Jonathan Litt

Glenn Research Center, Cleveland, Ohio

T. Shane Sowers

Vantage Partners, LLC, Cleveland, Ohio

Yuan Liu

N&R Engineering and Management Services, Parma Heights, Ohio

A. Karl Owen and Ten-Huei Guo

Glenn Research Center, Cleveland, Ohio

Validation of an Integrated Airframe and Turbofan 

Engine Simulation for Evaluation of Propulsion 

Control Modes

NASA/TM—2015-218715

May 2015

AIAA–2015–1476
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Jonathan Litt

Glenn Research Center, Cleveland, Ohio

T. Shane Sowers

Vantage Partners, LLC, Cleveland, Ohio

Yuan Liu

N&R Engineering and Management Services, Parma Heights, Ohio

A. Karl Owen and Ten-Huei Guo

Glenn Research Center, Cleveland, Ohio

Validation of an Integrated Airframe and Turbofan 

Engine Simulation for Evaluation of Propulsion 

Control Modes

NASA/TM—2015-218715

May 2015

AIAA–2015–1476

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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NASA/TM—2015-218715 1 

Validation of an Integrated Airframe and Turbofan Engine 

Simulation for Evaluation of Propulsion Control Modes 

Jonathan Litt 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

T. Shane Sowers 

Vantage Partners, LLC 

Cleveland, Ohio 44135 

Yuan Liu 

N&R Engineering and Management Services 

Parma Heights, Ohio 44130 

A. Karl Owen* and Ten-Huei Guo 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The National Aeronautics and Space Administration (NASA) has developed independent airframe 

and engine models that have been integrated into a single real-time aircraft simulation for piloted 

evaluation of propulsion control algorithms. In order to have confidence in the results of these 

evaluations, the integrated simulation must be validated to demonstrate that its behavior is realistic and 

that it meets the appropriate Federal Aviation Administration (FAA) certification requirements for 

aircraft. The paper describes the test procedures and results, demonstrating that the integrated simulation 

generally meets the FAA requirements and is thus a valid testbed for evaluation of propulsion control 

modes. 

Nomenclature 

A Inlet Area 

AEO All Engines Operational 

AGL Above Ground Level 

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

CD,wind Drag Coefficient Due to Windmilling 

CL Coefficient of Lift 

CFR Code of Federal Regulations 

D Drag 

DWind Drag Due to Windmilling 

FAA Federal Aviation Administration 

MCT Maximum Continuous Thrust 

MSL Mean Sea Level (ft) 

 

*

Distinguished Research Associate. 
NASA/TM—2015-218715 2 

NASA National Aeronautics and Space Administration 

OEI One Engine Inoperative 

q Dynamic Pressure 

S Wing Area (ft2

) 

TAS True Air Speed 

TCM Transport Class Model 

VFS Final Segment of a Departure Airspeed with One Engine Inoperative (OEI) 

VFTO Final Takeoff Airspeed 

VLOF Liftoff Speed 

VMC Minimum Control Airspeed—Clean Configuration (Flaps and Gear Retracted) 

VMCL Minimum Control Airspeed—Landing Configuration (Full Flaps, Gear Extended) 

VR Rotation Airspeed 

VS Stall Airspeed—Clean Configuration (Flaps and Gear Retracted) 

VS0 Stall Airspeed—Landing Configuration (Full Flaps, Gear Extended) 

VSR Reference Stall Airspeed—Takeoff/Approach Configuration at Half Flaps 

VX Best Angle of Climb Airspeed 

VY Best Rate of Climb Airspeed 

V1 Speed Above Which a Takeoff will Proceed with One Engine Inoperative (OEI) 

V2 Safe Takeoff Speed with One Engine Inoperative (OEI) 

α Angle of Attack 

1.0 Introduction 

The National Aeronautics and Space Administration (NASA) has developed a number of nonproprietary aerospace-related simulations for use by the public. These include several independentlydeveloped turbofan engine simulations (Refs. 1 to 5) and at least one airframe simulation (Ref. 6). While 

the engine simulations are freestanding, the airframe simulation contains a simplified engine model. Of 

these simulations, the Transport Class Model (TCM) (Ref. 6) airframe simulation and the Commercial 

Modular Aero-Propulsion System Simulation 40k (C-MAPSS40k) (Refs. 4 and 5) turbofan engine model 

are of particular interest for research in such areas as aviation safety and autonomous operation. The TCM 

represents a generic, mid-sized, single-aisle transport aircraft with two wing-mounted engines. 

C-MAPSS40k models a 40,000-lbf thrust class, high-bypass ratio turbofan engine with a representative 

commercial-type controller. These independently-developed simulations are deemed to be compatible 

based on their size. In fact, the baseline TCM utilizes first-order engine models derived from 

C-MAPSS40k, that have dynamics that vary as a function of altitude and fan speed. However, they are 

based on look-up tables (Ref. 6), so all physical and thermodynamic information is lost. The input to the 

TCM’s simplified engine models is throttle position, the outputs are fan speed and thrust. 

Using these simulations to perform meaningful research on integrated flight and propulsion control, 

or on the impact on flight dynamics resulting from propulsion control modification, requires replacing the 

simplified engine models with full nonlinear C-MAPSS40k simulations. This enables modification of the 

propulsion control system and allows access to internal engine variables within the integrated simulation. 

