Sanjay Garg

Glenn Research Center, Cleveland, Ohio

Aircraft Turbine Engine Control Research at

NASA Glenn Research Center

NASA/TM—2013-217821

April 2013
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Sanjay Garg

Glenn Research Center, Cleveland, Ohio

Aircraft Turbine Engine Control Research at

NASA Glenn Research Center

NASA/TM—2013-217821

April 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA/TM—2013-217821 1 

Aircraft Turbine Engine Control Research at

NASA Glenn Research Center

Sanjay Garg

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

This paper provides an overview of the aircraft turbine engine control research at the NASA Glenn 

Research Center (GRC). A brief introduction to the engine control problem is first provided with a 

description of the state-of-the-art control law structure. A historical aspect of engine control development 

since the 1940s is then provided with a special emphasis on the contributions of GRC. With the increased 

emphasis on aircraft safety, enhanced performance, and affordability, as well as the need to reduce the 

environmental impact of aircraft, there are many new challenges being faced by the designers of aircraft 

propulsion systems. The Controls and Dynamics Branch (CDB) at GRC is leading and participating in 

various projects to develop advanced propulsion controls and diagnostics technologies that will help meet 

the challenging goals of NASA Aeronautics Research Mission programs. The rest of the paper provides 

an overview of the various CDB technology development activities in aircraft engine control and 

diagnostics, both current and some accomplished in the recent past. The motivation for each of the 

research efforts, the research approach, technical challenges, and the key progress to date are summarized. 

Introduction

The modern dual spool turbofan engine used for commercial and military aircraft has evolved into its 

current form over the past 60 plus years, starting with the first U.S. jet engine, which was built in 1942 by 

General Electric based on the British Whittle turbojet engine. Controls technology has played a critical 

role in advancing the performance, reliability, operating life, and safety of modern aircraft engine. Jaw 

and Garg (2003) provides a broad historical overview of the propulsion control technology development 

in the United States. This paper will focus on the turbine engine control research done at the NASA Glenn 

Research Center (GRC).1

Simply stated, an engine control system determines the amount of fuel needed for the engine to 

produce a desired power (or thrust), based on pilot’s power request through a throttle (or a power lever); it 

then meters the right amount of fuel to the engine’s combustion chamber(s); and it maintains the engine 

power at the desired level in the presence of air flow disturbance and changes in flight conditions. The 

metering valve is usually called an actuator, whose position changes with the fuel flow command, and the 

fuel flow is called a control variable. Direct in-flight engine thrust measurements are not possible. So a 

good indicator of thrust, such as either engine shaft rotational speed (N) or engine pressure ratio (EPR, 

defined as the ratio of engine exhaust (nozzle) total pressure to engine intake (fan face) total pressure), is 

controlled to provide effective thrust regulation. An aircraft engine is designed to operate in a wide 

operating envelope (to support aircraft mission profiles). Typically, the altitude can vary from sea level to 

50,000 ft (or even higher), and for supersonic aircraft the air speed can go beyond Mach 3; furthermore, 

 

1 Although the center is referred to as “GRC” throughout this report, it has had different names throughout its 70-yr 

history. It began operations in 1942 under the National Advisory Committee for Aeronautics (NACA) as the Aircraft 

Engine Research Laboratory. The name was changed to the NACA Flight Propulsion Laboratory in 1947 and to the 

NACA Lewis Flight Propulsion Laboratory in 1948. When the NACA became part of the new National Aeronautics 

and Space Administration (NASA) in 1958, the name was changed to the NASA Lewis Research Center. Finally, 

the name was changed to the NASA John H. Glenn Research Center (GRC) in 1999.
NASA/TM—2013-217821 2 

the air temperature at the same altitude and airspeed can vary from that on a hot summer day to a cold 

winter night. Additionally, the performance of a turbine engine deteriorates with usage because of

component degradation caused by erosion, corrosion, foreign object ingestion, increased clearances 

between the rotating components and the seals, etc. The requirement for engine control is to ensure a 

minimum required level of performance, defined as the thrust achieved at a given throttle setting and 

given ambient conditions, throughout the operational envelope and certified operating life of the engine. 

This requirement for guaranteed performance with ambient condition variations and expected component 

degradations imposes severe challenges on control system design.

Engine operability constraints provide additional challenges for the control design. The control 

system has to ensure that the engine is operated safely; that is, without rotor overspeed, compressor stall, 

combustor blowout, or turbine overtemperature. This has to be done while maintaining acceptable 

acoustic and emissions levels. These considerations require nonlinear, often heuristic, control logic that 

complicates the problem further.

The main focus of discussion for this paper is the modern engine control referred to as “full authority 

digital engine control (FADEC). The term “full authority” implies that the computer-based controller has 

full authority over the command signals provided to the engine inputs with no manual override possible. 

The FADEC enables the pilot to move the throttle to any position during any operating phase of the 

airplane, and the control logic in the FADEC ensures that the engine will be able to operate safely. The 

term “full authority” stems from the fact that the early implementations of digital engine control had 

mechanical backup systems and/or capability for the pilot to override the digital control. This was to 

protect the engine from surging—a phenomenon during which the air flow in the compressor reverses, 

causing severe damage to the engine. Today’s modern aircraft engines are typically equipped with dualchannel FADECs, with extensive built-in test functions for the control electronics, and some basic form 

of embedded engine analytical models to perform sensor and actuator validation checks. 

This paper will focus on the control logic that is incorporated in the FADEC to provide safe and 

reliable operation of the engine throughout the operating envelope while also guaranteeing a certain 

operating life of the engine. Furthermore, this paper focuses on the high-bypass turbofans used in 

commercial aircraft. Typically an engine is certified to provide a minimum level of performance, thrust 

greater than a specified value for a given throttle setting at a given operating condition (ambient pressure 

and temperature), and a specified maximum thrust rise time for throttle movement from idle to maximum 

thrust setting. For commercial aircraft, a typical engine operating cycle consists of modulating the thrust 

from idle to maximum thrust for takeoff, setting the desired thrust levels for climb and cruise, and then 

setting the engine to idle for landing. As will be discussed later in this paper, the control logic plays a 

critical role in meeting these performance guarantees while ensuring safe operation of the engine. 

The paper is organized as follows. First a brief description of the state of the art of engine control is 

provided to familiarize the reader with the challenges of engine control design and how these are 

currently addressed. Historical contributions by GRC in the early stages of turbine engine technology 

development and then as technology advanced from hydro-mechanical to digital control are summarized. 

Major control technology developments recently completed by the Controls and Dynamics Branch (CDB) 

at GRC, and current ongoing research efforts are then described. A brief description of past GRC 

technology development for rocket engine control and current CDB technology development for fault 

management of space exploration systems is provided to give the reader an overall perspective of 

propulsion control technology development at GRC. The paper concludes with a summary of the 

discussed accomplishments and a look into the future of aircraft engine control. 

State of the Art of Engine Control

The intent of this brief overview of turbofan engine operation is to provide the reader with some context 

of the challenges associated with engine control. Spang and Brown (1999), DeCastro et al. (2008), and Jaw 

and Mattingly (2009) will enable the reader to gain a more complete understanding of this background 

information. The discussion in this section is based on the material in DeCastro et al. (2008). 
NASA/TM—2013-217821 3 

Figure 1.—High-bypass twin spool turbofan engine: N1 and N2 

are fan and core speeds; LPT, low-pressure turbine; LPC, 

low-pressure compressor; HPC, high-pressure compressor; 

and HPT, high-pressure turbine (image courtesy of NASA).

Figure 1 shows a schematic of a modern high-bypass dual-spool commercial aircraft turbofan engine. 

The air supplied by the inlet to the engine is compressed by the fan. A major portion of this compressed air 

is ducted to the outer gas flow path, referred to as “bypass air,” while the rest is routed to the engine core. 

The core air is further compressed by two multi-stage compressors: the low pressure compressor (LPC) and

high pressure compressor (HPC). The compressed air is mixed with fuel and burned in the combustor. The 

two turbines, high pressure turbine (HPT) and low pressure turbine (LPT), extract energy from the hot 

combustor air and convert it to mechanical (rotational) energy. The LPT drives the LPC and the fan while 

the HPT drives the HPC. The core air is expanded through the nozzle. In this engine configuration, the 

bypass air exhausted over the engine core creates most of the thrust. The bypass ratio is the ratio of the flow 

rate of the fan compressed air going through the bypass duct to the flow rate of the core air. The turbofan 

engine achieves fuel-efficient operation by accelerating a large amount of air through a small velocity 

change. Having a large proportion of thrust generated by the low-velocity air also makes the high-bypass 

turbofan much quieter. In Figure 1, N1 and N2 refer to the speeds of the two spools with N1 being fan speed 

and N2 being the core speed. Note that in general N1 is much lower than N2.

