Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Mark Robuck and Joseph Wilkerson

The Boeing Company, Philadelphia, Pennsylvania

Carl Nordstrom

Rolls-Royce Corporation, Indianapolis, Indiana

Summary of the Large Civil Tiltrotor (LCTR2)

Engine Gearbox Study

NASA/TM—2010-216908

November 2010
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Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Mark Robuck and Joseph Wilkerson

The Boeing Company, Philadelphia, Pennsylvania

Carl Nordstrom

Rolls-Royce Corporation, Indianapolis, Indiana

Summary of the Large Civil Tiltrotor (LCTR2)

Engine Gearbox Study

NASA/TM—2010-216908

November 2010

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

International Powered Lift Conference

cosponsored by the AIAA, SAE Aerospace, and RAeS

Philadelphia, Pennsylvania, October 5–7, 2010
Acknowledgments

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Mission Directorate, Fundamental Aeronautics Program for supporting this research.

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This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2010-216908 1

Summary of the Large Civil Tiltrotor (LCTR2) 

Engine Gearbox Study 

Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Mark Robuck and Joseph Wilkerson 

The Boeing Company 

Philadelphia, Pennsylvania 19142 

Carl Nordstrom 

Rolls-Royce Corporation 

Indianapolis, Indiana 46206 

Abstract 

In support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project, NASA is 

continuing to study the Large Civil Tiltrotor (LCTR) concept to help define/refine vehicle, system and 

subsystem attributes. These attributes can then be used to define performance requirements and identify 

new or advanced technologies to achieve an operational vehicle class. As part of this goal, NASA 

contracted with The Boeing Company and its subcontractor Rolls-Royce to perform an investigation of 

different combinations of engine and gearbox variability to achieve a maximum of 50 percent rotor tip 

speed reduction from hover to cruise conditions. Previous NASA studies identified the 50 percent rotor 

speed reduction minimized vehicle gross weight and fuel burn. The LCTR2 (LCTR—iteration 2) was the 

contracted study baseline for initial sizing. Rotor tip speed ratios (cruise to hover) of 100, 77, and 

54 percent were analyzed for each combination of engine and gearbox speed reduction to achieve the 

chosen rotor tip speed ratio. Three different engine and gearbox technology levels were assumed; 

commercial off-the-shelf (COTS), entry-in-service (EIS) in 2025 and EIS in 2035. These technology 

levels were applied to determine each particular effect on vehicle gross weight and fuel burn, while other 

vehicle technologies were assumed constant. This report summarizes the work performed that is being put 

together into a comprehensive NASA contractor report. Some background on the LCTR concept and 

baseline vehicle will be given and then a discussion concerning the technical approach utilized. Major 

study assumptions and results will be presented and discussed. Finally conclusions will be drawn as well 

as suggestions provided for future efforts. 

Introduction 

The NASA Heavy Lift Rotorcraft System Investigation (Ref. 1) identified a large tiltrotor as the best 

concept to meet commercial airspace requirements for the future, short-haul regional market. Through 

further analysis and refinement, the notional tiltrotor evolved into a conceptual vehicle designated as 

LCTR2 (Large Civil Tiltrotor—iteration 2) (Ref. 2) as seen in Figure 1. 

This vehicle was designed to carry 90 passengers at 300 knots with at least a 1,000 nautical mile (nm) 

range; powered by four turboshaft engines rated at 7,500 shaft horsepower (SHP) each (at Sea level Static 

(SLS) conditions). Other design features included a rotor tip speed of 650 feet per second (fps) in hover 

and 350 fps during cruise, enabled by a two-speed gearbox. This range of rotor tip speeds was needed to 

achieve the high level of performance and efficiency at two very different flight conditions for the 

reference design mission shown in Figure 2. Additional vehicle and mission details can be found in 

Reference 2. 
NASA/TM—2010-216908 2

Figure 1.—Conceptual view of LCTR2. 

Figure 2.—LCTR2 design reference mission. 

The LCTR2 currently serves as a representative vehicle and mission for the Fundamental Aeronautics 

Program (FAP) Subsonic Rotary Wing (SRW) project to further define performance levels required for 

these vehicle concepts to be practical. Gaps between these study performance levels and the state-of-theart would be used to identify research opportunities. Further studies were envisioned to delve into 

additional detail to identify specific research areas. A study (Ref. 3) was conducted to evaluate the net 

benefits of advanced technologies for two conceptual civil transport rotorcraft (single-main rotor 

compounds and civil tiltrotor), with different mission cruise speeds and payloads from the LCTR2. 