In order to have confidence that the results are representative of what could be expected in a fielded 

application, the integrated TCM/C-MAPSS40k simulation must be validated. There are several aspects to 

this integration, testing, and validation process. First, the interface between the airframe and propulsion 

system must be defined (fuel flow, Mach number, altitude, ambient temperature and pressure, throttle 

position, thrust, etc.). This is relatively straightforward, considering that these variables are standard in 

simulations of sufficient complexity. The more complicated aspect, once the two systems are combined, 
NASA/TM—2015-218715 3 

is to ensure that they are working together as they should. Simulated flying, especially when performed in 

a flight simulator by an experienced pilot, gives confidence that the models are integrated properly and 

that the physics captured in them is correct. However, successful integration does not mean the individual 

models are truly compatible. The Federal Aviation Administration (FAA) specifies minimum 

performance requirements for the combined system, and mandates that an aircraft meet them in order 

to be certified (Ref. 7). Therefore, for a simulation to be considered representative of a real aircraft, it 

should display similar behavior. Because these simulations were developed independently and then 

integrated, not unlike the actual hardware they represent, it is particularly important to validate their 

combined performance. 

The remainder of the paper describes the model integration, the test procedure used to validate the 

integrated simulation, and the results and analysis. The conclusions will discuss whether the certification 

testing justifies the combined simulation’s use for advanced propulsion control and integrated flight and 

propulsion control research. 

2.0 Background 

For this work, the TCM (Ref. 6), an airframe model with two wing-mounted engines, and two copies 

of the C-MAPSS40k (Refs. 4 and 5) turbofan engine model were integrated. The closed-loop 

C-MAPSS40k engine meets the relevant FAA thrust response criterion, FAR §33.73(b), which requires 

that, from the fixed minimum flight idle power lever position, the thrust should reach 95 percent of rated 

takeoff power in not over 5 sec (Figure 1) (Ref. 8). While important by itself, this is a stand-alone engine 

requirement. It is the resulting performance of the aircraft simulation with the engine models installed that 

will determine whether the integrated system is appropriate for use in meaningful piloted flight simulator 

testing. For example, an exemption may be granted for the engine response requirement (FAR §33.73(b)) 

on the basis of compliance with the climb performance for installed thrust requirement, FAR §25.119 

(Ref. 9), which states that the thrust achieved 8 sec after power application (starting from minimum flight 

idle) must allow a climb gradient (the ratio of the increase of altitude to horizontal distance travelled, 

expressed as a percentage) of not less than 3.2 percent for a go-around (Ref. 10). Previous testing has 

demonstrated go-around maneuvers from a variety of approaches (Ref. 11), both stabilized and 

Figure 1.—C-MAPSS40k response demonstrating compliance with §FAR 

33.73(b)—Power or thrust response. Rise time from the fixed minimum flight 

idle power lever position to 95 percent of rated takeoff thrust is not over 5 sec. 

Operating conditions are defined as sea-level static condition. 

18 19 20 21 22 23 24 25 26 27 28

40

50

60

70

80

Power Lever Angle

18 19 20 21 22 23 24 25 26 27 28

0

1

2

3

4

x 104

Thrust, lbf

Time, s

Thrust Response

95% Max Thrust

5 s Requirement
NASA/TM—2015-218715 4 

Figure 2.—Altitude loss during a go-around for the integrated TCM/C-MAPSS40k 

simulation. The runs comprise various initial airspeeds and glide path angles; 

in all cases the throttle is at or near idle. True Air Speed, TAS, in knots, and 

sink rate, in feet per minute, at initiation are given for each case. At the 

longest, the minimum altitude occurred about 5 sec after the maneuver was 

triggered, and in all cases, a relatively constant or increasing positive climb 

gradient was achieved in under 8 sec. 

unstabilized (Ref. 12). In all cases, the aircraft started to climb at a relatively constant rate within 8 sec 

(Figure 2), and much sooner in the more stabilized approaches (where the sink rate was less than 

1000 ft/sec at an appropriate approach speed). The actual requirement does not specify an initial speed 

or sink rate, meaning that starting the go-around from a stabilized approach, which is the best case, is 

acceptable. The integrated simulation should be suitable for non-real-time (i.e., faster than real time) 

desktop use, or for real-time piloted simulation. In either case, the integration and performance of the 

overall system is what is important. If the system is able to meet the FAA requirements for normal and 

emergency operation (e.g., one engine inoperative (OEI)), it will demonstrate that it is representative of a 

modern commercial aircraft. This will support the validity of the comparison of nominal and advanced 

control modes that could be evaluated using the simulation. 

The specific certification requirements that will be evaluated are limited to those that are relevant to 

the types of research and flight regimes that the integrated model is used for, and more precisely, to the 

performance features captured in the model. This testing does not claim to be comprehensive. Actual 

certification of modern aircraft can take years, requiring over 1500 hr for the flight testing alone (Ref. 13). 