Although fuel flow is the main control variable to set the engine operation for a desired level of 

thrust, typical modern engines have other control variables such as a variable bleed valve (VBV) and 

variable stator vanes (VSV) to allow for safe and efficient operation of the engine. The VBV is to prevent 

the LPC from entering stall at any operating condition by providing pressure relief at the LPC exit. The 

VSV are typically located at the forward stages of the HPC and are used as a means to maintain 

compressor stability and prevent choking of the downstream compressor stages. These other control 

variables are typically scheduled as a function of the engine operating condition and are not used for 

direct feedback control.

Since the turbofan engine is a thermo-mechanical device, there are various limitations on its 

operation. Some of the limits of concern from a control perspective are structural and operational limits. 

To maintain the structural integrity of the rotating components, there is a maximum limit on the allowable 

fan and core shaft speeds. Also, since the blades in the first stator stage of the HPT see the highest flow 

temperatures (at the exit of the combustor), the maximum temperature at turbine inlet (T41) has to be 

limited to prevent the blades from getting hot enough to lose structural integrity. Operational limits 

consist of ensuring that: (1) adequate stall margins are maintained for the LPC and HPC throughout the 

engine operating envelope to prevent the compressors from approaching stall during engine transients;

and (2) a minimum fuel flow is maintained at any operating condition to prevent combustor blowout. 
NASA/TM—2013-217821 4 

From an acceptable operating life perspective, the maximum T41 is further restricted to prevent excessive 

thermo-mechanical fatigue damage to the turbine components. 

The objective of the control logic design is to provide acceptable performance while ensuring that the 

engine operation remains within the structural and operational limits. The engine control logic should

provide smooth, stable, and stall-free operation of the engine via a single pilot input—throttle or power 

lever angle (PLA)—with no restrictions on throttle movement. Furthermore, it should provide reliable and 

predictable throttle movement to thrust response throughout the engine operating life. Some of the 

challenges in engine control design are

Thrust, the primary controlled variable of interest cannot be measured.

Changes in ambient condition and aircraft maneuvers cause distortion of flow into the 

fan/compressor, so the engine control has to be robust under these conditions. 

Parameters that need to be limited, such as turbine inlet temperature, minimum stall margin, etc., 

cannot be directly measured.

Engine components degrade with usage, so the control logic has to be robust to provide guaranteed 

performance in the presence of these significant component degradations. One approach to estimate 

the effect of these degradations on engine performance is to model the effects through a set of 

influence coefficients called health parameters (Jaw and Mattingly 2009; Urban 1974). 

Additionally, the harsh operating environment of the engine, high temperatures and large vibrations, 

places certain constraints on the control hardware. One constraint important for control logic development

is the limitation it places on the number and type of sensors that can be placed on the engine. A typical 

engine control system has 6 or 7 sensors including the two rotor speeds (N1 and N2), pressure and 

temperature measurements at 2 or 3 engine locations such as the fan inlet and LPC and HPC exits, and an 

exhaust gas temperature (EGT) sensor, which measures the gas temperature at the exit of the LPT. As an 

engine ages with use, the clearance between the turbine blade tips and the turbine casings open up, 

increasing the amount of hot air leaking through the turbine and resulting in an increase in EGT. The EGT 

then provides a direct indication of the remaining useful engine life. 

A typical modern engine control logic is shown in the block diagram of Figure 2. Since thrust cannot 

be directly measured, a variable which can be sensed and is directly correlated with engine thrust is 

selected to be the primary controlled variable through the PLA. Figure 2 shows the engine fan speed (N1) 

as the primary controlled variable. Note that since the majority of the thrust in commercial turbofan 

Figure 2.—Block diagram of typical modern engine control logic (image courtesy of NASA).
NASA/TM—2013-217821 5 

Figure 3.—Engine limit protection implementation philosophy: 

Wf/Ps30 is ratio of fuel flow to combustor inlet pressure and 

2

R N is corrected core engine speed (image courtesy of 

NASA). 

engines is created from the fan bypass flow, N1 is an appropriate variable to regulate to achieve the 

desired throttle to thrust response. Another variable that can be used as the primary controlled variable is 

the engine pressure ratio (EPR), defined as the ratio of the pressure at the LPT exit to the pressure at the 

fan inlet. In Figure 2, a PLA setting creates a reference command for N1 and a simple lead-lag 

compensator creates a fuel flow command based on the difference between the commanded and measured 

fan speed. However, this command cannot be implemented as is because, as discussed earlier, the engine 

structural and operational limits need to be taken into consideration. The acceleration and deceleration 

schedules are implemented primarily to prevent compressor stall during transients. Secondarily, the 

acceleration limit is an additional measure to prevent overstress of the rotor, and the deceleration limit 

protects against combustor blowout. These schedules are a function of the corrected core engine speed

( R N2 ; see Mattingly (1996) for details on how engine performance parameters are corrected to the sealevel standard day condition to take out the variability due to operating condition) and impose a limit on 

the ratio of fuel flow to combustor inlet pressure (Wf/Ps30) since this ratio, which can be computed from 

measurements, relates closely to surge and temperature margins. Figure 3 shows a typical example of how 

these schedules protect the engine operation from violating the structural and operational limits. 

Since the acceleration schedule imposes a maximum limit on the fuel flow, the lower of the fuel flow 

imposed by the acceleration limit and that required to track the regulated variable is selected through the 

“Select Low” block in Figure 2. The output of this block is compared with the minimum fuel flow 

command required to stay within the deceleration limit, and the higher value is selected through the 

“Select High” block. The output of the “Select High” block is further compared with the fuel flow 

required to stay within additional limits, such as the maximum core speed, etc., and the selected 

incremental fuel flow is sent as the command to the fuel pump and metering system. This Min-Max 

selection control architecture ensures that the fuel flow command sent to the fuel flow actuation system is 

such that the engine operation will avoid exceeding any of the limits. 

The engine acceleration and deceleration schedules are determined though simulation and analytical 

studies, and the engine set-point schedules are selected such that the desired performance setting can be 

achieved without operating on a limit in the steady state. The performance variable regulation and other 

limit variable regulation is typically implemented through PI (proportional + integral) control with the 

control gains designed using linear models and scheduled typically as a function of PLA and Mach no. 

and other additional variables that define the engine set point. Engine control design is developed and 
NASA/TM—2013-217821 6 

evaluated using dynamic engine models. The following provides a brief overview of the engine dynamic 

modeling. The interested reader can find details in DeCastro et al. (2008) and the references therein. 

Steady-state engine performance is obtained from cycle calculations derived from component 

performance maps obtained through detailed component modeling and component tests. Corrected 

parameter techniques are used to reduce the number of points that need to be evaluated to estimate engine 

performance throughout the operating envelope. Dynamics are modeled through inertia (the rotor speeds), 

combustion delays, heat soak and sink modeling, etc. This is a computationally intensive process since it 

is important to maintain mass/momentum/energy balance through each component. A detailed 

thermodynamic cycle deck is developed using this procedure and parameters within this model are 

adjusted based on experimental data to match the tested engine performance characteristics. Simplified 

models from this thermo-cycle deck are generated to develop and evaluate the control logic. The control 

design is evaluated throughout the operating envelope using a nonlinear aerothermal engine simulation 

and for engines at various levels of condition in the operating life, as modeled through the variation of 

health parameters. 

One of the typical transient test cases for engine control is “burst-chop” wherein the PLA is moved 

from idle to full (burst), and a few seconds later it is moved back to idle (chop). The burst-chop results 

from the nonlinear engine simulation evaluation for the control design discussed in DeCastro et al. (2008) 

are shown in Figure 4. In Figure 4, TRA stands for throttle resolver angle, which is analogous to PLA 

discussed earlier, Nf and Nc are fan and core speeds, respectively, and T48 is the exhaust gas temperature. 

Figure 4.—Engine response for burst and chop (from DeCastro et al. 2008; courtesy NASA): Nf and Nc are fan and 

core speeds, respectively; T48, exhaust gas temperature; and Wf/Ps30, ratio of fuel flow to combustor inlet pressure. 
NASA/TM—2013-217821 7 

The plots in Figure 4 show that the fan-speed compensator (in control the majority of the time) produces a 

smooth fan speed response that does not violate any limits. The VBV and VSV actuators follow schedules 

that are inversely proportional to the corrected fan and core speeds, respectively. The only limiter that 

comes into play is the Wf/Ps30 saturation limit; the maximum limit (Phi max to protect against stall and 

temperature limits) is encountered during burst, and the minimum (Phi min to protect against combustor 

blowout) is encountered during chop. 

Once the control design is determined to meet the requirements in terms of providing acceptable 

performance without violating any of the structural and operational limits, it is coded into software for 

implementation into the control hardware. The control gains are further adjusted based on engine ground 

and altitude tests and finally flight tests before releasing a product version of the control logic. 

Historical GRC Contributions

The year 2003 commemorated the 100th anniversary of the Wright brothers’ flight—the first powered

flight where both the airplane and the engine were designed and built by the Wright brothers. This first 

powered flight sparked enthusiasm about aviation around the world, leading to development of a mature 

industry that has revolutionized transportation. In the past 100 years, aviation has become a symbol of 

innovation with record-setting breakthroughs, many of which were contributed by advances in engines or 

propulsion systems. During this period, propulsion systems have evolved from piston engines (the first 

generation, 1903–1945) to jet engines (Garvin 1998; Mattingly 1996) (1942–present). GRC, since its 

inception in 1941, has made significant contributions to the development of engine control—both 

enhancing the fundamental understanding of engine dynamic performance and conducting research on 

concepts that improved the performance and safety of turbine engines. This section provides a brief 

overview of some of the most critical contributions. The section is divided into two subsections: early 

stages of turbine engine control (1945–1960s) and maturation of turbine engine control (1970–1990). 

