Link C. Jaw

Scientific Monitoring, Inc., Scottsdale, Arizona

Sanjay Garg

Glenn Research Center, Cleveland, Ohio

Propulsion Control Technology Development

in the United States

A Historical Perspective

NASA/TM—2005-213978

October 2005
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Link C. Jaw

Scientific Monitoring, Inc., Scottsdale, Arizona

Sanjay Garg

Glenn Research Center, Cleveland, Ohio

Propulsion Control Technology Development

in the United States

A Historical Perspective

NASA/TM—2005-213978

October 2005

National Aeronautics and

Space Administration

Glenn Research Center
Acknowledgments

We would like to express our gratitude to many people who have helped to gather information and review the

draft: Former chiefs, Controls and Dynamics Branch, NASA Glenn: Walter Merrell, Carl Lorenzo, and Fred Teren;

Ten-Hui Guo, Khary Parker, Lynn Anderson, and Bonita Smith, NASA Glenn; Bill Burcham, NASA Dryden;

Charles Skira and Al Behbahani, Air Force Research Laboratory, Wright-Patterson Air Force Base; Harold Brown

and Kiyoung Chung, General Electric Aircraft Engines; Al Levesque, Ramesh Rajagopalan, Matthew Schryver,

Dale Southwick, Alan Volponi, and Bruce Wood, Pratt & Whitney; Glen Schwent, Robert McCarty, and

Larry Yee, Honeywell Engines; and Hatti Wong, Scientific Monitoring, Inc.

This paper is dedicated to all those who have contributed to the advancement of air-breathing propulsion control

capabilities in the world.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http://gltrs.grc.nasa.gov
 

NASA/TM—2005-213978 1 

Propulsion Control Technology Development in the United States 

A Historical Perspective 

Link C. Jaw 

Scientific Monitoring, Inc. 

Scottsdale, Arizona 85258 

Sanjay Garg 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This paper presents a historical perspective of the advancement of control technologies for aircraft 

gas turbine engines. The paper primarily covers technology advances in the United States in the last 

60 years (1940 to 2002). The paper emphasizes the pioneering technologies that have been tested or 

implemented during this period, assimilating knowledge and experience from industry experts, including 

personal interviews with both current and retired experts. 

Since the first United States-built aircraft gas turbine engine was flown in 1942, engine control 

technology has evolved from a simple hydro-mechanical fuel metering valve to a full-authority digital 

electronic control system (FADEC) that is common to all modern aircraft propulsion systems. At the 

same time, control systems have provided engine diagnostic functions. Engine diagnostic capabilities 

have also evolved from pilot observation of engine gauges to the automated on-board diagnostic system 

that uses mathematical models to assess engine health and assist in post-flight troubleshooting and 

maintenance. Using system complexity and capability as a measure, we can break the historical 

development of control systems down to four phases: (1) the start-up phase (1942 to 1949), (2) the growth 

phase (1950 to 1969), (3) the electronic phase (1970 to 1989), and (4) the integration phase (1990 to 

2002). In each phase, the state-of-the-art control technology is described and the engines that have 

become historical landmarks, from the control and diagnostic standpoint, are identified. Finally, a 

historical perspective of engine controls in the last 60 years is presented—in terms of control system 

complexity, number of sensors, number of lines of software (or embedded code), and other factors. 

1.0 Introduction 

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 have been contributed by advances in 

engines or propulsion systems. During this period, propulsion systems have evolved from piston engines 

(the first generation, 1903 to 1945) (Ref. 1) to jet engines (1942 to present) (Refs. 2 and 3). This paper 

provides a historical perspective of the development of jet engine control technologies; specifically, it 

focuses on the technical advancements of turbofan and turbojet engines in the United States during the 

past 60 years. This historical period is further divided into four controls developmental phases as follows: 
 

NASA/TM—2005-213978 2 

(1) Start-up phase, 1942 to 1949 

(2) Growth phase, 1950 to 1969 

(3) Electronic phase, 1970 to 1989 

(4) Integration phase, 1990 to 2002 

In each phase, a brief description of the key technologies is first presented, then the engines that have 

made historical contributions are tabulated. Section 6 presents a global view of the evolution of the 

controls technologies in the last 60 years. Before we look back in history, a brief review of the gas turbine 

engine control principle may be helpful. 

A simple engine control system computes 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); then it 

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. Figure 1.1 

illustrates this fuel control system and identifies the various components needed to make the system work. 

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

fuel flow is called a controlled variable. 

If we change “power” in Figure 1.1 to “turbine temperature,” then the process represents a 

temperature control system, or a temperature governor. Likewise, if we change both “power” and “fuel 

flow” to inlet guide vane angle (IGV), then the diagram becomes an IGV control system. 

Figure 1.1.—Functional process diagram of a simple engine control system (from Ref. 2). 

Compute 

desired 

fuel flow 

Pilot’s 

power 

request 

Power 

desired? 

Deliver the 

desired 

fuel flow 

Pump fuel 

flow from 

fuel tank 

Inject fuel 

flow into 

combustor 

Measure 

produced 

power 

Determine 

operating 

condition 

Throttle 

Control 

Fuel Pump 

Metering 

Valve 

Fuel nozzle 

Sensor 

Control Logic 

No Yes 
 

NASA/TM—2005-213978 3 

In-flight engine thrust measurements are not possible. Good indicators of thrust can, however, be 

obtained from either engine shaft rotational speed (N) or engine pressure ratio (EPR) and these 

parameters have been used effectively in the engine control process. 

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), the air speed can go 

beyond Mach 3; furthermore, the air temperature at the same altitude and airspeed can vary from a hot 

summer day to a cold winter night. These ambient condition variations have imposed severe challenges 

on control system design. To a control engineer, these challenges are represented on a fuel flow (Wf) 

versus engine shaft speed (N) graph, or a fuel ratio unit (Wf/P3, where P3 is the compressor exit static 

pressure) versus speed graph, as shown in Figure 1.2. 

After engine start, the engine operating envelope (shaded in light green) is bounded by the maximum 

flow limit, the minimum flow limit, the idle power governor and the maximum power governor. The 

‘max. flow limit’ prevents the engine from surge and over-temperature. The ‘min. flow limit’ prevents the 

engine from flame-out. The idle power is typically set to produce a desirable ground thrust for taxiing the 

airplane, and the max. power produces the engine maximum-rated thrust. The max. and min. flow limits 

are also called the acceleration and deceleration schedules. These schedules change with the engine 

ambient condition. 