Engine technology was identified as giving more benefits to direct operating cost per available seat-mile 

than other technology groups included. This further motivated NASA to study engine and gearbox 

combinations and technologies to help guide engine and drive system research to maximize the gain from 

its technology investment. The subject of this report is to summarize the NASA-contracted effort 

(NNA06BC41C, Task Order 10) performed by The Boeing Company and Rolls-Royce Corporation to 

identify and evaluate engine and drive system concepts that could achieve the almost 50 percent rotor tip 

speed reduction identified in the previous NASA studies. A NASA contractor report (CR) is in process to 

give comprehensive details concerning the background, study methodologies and substantiation, results, 

and conclusions for that effort. 

Taxi, 4 min

Ground check, 

1 min @ 60% 

takeoff power

Climb to 25,000’ cruise 

altitude @ MCP

(25-30 mins)

Takeoff & convert, 2 min

(actual conversion is TBD)

Cruise @ design 

airspeed to mission range

Alternate 

Destination 

30 nm

Transfer Altitude

(conservative

estimate of fuel)

No credit for range.

Final approach to land

+ Reserve Fuel

30 min @ 10,000’, ISA

≥ 1000 nm @ ≈ 310 knots

(about 3 hours)

Vertical landing, 1 min

@ 5,000’, ISA+20C
NASA/TM—2010-216908 3

Study Overview 

Per the original statement of work for the contract: “The main emphasis of this effort is to identify the 

engine and gearbox/transmission technology barriers/challenges/needs for achieving the 50 percent rotor 

tip speed variation with a fixed rotor diameter, vehicle, and mission.” The contracted effort was divided 

into 6 tasks (5 engineering tasks and one for reporting and administration): 

 Task 1 is an evaluation of LCTR2 vehicle sizing and performance characteristics assuming COTS 

engine and drive system technologies. This task includes development of sizing methodology, 

baseline vehicle and system characteristics as well as initial sizing for 54 percent rotor cruise tip 

speeds. 

 Task 2 is executed by Rolls-Royce team members to generate engine information for advanced 

technology engines (EIS 2025 and 2035) that will be used in the performance analysis tasks that 

follow. 

 Task 3 generates information for the drive system configurations and performance (EIS 2025 and 

2035). 

 Task 4 is comprised of analysis tasks conducted by Boeing and Rolls-Royce to evaluate benefits 

for advanced technologies and operational scenarios conducted with 54 percent rotor cruise tip 

speeds. The primary criteria used to evaluate performance will be system weight and overall fuel 

burn. 

 Task 5 identifies technology challenges and needs for various parts of the overall system that will 

be addressed through subsequent research and development efforts. 

 Task 6 is for reporting and administration related efforts. 

To complement the NASA-derived cruise to hover rotor tip speed ratio of 54 percent, two additional 

cruise rotor tip speed ratios were included to get preliminary indications of system effects (changes in 

fuel, component and vehicle weights and performance, especially noting those changes in the engine and 

drive systems) that might offset the improvements in rotor efficiency achieved through the rotor cruise tip 

speed reduction. Finally, the rotor cruise tip speed reduction could be achieved through a multispeed 

gearbox alone, a gas turbine engine with a large operating speed range power turbine alone (requiring 

only a single-speed gearbox), or a combination of engine and multispeed gearbox technologies. The 

combination of parameters used in this study is shown in Table 1. 

TABLE 1.—ENGINE-GEARBOX STUDY MATRIX 

Rotor design cruise tip speed, 

percent 

Engine cruise/Normal rpm, 

percent 

Drive system cruise/Hover rpm, 

percent 

650 fps (100%) Reference 100 100% 

500 fps (77%) Mid-value 100 77% (2-speed) 

77 100% 

350 fps (54%) NASA goal 

100 54% (2-speed) 

77 70% (2-speed) 

54 100% 
NASA/TM—2010-216908 4

Analysis Methodology and Component Results 

Analysis tools for vehicle sizing and performance evaluation were devised from existing spreadsheetbased tools. These tools included many of the performance and sizing procedures from VASCOMP 

(Refs. 4 and 5) and were designed to facilitate “concept evaluation” over the matrix of engine-gearbox 

combinations mentioned previously for the three engine and drive system technology levels evaluated 

(COTS, EIS 2025, and EIS 2035). Some of the major assumptions/methodology used for the study are 

given below: 

 Resize vehicle with each engine/transmission/rotor tip speed to meet mission range and payload 

requirements. 