Early Stages of Turbine Engine Control (1945–60s) 

In 1948, General Electric (GE) tested the first afterburning turbojet engine in the world, the J47. The 

engine incorporated a hydromechanical fuel control for the main combustion chamber and an electronic 

(vacuum tube) fuel control for the afterburner. The original control law for the J47 was designed using 

only frequency response techniques. The engine underwent altitude testing at GRC. During the testing, it 

was found that the noise in the speed sensor, coupled with the high gain of the speed governor, caused the 

engine to limit cycle. To solve this problem, GE and NASA engineers worked together and applied the 

time-domain step response analysis method. The problem was soon fixed by reducing the control gain at 

altitude. This industry-government cooperation experience established a good foundation for building a 

knowledge base for engine controls that continues today. Some other major GRC contributions during 

this time period are briefly described below. 

Corrected Parameters for Model Simplification

Engine steady-state performance was obtained from cycle calculations using component performance 

maps derived from component test results, constant specific heat gas properties, and an iterative process 

for balancing internal engine flow and energy transfers. Performance calculations at each engine 

operating point could take several hours by slide rule or desk calculator. Generating an engine 

performance map for the entire operating envelope (altitude and Mach no.) would take months. Also, it 

was impractical to collect experimental data at the large number of possible operating points. The concept 

of corrected parameter techniques (corrected flows and speeds, temperature ratios, and pressure ratios) 

was developed by NASA engineers to reduce the number of operating points needed to be evaluated to 
NASA/TM—2013-217821 8 

define engine operation over the engine’s operating envelope (Mattingly 1996). The corrected parameter 

approach considerably reduced the amount of experimental data that needed to be collected and the time it 

took to perform the calculations to develop a complete engine performance map. The corrected parameter 

technique is being used to this day for efficient control design and analysis.

Simplified Dynamic Model of Engine Response

The early engine control was a hydromechanical governor, which metered the fuel flow going into the 

engine to be proportional to the difference between the set shaft speed and the actual speed of the turbine. 

The governor design was determined experimentally so that the proportional gain provided by the 

governor will provide a well-damped response. Dynamic behavior of the single-shaft turbojet engine was

first studied at GRC in 1948. The study focused on the dynamics of the shaft speed, because it is almost 

analogous to thrust and can be more easily and more accurately measured. The study showed that the 

dynamic characteristic of a turbojet engine (i.e., the transfer function from fuel flow to engine speed) can 

be represented by a first-order lag linear system with a time constant (Otto and Taylor 1950). The results 

from Otto and Taylor (1950) are summarized in Figure 5, which shows the experimental and calculated 

results for corrected engine time constant as a function of the corrected engine fan speed. This early 

analysis allowed the use of feedback control theory to design the control gains and significantly reduced 

the experimental iteration needed to achieve the desired closed-loop response. The first-order 

approximation of the engine response, with the time constant as a function of the operating condition, is 

still used today to generate simplified engine models that can be integrated with aircraft simulations for 

real-time piloted simulation evaluations of aircraft performance.

Real-Time Engine Dynamic Modeling Using Computers

In 1951, Pratt & Whitney (PW) flight-tested the first two-spool turbojet engine in the world, the J57, 

which helped achieve supersonic speed on a YF-100 aircraft. As engine capabilities advanced in the 

1950s, engine control technologies also advanced to deliver the new capabilities. As the engine 

technology matured to high compression ratio, by-pass flow turbofans during this period, the control 

technology also matured to variable geometry controls (e.g., the compressor stator control and intake and 

nozzle controls). With multiple control effectors, the design of control logic became too complicated to be 

performed and evaluated entirely through testing. The simultaneous emergence of computer technology 

provided the opportunity to develop computer-based real-time dynamic simulations of engines, which 

could then be used for control design and analysis. Initial development of simulations of the dynamic 

Figure 5.—Corrected engine time constant as a function of corrected 

engine speed (from Otto and Taylor 1950; courtesy NASA). 
NASA/TM—2013-217821 9 

Figure 6.—Comparison of simulated and experimental 

responses of turbojet engine (from Ketchum and Craig 

1952; courtesy NASA). 

performance of gas turbine engines was achieved through the use of electronic analog computers at GRC 

in the early 1950s. During this time, the response of a turbojet engine to a step change in fuel flow was 

simulated and the results were validated experimentally (Ketchum and Craig 1952). Figure 6 shows an 

example result from Ketchum and Craig (1952). The confidence generated through this initial effort in the 

capability to be able to accurately simulate complex engine dynamic behavior led to further advancements 

in the simulation methods and helped keep the time and cost of control design development and validation 

commensurate with that for other technologies. 

Maturation of Turbine Engine Control (1970–1990) 

As engine control functionality expanded, control system complexity also increased. Control 

complexity is reflected by the number of controlled variables managed by a control system. As control 

complexity increased, the traditional hydromechanical fuel control and servo components, although 

having been proven as highly reliable, were quickly reaching a practical limit; that is, they became 

increasingly bigger, heavier, and more expensive to be practical for aircraft engine controls. The 

emerging electronics and computer industry provided an opportunity to transition engine control from 

hydromechanical to computer based. In the early 1970s, analog and digital electronic control units (ECU) 

were designed to provide high-level supervisory functions. These supervisory control units primarily 

calculated speed or temperature set-points throughout the engine operating envelope. For some noncritical 

controlled variables, they would have sole control responsibilities. Early production supervisory control 
NASA/TM—2013-217821 10

units included: the digital electronic engine control (DEEC) unit for the PW F100 and the analog 

augmentor fan temperature (AFT) control for the GE F101. With the confidence built in using ECUs for 

engines, there was interest in adding additional capability to further enhance engine performance and 

provide fault tolerance capabilities. Some of the key GRC contributions for engine control and 

diagnostics during this period are summarized below. 

Digital Electronic Engine Control (DEEC)

In the early 1980s, NASA in collaboration with the U.S. Air Force and Pratt & Whitney conducted 

flight investigation of a DEEC system on an F100 engine in an F-15 airplane (Burcham et al. 1984). The 

DEEC was developed under PW Internal Research and Development Funds, altitude facility tests were 

conducted at GRC, and flight tests were conducted at NASA Dryden Flight Research Center (DFRC). 

The following is a summary of testing at GRC from paper No. 4, “NASA Lewis F100 Engine Testing,” in 

Burcham et al. (1984). Two builds of an F100 engine model derivative were evaluated in the Propulsion 

System Laboratory (PSL) altitude facility for improvements in engine components and DEEC control 

logic. Two DEEC flight logics were verified throughout the flight envelope in support of flight clearance 

for the tests on the F-15: a nozzle instability and a faster augmentor transient capability were successfully 

investigated in support of the F-15 DEEC flight program. Also included in the testing were identification 

of an off-schedule coupled-system mode fan flutter, DEEC noseboom pressure correlation, DEEC station 

6 pressure comparison, and a new compressor inlet variable vane (CIVV) schedule. Following the altitude 

testing at GRC, a 30-flight test program was conducted at DFRC. The results of the testing showed the 

DEEC to be a very capable controller for the F100 engine. This test program led to the introduction of 

digital engine control on military engines and ultimately to the introduction of the current-generation 

FADECs on commercial aircraft engines with the control logic as described in the earlier section of this 

paper. As mentioned earlier, the term “full authority” implies that the pilot has no restrictions on throttle 

movement throughout the aircraft flight envelope, and the engine controller maintains full authority to 

ensure safe operation of the engine. Prior to the introduction of the FADECs, the pilot typically needed to 

keep the throttle movements within cleared safety limits for a given flight condition in order to prevent 

engine stall.

Multivariable Engine Control Development

The successful DEEC testing on the F100 engine led to various research efforts to improve the 

functionality of digital engine control. One of these efforts focused on application of emerging 

multivariable control design techniques to the engine control problem. Air Force Wright Aeronautical 

Laboratory (AFWAL), now part of Air Force Research Laboratory (AFRL), and GRC jointly sponsored 

the Multivariable Control Synthesis (MVCS) program from 1975 to 1978. The program developed a 

series of Linear Quadratic Regulators (LQRs) and connected them with simple transition logic and several 

operating limits to achieve effective large-transient controls. The control design (DeHoff et al. 1977) 

(see Fig. 7) was verified extensively at NASA’s hybrid simulation facility (Szuch et al. 1977), and 

followed by a successful testing on an F100 engine in the GRC altitude test facilities (Lehtinen et al. 

