DeCastro, J.A., Litt, J.S., and Frederick, D.K., “A Modular Aero-Propulsion System Simulation of a 

Large Commercial Aircraft Engine,” AIAA–2008–4579, 44th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference & Exhibit, Hartford, CT, July 21–23, 2008. 

7. Jaw, L., and Mattingly, J.D., Aircraft Engine Controls: Design, System Analysis, and Health 

Monitoring, American Institute of Aeronautics and Astronautics, Inc., Virginia, 2009. 

8. Litt, J.S., Sharp, L.M., and Guo, T.-H., “A Risk Assessment Architecture for Enhanced Engine 

Operation,” AIAA–2010–3469, AIAA Infotech@Aerospace 2010, April 20–22, 2010, Atlanta, 

Georgia, July 2010. 

9. Guo, T., and Litt, J.S., “Risk Management for Intelligent Fast Engine Response Control,” 

AIAA–2009–1873, AIAA Infotech@Aerospace Conference, Seattle WA, June 6–9, 2009. 

10. National Transportation Safety Board, “Aircraft Accident Report, United Airlines Flight 232, 

McDonnell Douglas DC-10-10, Sioux Gateway Airport, Sioux City, Iowa, July 19, 1989,” 

PB90-910406, NTSB/AAR-90/06, 1990. 
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A Sensitivity Study of Commercial Aircraft Engine Response for Emergency Situations 

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14. ABSTRACT

This paper contains the details of a sensitivity study in which the variation in a commercial aircraft engine’s outputs is observed for 

perturbations in its operating condition inputs or control parameters. This study seeks to determine the extent to which various controller 

limits can be modified to improve engine performance, while capturing the increased risk that results from the changes. In an emergency, the 

engine may be required to produce additional thrust, respond faster, or both, to improve the survivability of the aircraft. The objective of this 

paper is to propose changes to the engine controller and determine the costs and benefits of the additional capabilities produced by the 

engine. This study indicates that the aircraft engine is capable of producing additional thrust, but at the cost of an increased risk of an engine 

failure due to higher turbine temperatures and rotor speeds. The engine can also respond more quickly to transient commands, but this action 

reduces the remaining stall margin to possibly dangerous levels. To improve transient response in landing scenarios, a control mode known 

as High Speed Idle is proposed that increases the responsiveness of the engine and conserves stall margin. 

15. SUBJECT TERMS

Turbofan; Control; Risk; Sensitivity analysis; Engine control 

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