NASA/CR–2012-216038

One Engine Inoperative (OEI) and Autorotation for 

Heavy Lift Rotorcraft System

Tom Wood

Bell Helicopter Textron, Inc.

Hurst, Texas

August 2012

Prepared for Ames Research Center under 

Contract No. NAS2-01063, Task Order #15
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NASA/CR–2012-216038

One Engine Inoperative (OEI) and Autorotation for 

Heavy Lift Rotorcraft System

Tom Wood

Bell Helicopter Textron, Inc.

Hurst, Texas

August 2012

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000
Available from:

NASA Center for AeroSpace Information National Technical Information Service

7115 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161

(301) 621-0390 (703) 487-4650
iii 

TABLE OF CONTENTS 

LIST OF FIGURES ............................................................................................................................. iii 

LIST OF TABLES ............................................................................................................................... iii 

NOMENCLATURE ............................................................................................................................ iv 

LIST OF FIGURES 

Figure 1. Normal vertical takeoff. .................................................................................................... 4

Figure 2. Normal vertical landing. .................................................................................................... 5

Figure 3. OEI completed rejected takeoff/balked landing profiles for helipad. ............................... 6

Figure 4. OEI completed landing profile for helipad. ...................................................................... 6

Figure 5. Demonstrated values of power ratio to achieve vertical helipad performance. ................ 7

Figure 6. Current engine rating structure for OEI operation. ........................................................... 7

Figure 7. Autorotation sink rate as a function of disk loading. ........................................................ 8

Figure 8. XV-15 demonstrated windmilling. .................................................................................... 9

Figure 9. XV-15 AEI reconversion. ................................................................................................. 9

Figure 10. XV-15 AEI flare effectiveness. ....................................................................................... 10

LIST OF TABLES 

Table 1. Comparison of Category A Performance Requirement Between Helicopter 

and Tiltrotor ....................................................................................................................... 4
iv 

NOMENCLATURE 

AEI All engines inoperative 

FAA Federal Aviation Administration 

FADEC full authority digital engine control 

HOGE hovering out-of-ground-effect 

LDP landing decision point 

MCP allowable aircraft maximum continuous power operation

OEI one engine inoperative 

rpm revolutions per minute 

TDP takeoff decision point 

Vy velocity for best climb speed 

Vtoss velocity for takeoff safety speed 
1 

ONE ENGINE INOPERATIVE (OEI) AND AUTOROTATION 

FOR HEAVY LIFT ROTORCRAFT SYSTEM 

Tom Wood1

Ames Research Center 

The Federal Aviation Administration (FAA) will certainly require the Heavy Lift Rotorcraft to be 

operated under Category A performance and operations requirements. Because of the weight, no 

operation will be allowed except Category A according to FAA Part 29.1(c). This means that 

anywhere along the flight path, the aircraft must be able to either land safely following an engine 

failure or continue flight. A repeatable flight profile must be developed and executed to ensure that 

the aircraft can be safely landed, or flown away, depending on its location on the flight profile. This 

means that there will be no Height-Velocity testing as is currently required for Part 29 Category B. 

Because all the configurations shown to date are different than existing rotorcraft, each type would 

have to develop their individual requirements under existing special conditions, FAA Part 21.17(b). 

This means the FAA will take the opportunity to negotiate additional requirements or change 

requirements to ensure safety. For example, because the tiltrotor did not fit normal rotorcraft 

category, new rules were negotiated between the applicant and the FAA. As a result of this 

negotiation, performance requirements for Category A were increased. The rules were written in 

terms of guaranteed performance instead of Category A requirements. Detailed discussion will 

follow later. The proposed tiltrotor would likely follow along with the current tiltrotor rules with the 

possibility of increased Category A performance requirements. Compounding with the addition of 

wing and auxiliary thrust to both the tandem and coaxial rotor would result in new special condition 

aircraft. To my knowledge, no compound tandem or compound coaxial rotor has ever been certified 

by the FAA. 

Recent experience in certification of rotorcraft for Category A performance (M430, M412, and 

M427), along with analysis conducted during the design of the first commercial tiltrotor (BA609), 

has provided some parametrics that can be used to estimate the Category A performance capability. 