Although this effort is, out of necessity, significantly scaled down, it will sample enough representative 

aspects of the requirements to justify confidence in the simulation. The regulations not tested include 

those related to: weather, including ice, rain, wind, and wind gusts; aerodynamic effects such as buffeting; 

fire hazards; passenger or cargo; aircraft to ground interaction during taxiing, takeoff and landing; 

changes in performance due to runway gradients; cockpit arrangement and instrumentation; control 

surface physics, including loads, range of motion, gust locks, or jamming; structural issues such as load 

factors; crosswind performance; and weight and balance (center of gravity). 

 

0 1 2 3 4 5 6 7 8 9 10

-150

-100

-50

0

50

100

150

200

250

300

350

Time Since Trigger, s

Altitude Change From Trigger, ft

TAS: 156 kts; Sink Rate: 3128 fpm

TAS: 108 kts; Sink Rate: 857 fpm

TAS: 109 kts; Sink Rate: 1262 fpm

TAS: 116 kts; Sink Rate: 664 fpm

TAS: 124 kts; Sink Rate: 761 fpm
NASA/TM—2015-218715 5 

2.1 Integration of the Simulations 

The two default engine modules within the TCM were each replaced by a copy of the closed-loop 

C-MAPSS40k engine model. The inputs to the engine models are pilot commands (e.g., throttle position) 

and flight conditions (e.g., altitude, Mach number). The engine models calculate net thrust, which is 

transferred back to the TCM and converted into body-axis force and moment components. Within the 

integrated simulation, application of thrust reversers on landing slows the aircraft by causing the thrust to 

change direction, that is, point forward, at a level equal to 10 percent of the engines’ net thrust; this ad 

hoc reduction loosely approximates the cancellation of forces since core thrust is not reversed, and bypass 

thrust is not redirected a full 180°. Other outputs, such as spool speeds and exhaust gas temperature, are 

used for cockpit instrumentation display when the integrated model is implemented in a flight simulator. 

Fuel usage is another potential input to the TCM from C-MAPSS40k. The TCM can model the flow 

of fuel to the engines, where it is consumed, thus reducing the aircraft weight (Ref. 6), but the amount of 

fuel used is determined by the engines. Since, in this implementation, the existing simplified engine 

models, which were replaced by copies of C-MAPSS40k, did not account for fuel at all, no attempt was 

made to make this connection in the integration, i.e., fuel flow is all internal to the engine models. This 

design choice is justified by the fact that the integrated simulation is planned to be used for short tests, 

lasting several minutes at most, where changes in overall aircraft weight would be negligible. Although 

the connection itself could be made with little effort, the accompanying changes in weight and inertias 

(of which only the baseline value of each is provided with the TCM simulation), and the resulting 

modification of the flight dynamics, would have to be incorporated. This modification and subsequent 

validation would be beyond the scope of this effort. However, it should be noted that the TCM is stated to 

represent a mid-value weight for a mid-sized transport aircraft (185 000 lb) (Ref. 6). So, although the 

certification requirements for an aircraft are meant to cover its entire weight range, as long as all of the 

certification testing and research efforts are performed at a consistent aircraft weight, the research results 

can be deemed valid. 

2.2 Modification to the Models 

In addition to the integration, two main modifications were made to the TCM: the landing gear model 

was modified and an engine out/windmilling model was added. 

2.2.1 Landing Gear 

The landing gear model modification was intended to improve the aircraft’s interaction with the 

ground. It is not related to the certification testing that is being reported in this paper. 

2.2.2 Windmilling 

Validation testing of aircraft performance under specific engine out scenarios is required for 

certification. To account for the additional drag contribution due to an engine shutdown during flight, 

windmilling (or unpowered spinning) models were implemented in each of the two C-MAPSS40k engine 

simulations within the TCM. The windmilling models are based on empirical data from the literature 

(Ref. 14). The models provide engine spool speeds and thrust/drag in the event of a shutdown. 

Windmilling drag was obtained as the sum of two values: a baseline drag coefficient due to the thrust 

capability of the engine, and an additional contribution represented as a function of Mach number. 

 
NASA/TM—2015-218715 6 

Figure 3.—Additional engine drag coefficient due to windmilling. 

Figure 4.—Modeled rotational speed of the engine spools due to windmilling. 

Using data from Reference 14, the baseline windmilling drag coefficient for the C-MAPSS40k engine 

was estimated to be 0.31. The additional drag due to Mach number is shown in Figure 3. Note that the 

data displayed in Figure 3 begin at Mach 0.2, which is about the takeoff speed of the TCM. Any engine 

out testing below this speed uses the minimum value from Figure 3. The drag force is then calculated as: 

DWind  CD,wind  q  A (1) 

where DWind is the drag force due to windmilling; CD,wind is the drag coefficient due to windmilling, which 

equals the sum of the baseline value, 0.31, and the contribution due to windmilling from Figure 3; q is the 

freestream dynamic pressure; and A is the inlet area. 

In the event of an engine failure, the TCM inputs switch from spool speeds and thrust values 

calculated by the C-MAPSS40k engine model to the speeds and drag values calculated by this 

windmilling model. When this switch occurs, a low-pass filter is applied to each parameter to simulate a 

more realistic, smooth transition between C-MAPSS40k and the windmilling model. Windmilling spool 

speeds are calculated as functions of flight Mach number (Figure 4). These parameters are used for flight 

simulator cockpit instrumentation displays and do not affect the physics of the aircraft model. 