1979). As discussed in an earlier section, even the current commercial engine control logic consists of 

controlling just the fuel flow (WF) in a “closed-loop” (feedback) manner, with the other variable 

geometry effectors such as inlet guide vanes on an “open-loop” schedule based on measured parameters 

that define the operating conditions. The F100 engine had considerable variable geometry including 

compressor inlet variable vanes (CIVV), rear compressor variable vanes (RCCV), nozzle jet area (Aj), and 

also customer bleed flow (CBF), which affected engine performance. Under the multivariable control 

design approach, all the engine sensors are used as input to control logic to generate commands for all the 

control effectors. The objective is to keep the engine operating at its maximum possible efficiency at each 

operating point. Testing on the F100 engine in the GRC altitude facility indicated that the engine could be 

successfully and safely controlled using MVCS. However, the performance benefits of the MVCS were 
NASA/TM—2013-217821 11

Figure 7.—F 100 LQR control block diagram (from Lehtinen et al. 

1979; courtesy NASA; acronyms are defined in the reference). 

not commensurate with the complexity of the control design and implementation. So MVCS did not make 

the transition to operating F100 engines. Since this MVCS study, multiple research studies have been 

done on multivariable control of engines (e.g., see Polly et al. 1988); however, this technology has not yet 

made a transition to currently flying engines except for the two-input two-output multivariable control on 

the F414-GE-400 engine for the U.S. Navy F18 aircraft. The reason is that there is sufficient bandwidth 

separation between the control effectors to negate any benefits of using multivariable control. The author 

has been told in verbal communications by industry researchers that PW has implemented an advanced 

multivariable control for the F135 engine; however, no detailed information is publicly available. Also, to 

the author’s best knowledge, no detailed information about the multivariable control on the F414-GE-400 

engine is publicly available.

Sensor Fault Detection, Isolation, and Accommodation

In 1970, the F100 was the first engine equipped with an onboard event history recorder (EHR). This 

recorder was the predecessor of the engine diagnostic unit (EDU) installed on later models of the engine 

in 1980s. The EDU and the DEEC together perform the engine monitoring system (EMS) function. The 

EMS provides the user with information on engine condition and engine life usage data. The engine 

condition data is used to isolate faulty components or systems during engine troubleshooting. The initial 

success in applying multivariable control design methods to jet engines prompted the AFWAL to explore 

failure detection and accommodation techniques (Small 1982). In the early 1980s, GRC explored an 

advanced approach to sensor failure detection under the Advanced Detection, Isolation and 

Accommodation (ADIA) program (Merrill et al. 1988). The objective of the program was to improve the 

overall demonstrated reliability of digital electronic control systems for turbine engines by detecting, 

isolating, and accommodating sensor failures using analytical redundancy methods. The program focused 

on sensor failures because the engine sensors are among the least reliable control components. The ADIA 

algorithm consisted of three elements: (1) hard sensor failure detection and isolation logic, (2) soft sensor 

failure detection and isolation logic, and (3) an accommodation filter. Hard failures are defined as out-ofrange or large bias errors that occur instantaneously in the sensed values. Soft failures are defined as small 

bias errors or drift errors that increase relatively slowly with time. The accommodation filter incorporates 

an engine model along with a Kalman filter to generate estimates of the engine states and outputs. The 

ADIA algorithm was integrated with an existing microcomputer implementation of the F100 

multivariable control algorithm, and a real-time simulation evaluation of the algorithm was performed 

using a hybrid computer simulation of a F100 engine (Merrill et al. 1988). The success of the real-time 

evaluation provided the confidence to proceed with testing and evaluation of the ADIA algorithms on an 

F100 engine in the GRC altitude test chamber. Engine sensor failure detection and accommodation were 
NASA/TM—2013-217821 12

demonstrated at two power conditions (Merill et al. 1987). Transient engine operation over the full power 

range with a single sensor failed and accommodated was successfully demonstrated. The GRC team and 

industry partners in the ADIA program received the prestigious R&D 100 award for this technology 

development and demonstration. Although the ADIA approach is not yet fully implemented in currently

operating engine control systems because of the processing power required for full envelope 

implementation as well as certification issues with using a model-based sensor estimate for engine 

control, some elements of this technology were incorporated in the FADEC logic for sensor failure 

detection and isolation.

Advanced Engine Control

With the success of the F100 MVCS and ADIA programs, there was renewed interest at GRC in 

investing in engine control technologies. The Controls group grew significantly during the 1990s with a 

multitude of technology development efforts under various NASA programs and in collaboration with Air 

Force research programs. The current Controls and Dynamics Branch (CDB) started with a modest 

complement of 5 staff members in the mid-1980s, and currently consists of 14 civil servants and 8 

contractors with additional support being provided by various employees from other GRC organizations. 

The research emphasis was on how to apply emerging multivariable and intelligent control approaches to 

provide improved functionality and performance for engines. The number of sensors and actuators on the 

engines continued to increase, thus providing an opportunity for advanced controls research. The 

continued advancement in computer technology also opened the possibilities of incorporating more 

complexity into the control logic in the FADECs. The idea of using active control of engine components 

to maximize the achievable performance throughout the operating envelope, and not be constrained to 

accept the suboptimal performance at other than design points, began to emerge. With the reorganization 

of NASA Aeronautics Research Mission programs in 2006, there was additional opportunity to 

investigate nontraditional uses of the engines, enabled through advanced controls, to increase aviation 

safety, and to start formulating engine control architectures that will help meet the challenging goals of 

increased efficiency and reduced environmental impact. Some of the significant accomplishments of CDB 

in the mid-1990s to early 2000s are described in this section, followed by a brief description of the main 

current research efforts.

Recent Significant Accomplishments

Integrated Flight/Propulsion Control

In the mid-1980s to 1990s, the trend in future military fighter/tactical aircraft design was towards 

aircraft with new/enhanced maneuver capabilities such as short takeoff and vertical landing (STOVL) and 

high-angle-of-attack performance. An integrated flight/propulsion control (IFPC) system is required in 

order to obtain these enhanced capabilities with reasonable pilot workload. Under an AFWAL sponsored 

“Design Methods for Integrated Systems” (DMICS) program in the mid-1980s, two contractor teams had 

developed two very different approaches to IFPC design: (1) a centralized approach based on the Linear 

Quadratic Gaussian/Loop Transfer Recovery (LQG/LTR) method (Smith 1986) and (2) a decentralized, 

hierarchical approach using LQR-based explicit model-following for control synthesis (Shaw et al. 1988). 

In the late 1980s, NASA in collaboration with the U.S. Air Force started a concerted effort on 

investigating IFPC designs for advanced short takeoff vertical landing (ASTOVL) aircraft with 

supersonic cruise capability. Two contractor teams, each consisting of an airplane and an engine 

company, developed IFPC designs for two different conceptual ASTOVL aircraft with a view towards 

evaluating them in piloted simulations in the Vertical Motion Simulator (VMS) at the NASA Ames 

Research Center (ARC). In parallel, GRC started an in-house STOVL/IFPC program to develop advanced 

IFPC design methods based on the lessons learned from the DMICS studies. The GRC-developed IFPC 

design approach combined the best aspects of the two DMICS studies and was referred to as Integrated 
NASA/TM—2013-217821 13

Methodology for Propulsion and Airframe Control (IMPAC) (Garg et al. 1991). The major IMPAC 

design steps are (1) generation of integrated airframe/engine models for control design; (2) centralized 

control design, considering the airframe and engine system as an integrated system; (3) partitioning of the 

centralized controller into separate airframe and engine subcontrollers; (4) Operational flight envelope 

expansion through scheduling of the partitioned subcontrollers; (5) nonlinear design such as incorporation 

of limit logic for operational safety; and (6) full system controller assembly and evaluation. 

The IMPAC approach was applied for IFPC design for a conceptual ASTOVL aircraft in the landing 

approach to hover transition flight phase (Garg and Mattern 1994). During this phase the control of the 

aircraft transitions from forces and moments generated by aerodynamic control effectors to purely 

propulsion system generated forces and moments. The ASTOVL aircraft concept had multiple control 

effectors including left and right elevons used collectively as elevator and differentially as ailerons; a 

rudder; ejectors on the wing to provide propulsive lift at low speeds and hover; a two-dimensional 

convergent-divergent (2D-CD) vectoring aft nozzle; a vectoring ventral nozzle for pitch control and lift 

augmentation during transition; and jet reaction control systems (RCS) for pitch, roll, and yaw control 

during transition and hover. During the transition flight phase, all these control effectors are active, and it 

is quite a complex problem to coordinate their usage in a way that is transparent to the pilot and also 

provides the aircraft response characteristics, which allow the pilot to complete landing-related tasks 

successfully within acceptable workload. The IFPC design developed using the IMPAC approach was 

evaluated in a fixed-base piloted simulator at GRC (Bright et al. 1994). In the simulation, the pilots 

successfully completed typical transition phase tasks such as combined constant deceleration with flight 

path tracking and constant acceleration wave-off maneuvers. The pilots could successfully perform abort 

sequences and maneuverability tasks without loss of control predictability or excessive workload.