The figure also shows four straight lines intersecting the idle power point and the max. power point 

on the steady state operating curve. These lines are called droop lines, which represent the proportional 

control gains (∆Wf/∆N) at the two power points, respectively. In reality, the trajectories between a power 

set-point (e.g., idle or maximum) and a fuel control limit (either the max. limit or the min. limit) is not a 

straight line, because all engine fuel controls use at least one additional control law, the integral control, 

to improve control system robustness. Many modern engines use derivative control and complex control 

compensations to shape the transient performance. 

Figure 1.2.—Gas turbine engine fuel control system operating envelope. 

Engine shaft speed 

Fuel flow 

rate (Wf) 

or fuel 

ratio unit 

(Wf/P3) 

Required fuel flow 

@ steady state 

Max. flow limit

Min. flow limit

Idle 

power 

Max. 

power 

Proportional 

control gain or 

droop slope 

Droop 

slope 

Safe operating 

region 
 

NASA/TM—2005-213978 4 

 Additional limit-protection capabilities are provided in engine controls, such as the temperature 

limiting governor, the over-speed governor, and the stator vane scheduler. These additional limit 

protections together with the basic control schedules are designed to support unrestricted throttle 

movement by the pilot. Consequently, the time history of the fuel flow controlled variable of a jet engine 

is characteristically nonlinear. 

2.0 The Start-up Phase (Before 1950) 

2.1 Control System Description 

The first U.S. jet engine was built in 1942 by General Electric (GE). This engine, GE I-A, was 

designed after the British Whittle turbojet engine. The control for the GE I-A was a hydro-mechanical 

governor, which metered the fuel flow going into the engine to be proportional to the difference between 

the set speed and the actual speed of the turbine. To prevent the engine from flame-out, a minimum-flow 

stop was added to the fuel metering valve. To prevent the engine from over-temperature, a maximumflow schedule was also incorporated. This system possessed the basic functionality of controlling a 

single-spool turbojet engine. 

In 1948, GE tested the first afterburning turbojet engine in the world, the J47. The engine 

incorporated a hydro-mechanical fuel control for the main combustion chamber, and an electronic 

(vacuum tube) fuel control for the afterburner; unfortunately, the reliability of the electronic control was 

poor due the problem-prone vacuum tube technology in a harsh operating environment like a jet engine. 

2.2 Control Design Methodology 

The original control law for the J47 was designed using only frequency response techniques. The 

engine underwent altitude testing at NACA (National Advisory Committee for Aeronautics) Lewis 

Laboratory (now NASA Glenn Research Center). 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 behave like a limit 

cycle. To solve this problem, GE and NACA engineers worked together and applied the time-domain step 

response analysis method. The problem was soon fixed by reducing the control gain at altitude (Ref. 4). 

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

knowledge base for engine controls. 

2.3 Modeling Capability 

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. Since each engine operating point could take 

several hours by slide rule or desk calculator, corrected parameter techniques (corrected flows and speeds, 

temperature ratios, and pressure ratios) were used to reduce the number of operating points which needed 

to be evaluated to define engine operation over the engine's operating envelope. The corrected parameter 

approach was derived from dynamic similarity techniques based on the Buckingham π-theorem. This 

corrected parameter approach is still in use today as it simplifies the process of control law development 

and implementation. 

Dynamic behavior of single-shaft turbojet engine was studied at NACA Lewis Laboratory 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 (Ref. 5). 
 

NASA/TM—2005-213978 5 

3.0 The Growth Phase (1950 to 1969) 

3.1 Control System Description 

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

and achieved supersonic speed on a YF-100. The control system of the J57 consisted of a hydromechanical fuel control (HMFC or HMC) for dry-engine combustion chambers, a HMFC for the 

afterburner, and separate anti-ice and ignition controls. 

As engine capabilities advanced in the 1950s, engine control technologies also accelerated to deliver 

the new capabilities. By 1969, a number of well-known and long-service engines had been tested, such as 

GE’s J79 and F101, and PW’s TF30 and F100. As the engine technology had matured to high 

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

variable geometry controls, e.g., the compressor stator control, intake and nozzle controls. This period 

exemplified several successful transfers of military engine technologies into commercial engines (Refs. 6 

and 7). The most notable examples are the PW JT3 (originated from the J57) for the Boeing B-707 and 

the Douglas DC-8, PW JT8D (from the J52) for the B-727 and the DC-9, PW JT9D (from the TF30) for 

the B-747, GE CF6 (from the TF-39) for the DC-10, and CFM56 (from the F101) for the B-737-300. 

3.2 Control Design Methodology 

Control designs during this phase were predominantly the frequency response method through which 

the gain and the phase margins were analyzed as the leading design parameters for closed-loop system 

stability. The step response method was also used to assure stability and performance in the entire 

operating envelope. Successive loop-closure was the technique used invariably in the industry to decouple 

the effects among multiple controlled variables. 

 The early 1960s marked the birth of the modern control era. Linear state-space methods provided a 

new paradigm for modeling, control, and estimation (Ref. 8). The multivariable control design method 

became an intellectual pursuit, although the traditional single-input, single-output (SISO) approach and 

the frequency response design method were used for all production engines. Although theoretically 

matured in the 60s, the modern control design paradigm did not find its way into the turbine engine 

industry until mid-1970s, and the first implementation of a multivariable-control-designed engine did not 

happen until almost 30 years later. 

3.3 Modeling and Simulation Capability 

The first computerized, component-level, steady-state engine performance model was developed in 

1953 for constant specific heat gas properties. It was extended to actual variable gas properties in 1955, 

and from turbojet to turbofan engines in the late 1950s. 

Initial development of simulations of the dynamic performance of gas turbine engines was achieved 

through the use of electronic analog computers at NACA Lewis 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 by 

experimental results (Ref. 9). 

A more efficient numerical solution, commonly called the balancing or matching technique, or more 

theoretically, the Newton-Raphson method, for component-level, steady-state engine performance models 

were developed in the mid-1960s by the Air Force Aero Propulsion Laboratory (AFAPL). This new 

solution technique significantly reduced the time to compute a design point from weeks to days (Ref. 10). 

By the mid-60s, 90 percent of the work involved in dynamic (or transient) performance analysis, 

which was essential for control system design, was already handled by either analog or digital computer 

analysis, while the remaining 10 percent was handled by hand analysis. Almost every known jet engine 

transient was analyzed through simulations. A dynamic simulation typically consisted of engine models 
 

NASA/TM—2005-213978 6 

and control system models connected in closed-loop. The analysis typically included frequency response, 

transient response, large accelerations and decelerations, starting cycles, afterburner light-off and shut-off, 

compressor stalls, and combustor blow-out (Ref. 11). 