 

0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

0.17

0.175

0.18

0.185

0.19

0.195

0.2

Mach Number

Additional Drag Coefficient

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9

0

5

10

15

20

25

30

35

40

Mach Number

Rotational Speed (%)

LP Spool

HP Spool
NASA/TM—2015-218715 7 

3.0 Code of Federal Regulations Part 25 

The Federal Aviation Administration’s Code of Federal Regulations (CFR) Part 25 (Ref. 7), 

“Airworthiness Standards: Transport Category Airplanes,” is to be applied to aircraft in the category of 

the TCM (Ref. 6). Although the standards apply to real aircraft, they are useful in gauging the realism of 

TCM performance. A systematic review was utilized to identify regulations that may be applied to the 

TCM. The process removed requirements that were not relevant to nor modeled within the TCM 

(e.g., loads, structure, wind, turbulence). The remaining regulations relate primarily to aspects of the 

TCM relevant to its use in research applications. That is not to say that performance-related features could 

not be incorporated in the future; wind and turbulence are worthy additions, and the effect of ice 

accretion, both on the structure and in the engines, is certainly a topic of interest. However, the scope of 

this work is limited to those regulations that specifically address the existing airframe and propulsion 

system model integration from a performance perspective. 

The testing was further restricted to the regime in which the TCM is used. Flight test conditions were 

kept consistent when examining the TCM performance characteristics. Flight was executed at or below 

5,000 ft MSL (mean sea level), the weight and balance of the aircraft model were kept constant, and a 

standard atmosphere was used. These test constraints go against the spirit of CFR Part 25, which seeks to 

confirm compliance across all operational weights, balances, altitudes and temperature variations. 

However, full compliance testing is beyond the scope of this paper, which has a goal of simply 

confirming the reasonableness of the TCM for research applications. 

The regulations were reduced and grouped into four performance areas: stalls, takeoff, climb, and 

stability. For each, there is an introduction and results section. However, before the simulated certification 

process began, analysis was performed to determine the common aircraft reference speeds, called “V” 

speeds (Ref. 15). Table 1 lists these speeds for the TCM. 

Values for the common “V” speeds given in Table 1 were determined analytically and empirically. 

The final takeoff speed, VFTO, was determined experimentally in the flight simulator by noting when the 

TCM consistently lifted off from the runway at takeoff. The minimum control airspeeds, VMC and VMCL, 

were determined experimentally in the flight simulator by the pilot flying maneuvers with one engine 

inoperative. The objective of this test was to determine the airspeed above which the rudder authority is 

sufficient to counteract the thrust asymmetry. The rotation airspeed, VR, was determined by trial-anderror. The selected VR is somewhat subjective because the aircraft is able to rotate at a lower speed, but 

TABLE 1.—COMMON “V” SPEEDS USED WITH THE TCM 

Speed, kn Description 

VFTO 127 Final takeoff airspeed 

VMC 115 Minimum control airspeed—clean configuration 

VMCL 90 Minimum control airspeed—landing configuration 

VR 100 Rotation airspeed 

VS 130 Stall airspeed—clean configuration (flaps and gear retracted) 

VS0 96 Stall airspeed—landing configuration (full flaps, gear extended) 

VSR 110 Reference stall airspeed—takeoff/approach configuration at half flaps 

VX 176 Best angle of climb airspeed 

VY 358 Best rate of climb airspeed 

 
NASA/TM—2015-218715 8 

the value consistently gave good performance. The stall speeds, VS, VS0, and VSR, were determined 

experimentally and analytically. Using the TCM aerodynamic database, the maximum lift coefficient, 

CL,max, was found; there is a different maximum lift coefficient for each flap setting. The stall speed is 

determined using the equation, 

,max

2

L

stall density S C

weight V    (2) 

where the aircraft weight and air density are known constant values, and S is the area of the wing, which 

is about 1950 ft2

. The stall speeds calculated for each flap configuration at sea level were confirmed 

against experimental results from the flight simulator. 

The best angle and rate of climb speeds were determined analytically. A velocity range was selected 

(i.e., 150 to 450 kn) and for each point in the velocity profile, the dynamic pressure, q, was calculated using 

2

2

1

q   density  velocity (3) 

From the dynamic pressure, the required coefficient of lift, CL, was determined for the given aircraft 

weight using 

q S

weight CL   (4) 

Using the aerodynamic database, the angle of attack, α, for the given CL was determined. Then, given α, 

the corresponding drag coefficient, CD, was found from the aerodynamic database. Finally, the drag, D, 

was determined using 

D  q S CD (5) 

The thrust required for steady flight at the given velocity is equal to the drag. Using this information, the 

thrust available from the two engines at full power and the thrust required for steady flight can be plotted 

versus velocity (Figure 5). The best angle of climb speed, VX, is the airspeed with the greatest climb per 

horizontal distance; it is the airspeed with the maximum excess thrust, which was found to be 176 kn. 

Because of the variation in the thrust required curve, there are two secondary speeds (local maxima), 

194 and 268 kn, which will maximize climb angle performance for higher airspeeds. These secondary 

angle of climb speeds are also indicated in Figure 5. 