Various technologies that are relevant to practical use of multivariable control design techniques were 

developed as part of the IMPAC-based IFPC design study (Garg and Mattern 1994). These are (1) a 

generic command tracking problem framework for robust control law synthesis using H-infinity control 

design techniques with rules of thumb for selecting various frequency weights in the design procedure; 

(2) a systematic procedure for partitioning a centralized controller into decentralized, hierarchical 

subsystem controllers; (3) a simplified scheme for controller scheduling, which exploits the robustness 

properties of centralized/partitioned control designs; and (4) a modified scheme for designing integrator 

wind-up protection (IWP) gains, which guarantees closed-loop system stability for single-actuator 

saturation. An example result which shows application of the optimized IWP technique to the propulsion 

control portion of the STOVL IFPC design (Mattern and Garg 1992) is shown in Figure 8. 

Figure 8.—Integrator wind-up protection for engine control (from Mattern 

and Garg 1992; courtesy NASA).
NASA/TM—2013-217821 14

Shown are the responses of the aft nozzle thrust (FG9) to a step command and the corresponding fuel 

flow (WF36) requirement with and without the IWP. Without IWP (dashed lines), when the fuel flow 

limit is encountered, the fuel flow command continues to grow because of integrator wind up. As the FG9 

command is reduced to zero, the integrator takes time to wind down and the FG9 response is degraded. 

With IWP active (solid lines), the fuel flow command tracks the fuel flow limit, and good FG9 tracking 

performance is obtained when the step command is removed. Although IMPAC methodology was 

successfully demonstrated by NASA, its application to current and future aircraft configurations was 

perceived as too complex during industry evaluation. However, the technologies developed for practical 

application of multivariable control design techniques were perceived to be of value by the industry and 

were transferred to industry through joint studies as documented in Watts and Garg (1995) and Frederick 

et al. (1996). 

Intelligent Life Extending Control

With the desire to reduce engine operating cost, the industry is interested in developing technologies 

that will allow the engine and its components to operate (remain on wing) longer, thus increasing the time 

between engine overhauls. How the engine is controlled has a significant impact on the life of the 

components. Typically, the propulsion system control design engineer attempts to get the maximum 

performance out of the system while maintaining safe operation. Recent studies have shown that small 

changes in engine operating parameters, such as turbine inlet temperature, can have a significant impact 

on the damage accrued by engine components while having little to no effect on engine performance. 

GRC developed the concept of Life Extending Control where the engine control system is designed to 

achieve the desired performance while minimizing the damage accrued in engine components, hence 

maximizing the usable engine life. The feasibility of this concept was demonstrated for the Space Shuttle 

Main Engine through simulation. Efforts in collaboration with industry led to the development of 

Intelligent Life Extending Control (ILEC) (Guo et al. 2004) for gas turbine engines with a focus on 

optimizing the engine acceleration schedule to minimize the damage accumulation in the hot-gas-path 

components, while meeting the Federal Aviation Administration (FAA) thrust response rise time 

requirement for PLA command from idle to maximum power. 

During takeoff, when the pilot pushes the throttle to move the engine from idle to maximum power, 

the engine control generates a fuel flow command based on acceleration logic that ensures that the 

maximum thrust is achieved within a time limit based on FAA requirements. The engine components 

accumulate damage during this transient because of the quick and large changes in temperature and 

aerodynamic loads. By adjusting the acceleration logic such that the time to achieve maximum thrust is 

just within the FAA requirements, the temperature and load changes on the engine components can be 

kept to the minimum required to meet the performance. This will result in reduced damage accumulation 

for each takeoff and hence increase the amount of time the engine can stay on the aircraft before a major 

overhaul is required. The idea of ILEC is then to design a smart acceleration logic for engine control that 

will minimize the thermomechanical fatigue damage accumulated during a typical engine acceleration 

transient from idle to full power without any noticeable loss in engine performance. A typical baseline 

acceleration schedule for a commercial turbofan engine, as well as schedules optimized to achieve 

acceptable takeoff transients, while minimizing life consumption for a specific hot section engine 

component, are shown in Figure 9. The acceleration schedule is defined as the limit on the core speed rate 

of change as a function of the core speed (NH in Fig. 9). The thrust transients that correspond to these 

schedules are shown in Figure 10. In Figure 10, Trise is defined as the time for the thrust to reach

95 percent of the maximum power level. The optimized schedules were developed for Trise ranging from 

the value for the baseline schedule to the maximum allowable as per FAA certification requirements. 

Simulation results demonstrated that the optimized acceleration schedule decreased component life 

consumption during takeoff significantly for the same rise time, and extending the rise time slightly 

reduced the life consumption even more (Gou et al. 2004). 
NASA/TM—2013-217821 15

Figure 9.—Baseline and optimized acceleration schedules for 

takeoff acceleration process (from Guo et al. 2004; courtesy 

NASA); Trise is time for thrust to reach 95% of commanded thrust.

Figure 10.—Thrust response curves for the nominal and optimized 

acceleration schedule (from Guo et al. 2004; courtesy NASA): 

Trise is time for thrust to reach 95 percent of commanded thrust.
NASA/TM—2013-217821 16

Figure 11.—Distortion tolerant control architecture developed 

under the HISTEC program (image courtesy of NASA).

High Stability Engine Control (HISTEC)

For aircraft engines, a safety margin, called the “stability margin,” is built into the operation of the 

engine to prevent inception of fan/compressor stall due to inlet distortions caused by aircraft maneuvers 

or atmospheric disturbances. This stability margin results in a performance penalty being paid even at low 

distortion operating conditions such as cruise. Engine companies have estimated that being able to 

actively control the engine to safely maintain low stability margins under such low distortion operating 

conditions can result in reduction of 2 percent or more in specific fuel consumption. 

GRC in partnership with PW, U.S. Air Force, McDonnell Douglas Aerospace, and DFRC developed 

and demonstrated technologies to allow for online active management of the engine stability margin. The 

distortion-tolerant control approach developed for the High Stability Engine Control (HISTEC) program 

(DeLaat et al. 1998) used a small number of engine-face pressure measurements combined with flight 

control data based predictions of angle of attack ( ) and sideslip ( ) to accurately estimate the actual 

distortion present. From this pressure-based distortion estimate, an onboard stability audit generated a 

time-varying stall margin requirement. The engine controller (see Fig. 11) then accommodated the 

distortion by acting upon the current stall margin requirement supplied by the onboard stability audit.

These technologies were demonstrated and evaluated in flight tests on a modified F-15 aircraft referred to 

as the Advanced Control Technologies for Integrated Vehicles (ACTIVE) aircraft. The flight 

demonstration was carried out in two parts: the first to show estimation and the second to show distortion 

accommodation. Postflight analysis showed that the HISTEC technologies were able to successfully 

estimate and accommodate distortion, transiently setting the stall margin requirement online and in real 

time. Flight demonstration of the HISTEC technologies significantly reduced the risk of transitioning 

these technologies to tactical and commercial aircraft engines.

Active Stall Control

The peak efficiency operating point of a turbine engine compressor is typically very near the 

compressor stall line as shown in Figure 12. In order to prevent catastrophic stall from occurring during 

large transients, the engine is operated with a sufficiently large safety margin. If the compressor can be 

safely operated closer to the designed compressor peak efficiency, then it will result in increased engine

efficiency leading to significant savings in aircraft fuel costs. Even if compressor efficiency drops, cycle 

efficiency can be higher at higher overall pressure ratios. As shown in Figure 12, the compressor stall line 
NASA/TM—2013-217821 17

Figure 12.—Active stall control extends operating region for 

engines (image courtesy of NASA).

can be moved up through active control thus allowing safe operation at peak efficiencies. Note that the 

concept of active stall control differs from the distortion-tolerant control demonstrated under the HISTEC 

program. The distortion-tolerant control moves the engine operating line up or down using the variable 

geometry control effectors in the engine, such as the nozzle throat area in the F-15 ACTIVE aircraft 

demonstration, to adjust the stall margin. Under the active stall control concept, however, the engine stall 

line is moved up through some active flow control. The active stall control is obtained by sensing 

pressure changes at the inlet face of the compressor, which will indicate flow distortion that is the

precursor to stall, and activating high bandwidth flow valves located around the circumference of the 

compressor that blow high-pressure air to counter the flow distortion before it builds up to stall. 

The challenges for implementation of active stall control are developing accurate models of the stall 

phenomenon that can be used for control design, developing high-bandwidth (on the order of 500 Hz) 

actuators for controlling flow valves, and understanding the effectiveness of different actuating schemes 

for stall control. The active stall control program at GRC was a cooperative program with industry and 

academia with strong participation from Massachusetts Institute of Technology (MIT). Active stall

control was demonstrated in the laboratory environment at GRC for a single-stage high-speed compressor 

Weigl et al. 1998). One of the significant GRC contributions was in developing methods for early 

detection of stall precursors to enable timely application of the stall control (Bright et al. 1997). 

Experimental investigations were later completed in collaboration with industry under a program 

sponsored by the U.S. Air Force to demonstrate the technology on a multiple-stage compressor in the 

presence of inlet distortion. The effects of bleeding air from rear stages of the compressor for injection in 

the front stages was simulated during this experimental effort. These results indicated that substantial 

extension in safe operating range of the compressor can be achieved by active stall control technology.