Many factors can influence these data, but they do represent demonstrated performance. Parametrics 

for the BA609 were developed using man-in-the-loop simulation. NASA also conducted 

considerable simulation efforts along these lines over the past several years. The ground-level 

helipad in confined areas turned out to be the most difficult even though the elevated helipad is very 

demanding due to the surrounding environment. The primary difference is in the height of the 

takeoff decision point (TDP). The elevated helipad can use drop-down height below the helipad 

while ground-level cannot. Examples of the ground-level helipad flight profiles are shown in figures 

1–4. Normal vertical takeoff is shown in figure 1. Here the pilot moves the aircraft up and back 

using a slow rate of climb while maintaining the helipad in his window. This is referred to as the site 

picture. Once the TDP is reached, the pilot accelerates the aircraft to velocity for takeoff safety 

speed (Vtoss) and initiates a climb. At 200 feet above the takeoff surface the pilot accelerates to 

 

1

 Bell Helicopter Textron, Inc., 600 E. Hurst Blvd., Hurst, TX 76053 
2 

velocity for best climb speed (Vy) where he climbs until he is 1,000 feet above the takeoff surface. 

Normal vertical landing flight profile is presented in figure 2. The pilot enters the approach window 

at 200 feet and 30 KIAS while decelerating to the landing decision point (LDP) where he acquires 

the correct site picture out his window. The one engine inoperative (OEI) flight profiles for rejected 

takeoff and balked landing are shown in figure 3. Rejected takeoff can occur anywhere between 

liftoff and the TDP. If power is lost along this flight profile, the pilot must be able to safely land the 

aircraft back on the helipad. For this approach, the TDP and LDP are the same values. If power 

failure occurs at LDP, the pilot must be able to fly away and clear the helipad by 15 feet while 

accelerating to Vtoss. The OEI completed landing flight profile is shown in figure 4. If power failure 

occurs at LDP or later, the pilot can continue the landing to the helipad safely. These flight profiles 

were developed to allow very little, if any, sliding on the helipad following an OEI rejected takeoff 

or OEI completed landing. This approach is considered to be the limiting Category A case and, 

therefore, would become a criteria for use in this conceptual design study. It should be noted that 

these flight profiles must be flown on each takeoff and landing to helipad to ensure safe takeoffs and 

landings. 

The flight data and simulation data were reviewed to develop simplified criteria for use in 

conceptual studies. Category A analysis is generally broken into three segments. The segments are 

takeoff and landing, climb at velocity for takeoff safety speed (Vtoss), and climb at best climb speed 

(Vy). A weight is established for the requirements of each segment at given atmospheric condition 

and appropriate power setting, and then the minimum weight of the three segments becomes the 

limiting weight. The limiting segment for helipad was the landing and takeoff phase for the majority 

of the atmospheric conditions. The data is presented as power ratio versus disk loading as shown in 

figure 5. Power ratio is defined as the OEI power available at the given atmospheric condition 

divided by the hovering out-of-ground-effect (HOGE) power required at the weight for the given 

atmospheric conditions. The OEI power available may be either 30-second or 2.5-minute power. 

From figure 5, the demonstrated performance is in the area of 6–7 pounds per square foot of disk 

loading. Simulation results are presented at 15 pounds per square foot. A linear curve is suggested as 

shown. Scatter exists in the measured data that suggests there are other factors involved. However, 

as an initial criterion, the power ratio appears to be a reasonable approach. For configurations with 

more than two engines, the power ratio would be the sum of the remaining engines’ OEI power 

divided by HOGE power required. For this study, a power ratio of 0.9 is recommended to account 

for the uncertainty resulting from lack of demonstrated helipad data for tiltrotors. 