The best rate of climb speed, VY, is the airspeed that produces the fastest rate of climb (i.e., shortest 

time to climb) and is the airspeed with the maximum excess power. Power is determined using, 

 power = thrust × velocity (6) 

From this equation, the power required and the power available from the two engines at full power are 

computed across the velocity profile and plotted in Figure 6. From these data, the maximum excess power 

airspeed (VY) is found to be 358 kn. 

Knowing the maximum excess power allows for the maximum rate of climb to be calculated. To find 

the maximum sustainable rate of climb, the excess power is divided by the airplane weight as shown by 

weight

Power Power

R C avail  req max /  (7) 

This gives a maximum rate of climb of 127 ft/sec or 7,619 ft/min. 
NASA/TM—2015-218715 9 

Figure 5.—Maximum excess thrust airspeed at sea level (best angle of climb airspeed, VX). 

Figure 6.—Maximum excess power airspeed (best rate of climb airspeed, VY). 

Ideally, the thrust required and power required curves should decrease to a minimum and then start 

increasing smoothly over the velocity range of interest. In Figure 5 and Figure 6, however, the slopes 

change sign several times from about 170 kn to about 265 kn. This variation is due to the CD profile in the 

TCM aerodynamic database. In the region specified by angle of attack (α) values from around 4° to 10°, 

150 200 250 300 350 400 450

1

2

3

4

5

6

7

x 104

Airspeed (knots)

Thrust (lbf)

Thrust Required

Thrust Available

268 knots

194 knots

176 knots

Maximum Excess Thrust

150 200 250 300 350 400 450

0

0.5

1

1.5

2

2.5

3

3.5

4

x 107

Power (lbf ft/s)

Airspeed (knots)

Power Required

Power Available

Maximum

Excess

Power

358

knots
NASA/TM—2015-218715 10 

there are multiple increases and decreases in the slope of the CD versus α curve (Figure 7) (Ref. 6), that, 

when carried through the required thrust/power calculations, produce non-convex regions on the plots 

(Figure 5 and Figure 6). Because of this jaggedness, the climb performance of the TCM is not as sensitive 

to variations in airspeed as other aircraft with more definitive VX and VY speeds. This is especially true for 

the best angle of climb speed, where local maxima produce two secondary angle of climb speeds. 

Attempting to use the calculated climb performance speeds, particularly VY, would lead to operational 

challenges. On takeoff, it is difficult to accelerate to these speeds with a VFTO of 127 kn. To achieve a high 

performance takeoff climb, the airplane would be required to lift off, fly level to accelerate, and then 

climb when the desired airspeed is attained. The acceleration portion of this maneuver would impose a 

penalty in the climb gradient and would expose the airplane to risk in the event of an engine failure. 

Another consideration is that the VY speed would not be permissible for an actual aircraft taking off at a 

sea level airport, dictated by CFR §91.117(a), which states that there is a 250 kn indicated airspeed limit 

while flying below 10 000 ft MSL. 

In the evaluations presented in the following sections, four airplane configurations were standardized 

for consistency, as shown in Table 2. It is assumed that the takeoff and approach configurations for the 

TCM are the same. Figure 8 shows the flight segments used in this analysis, along with the various 

speeds, power settings, and aircraft configurations. Because of modeling limitations, takeoff examinations 

begin at 35 ft above ground level (AGL), where it is assumed that the landing gear has been retracted, but 

the flaps are still deployed (Second and Third Segments). At the end of the initial takeoff climb, flaps are 

raised and the airplane is then in a cruise configuration as it continues its climb out to its desired altitude 

at a reduced power setting (Maximum Continuous Thrust or MCT) (Final Segment). The speeds shown in 

Figure 8 are defined in the nomenclature and important values for the testing described here are provided 

in Table 1. 

Figure 7.—TCM coefficient of drag, CD, versus angle of attack, α, for no flaps (recreated from Ref. 6). 

TABLE 2.—ASSUMED AIRPLANE CONFIGURATION 

DURING FLIGHT PHASES 

Flight phase Landing gear Flaps 

Takeoff Retracted Half flaps 

Approach Retracted Half flaps 

Landing Extended Full flaps 

Cruise Retracted No flaps 

-5 0 5 10 15 20

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

 (degrees)

CD
NASA/TM—2015-218715 11 

Figure 8.—Flight path showing flight segments used in certification, along with the various speeds, power settings, 

and aircraft configurations. 

Figure 9.—The procedure used to stall the aircraft in a turn. 

3.1 Stalls 

Stall occurs when the angle of attack is too great for the wing to produce lift. The airflow separates 

from the wing causing loss of lift. This typically occurs during approach due to slow airspeed, or during a 

takeoff climb that is too steep. 

Due to its hazardous nature, airplane stall behavior is managed through several regulations in 

CFR 25: §25.103, §25.201, and §25.203. These regulations stipulate that stalls are to be demonstrated 

from a stabilized trim condition in both straight flight and in a 30° banked turn. When encountered, the 

airplane is to give a clear and distinctive indication of the stall, and it is to react with no excessive 

banking or pitching so that an average pilot may regain control. 