4.0 The Electronic Phase (1970 to 1989) 

4.1 Control System Description 

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 hydro-mechanical fuel control and servo components, although 

having been proven as highly reliable, were quickly reaching a practical limit, i.e., they became 

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

In the early 1970s, analog and digital electronic control units (ECU) were designed to provide highlevel supervisory or trim functions. These supervisory control units primarily calculated speed or 

temperature set-points throughout the engine operating envelop. For some non-critical controlled 

variables, they would have sole control responsibilities. Early production supervisory control units 

included: the digital Electronic Engine Control (EEC) unit for the PW F100 and the analog Augmentor 

Fan Temperature (AFT) control for the GE F101. The F100 EEC is the first supervisory control for 

aircraft engines. 

An ECU is considered full-authority if it controls the entire operation of an engine from start-up to 

shutdown according to pilot’s throttle command (or power request). The first full-authority ECU was 

introduced into service in 1972 on the Garrett/Air Research (now a division of Honeywell) TFE731 

engine. The TFE731 ECU was a single-channel analog control. In case of a failure or pilot over-ride, the 

electronic control switched to a back-up HMC with conservative, fixed fuel-ratio max. and min. 

schedules. The electronic control provided the following key functions (Ref. 12): 

 

(1) Minimal pilot attention and fail-safe operation. 

(2) Unrestricted power lever movement, i.e., the electronic control computes all control schedules, 

power setting, and engine operating limits. 

(3) No throttle dead-band. 

(4) Maximum power lever always gave the maximum allowable thrust. 

(5) Automatic flat rating for thrust. 

(6) Rapid surge-free acceleration through a modulated on-off bleed valve. 

 

Improving the mean time between failure (MTBF), especially for a given mean time between 

unscheduled removal (MTBUR), and the ease of field maintenance were the prime motivations behind the 

initial transition from analog to digital controls. Other advantages were quickly discovered and set the 

trend for digital electronic controls such as the flexibility of the digital software for complex computation 

and software revision, the capability of fault isolation to LRU (line replaceable unit) level through selftesting, and simplified and more accurate control adjustments. 

The F100 DEEC (Digital Electronic Engine Control) was the first full-authority digital electronic 

control (FADEC) that was flight-tested. The DEEC evolved from the supervisory EEC of the F100 engine 

HMC and from a digital electronic breadboard for automating the design of the cam schedule used in the 

HMC. This electronic breadboard significantly reduced the development time from months to weeks, 

demonstrating the capability of electronic controls to improve the overall engine development cycle time. 
 

NASA/TM—2005-213978 7 

During the past 30 years, the ECU has proven the following undisputable benefits: 

(1) Significant size and weight reduction for the same control functionality. 

(2) Unrestricted power lever movement, especially in military applications, thus reducing pilot 

workload. 

(3) Elimination of engine trim after any control component change. 

(4) Easy software update (through a laptop computer). 

(5) Compatibility with flight control, fire control, and other on-board and monitoring systems. 

(6) Power to record engine data. Better maintainability due to increased diagnostic capabilities from 

electronics (fault flags and fault isolation). 

4.2 Control Design Methodology 

The trend in this phase was the multivariable control design method. Although a lot of interest was 

shown in advanced control design technologies; however, all production engine controls were still 

designed by SISO and successive loop-closure techniques. Multivariable control design methods had 

gradually matured during this phase since the early 1970s. Both frequency-domain and time-domain 

methods were studied extensively (Ref. 13). Frequency-domain methods were led by the Inverse Nyquist 

Array (INA) method (Ref. 14). Time-domain methods were led by the Linear Quadratic Regulator (LQR) 

method. 

In the area of time-domain control design, the United Aircraft Research Laboratories (now the United 

Technologies Research Center) appeared to be the first to apply a modern control design method, or more 

specifically, the LQR design, to an aircraft engine (Ref. 15) in early 1970s. The U.S. Air Force Wright 

Aeronautical Laboratory (AFWAL, now the Air Force Research Laboratory) started to apply 

multivariable control design methods to a J85 engines in 1973 (Ref. 16). AFWAL and NASA Lewis then 

jointly sponsored the Multivariable Control Synthesis (MVCS) program from 1975 to 1978. The program 

developed a series of LQRs and connected them with simple transition logic and several operating limits 

to achieve effective large-transient controls. The control design (Ref. 17) was verified extensively at 

NASA’s hybrid simulation facility (Ref. 18), and followed by a successful testing on an F100 engine in 

NASA’s altitude test facilities (Ref. 19). 

The multivariable control methodology lends itself to complex, integrated control problems. One of 

these problems is the integrated flight and propulsion control (IFPC). IFPC relieves pilot workload and 

improves aircraft maneuverability. Improving maneuverability can be achieved by thrust vectoring and jet 

diversion. In the mid-1980s, AFAPL sponsored the DMICS (design methods for integrated control 

systems) program to study different design methods. The natural approach was to combine the flight 

control and the propulsion control into one integrated system and to design an integrated control law for 

the combined system. This approach had some practical concerns: (1) airframe and engine manufacturers 

traditionally design their control systems separately before system integration, and (2) early and close 

coordination and cooperation between flight and propulsion system designers could drive the design cycle 

longer (hence cost). A less intuitive approach was to treat the IFPC control design problem as a 

hierarchical, de-centralized system and integrate the functions of the flight and propulsion subsystems 

through cross-coupling effects and flight command translation (Ref. 20). 

4.3 Modeling and Simulation 

By early 1970s, steady-state and transient performance simulations (commonly called the steady-state 

cycle deck and the transient cycle deck, respectively) of jet engines were routinely used in engine and 

control system designs. Furthermore, the combined cycle deck proved effective in improving engine test 

quality by ensuring engine and test cell compatibility, reducing test requirements, and extending the 

knowledge of the engine performance (Ref. 21). 
 

NASA/TM—2005-213978 8 

In the early 1980s, NASA Lewis initiated an effort to build an engine real-time simulator. The 

simulator consisted of both hardware and software. The hardware was designed with parallel, multiprocessors. The software included a high fidelity, aerodynamic model of a turbofan engine. At the same 

time, an integrated real-time simulation of V/STOL (vertical/short takeoff and landing) aircraft was 

developed, combining both airplane and engine simulations models and setting the foundation for 

integrated flight and propulsion control studies (Ref. 22). 