Current Research

The CDB is currently conducting research in advanced engine control technologies under the 

Fundamental Aeronautics Program and the Aviation Safety Program in the NASA Aeronautics Research 

Mission Directorate. Some of the most significant current efforts are described below. An overview of all 

the current CDB research efforts is provided in Garg (2010), which includes additional references for the 

research areas not covered in this paper.
NASA/TM—2013-217821 18

Model-Based Engine Control and Diagnostics 

So far the focus has primarily been on the engine control challenges without directly taking into 

consideration any information about the condition or health of the engine. Engine health monitoring can 

provide valuable information to enhance the engine control performance. The various on-wing health 

monitoring systems of today, which are a collection of separate, unrelated technologies, provide a basic 

level of monitoring. Their capabilities are relatively limited, and the information they provide is used 

mostly to initiate maintenance actions, not for real-time decision-making. One instance in which the 

information is used on wing is for sensor validation. The controller has some simple logic to perform 

checks of basic limit or rate of change on engine sensors and actuators. In some cases, on-board engine 

models are used in conjunction with the controller’s own sensor voting scheme to help determine which 

sensor is correct when redundant sensors disagree. Whereas the current engine control architecture, as 

discussed in an earlier section, and the traditional diagnostic techniques are time tested and reliable, 

advanced techniques provide the promise to meet the challenging requirements of improved fuel 

efficiency, increased durability and life, decreased life cycle costs, and improved operations. The use of 

an onboard engine model to meet the challenging control and diagnostics requirements has emerged as 

the most viable approach. The continuing increase in computer processing capability has reached the 

point where the use of model-based algorithms for diagnostics and control of aircraft engines has become 

practical. Model-based technologies offer the potential for creating intelligent propulsion systems—

defined as self-diagnostic, self-prognostic, self-optimizing, mission–adaptable, and, inherently robust—

that far exceed current systems in performance, reliability, and safety (Litt et al. 2004). 

The concept of model-based engine control and diagnostics is shown in Figure 13. The onboard realtime engine model is driven with the engine control inputs and contains a tracking filter that uses the 

sensed measurements to update the health parameters in the engine model. The health parameters reflect 

the effects of engine degradation with usage and any gas path component faults. The outputs of the 

onboard model are the estimates of the sensed values and other performance parameters of interest. 

Transition to model-based control can occur in several ways. First, faults can be accommodated by 

changing the control laws, in a predetermined way, when a fault is detected. The changes are designed, at 

a minimum, to take the engine to a safe state, and preferably allow the engine to operate safely with the 

best, although probably degraded, performance. Secondly, the model allows the loop to be closed on 

unmeasured values (e.g., thrust, stall margin, etc.) for which there is no sensor (i.e., virtual sensors). 

Finally, in its most advanced form, the model is used directly in the control enabling the control to 

automatically adjust as the model adapts to the mission, deterioration, faults, weather, etc. Here the 

control can be designed to maximize performance without excessive conservatism. 

Figure 13.—Model-based engine control and diagnostics architecture (image 

courtesy of NASA). 
NASA/TM—2013-217821 19

Adibhatla and Lewis (1997) provides an excellent example application of model-based control to 

aircraft engines. However, as discovered during the early research on model-based engine control, one of 

the challenges in engine application is that the number of health parameters to be estimated generally 

exceeds the number of sensors available for the tracking filter design, which poses an underdetermined 

estimation problem. A common approach to address this shortcoming is to estimate a subset of the health 

parameters, referred to as “model tuning parameters.” Although this approach enables the Kalman filter to 

tune the onboard model so that its outputs track measured (sensed) engine outputs, there can be 

significant error in estimation of unmeasured engine outputs because the impact of the entire set of health 

parameters will not be accurately represented within the Kalman filter model. 

Recently an innovative methodology (Simon and Garg 2010) has been developed at GRC that creates 

a tuning parameter vector defined as a linear combination of all health parameters and of appropriate 

dimension to enable Kalman filter estimation. Selection of this tuning parameter vector is performed 

using a multivariable iterative search routine that minimizes the theoretical mean-squared estimation error 

in the parameters of interest. The new methodology was validated in simulation using an aircraft turbofan 

engine model. The simulation results demonstrated that applying the enhanced tuning parameter selection 

methodology resulted in a significant reduction in average estimation error compared to the conventional 

approach of selecting a subset of health parameters as tuners. Figure 14 shows a comparison of engine 

thrust estimates from an engine model using the conventional approach of selecting a subset of health 

parameters for tuning the onboard model versus the recently developed optimal tuner selection approach. 

With the optimal tuner selection approach, both the variance and the steady-state error in the thrust 

estimation are significantly reduced.

With the optimal tuner selection approach developed in (Simon and Garg 2010), sufficient estimation 

accuracy of the unmeasured variables can be obtained to allow for implementation of direct control of 

thrust and stall margins using the model estimated value. Research on this application is currently 

ongoing at GRC. Figure 15 shows a preliminary result from this research. The optimal tuner approach 

was applied to the simulation of a modern commercial high bypass turbofan engine. Response of the 

engine from throttle input to thrust was compared for the baseline case of controlling the (EPR) using 

measured pressures versus closing the loop on the estimated thrust for direct thrust control. In Figure 15, 

the dashed line shows the bounds for thrust response with the EPR control as the engine degrades with 

usage. With direct thrust control, the thrust response is maintained within a very tight bound regardless of 

the engine condition. 

Figure 14.—Thrust estimation accuracy comparison—conventional 

versus enhanced (optimal) tuning parameter selection (image 

courtesy of NASA).
NASA/TM—2013-217821 20

Figure 15.—Thrust response comparison—EPR control versus direct 

thrust control (image courtesy of NASA).

The optimal-tuner-based onboard modeling approach has also proven to be extremely useful for gas 

path diagnostics. A full envelope self-tuning model for a typical commercial aircraft type engine 

simulation has been developed as documented in Simon et al. (2011). A technique for selecting a single 

globally optimal set of tuners applicable throughout an engine’s operating envelope has been 

demonstrated. This is necessary to enable interpolation between operating points within a piecewise linear 

design implementation. Theoretical evaluation results using an aircraft engine model showed that the 

application of a single globally optimal set of tuners results in only a slight increase in estimation errors 

for parameters of interest over using a set of tuners that are optimal for a particular operating point. The 

methodology developed in Simon et al. (2011) also takes into account the closed-loop operation of the 

engine. Most recently, a Kalman-filter-based approach for integrated online aircraft engine performance 

estimation and gas path fault diagnostics has been developed using the optimal tuner approach (Simon 

and Armstrong 2012). The Kalman filter is designed to estimate the vector of tuning parameters, 

appropriately sized to enable estimation. The estimated tuning parameters are then transformed into a 

larger vector of health parameters representing system performance deterioration and fault effects. The 

results of this study showed that basing fault isolation decisions solely on the estimated health parameter 

vector does not provide ideal results. Furthermore, expanding the number of the health parameters to 

address additional gas path faults causes a decrease in the estimation accuracy of those health parameters 

representative of turbomachinery performance deterioration. However, improved fault isolation 

performance was demonstrated through direct analysis of the estimated tuning parameters produced by 

the Kalman filter. This was found to provide equivalent or superior accuracy compared to the 

conventional fault isolation approach based on the analysis of sensed engine outputs, while also 

simplifying online implementation requirements. Results from the application of these techniques to an 

aircraft engine simulation are presented and discussed in Simon and Armstrong (2012). 

Engine Dynamic Models for Control and Diagnostics Technology Development

Development and validation of the various controls and diagnostics technologies discussed so far 

requires the availability of a transient engine simulation representing the dynamic behavior of the engine. 

Typically, the engine simulation cycle decks developed by industry are proprietary in nature and are not 

available for NASA to share with universities or small businesses conducting controls research. NASA 

has developed the Numerical Propulsion Simulation System (NPSS), which has now become the default 

simulation system being used by the aerospace industry. However, the NPSS system is geared more 
NASA/TM—2013-217821 21

towards doing cycle analysis or detailed engine component design and is not “user-friendly” for controls 

research. To overcome these shortcomings and meet the need of the engine controls and diagnostics 

researchers to have realistic engine simulation platforms to evaluate their ideas and concepts, the CDB

started an effort in the early 2000s to develop dynamic engine simulations in the Matlab/Simulink 

environment that could be made publicly available without any industry proprietary issues. 

In 2003, GRC released the Modular Aero-Propulsion System Simulation (MAPSS) software (Parker 

and Guo 2003), which provides a model of a conceptual advanced military turbofan engine with a 

research-type controller. The MAPSS code was made available through the GRC software repository and 

was downloaded by many researchers. Building on the success of MAPSS, CDB developed the 

Commercial Modular Aero-Propulsion Simulation (C-MAPSS) software code representing a commercial 

40,823 kg-f (90,000 lbf) thrust engine, originally modeled in NPSS. The C-MAPSS simulation (DeCastro 

et al. 2008) included an extensive graphical user interface (GUI) capability, offering the capability to 

generate linear models, design controllers using the KQ (or Edmunds) model-matching approach, and 

perform closed-loop system evaluations. The C-MAPSS was also the first publicly available (limited to 

U.S. citizens because of International Traffic in Arms Regulations (ITAR) restrictions) dynamic engine 

simulation with a realistic control logic representative of what is currently flying on modern commercial 

aircraft engines. The discussion and the results shown in the earlier section on the “State of the Art 

Engine Control” are based on the C-MAPSS simulation.