For level flight, the stall speeds were determined empirically. The procedure required the pilot to 

retard the throttles to idle and increase α in an attempt to maintain altitude as the airspeed decreased. Once 

the aircraft stalled, it began to descend and the nose dropped. The pilot applied full power and pulled up 

to recover. The reference stall speed (VSR) was selected in the takeoff/approach configuration with flaps 

set to one half. In this configuration, the airplane was found to stall at 110 kn. In the landing 

configuration, the airplane stalled at 96 kn (VS0). In a cruise configuration, the airplane stalled at 130 kn 

(VS). Stalls in level flight were stable with easy recovery. 

The procedure to stall the aircraft in a turn is shown in Figure 9. The plane was put into a 30° banked 

turn with half flaps, the throttles were retarded to idle, and the nose was slowly pulled up to increase α

while maintaining altitude. When stall occurred, the nose dropped as the aircraft yawed toward the low 

Ground Roll First Segment Second Segment Third Segment Final Segment 35’ 400’

1500’

All

Operating Engines

Landing Gear

Airspeed

Flaps

Thrust

V1 VR VLOF

V2

Down

One

Inoperative

Retracted

Accelerate

Down

V2 Accelerate VFS

Up

Takeoff MCT
NASA/TM—2015-218715 12 

Figure 10.—Stall during 30° banked turn. Airspeed and angle of attack, α, are shown versus time. 

wing. Altitude loss was gradual and, as shown in Figure 10, the aircraft began to accelerate slowly. The 

airplane was stable throughout the maneuver with little change in bank angle being noted. Recovery was 

achieved by rolling wings level, applying power, and pulling up. It was prompt and easily executed. Stall 

speed was approximately 123 kn, at which point α started to decrease, as indicated in Figure 10. 

3.2 Takeoff 

TCM takeoff performance was evaluated against portions of CFR §25.111. Because of modeling 

limitations, takeoff evaluations did not involve requirements on runway distance or the immediate runway 

environment. Therefore, the flight simulator studies began when the airplane was 35 ft above the runway 

to align with CFR regulations for initial climb out altitude. Per regulations, one engine is made 

inoperative during takeoff and, at 35 ft above the runway, the airspeed is to be at or above V2, the safe 

takeoff speed. As the airplane climbs, the slope of the climb gradient is to be positive at each point. Once 

the airplane climbs above 400 ft, the climb gradient is to be not less than 1.2 percent. The only change in 

airplane configuration permitted during this segment is landing gear retraction.

In the takeoff evaluation, the left engine was failed during rotation with flaps set at one half, and 

landing gear was retracted once a positive rate of climb was achieved. From Figure 11, it can be seen that 

the slope of the takeoff path remained positive and was well above the 1.2 percent gradient of climb 

requirement throughout the profile. At 400 ft, the airspeed was approximately 140 kn, well above the 

assumed V2 speed of 127 kn (V2 was assumed to be the same as VFTO for this work). 

3.3 Climb 

The TCM was assessed for its ability to climb with all engines operating and one engine inoperative, 

as stipulated in CFR §25.119 and §25.121. Evaluation considered the five test scenarios listed in Table 3, 

with gradient of climb being the primary measure of interest. In Test Scenario 1, the takeoff performance 

is measured with all engines operating, half flaps and the landing gear extended. This is the only climb 

evaluation scenario where engine response time comes into play. Per regulation, the aircraft is to have 

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5

17.5

18

18.5

19

19.5

Time (seconds)

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5

17.5

18

18.5

19

19.5

 (degrees)

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5

120

122

124

126

128

Airspeed (knots)

 (degrees)

Airspeed (knots)

123 knots
NASA/TM—2015-218715 13 

Figure 11.—Gradient, altitude, and true airspeed during takeoff with one engine inoperative. 

TABLE 3.—CLIMB PERFORMANCE EVALUATION SCENARIOS 

[AEO, all engines operational; OEI, one engine inoperative.] 

Test 

scenario 

Engine 

status 

Airplane configuration Gradient of 

climb required 

1 AEO Landing—half flaps with landing gear extended 3.2 percent 

2 OEI Takeoff—half flaps with gear extended Positive 

3 OEI Takeoff—half flaps with landing gear retracted 2.4 percent 

4 OEI Final takeoff—flaps up with landing gear retracted 1.2 percent 

5 OEI Approach—half flaps with landing gear retracted 2.1 percent 

achieved, for go-around from minimum flight idle, a minimum climb gradient of 3.2 percent within 8 sec 

of power application to the go-around thrust setting. This is also the only climb evaluation with all 

engines operating, as the remaining test scenarios specify that one engine is inoperative. Test Scenarios 2 

and 3 have flaps deployed and the landing gear extended and retracted, respectively. Test Scenario 4 

specifies that the flaps and landing gear are retracted in a cruise configuration, while Test Scenario 5 has 

the same configuration as Test Scenario 3 but with a reduced gradient of climb requirement. 

The five test scenarios were flown with the TCM travelling at airspeeds of around 130 to 145 kn. 

These speeds were used because they give good performance and are readily achievable during the 

takeoff climb out. They are also in the neighborhood of the speeds called for in the regulations, although 

in some cases are several knots higher than specified. 