4.4 Failure Detection, Isolation, and Accommodation (FDIA) 

In 1970, the F100 was the first engine equipped with an on-board 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 DEEC 

continuously performs integrity test on itself and system signals (i.e., built-in test, BIT). These tests 

enable the DEEC to detect and accommodate system malfunctions while maintaining normal engine 

operation. The DEEC transmits diagnostic information and selected control parameters to the EDU to 

identify control system faults and facilitate data recording. The EDU detects engine events, records data, 

and handles engine-airframe communication. The EDU also receives data through the engine-airframe 

interface and engine sensors. This information permits the EDU to record engine data and perform fault 

detection independent of the DEEC (Ref. 23). 

The initial success in applying multivariable control design methods to jet engines in the mid 1970s 

prompted the AFWAL to explore failure detection and accommodation techniques (Ref. 24). In early 

1980s, motivated by improving the reliability of FADECs, NASA Lewis explored the analytical 

redundancy approach to sensor failure detection, because the engine sensor is one of the least reliable 

control components. NASA later conducted a real-time simulation of an advanced detection, isolation, 

and accommodation algorithm for sensor failures using analytical redundancy. The simulation 

demonstrated the feasibility of real-time detection and accommodation of both large-magnitude (or hard) 

and small-magnitude (or soft) failures (Ref. 25). 

4.5 Flight Research and Advanced Control 

Propulsion flight research with the purposes of investigating engine control-related problems and 

testing new control concepts began in late 1960s as a joint effort between NASA and U.S. Air Force. 

Since that time, propulsion flight research has been conducted at Edwards Air Force Base and the NASA 

Dryden Flight Research Center in Edwards, California (NASA Dryden), with some level of support from 

NASA Glenn. The U.S. Air Force asked NASA in 1973 to assist in developing and flight-testing a digital 

IPCS (Integrated Propulsion Control System). The IPCS, installed on an F-111E airplane, was the first 

flight test of an integrated engine control system (Ref. 26). The afterburner ECU of the PW TF-30 engine 

was integrated with the control of the variable-geometry external compression inlet. In late 1970s, NASA 

Dryden continued its pioneering research in integrated control with the flight test of a cooperative control 

system, where all functions of inlet control, autopilot, auto-throttle, and navigation systems were 

combined in a single digital computer on a YF-12 airplane. 

In 1981, flight test of the F100 DEEC began at NASA Dryden on the NASA F-15 aircraft. The 

success of the flight test program allowed the Air Force to elect to put the DEEC into full-scale 

development and production. The success also stimulated a series of flight research projects in the midand late 1980s to study the advanced control modes. These projects, centered on the HIDEC (Highly 

Integrated Digital Electronic Control modes) program, included: integrated flight and propulsion control, 

adaptive engine control, extended engine life control, and inlet integration. 
 

NASA/TM—2005-213978 9 

In the advanced control research area, the study of the characteristics of compressor surge and 

rotating stall during this phase enabled the blossoming work in active compressor stability controls of the 

1990s. In the late 1970s, theoretical and experimental works on surge and rotating stall were summarized 

for compressors (Ref. 27). These works prompted NASA to conduct research in surge and rotating stall. 

A hybrid computer simulation was developed at NASA Lewis to extract the in-stall pumping 

characteristics of a multi-stage axial flow compressor from engine surge data (Ref. 28). 

The advances in multivariable control design and FADEC had stimulated the concept of model-based 

controls. Since the engine model typically represents the nominal performance of the engine fleet, on-line 

estimation of model parameters allows the model to be adjusted for each specific engine (i.e., off-nominal 

performance), even with engine deterioration. This on-line tuning (or adaptive) model is then able to 

improve both the controlled engine performance and the diagnostic accuracy. In late 1980s, the Kalman 

filtering approach to adaptive modeling was first applied to jet engines. A piecewise linear, state-variable 

model of a turbofan engine was adjusted by five estimated parameters, representing five implicit engine 

component variations (e.g., efficiency and flow variations), using a Kalman filter with in-flight engine 

measurements (Ref. 29). This work paved the way for the development and flight test of the performanceseeking control (PSC) in the early 1990s (Ref. 30). 

5.0 The Integration Phase (1990 to 2002) 

 5.1 Control System Description 

Dual-channel FADECs became the standard control system for jet engines in this phase. Compared to 

the “first generation” FADECs in the previous phase, they have more built-in test functions; they have 

some form of embedded engine models to improve engine performance and diagnostics; they have some 

type of usage and life tracking algorithms; and they are much lighter and smaller per controlled variable 

due to spectacular improvements in microprocessor speeds and memory technologies. 

Electric actuation systems gained wider acceptance. In addition to stepper motors, electro-mechanical 

actuators have been installed on engines. New sensors, especially diagnostic-specific sensors, are being 

tested for enhanced engine health management capabilities. 

5.2 Control Design Methodology 

Subsystem integration for improved overall performance was a driving force for engine control 

design. Integration of engine control with inlet and exhaust controls, as well as the integration of 

propulsion and flight controls, made significant advancement in this phase. The issues facing the design 

of integrated flight and propulsion control systems were analyzed in a NASA-sponsored STOVL (short 

take-off and vertical landing) controls technology program. A modified F110 engine with ejectoraugmentor was assumed as the propulsion system to provide thrust during cruise, and provide lift and 

reaction (jet) controls during hover and landing (Ref. 31). This work helped the development and piloted 

simulation of an advanced IFPC design methodology at NASA. The methodology included the H∞ control 

synthesis technique, controller partitioning between the airplane and the engine, controller scheduling 

over a wide operating regime, and nonlinear design techniques such as rate limiting and integrator windup protection (Ref. 32). 

Much of the modern control design methods described for the Electronic Phase and the Integration 

Phase was theoretical and experimental, i.e., most of the advanced designs were only validated in 

simulations or laboratory environment, they were not implemented on real engines. However, the 

following two examples offer some encouragement for more practical adoption of advanced design 

methods in the future. 

System identification techniques were applied to engine control development in early 1990s. Several 

operating point models of the Honeywell TFE731-5 engine were identified from ground test and in-flight 
 

NASA/TM—2005-213978 10 

data. These input-output models were then used as design models to derive the control gain values for the 

control. After the control laws were implemented in the FADEC, the engine was tested in a test cell. It 

was found that all the control gains were adequately set. This was the first time that a designed gain value 

did not have to be adjusted (usually several-fold) in the test cell during control development testing. The 

engine was then flown with the new control law to 45,000 feet. The altitude-compensated control laws 

performed exceedingly well, again without the need for adjustment. 

After almost 30 years of theoretical development, the multivariable control design methodology was 

applied to the design of the PW JSF F119 (now designated F135) control law, which was implemented in 

the engine’s FADEC and tested in flight in 2000. 