The one engine inoperative rating structure varies according to the manufacturer and the time frame 

that the engine was certified. The addition of Full Authority Digital Engine Control (FADEC) has 

allowed an improved OEI rating structure such as 30-second OEI power. Engines that Bell has 

certified in aircraft that have Category A performance are summarized in figure 6. The engine rating 

structure is presented in terms of a ratio of OEI power to allowable aircraft maximum continuous 

power operation (MCP). This is in keeping with the NASA study, which uses OEI as a percent of 

MCP. For engines with a 30-second OEI rating, the power ratio varies between 1.3 to 1.4 times 

MCP. A 2-minute or 2.5-minute OEI varies between 1.25 to 1.35 times MCP. A 31-minute OEI 

varies between 1.2 to 1.3 times MCP. It is recommend that engine manufacturers provide a 

30-second OEI rating structure to minimize impact on sizing of MCP for this study. 
3 

As mentioned earlier, the FAA will negotiate different requirements for Special Condition 

certification as opposed to Part 29 rotorcraft. A comparison between helicopter and tiltrotor 

Category A requirements is shown in table 1. For tiltrotors, the takeoff path extends to 1500 feet. 

Rate-of-climb requirements at Vtoss and Vy are increased. Also, an enroute OEI paragraph was added 

to provide the pilot with single-engine climb or descent gradients. For example, if the criteria for this 

study is takeoff at maximum gross weight at 5000 feet on ISA +20 deg C day, then one must 

consider the ability of the aircraft to stay up in the event an engine failure occurs in cruise. Because 

these atmospheric conditions could represent Denver in the summer, stay-up capability OEI in the 

order of 12,000 feet would be desirable to miss the mountains by at least 1,000 feet. 

The question of autorotation in the nonconventional rotorcraft that are being considered in this study 

needs to be discussed. The FAA requires a full auto touchdown from cruise for all Part 29 

helicopters according to FAA Part 29.79(b). The addition of the wing for all configurations along 

with high speed complicates the issue and requires special attention to evaluate. The FAA has taken 

the position to date that autorotation must be demonstrated in tiltrotors. This demonstration requires 

a full auto touchdown at maximum gross weight at sea level standard from a straight-in approach, 

which is the same requirement for helicopters (TR.79 (b)). The use of a large wing on all the 

configurations complicates the control of the rotor rpm during descent due to the load sharing 

between the rotor and the wing. The rate of descent that must be overcome in all engine inoperative 

(AEI) is a function of disk loading as shown in figure 7. Here a clear trend can be observed between 

the measured sink rate for autorotation and disk loading. For configurations that use four engines, 

the position that full auto touchdown should not have to be demonstrated should be negotiated with 

the FAA. It should be debated that a demonstration of flare effectiveness AEI followed by power 

recovery would be adequate. However, AEI controllability should have to be demonstrated. 

Autorotation was investigated using the XV-15. Figure 7 shows the value of the sink rate required 

for autorotation for the XV-15. In airplane mode, the proprotors can windmill. Results of 

windmilling of the proprotors for the XV-15 are shown in figure 8. However, the proprotors could 

not be stopped from turning in order to drive the generators on the transmission. An AEI 

reconversion was flown as shown in figure 9. Finally an autorotative flare was flown to demonstrate 

that the autorotative sink rate could be arrested as presented in figure 10. 

In summary, a power ratio (OEI power/HOGE power) of 0.9 is recommended for this study to 

provide ground-level-helipad Category A performance. It is recommend that the engine 

manufacturers provide an engine rating structure that provides 30-second OEI power. Attention 

should be paid to the OEI power requirements because this aircraft will only be allowed to operate 

Category A. Recognize that the FAA will negotiate higher Category A requirements than are 

currently in Part 29 during development of special condition rules. An attempt should be made to 

negotiate with FAA to demonstrate flare effectiveness followed by power recovery instead of a full 

auto touchdown. 
4 

TABLE 1. COMPARISON OF CATEGORY A PERFORMANCE REQUIREMENT BETWEEN 

HELICOPTER AND TILTROTOR 

Description Part 29 Helicopter Part TR Tiltrotor 

Takeoff Path 29.59 1000 ft TR.59 1500 ft

VTOSS Altitude 29.59(b) 200 ft TR.59 (b) 400 ft 

VY Altitude 29.59(c) 1000 ft TR.59(c) 1500 ft 

OEI Climb at VTOSS 29.67(b) 100 fpm TR.67 (b) 200 fpm

OEI Climb at VY 29.67(c) 150 fpm TR.67(c) 1.20% 

Enroute Flight Paths n/a n/a TR.69 (b) Show gradient 

Figure 1. Normal vertical takeoff. 
5 

Figure 2. Normal vertical landing. 
6 

Figure 3. OEI completed rejected takeoff/balked landing profiles for helipad. 