Most recently, CDB has developed the C-MAPSS40k software package, which is a high-fidelity 

dynamic simulation of a generic 18,142 kg-f (40,000 lbf) thrust class commercial engine with a 

representative controller (May et al. 2010). Flight test data collected from a highly instrumented engine 

was used to provide realism in the C-MAPSS40k model engine performance. This engine simulation was 

created especially for the development and evaluation of control strategies to use the engine as a flight 

control effector during emergencies. C-MAPSS40k incorporated a number of new features including an 

iterative solver and detailed compressor stall models. The iterative engine solver enforces mass flow 

balance through the engine at every time step. This feature produces a realistic simulation while allowing 

the user the ability to configure the solver for the specific test, such as small throttle steps, and still 

maintain faster-than-real-time operation. The detailed compressor stall models include the compressor 

map stall line as well as a number of stall margin debits such as stall line change due to engine 

deterioration, stall margin change due to rotor tip clearance changes, and the transient debit due to heat 

transfer between the mass flow and rotor blades and compressor casing. The inclusion of these debits 

allows control developers to have a good estimate of the required stall margin stack-up by eliminating 

some of the uncertainty. C-MAPSS40k also includes an “industry standard” controller for use as a 

baseline against which to compare new controllers and architectures. This baseline controller follows the 

most common design principles in current commercial engine controllers and produces realistic transient 

behavior at conditions across the complete engine flight envelope. With these capabilities, users in 

academia, industry, and government have access to a generic, high-fidelity engine simulation that can be 

used to quickly develop, simulate, and test new control system components and architectures and 

diagnostic algorithms. The C-MAPSS40k simulation is being used extensively in NASA-sponsored 

research programs under the Aviation Safety Program and the Fundamental Aeronautics Program and has 

been downloaded by multiple organizations including academia and industry. 

Distributed Engine Control 

Presently, engine control system architecture is based on a centralized design in which discrete 

sensors and effectors are directly wired to a single engine-mounted electronics package (see Fig. 16). This

avionics unit, called the FADEC or ECU contains all the necessary circuitry to properly interface with 

engine control devices as well as cockpit command and data communications. The major driving factor 

for centralized engine control architecture is that commercially available electronics cannot withstand the 

high operating temperatures around the engine. Putting all the electronics needed for sensor signal 

processing, actuator loop closures, control law processing, etc., in one box, which is hardened to 
NASA/TM—2013-217821 22

Figure 16.—Centralized engine control architecture: CPU is 

central processing unit (image courtesy of NASA).

withstand the harsh engine operating environment (high pressures and vibrations) and also provides 

means of cooling the electronics, has emerged as the most weight-efficient approach to implement digital 

engine control. For commercial engines, the FADEC is placed on the fan casing and is air cooled; for 

military engines, the FADEC is placed on the casing of the gas generator and is fuel cooled.

The centralized control architecture imposes many constraints which limit the opportunities to 

introduce new control capabilities and also add to the engine life cycle cost. The first issue is that of 

electronics obsolescence management. Typically a FADEC design is expected to operate over a long 

period of time since any changes in the FADEC design require recertification of the whole engine control 

system. A number of the electronic components, especially those associated with control law processing 

and communications are subject to rapid obsolescence, driven by a high turn-over rate for the consumer 

electronics. Adequate inventories of these components have to be maintained to last over the FADEC 

design life time. Second, the sensors and actuators are connected to FADEC through a bundle of wires 

that need to be routed around the rotating machinery, resulting in multiple connectors. The wiring bundle 

not only adds weight, but also the multitude of connectors poses reliability issues. The FADEC design is 

highly individualized for a specific engine, and the lack of modularity prevents the opportunity for 

reducing costs. Finally, the size of the FADEC itself poses a problem for the engine configurations such 

as the geared turbofan (GTF) and open rotor that are emerging to meet the challenges of increased fuel 

efficiency and reduced emissions. The large fan on the GTF poses challenges for accommodating the 

current size FADECs on the fan casing—both from the perspective of keeping the engine frontal area to 

the lowest possible to reduce drag and incurring weight penalties on the fan casing for the structural 

strengthening required for the FADEC. For open rotor configurations, the FADEC will need to be placed 

on the gas generator casing, which is a much harsher environment and will result in weight penalties 

associated with further ruggedizing of the FADEC and the need to provide fuel cooling. 

To keep the engine control architecture from being a limiting factor in the realization of technologies 

that enable highly efficient engines with low life cycle costs, there is extensive interest among the 

aeropropulsion industry to develop technologies that will enable a distributed engine control (DEC) 

architecture. In a DEC system architecture, as shown in Figure 17, any number of control elements are tied 

together through a common, standardized communication interface. Sensors and effectors are replaced by 

control nodes which may provide sensor data, operate actuators, or perform combinations of both. The 

massive wiring harness, which previously tied together the control element to interface circuitry in the 

engine-mounted avionics package, is replaced by a simple but robust digital communication structure.
NASA/TM—2013-217821 23

Figure 17.—Distributed engine control architecture (image courtesy of NASA).

Potential benefits of DEC include reduced control system weight, improved reliability, reduced operating 

cost, reuse for multiple applications, and flexibility to add new capability. Additionally, DEC is critical to 

integrate active component control technology with the overall engine system control. 

The CDB is working in collaboration with AFRL, the aeropropulsion industry, and major providers of 

FADEC technology by way of the Distributed Engine Control Working Group (DECWG) to help identify 

the key challenges for enabling DEC and provide directions for overall research. This partnership is 

essential for leveraging a common interest in developing precompetitive controls technologies, establishing 

standards for continued collaboration, and identifying requirements for electronics. Working in 

collaboration with the DECWG, NASA has developed a roadmap for maturing critical technologies to move 

the engine control architecture from the current centralized to a future fully distributed one (Culley 2010). 

Apart from developing the technologies to enable distributed engine control, it will also be important to 

develop a hardware-in-the-loop facility that can be used to test and investigate integration aspects of various 

components of the distributed control. GRC currently has a contracted effort in place with an engine 

manufacturer, which will develop such a capability and have it available at GRC by late 2014.

Active Combustion Control 

As the aircraft engine industry moves towards lean-burning combustors in order to meet stringent 

low-emissions requirements, maintaining stability for these combustors can become problematic. The 

thermoacoustic combustion instability issues have already been encountered in lean-burning groundbased power generation gas turbine systems. For ground-based systems, these issues have been overcome 

by making ad-hoc design changes or by applying active control. However, dealing with these issues in 

aeroengines is more challenging since aeroengines operate over a wide range of conditions. In early 2003, 

GRC started working in collaboration with PW and United Technologies Research Center (UTRC) to 

develop and demonstrate technologies for the active suppression of thermoacoustic instability. As part of 

this effort, a single-nozzle combustor rig was developed at UTRC which had the capability to duplicate 

the thermoacoustic instability observed in an actual engine test and also to exhibit a larger amplitude, 

lower frequency instability. In order to suppress instabilities, it is necessary to modulate the fuel flow 

entering this representative combustor at roughly 500 Hz (instability frequency, which is a function of 

combustor length). A valve was developed in conjunction with Georgia Institute of Technology, which is 

capable of generating the required high-frequency modulations in fuel flow. An actuator characterization 

rig was built up at GRC to be able to identify the dynamic characteristics of the valve so that resulting 

actuator models can be used for control design development. 
NASA/TM—2013-217821 24

In order to achieve closed-loop suppression of the combustion instability, two alternative control 

methods were developed (Kopasakis et al. 2004; Le et al. 2005). These control methods were formulated 

to deal with the large wideband combustor noise, severe time delay, and randomness in phase associated 

with the combustor thermoacoustic pressure oscillations. Both controllers use the sensed combustion 

pressure as input and fuel modulations as output. Both control methods were initially evaluated against 

reduced-order oscillator models of the combustor pressure in order to verify basic functionality. To 

provide better fidelity validation of controller performance prior to rig testing, both controllers were then 

tested against a sectored one-dimensional (1D) model of the combustor rig. The controllers were 

evaluated on the NASA combustor rig at UTRC and at GRC, and they demonstrated significant reduction 

in instability magnitude for both the high-frequency (~500 Hz) engine-like instability and the lower 

frequency (~300 Hz) large-amplitude instability. This was the first time such instability suppression had 

been demonstrated in an aeroengine-like environment. 

GRC has continued to investigate the capability to suppress thermoacoustic instabilities in advanced 

ultra-low-emissions combustors being designed by NASA and the aerospace industry. Key to the success 

of this effort are simulations that can capture the instability behavior of these advanced combustors. A 

simulation has been developed, which captures the thermoacoustic instability behavior of an advanced, 

low-emissions combustor prototype that was being tested for emissions performance in a rig at GRC. 