5.3 Modeling and Simulation 

Block diagrams have been commonly used to represent control logics and diagnostic procedures. In 

this phase, GUI-based modeling and simulation tools reached a high level of maturity, and were widely 

adopted by control engineers to build and analyze engine models. Among all commercial-off-the shelf 

(COTS) design tools, the MatrixX/SystemBuild and the Matlab/Simulink were most widely adopted. 

Automatic software (or code) generation tools became a new standard for real-time simulation and 

embedded applications. Aside from the auto-code generation tool offered by SystemBuild and Simulink, 

proprietary tools such as PW’s Pictures-to-Code and GE’s BEACON were both proven to reduce control 

and diagnostic software development time significantly. However, the overwhelming cost of maintaining 

these in-house tools has limited the use of proprietary code-generation tools. 

Control system rapid prototyping capabilities were significantly enhanced. Vertically integrated 

process starting at the top from model-building, control analysis and simulation, automatic generation of 

control code, code compilation, downloading the code to real-time control hardware (typically a single 

board computer), and running the control system with the plant, actuator, and sensor hardware in closedloop has been a standard procedure in a design process to reduce development cost and risk. 

5.4 Advanced Control and Flight Research 

Flight test of the performance-seeking control (PSC) on the NASA F-15 aircraft started in the late 

1980s and finished in the mid-1990s. The PSC took the HIDEC to the next step—onboard real-time 

adaptive optimization of engine and airplane parameters. A Kalman filter was used to estimate key engine 

parameters in near-steady state conditions. This program provided valuable information for the selftuning, onboard engine model that was implemented in the ECU of the F119 engine on the F-22 airplane. 

A distortion-tolerant control system was developed and flight-tested in 1997 on the High Stability 

Engine Control (HISTEC) program (Ref. 33). The control system included a distortion estimation system, 

which assimilated information from engine face pressure sensors and flight control estimated angle-ofattack and side-slip angle. The estimated distortion level was then used to trim engine actuator commands 

for stability management. 

The propulsion-controlled aircraft (PCA) research represented another contribution in this phase. The 

PCA used computer-controlled engine thrust to provide a safe landing capability in case of major flight 

control system failure. Later tests verified PCA operation over the full flight envelope, even in upset 

conditions with all hydraulic systems turned off, and automatic landing with the ILS (instrument landing 

system). 

The importance of flight test in the validation and verification of control technologies has now been 

extended to health management technologies. NASA Dryden, with the support of NASA Glenn and 

NASA Ames, is leading the vehicle health management program to test a number of engine and vehicle 

health management technologies on an Air Force C-17 airplane with its four PW F117 engines. This 

program promises to advance the health management capabilities significantly as the flight tests did to 

engine controls in the past. 
 

NASA/TM—2005-213978 11 

Life-extending controls (LEC) were investigated for both large and small engines. NASA pioneered 

the concept of LEC and demonstrated, in simulation, the objectives of achieving high performance in 

transient responses and damage mitigation with increased structural durability for rocket engines 

(Ref. 34). The LEC concept was applied to aircraft turbine engines in mid 1990s. GE, Honeywell, and 

PW were investigating different methods to reduce damage and increase engine life. One method showed, 

in a production-quality FADEC-in-the-loop simulation, that thermal-mechanical fatigue and creep/rupture 

life can be extended significantly by small increases in acceleration transient time and cruise flight time, 

respectively. The small increase in acceleration time would not violate FAA’s (Federal Aviation 

Administration) engine acceleration time requirement, because the time increase, resulting from reduced 

fuel flow and turbine temperature, occurred in the final 10 percent of the acceleration (Ref. 35). 

Active controls represented new opportunities to improve engine performance, operability, and 

reliability. Research programs in high-response controls (e.g., hundreds of cycles per second) of surge and 

stall, tip clearance, and combustion instabilities were sponsored by Air Force, Army, Navy, and NASA. 

Although the technology readiness levels (TRL) of active control technologies have not reached the 

implementation stage, they could significantly increase the number of controlled variables and complexity 

due to their nature of distributed control system. 

6.0 Evolution of Controls Complexity 

6.1 Number of Controlled Variables 

One way to assess the complexity of an engine control system is by the number of controlled 

variables (Ref. 35). These controlled variables, or actuators, typically include: fuel flow, stator/vane 

position, valve position, nozzle area, etc. In addition to these multi-valued control variables, there are also 

discrete, or on/off, control variables, such as engine ignition and anti-ice air. Table 6.1 shows the 

evolution of engine control variables of the historically important engines (from the control’s standpoint). 

Note that the anti-ice and ignition control variables are not included in the table. Also note that the trend 

of the number of controlled variables for military fighter engine has outpaced the trend for commercial 

engines. 

TABLE 6.1.—EVOLUTION OF THE NUMBER 

OF ENGINE CONTROL VARIABLES 

Engine model Year (flt) tested Number of control variables 

GE I-A 1942 1 

GE J47 1948 3 

GE J79 1954 4 

GE TF39 1966 2 

GE CF6 1968 3 

PW F100 1970 8 

AR TFE731 1972 2 

GE F101 1972 5 

CFM56 1974 2 

GE YF120 1989 11 

PW F119 1990 8 

GE90 1993 3 
 

NASA/TM—2005-213978 12 

Figure 6.1.—Trends of engine control complexity. 

Figure 6.1 graphically depicts the trends for military and commercial engines of control complexity. 

6.2 Number of Engine Sensors 

In a feedback control system, measurement is made to infer how the physical system under control is 

doing. Hence, the number of measurements, or sensed variables, is another indication of the complexity 

of a control system. Table 6.2 shows that the number of engine sensors, not including the power lever 

position or throttle, has increased with the increasing complexity of the engine controls. The number of 

sensors is expected to increase rapidly as additional sensors are added in future engines for health 

management purposes. Figure 6.2 shows the trends in the number of engine sensors. 

TABLE 6.2.—EVOLUTION OF THE NUMBER OF ENGINE SENSORS 

Engine model Year (flt) tested Number of engine sensors 

GE I-A 1942 3 

GE J47 1948 5 

GE J79 1954 5 

GE TF39 1966 6 

GE CF6 1968 6 

PW F100 1970 8 

AR TFE731 1972 5 

GE F101 1972 7 

CFM56 1974 6 

GE YF120 1989 11 

PW F119 1990 10 

GE90 1993 7 

Military Commercial
 

NASA/TM—2005-213978 13 

Figure 6.2.—Trends in the number of engine sensors. 

Start-up Phase Growth Phase Electronic Phase Integration Phase 

(Before 1950) (1950 to 1969) (1970 to 1989) (1990 to 2002) 

 

Figure 6.3.—Evolution of engine controls. 