0 0.5 1 1.5 2 2.5

x 104

0

2

4

6

8

10

Gradient of Climb (%)

Distance (feet)

0 0.5 1 1.5 2 2.5

x 104

0

500

1000

1500

2000

2500

Altitude (feet)

Distance (feet)

0 0.5 1 1.5 2 2.5

x 104

130

135

140

145

150

155

True Airspeed (knots)

Distance (feet)

Observed Gradient of Climb

1.2% Gradient of Climb Limit

35'

400'
NASA/TM—2015-218715 14 

For Test Scenario 1, an approach to landing was flown with a go-around being executed at 50 ft AGL. 

The resulting gradient of climb is shown in Figure 12. In this situation, the TCM had no difficulty in 

satisfying the 3.2 percent climb gradient requirement. In fact, by 8 sec after power was applied, the 

gradient requirement was greatly exceeded with values in the mid-20 percent range. 

In Test Scenario 2, a takeoff was flown with the TCM configured with half flaps and landing gear 

extended; one engine was inoperative. From the gradient of climb plot shown in Figure 13, it can be seen 

that the TCM met the positive gradient of climb requirement. 

Figure 12.—Gradient of climb for all engines operating in landing configuration (go-around) (Test Scenario 1). 

Figure 13.—Gradient of climb for one engine inoperative in landing configuration (Test Scenario 2). 

 

0 10 20 30 40 50 60 70

-5

0

5

10

15

20

25

30

Gradient of Climb (%)

Time from Throttle Up (sec)

All Engines Operating, 1/2 Flaps, Landing Gear Extended (Landing Configuration)

Observed Gradient of Climb

3.2% Gradient of Climb Limit 8 Second

Response

0 1000 2000 3000 4000 5000 6000 7000 8000

0

2

4

6

8

10

12

Gradient of Climb (%)

Distance (feet)

Observed Gradient of Climb
NASA/TM—2015-218715 15 

Test Scenarios 3 and 5 utilized the same TCM configurations. Test Scenario 3 is a takeoff climb 

while Test Scenario 5 is an approach go-around. In each of these cases, the TCM is in the same 

configuration with one engine inoperative. As shown in Figure 14, the TCM was easily able to satisfy the 

2.4 percent (Scenario 3) and 2.1 percent (Scenario 5) gradient of climb specifications. 

In Test Scenario 4, climb performance at the end of the takeoff climb out is measured. The TCM has 

landing gear and flaps retracted in the cruise configuration with one engine inoperative. As shown in 

Figure 15, the TCM exceeded the 1.2 percent gradient of climb requirement. 

In all of the test scenarios, the TCM was able to fulfill all of the climb requirements evaluated even 

though the best climb performance speeds were not flown. 

Figure 14.—Gradient of climb for one engine inoperative in takeoff/approach configuration (Test Scenarios 3 and 5, 

respectively). 

Figure 15.—Gradient of climb for one engine inoperative in cruise (en route) configuration (Scenario 4). 

0 0.5 1 1.5 2 2.5

x 104

0

1

2

3

4

5

6

7

8

9

10

Gradient of Climb (%)

Distance (feet)

Observed Gradient of Climb - Takeoff

Observed Gradient of Climb - Approach

2.4% Gradient of Climb Limit - Takeoff

2.1% Gradient of Climb Limit - Approach

0 2000 4000 6000 8000 10000 12000 14000 16000

1

2

3

4

5

6

7

8

9

10

Gradient of Climb (%)

Distance (feet)

Observed Gradient of Climb

1.2% Gradient of Climb Limit
NASA/TM—2015-218715 16 

3.4 Stability 

The stability of the TCM was evaluated using the standards in CFR §25.171, §25.173, §25.177, and 

§25.181. These regulations state that the airplane must be stable in all axes, longitudinally, directionally, 

and laterally, in any condition encountered during flight. Once trimmed for level flight, when a 

disturbance is experienced, the airplane must dampen any resulting oscillation and return to its trim 

condition. A pull on the control yoke or stick must be required to maintain speeds below the trim speed 

and a push to maintain speeds above. 

To gauge the longitudinal stability, a perturbation is introduced to the flight controls to measure the 

TCM’s ability to return to its trim condition. Per regulations, the airspeed is to return to within 10 percent 

of its value prior to the perturbation for the climb, approach and landing configurations, and to within 

7.5 percent for the cruise configuration. Stability was tested by initiating a 10° pitch up command 

followed by a release of the flight controls. The TCM was then allowed to settle back to its trim condition 

without any pilot input. The test was performed with the model in both the takeoff/approach and cruise 

configurations; the maneuver was initiated at airspeeds of 177 and 250 kn. From Figure 16, in the 

takeoff/approach configuration, the TCM appears to be dampening out the oscillations at both airspeeds. 

When the pilot input the 10° pitch up command, the nose remained high. As the airspeed dropped off, the 

Figure 16.—Response from flight control release after 10° pitch-up (takeoff/approach configuration). 