Figure 4. OEI completed landing profile for helipad. 
7 

Figure 5. Demonstrated values of power ratio to achieve vertical helipad performance. 

Figure 6. Current engine rating structure for OEI operation. 

4 6 8 10 12 14 16

Disk Loading - Lb/Ft^2

0.5

0.6

0.7

0.8

0.9

1

Power Ratio -

OEI Power / HOGE Power

CERTIFICATION TEST DATA

PILOTED SIMUALTION

1

1.05

1.1

1.15

1.2

1.25

1.3

1.35

1.4

1.45

1.5

#1 #2 #3 #4 #5 #6 #7

POWER RATING/MCP

TAKEOFF/MCP

30 SEC OEI/MCP

2/2.5 MIN OEI/MCP

30 MIN OEI/MCP
8 

Figure 7. Autorotation sink rate as a function of disk loading.

0 4 8 12 16 20 24

DISK LOADING - LB/FT^2

-6000

-5000

-4000

-3000

-2000

-1000

0

RATE OF DESCENT - FPM

M214ST

YAH-63

YUH-60

YCH-53

CH-47D

XV-15

V-22

OH-58C

AH-IJ

YUH-ID

MEASURED DATA

AUTODL.GRF
9 

Figure 8. XV-15 demonstrated windmilling. 

Figure 9. XV-15 AEI reconversion.

0

1000

2000

3000

4000

5000

6000

100 120 140 160 180 200

Calibrated Airspeed (knots)

Rate of Descent (fpm)

XV15 98% RPM

XV15 85% RPM

S/N 702 TEST 340 

GW = 13576 LB 

ALT = 7000 - 4000 FT

0

1000

2000

3000

4000

5000

6000

100 120 140 160 180 200

Calibrated Airspeed (knots)

Rate of Descent (fpm)

XV15 98% RPM

XV15 85% RPM

S/N 702 TEST 340 

GW = 13576 LB 

ALT = 7000 - 4000 FT

PYLON

ANGLE

(DEG)

TRUE

AIRSPEED

(KN)

TIME (SECONDS)

100

0

50

160

120

80

0 30

SHIP 702

FLIGHT 339:2556

GW = 13,576 LB

CG = 200.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS

PYLON

ANGLE

(DEG)

TRUE

AIRSPEED

(KN)

TIME (SECONDS)

100

0

50

160

120

80

0 30

SHIP 702

FLIGHT 339:2556

GW = 13,576 LB

CG = 200.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS
10 

Figure 9. (cont.) 

Figure 10. XV-15 AEI flare effectiveness. 

SHIP 702

FLIGHT 339:2556

GW = 13,576 LB

CG = 200.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS

SHIP 702

FLIGHT 339:2556

GW = 13,576 LB

CG = 200.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS

ROTOR

SPEED

(RPM)

650

550

450

20

0

–20

120

80

40

2000

–2000

–6000

0 30

TIME (SECONDS)

PITCH

ATTITUDE

(DEG)

TRUE

AIRSPEED

(KN)

RATE OF

CLIMB

(FT/MIN)

SHIP 702

FLIGHT 340A:2584

GW = 13,576 LB

CG = 299.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS

ROTOR

SPEED

(RPM)

650

550

450

20

0

–20

120

80

40

2000

–2000

–6000

0 30

TIME (SECONDS)

PITCH

ATTITUDE

(DEG)

TRUE

AIRSPEED

(KN)

RATE OF

CLIMB

(FT/MIN)

SHIP 702

FLIGHT 340A:2584

GW = 13,576 LB

CG = 299.2 IN

COMPARISON BETWEEN 

MEASUREMENT AND ANALYSIS