6.3 Types of Control Systems 

In a control system, the computational unit and algorithms, which “figure out” what values the 

corresponding controlled variables should be at any operating condition, are called controls. Engine 

controls have evolved from a simple hydro-mechanical fuel control (HMFC) in 1940s to the full-authority 

digital electronic control used in all modern gas turbine engines. Figure 6.3 shows the evolution of engine 

controls of the historically important engines. 

Backup HMC 

UFC

Elec. AB 

control 

Electronic 

supervisory 

control Full-authority 

elec. ctrl. 

HMFC 

Dual 

channel 

FADEC 

Fault code 

recording 

On-board 

engine 

diagnostic 

Integrated 

prop. ctrl. 

Military Commercial 
 

NASA/TM—2005-213978 14 

7.0 Conclusions 

This paper has presented a historical perspective on the development of jet engine control 

technologies in the U.S. aircraft propulsion industry. The paper focuses on the control technologies 

developed from 1940 to 2002. The paper also focuses on the control systems of turbojet and turbofan 

engines, with the belief that these control systems provide a broad representation of the historical 

development of engine controls in general. 

In preparing this paper, we have been challenged with selecting the adequate amount of information 

from a vast collection of technical papers, books, personal notes, and expert-interview transcripts. We 

believe that choosing a high-level overview of the technologies that have been implemented on engines, 

tested in flight (or tested on ground full-scale models), and identifying the technologies that have made 

significant shifts in paradigms, like the multivariable control, are a good compromise. We may have 

omitted some technologies that readers think could be included in this paper; we encourage the reader to 

contact us so that we can continue to update this legacy. 

In the process of designing engine control systems, we have taken for granted many common 

practices like altitude corrections, gain compensation, and limit protections. As we look back at the 

historical development of these common practices, we begin to recognize their roles in advancing the 

state-of-the-art of gas turbine engines. We have also gained a deeper appreciation of their contributions to 

the increased performance, reliability, and safety of jet engines over the last 60 years. We hope this paper 

can remind us of the dedicated engineers who persevered through difficult challenges. We also hope that 

the paper can encourage us to continue the honorable tradition of engine controls. 
 

NASA/TM—2005-213978 15 

8.0 Appendix—Historically Important Engines 

8.1 Start-up Phase 

The table is organized by the engine model, the year when the engine was first flight-tested (or first 

run on the ground when the flight test date was not confirmed), the aircraft model, and the technology 

advancements made by the engine. 

TABLE 8.1.—MAJOR MILESTONES FOR PROPULSION CONTROLS IN THE START-UP PHASE (BEFORE 1950) 

Year Engine model Aircraft model Technology advancements 

1942 GE I-A Bell XP-59A (1st flt) Design based on British Whittle engine (single-stage 

centrifugal compressor with canned combustors) 

Hydro-mechanical governor. 

Flame-out protection (min. flow stop). 

Maximum-flow schedule. 

1943 GE TG-100/T31 World’s first turboprop engine tested (Ref. 6). 

First two-spool turbine engine with the free turbine driving 

the propeller at the front. 

First axial-flow compressor aircraft gas turbine. 

1943 GE I-16 Bell P-59 (1st flt) First afterburner (AB) experiment on this engine. 

1944 GE J-33 Lockheed XP-80 

(1st flt) 

A direct descendant of the British Whittle engine of the early 

1940s. 

November 1945, the Allison Division of General Motors 

assumed complete responsibility for the development and 

production of J33 series engines. 

1945 Westinghouse J30 McDonnell FH-1 

(1st flt) 

The first pure jet fighter for the Navy (Ref. 7). 

1946 GE J35 Republic XP-84 

(1st flt) 

First axial-flow compressor turbojet. 

Late in 1947, complete responsibility for the production of 

the engine was transferred to the Allison Division of General 

Motors. 

1948 GE J47 F-86, F-86D 

(1st flt with AB) 

Production started in 1948, the most-produced aircraft turbine 

engine in history. 

First electronically controlled afterburner (AB), AB control 

was linked to main engine control. 

First anti-ice control system. 

1949 PW J52 A-4, A-6 First two-spool, aerodynamically connected turbine engine 

design.J52-P3 production started in 1959. 

Predecessor of JT8D. 
 

NASA/TM—2005-213978 16 

8.2 Growth Phase 

TABLE 8.2.—MAJOR MILESTONES FOR PROPULSION CONTROLS IN THE GROWTH PHASE (1950 TO 1969) 

Year Engine model Aircraft model Technology advancements 

1951 PW J57 B-50 (1st flt) First two-spool, aerodynamically connected turbine engine 

in production. Production started in 1953. 

1955 GE J79 NA B-45 (1st flt) 

XF-4D, F-104 

First high compression ratio engine flight-tested. 

First variable geometry engine, variable-stator compressor. 

First engine to reach Mach 2 in late 1955. 

1956 

1959 

GE J85 GAM-72 drone 

T-38 

First small turbojet engine. 

Afterburning version first flown in T-38. 

1957 PW JT3 B-707, DC-8 Commercial version of J57 in production. 

First U.S. commercial turbine engine. 

1958 PW J58 SR-71 (1st flt) First engine to reach Mach 3. 

Variable intake and nozzle controls. 

First variable cycle engine (var. comp. bypass). 

First electronic temperature trim, predecessor of 

supervisory control. 

First engine with a controlled bleed valve (during engine 

start). 

1958 GE CJ805 Convair 880 (1st flt) 

Convair 990 

Commercial version of J79. 

CJ805-23 was the first U.S. turbofan (free turbine driven 

aft fan); it first flew on the 990. 

1958 PW TF30 F-111, F-14 First afterburning turbofan engine. 

Hydro-mechanical fuel control (HMFC) and separate 

controls for AB and compressor stators. 

1959 GE J85/Lift Fan XV-5A First lift fan system powered by two J85’s tested. 

Twin lift fans with controllable louvers. 

First fixed-wing VTOL flight on Republic XV-5A. 

1960 GE CJ610 Learjet Business jet/general aviation version of J85. 

1961 PW JT8D DC-9, B-727, B737 First engine run of commercial version of J52 in 1961. 

Production started in 1963. 

Annular full-length bypass duct. 

1964 GE J93 NA XB-70 (1st flt) Predecessor of GE’s SST engine, the GE4. 

2,300 F turbine inlet temp. with turbine cooling. 

Mach 3 cruise for a large bomber/transport demonstrated in 

a 1965 flight. 

1966 GE TF39 C-5, B-52 First derivative of the core engine concept from the GE1 

demonstrator engine. 

First GE high bypass ratio (8:1) turbofan. 