 

0 50 100 150 200 250

4000

5000

6000

7000

8000

9000

Altitude (feet)

Time from Control Release (second)

0 50 100 150 200 250

-20

-10

0

10

20

Pitch (degree)

Time from Control Release (seconds)

250 knot Entry Speed

177 knot Entry Speed
NASA/TM—2015-218715 17 

nose lowered and airspeed increased. As this repeated, the airplane began to settle back to its trim 

condition. When this was performed in the cruise configuration, shown in Figure 17, the nose remained 

high until the airspeed decreased. The airplane oscillated as it did in the previous scenario, but did not 

dampen out. In fact, for the case with the 177 kn entry speed, the airplane began to stall at the top of the 

climb, causing the nose to drop. Once it began to stall, the airplane became unstable, rolled to the left and 

nosed downward toward the ground. For the case with the 250 kn entry speed, the airplane appeared to be 

neutrally stable. The oscillations did not seem to increase nor decrease though altitude was lost with each 

pitch cycle. 

For directional and lateral stability evaluation, the TCM, configured for cruise (no flaps, landing gear 

retracted), was placed in a skidding or uncoordinated turn and the flight controls were released. Per 

regulations, recovery from this condition requires the low wing in the turn to rise and the nose to return to 

the direction of flight. As with longitudinal stability, any combined lateral-directional oscillation must be 

dampened by the airplane whether the primary controls are free or fixed. When tested with the TCM, the 

nose turned back toward the direction of flight, but the low wing remained low. The airplane remained in 

the turn and oscillated in the lateral axis, as can be seen in Figure 18. The airplane had an undamped pitch 

oscillation that was becoming divergent. The TCM became unstable without intervention, although the 

pilot was able to regain control easily. 

It is important to note that the version of the TCM used for this work did not include a feedback flight 

control system. Namely, yoke and pedal inputs were directly mapped to control surface deflections. A 

control and stability augmentation system, permissible under §25.672, would be able to stabilize and 

dampen the oscillatory responses seen in Figure 16 through Figure 18. 

Figure 17.—Response from flight control release after 10° pitch-up (cruise configuration). 

 

0 20 40 60 80 100 120 140

2000

3000

4000

5000

6000

7000

Altitude (feet)

Time from Control Release (second)

0 20 40 60 80 100 120

-40

-20

0

20

40

Pitch (degree)

Time from Control Release (seconds)

250 knot Entry Speed

177 knot Entry Speed
NASA/TM—2015-218715 18 

Figure 18.—Response from flight control release during uncoordinated turn (cruise configuration). 

4.0 Discussion 

FAA-type certification testing was performed on the relevant aspects of the integrated 

TCM/C-MAPSS40k aircraft simulation, to evaluate its appropriateness as a research testbed. Overall the 

model performed well, exceeding most of the requirements easily. In fact, it generally appears to be a 

little overpowered, as evidenced by the ease with which it exceeded the takeoff and go-around climb 

requirements. On the other hand, the rudder seemed to very effectively compensate for the thrust 

imbalance incurred during the engine out tests. Together these observations seem to point to the aircraft 

being underweight for its size. In fact, it was previously stated that the weight is set to the middle of its 

empty to fully loaded range. Setting the weight to the upper limit of this range would certainly make it 

more challenging to meet the certification requirements, although whether it would still exceed the 

requirements is unknown. In order to perform the certification testing at the higher weight, the inertias 

would have to be modified correspondingly, and information regarding this is not provided with the TCM 

simulation. However, it is reasonable to assume that the excess power is due to the reduced aircraft 

weight rather than the engines being too large. 

A somewhat more serious concern is the results of the stability testing. Undamped oscillations 

resulting from a pilot input under certain conditions might make the airframe fail the certification. 

However, the use of flaps counteracts this neutral stability, which, at least for research purposes, provides 

a workaround should testing in that regime be required. Additionally, these dynamics represent the 

inherent stability qualities of the airframe without a feedback control system. Given the slowly divergent 

nature of the responses observed, an appropriately designed control and stability augmentation system 

would improve the aircraft’s stability characteristics. Furthermore, the pilot was able to easily maintain 

control despite these instabilities. 

0 50 100 150 200 250

2000

4000

6000

8000

Altitude (feet)

Time from Control Release (seconds)

0 50 100 150 200 250

-20

0

20

40

Pitch (degree)

Time from Control Release (seconds)

0 50 100 150 200 250

0

20

40

60

Roll (degree)

Time from Control Release (seconds)
NASA/TM—2015-218715 19 

5.0 Conclusions 

An integrated simulation consisting of the Transport Class Model (TCM) twin engine airframe 

simulation and two copies of the Commercial Modular Aero-Propulsion System Simulation 40k 

(C-MAPSS40k) turbofan engine simulation was evaluated against relevant aspects of the Federal 

Aviation Administration’s Code of Federal Regulations (CFR) Part 25, “Airworthiness Standards: 

Transport Category Airplanes.” The objective was to determine if the behavior of the combined model is 

representative enough of an actual aircraft to make it an appropriate testbed for advanced propulsion 

control and integrated flight and propulsion control research. Evaluation against the relevant certification 

criteria demonstrated that the integrated TCM/C-MAPSS40k simulation behaves like a lighter-than-fullweight commercial transport aircraft. Any potential problems exposed are relatively minor and can be 

ignored or worked around for research purposes. Therefore, the combined model is an appropriate testbed 

for use in such research. 

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