During testing, the combustor exhibited thermoacoustic instabilities that are related to increasing fuel 

flow and that potentially prevent full-power operation. Comparison of the advanced, low-emissions 

combustor rig experimental data and the simulation data showed that the simulation captures the 

essentials of the dynamic behavior of the rig (DeLaat and Paxson 2008). Utilizing this simulation, the 

NASA Adaptive Sliding Phasor Average Control (ASPAC) instability control method was updated for the 

low-emissions combustor prototype. Active combustion instability suppression using the ASPAC control 

method has recently been demonstrated experimentally with this combustor prototype in a NASA 

combustion test cell operating at engine pressures, temperatures, and flows (DeLaat et al. 2012). 

Successful instability suppression was shown using a dynamic pressure sensor in the combustor with 

controller feedback driving a high-frequency fuel valve to perturb the combustor fuel flow. It was also 

demonstrated that the controller can prevent the instability from occurring while combustor operation is 

transitioning from a stable, low-power condition to a normally unstable high-power condition, thus 

enabling the high-power condition. Figure 18 shows some of the results discussed in DeLaat et al. (2012). 

Figure 18.—Thermoacoustic instability suppression and prevention with GRC-developed combustion control

(image courtesy of NASA).
NASA/TM—2013-217821 25

The two sets of results show the pressure with and without the controller being on. For the plots on 

the left, the top plot shows the pressure oscillations at a particular set point; the bottom plot with the 

controller on shows that the pressure oscillations are suppressed significantly and the pressure variation is 

within the noise level. For the plots on the right, the top plot shows that pressure oscillations grow in 

magnitude as the engine is throttled up (fuel flow is increased to the combustor); the bottom plot shows 

that with the controller on, the instability can be prevented from occurring as the engine is throttled up. 

These results are very promising in demonstrating the capability of GRC-developed combustion control 

techniques to enable safe operation of ultra-low-emissions combustors throughout the operating envelope. 

Research is currently ongoing to develop fuel actuation systems that can be used with the emerging 

industry and NASA ultra-low-emissions combustor concepts.

High-Speed Propulsion System Dynamic Modeling and Control

The CDB is developing dynamic modeling and control methods for high-speed propulsion systems to 

support the objectives of the Supersonic (SUP) and Hypersonic (HYP) projects of the NASA 

Fundamental Aeronautics Program. For the supersonic propulsion system area, the objective is to develop 

the models and design the controls so that the inlet and engine perform as desired, especially as the inlet is 

the most crucial component of a supersonic engine. The integrated inlet/engine control design should be 

able to suppress upstream flow disturbances such as those due to atmospheric wind gusts, aircraft 

maneuver, as well as flexible mode excitation from the slender body aircraft structure. The propulsion and 

integrated engine and aero-servo-elastic structure should not produce thrust variations that impact ride 

quality and aircraft stability. The CDB approach is to develop high-fidelity propulsion system models 

(1D computational fluid dynamics (CFD) for the inlet and stage-by-stage volume dynamics for the 

engine) that can be used for controls methodology development and application. To date, both a turbojet 

and a turbofan engine have been modeled, control designs and schedules have been developed to operate 

the engine throughout its expected operating envelope, atmospheric turbulence models have been 

developed, and a quasi-1D CFD supersonic inlet model has been developed for internal compression 

inlets. Also, an innovative loop shaping feedback controls design methodology has been developed that 

maximizes system performance based on hardware capability. This controls approach has been applied to 

design engine speed control as well as inlet shock position controls in the presence of atmospheric 

disturbances (Kopasakis and Connolly 2009). Research is ongoing in developing a dynamic model of the 

N+3 propulsion system configuration that has been established as the baseline for the SUP project.

For the HYP project, the CDB effort has focused on developing dynamic models and tools to aid the 

design of control algorithms to manage a turbine-based combined cycle (TBCC) airbreathing propulsion 

system during the critical operating period when the propulsion system transitions from one cycle to 

another, “referred to as mode transition.” The TBCC consists of a dual flow path: a low-speed flow path 

with a turbine engine to accelerate to supersonic speeds and a high-speed flow path with a scramjet to 

accelerate to and maintain hypersonic speeds. During the mode transition, the control challenge is to 

maintain the shock position for the low-speed flow path in such a way as to maintain the required flow for 

an orderly safe shut-down of the turbine engine. Current research is focused on developing the 

computational models to simulate an airbreathing TBCC propulsion system inlet. The TBCC inlet 

aerodynamic design being modeled is that of the Combined-Cycle Engine (CCE) Testbed. The CCE 

testbed is a large-scale hardware model of an aerodynamic design that was verified in a small-scale

screening experiment. The CCE testbed has been designed for experiments in the GRC 10- by 10-Foot 

Supersonic Wind Tunnel (SWT). This mixed compression inlet system is suitable for experiments 

focusing on mode transition studies. The modeling approach includes employing existing state-of-the-art 

simulation codes, developing new dynamic simulations, and performing system identification 

experiments on the CCE testbed in the SWT (Stueber et al. 2010). The system identification experiments 

were recently completed, and the extensive amount of data collected is currently being analyzed. The 

original research plan was to develop inlet control logic using the dynamic models based on the system 
NASA/TM—2013-217821 26

identification data, and then test the control on the CCE testbed in the SWT. However, future testing on 

the CCE testbed has been cancelled due to funding cuts in the HYP project. The current plan is to 

demonstrate the control logic in a simulation of the CCE propulsion system, which will include updated 

inlet model from the experimental data. 

Rocket Engine Control

Because of space limitations, this paper has focused on the aircraft engine control work done by the 

controls group at GRC. This group has also been active in the area of rocket engine dynamic modeling 

and control and has made various significant contributions. These contributions include the theory to 

understand and thus avoid rocket engine chugging. This theory was applied to stabilize the Lunar Lander 

engine. Also hot-firing studies of the Centaur vehicle in the Plum Brook B-2 altitude test chamber were 

used to determine the feed-line dynamics to assure POGO stability of that vehicle. POGO instability 

refers to a coupling between the fuel flow feed-lines and the thrust generated by the rocket which results 

in a pogo like behavior for the rocket due to thrust oscillations. Further studies, in the 1970s, supplied 

advanced technology for possible active POGO suppression on the Space Shuttle Main Engine. The 

concept of life-extending control was initially developed for rocket engines (Lorenzo et al. 1992) and then 

applied to aircraft engines as discussed in an earlier section. In the late 1980s to mid-1990s extensive 

work was done on applying emerging health management technologies to the Space Shuttle Main Engine 

(Duyar et al. 1992). The CDB is continuing the technology development effort in support of the current 

NASA Human Exploration program through development of fault management technologies as described 

in Melcher (2011). 

Conclusion

In conclusion, this paper provided a broad overview of the state-of-the-art of engine control, some 

significant past contributions of the controls group at the NASA Glenn Research Center (GRC) that have 

resulted in increased efficiency and performance of aircraft engines, and some recent accomplishments 

where elements of the technologies developed can be expected to find their way into future aircraft 

engines. The Controls and Dynamics Branch (CDB) continues to perform cutting-edge engine control and 

diagnostics research to meet the challenging goals of the NASA Fundamental Aeronautics Program and 

the Aviation Safety Program. From a control logic perspective, the CDB is leading research in modelbased engine control and diagnostics, which will enable highly efficient and reliable operation of future 

engines. From a control architecture perspective, the CDB is leading research to enable transition from 

the current centralized control architecture to a future distributed engine control architecture, which will 

provide increased modularity, capability to incorporate new control component technologies as these 

mature, and reduced life cycle cost. The CDB is also leading research in developing active component 

control technologies, such as active combustion control, which are critical to realizing the benefits of 

high-efficiency components with safe operation throughout the flight envelope. 

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Aircraft Turbine Engine Control Research at NASA Glenn Research Center 

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14. ABSTRACT

This paper provides an overview of the aircraft turbine engine control research at the NASA Glenn Research Center (GRC). A brief 

introduction to the engine control problem is first provided with a description of the state-of-the-art control law structure. A historical aspect 

of engine control development since the 1940s is then provided with a special emphasis on the contributions of GRC. With the increased 

emphasis on aircraft safety, enhanced performance, and affordability, as well as the need to reduce the environmental impact of aircraft, 

there are many new challenges being faced by the designers of aircraft propulsion systems. The Controls and Dynamics Branch (CDB) at 

GRC is leading and participating in various projects to develop advanced propulsion controls and diagnostics technologies that will help 

meet the challenging goals of NASA Aeronautics Research Mission programs. The rest of the paper provides an overview of the various 

CDB technology development activities in aircraft engine control and diagnostics, both current and some accomplished in the recent past. 

The motivation for each of the research efforts, the research approach, technical challenges, and the key progress to date are summarized. 

15. SUBJECT TERMS

Gas turbine engines; Engine control; Engine performance; Robust control; Fault detection 

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