Improved foreign object damage (FOD) rejection. 

1968 PW JT9D B-747 (1st flt) First PW high bypass ratio (5:1) turbofan. 

1968 GE CF6 DC-10 First ground run of the commercial version of TF39 in 

1968. 

Significant increase in fuel efficiency. 
 

NASA/TM—2005-213978 17 

8.3 Electronic Phase 

TABLE 8.3.—MAJOR MILESTONES FOR PROPULSION CONTROLS IN THE ELECTRONIC PHASE (1970 TO 1989) 

Year Engine model Aircraft model Technology advancements 

1970 PW F100 F-15 First modular concept engine tested on ground. 

First unified fuel control (UFC). 

First supervisory electronic engine control (EEC) - digital. 

First engine with event history recorder (EHR), 

predecessor of engine diagnostic unit (EDU). 

1972 Garrett TFE731 Falcon, Learjet First gas turbine engine with full-authority electronic 

control—analog, single channel. 

A reduced capability hydro-mechanical control (HMC) as a 

back up. 

First modulated inter-stage bleed valve. 

1972 GE F101 Rockwell B-1 Hydro-mechanical control (HMC) with full-authority 

electronic AB control—analog. 

1974 CFMI CFM56 DC-8 Series 70, 

KC-135, B-737 

GE F101 core and SNECMA M56 fan, turbine, and 

gearbox. 

Better fuel efficiency, lower noise and emissions. 

1981 PW F100 F-15, F-16 First flight test of a full-authority digital electronic engine 

control (DEEC). 

Single-channel DEEC with a back-up HMC. 

First FADEC-controlled military engine. 

1982 PW JT9D-7R4 B-747-400 Entry into service in 1982. 

Supervisory electronic control with HMC. 

1984 GE F110-100 F-16 Entry into service. 

Set single-engine reliability and safety record. 

1984 PW2037 B-757 First commercial service of advanced turbofan. 

Dual-channel FADECs. 

Active tip clearance controls for HPC, HPT, LPT. 

Multiple bleed valve controls. 

Engine program plug (EPP) in the FADEC to select 

multiple engine models. 

Integrated thermal management for oil and fuel. 

Advanced thrust reverser. 

Predecessor of F117 for C-17. 

1989 IAE V2500 A320, MD-90 Entry into service. 

First engine with auto-start (ignition and re-start). 

Model-based tip clearance control. 

Central maintenance features built in the FADEC. 
 

NASA/TM—2005-213978 18 

8.4 Integration Phase 

TABLE 8.4.—MAJOR MILESTONES FOR PROPULSION CONTROLS IN THE INTEGRATION PHASE (1990 TO 2003) 

Year Engine model Aircraft model Technology advancements 

1990 PW F119 F-22 First flight in 1990. 

Dual-channel, active FADECs. 

2-D thrust vectoring nozzle. 

Self-tuning on-board real-time model (STORM) for 

diagnostics. 

Analytical redundancy for gas-path sensors. 

Separate diagnostic unit (CEDU) from FADEC. 

1990 GE YF120 First flight in 1990. 

Variable cycle engine (variable bypass). 

Flight-qualified engine that has the highest number of 

controlled variables. 

1992 GE F110-129 F-16, F-14 Entry into service on F-16 in 1992. 

Full-authority digital electronic control (FADEC). 

1992 PW F100-220 F-15, F-16 First life extension control implementation by limiting 

rotor speed excursions with varying nozzle area and inlet 

guide vanes. 

1995 PW F100-229 F-15 Active First supersonic-yaw thrust vectoring nozzle flight 

demonstration (pitch yaw balanced beam nozzle, PYBBN). 

2000 PW JSF F119/ F135 JSF First flight in September 2000. 

Increased integration of flight-propulsion controls (for 

short take-off and vertical landing, STOVL). 

Integrated lift fan and jet propulsion system. 

First flight-critical propulsion system. 

First implementation of multivariable control. 

3-bearing swivel duct (3BSD) nozzle. 

2002 PW 6000 A-318 First flight in January 2002. 

First commercial application of PW’s self-tuning on-board 

real-time model, STORM. 
 

NASA/TM—2005-213978 19 

9.0 References 

 1. Garvin, R.V.; Starting Something Big—The Commercial Emergency of GE Aircraft Engines; AIAA 

Publisher, 1998. 

 2. Mattingly, J.D.; Elements of Gas Turbine Propulsion; McGraw-Hill Science/Engineering/Math 

Publisher, 1996. 

 3. Mattingly, J.D.; “Pictures of Aircraft Gas Turbine Engines,” http://www.aircraftenginedesign.com/ 

enginepics.html. 

 4. Schwent, G.; narrated personal experience, interviewed in 2003. 

 5. Otto, E.W. and Taylor, B.L., III; “Dynamics of a Turbojet Engine Considered as a Quasi-static 

System,” NACA Report 1011, 1950. 

 6. General Electric Company; Seven Decades of Progress—A Heritage of Aircraft Turbine Technology; 

Aero Publisher, 1979. 

 7. Swain, C.; “The Navy’s Jets,” http://www.microworks.net/pacific/aviation/ jetplanes.htm . 

 8. Bryson, A.E., Jr. and Ho, Y.C.; Applied Optimal Control, 1966. 

 9. Ketchum, J.R. and Craig, R.T.; “Simulation of Linearized Dynamics of Gas-turbine Engines,” 

NACA–TN–2826. 

10. McKinney, J.S., Captain, USAF; “Simulation of Turbofan Engine—Part I: Description of Method 

and Balancing Technique,” technical report for Air Force Aero Propulsion Laboratory, Air Force 

Systems Command, Wright-Patterson Air Force Base, AFAPL–TR–67–125, November, 1967. 

11. Durand, H.P.; “Simulation of Jet Engine Transient Performance,” paper presented at the Winter 

Annual Meeting of the Machine Design Division of ASME, ASME 65–WA/MD–16, 1965. 

12. Keck, M.F., Fredlake, J.J. and Schwent, G.V.; “A Control Concept Combining the Best of the Current 

Hydromechanical and Electronic Technologies,” SAE Paper 740380, 1974. 

13. Merrill, W., Lehtinen, B., and Zeller, J.; “The Role of Modern Control Theory in the Design of 

Controls fro Aircraft Turbine Engines,” AIAA Journal of Guidance, Control, and Dynamics, Vol. 7, 

No. 6, pp. 652–661, 1984. 

14. McMorran, P.D.; “Design of Gas Turbine Controllers Using Inverse Nyquist Array Method,” 

Proceedings of Institute of Electrical Engineers, Vol. 117, pp. 2050–2056, 1970. 

15. Michael, G.J. and Farrar, F.A.; “Development of Optimal Control Modes for Advanced Technology 

Propulsion Systems,” United Aircraft Research Laboratories internal report, East Hartford, CT, 1974. 

16. Stone, C.R., Miller, N.E., et al.; “Turbine Engine Control Synthesis—Vol. I, II, III,” 1975, AFAPL–

TR–75–14. 

17. DeHoff, R.L. Hall, W.E., et al.; “F100 Multivariable Control Synthesis Program—Volume I: 

Development of F100 Control System,” 1977, AFAPL–TR–77–35. 

18. Szuch, J.R., Soeder, J.F., et al.; “F100 Multivariable Control Synthesis Program—Evaluation of 

Multivariable Control Using a Real-time Engine Simulation,” NASA Lewis Research Center 

Technical Paper 1056, 1977. 

19. Lehtinen, B., DeHoff, R.L., and Hackney, R.D.; “Multivariable Control Altitude Demonstration on 

the F100 Turbofan Engine,” paper presented at the 15th Joint Propulsion Conference, 1979, AIAA–

79–1204. 

20. Joshi, D.S., Shaw, P.D., Rock, S.M., et al.; “A Design Approach to Integrated Flight and Propulsion 

Control,” SAE Paper 831482, 1983. 

21. Przybylko, S.J., Hutcheson, L., et al.; “Advanced Integrated Digital Engine Simulation—A Step 

Forward for Propulsion System Testing,” paper presented at the 6th AIAA Propulsion Joint Specialist 

Conference, 1970, AIAA–70–633. 

22. Teren, F.; “NASA Propulsion controls Research,” paper presented at the Sixth International 

Symposium on Air Breathing Engines, 1983, NASA TM–83343. 

23. F100-PW-220 Engine Monitoring System Reference Guide; Pratt & Whitney. 
 

NASA/TM—2005-213978 20 

24. Small, L.L.; “AFWAL Turbine Engine Controls Research and Development—Present and Future,” 

paper presented at the AIAA 20th Aerospace Sciences Meeting, January 1982. 

25. Merrill, W.C., DeLaat, J.C., and Burton, W.; “Advanced Detection, Isolation, and Accommodation 

of Sensor Failures—Real-time Evaluation,” J. of Guidance, Vol. 11, No. 6, Nov.–Dec. 1988, 

pp. 517–526. 

26. Burcham, Jr., F.W., Ray, R.J., et al.; “Propulsion Flight Research at NASA-Dryden from 1967 to 

1997,” NASA/TP—1998-206554. 

27. Greitzer, E.M.; “The Stability of Pumping Systems—The 1980 Freeman Scholar Lecture,” Journal of 

Fluids Engineering, Vol. 103, June 1981, pp. 193–242. 

28. Lorenzo, C.F., Chiaramonte, F.P., and Mehalic, C.M.; “Determination of Compressor In-stall 

Characteristics from Engine Surge Transients,” paper presented at the 20th AIAA/SAE/ASME Joint 

Propulsion Conference, June 1984, 84–1206. 

29. Luppold, R.H., et al.; “Estimating In-flight Engine Performance Variations Using Kalman Filter 

Concepts,” paper presented at the 25th Joint Propulsion Conference, July 1989, AIAA–89–2584. 

30. Maine, T.A. et al.; “A Preliminary Evaluation of an F100 Engine Parameter Estimation Process Using 

Flight Data,” NASA TM–4216, 1990 (AIAA paper no. 90–1921). 

31. Adibhatla, S.; “Propulsion Issues in Design of Integrated Flight and Propulsion Control Systems—

Lessons Learned from a STOVL Design Study,” AIAA–94–3610–CP. 

32. Garg, S. and Mattern, D.; “Application of an Integrated Methodology for Propulsion and Airframe 

Control Design to a STOVL Aircraft,” AIAA–94–3611–CP. 

33. Garg, S.; “Propulsion Controls and Health Management Research at NASA Glenn Research Center,” 

paper presented at the JANNAF Inter-agency Propulsion Committee Meeting, April, 2002. 

34. Lorenzo, C.F. and Merrill, W.C.; “Life Extending Control—A Concept Paper,” American Control 

Conference, June 1991, NASA TM–104391. 

35. Jaw, L.C., Wu, D.N., and Guo, T.H.; “Intelligent Life Extending Control for Gas Turbine Engines,” 

American Control Conference, June 2001. 

36. Drain, D.I.; “Propulsion Control and Control Theory—A New Research Focus,” paper presented at 

Propulsion Control Symposium, NASA Lewis Research Center, 1979. 

37. Anderson, Lynn; “Engine Control Systems Development at NASA,” paper being prepared as a NASA 

internal research effort, 2003. 
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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26

Propulsion Control Technology Development in the United States

A Historical Perspective

Link C. Jaw and Sanjay Garg

Propulsion; Engine control; Control systems design; Adaptive control

Unclassified -Unlimited

Subject Category: 07

Link C. Jaw, Scientific Monitoring, Inc., 8777 E. Via De Ventura, Scottsdale, Arizona 85258–3343; and Sanjay Garg,

NASA Glenn Research Center. Responsible person, Sanjay Garg, organization code RIC, 216–433–2685.

This paper presents a historical perspective of the advancement of control technologies for aircraft gas turbine engines.

The paper primarily covers technology advances in the United States in the last 60 years (1940 to approximately 2002).

The paper emphasizes the pioneering technologies that have been tested or implemented during this period, assimilating

knowledge and experience from industry experts, including personal interviews with both current and retired experts. Since

the first United States-built aircraft gas turbine engine was flown in 1942, engine control technology has evolved from a

simple hydro-mechanical fuel metering valve to a full-authority digital electronic control system (FADEC) that is common

to all modern aircraft propulsion systems. At the same time, control systems have provided engine diagnostic functions.

Engine diagnostic capabilities have also evolved from pilot observation of engine gauges to the automated on-board

diagnostic system that uses mathematical models to assess engine health and assist in post-flight troubleshooting and

maintenance. Using system complexity and capability as a measure, we can break the historical development of control

systems down to four phases: (1) the start-up phase (1942 to 1949), (2) the growth phase (1950 to 1969), (3) the electronic

phase (1970 to 1989), and (4) the integration phase (1990 to 2002). In each phase, the state-of-the-art control technology is

described and the engines that have become historical landmarks, from the control and diagnostic standpoint, are

identified. Finally, a historical perspective of engine controls in the last 60 years is presented—in terms of control

system complexity, number of sensors, number of lines of software (or embedded code), and other factors.