 Fix vehicle system technology levels at 2025 technology, to isolate effect of engine/drive system 

technology changes. 

 Maintain engine and drive systems at similar technology levels (COTS, 2025 or 2035). 

 Scale engine size to meet most stringent requirement of hover, one-engine inoperative (OEI), or 

cruise power levels (generally hover requirements set engine power). 

 Size drive system for maximum torque condition for either hover or cruise, recognizing that this 

could limit torque at other flight conditions. 

 Assume rotor speed reduction occurs right after transition from hover to climb/cruise phase. This 

assumption resulted in torque-limited climb power for 500 and 350 fps cases. This torque-limit 

could have been alleviated if the study assumed rotor speed reductions occurred later in flight. 

Review of results determined that this limit would only have minimal effects on the overall 

vehicle fuel burn or mission time (since the 1000 nm mission is cruise-dominated). 

Engine Performance 

The NASA, Boeing and Rolls-Royce team defined the engine technology strategy that would be used 

by Rolls-Royce to develop engine models and provide scalable engine data consistent with technology for 

COTS, EIS 2025, and EIS 2035. The COTS engine was based on a conventional (single-spool), all-axial 

turbofan core modified to a turboshaft engine (the power turbine would be on its own shaft), with an 

overall pressure ratio equivalent to current engines. A representative image of the COTS engine is given 

in Figure 3. Engine performance is shown in Figure 4 assuming a conventional power turbine, optimized 

for a limited operating speed range around 95 percent (typical for conventional aircraft mission and duty 

cycle). It shows some power and efficiency losses for power turbine operation at 77 percent speed (which 

were the most significant at high power levels). At 54 percent speed, there were significant losses in 

maximum power available as well as fuel efficiency, the subsequent effects on vehicle sizing will be 

shown later. 

Figure 3.—Representative image of the COTS engine. 
NASA/TM—2010-216908 5

Figure 4.—COTS engine performance. 

Figure 5.—2025 engine performance with wide speed range power 

turbine optimized for 77 percent rpm. 

For the EIS 2025 technology engine, it was assumed that a significantly improved, new engine core 

would not be available. Enhancements to the core engine would be limited to material and cooling 

updates applied to the COTS engine and the core would be similar to the COTS engine already shown. 

With a standard power turbine, performance would be similar to the COTS engine, although there would 

be additional margins available that could be used to improve maintainability or some additional power 

capability. The EIS 2025 timeframe would be sufficient for incorporation of a wide-speed range power 

turbine (employing variable geometry), which could be used to improve performance at reduced power 

turbine speeds. Since the LCTR2 mission is cruise dominated, power turbine design was optimized 

around the 77 percent speed point (which would improve cruise performance at reduced engine speed). 

Engine performance for the EIS 2025 engine with a wide speed range power turbine is shown in Figure 5. 

The wide-speed range turbine gave excellent performance at its design speed and almost as good at 

54 percent rpm. Even with variable geometry, performance fell off for the 100 percent rpm condition. The 

advances in materials and cooling assumed available for the core would give some increased capability in 

turbine inlet temperature, which was used to maintain power for the hover condition, to prevent over 

sizing the 2025 engine and further penalizing the vehicle. 

COTS Engine Technology

Shaft Horsepower

Absolute Fuel Flow, lb/hr

100% Engine RPM

77% Engine RPM

54% Engine RPM

2025 Engine Technology

Shaft Horsepower

Absolute Fuel Flow, lb/hr

100% Engine RPM

77% Engine RPM

54% Engine RPM
NASA/TM—2010-216908 6

For the EIS 2035 engine, it was assumed that a new, advanced core would be available. For this 

advanced core, it was further assumed that compression would be accomplished in two spools, increasing 

maximum temperatures and pressures present in the cycle (versus the COTS and 2025 engines), and 

significantly improving engine efficiency and weight. To maintain compressor performance at low exit 

corrected flows, the high pressure compressor was assumed to be an axi-centrifugal design. It also 

included a wide-speed range power turbine (three-spool engine overall), to maximize fuel efficiency at an 

engine speed of 77 percent. A representative image of this engine is shown in Figure 6. Engine 

performance for the EIS 2035 engine with a wide speed range power turbine is shown in Figure 7. The 

advanced core combined with the wide speed range power turbine was able to achieve high power and 

efficiency over the entire rpm operating range, minimizing those losses for operation at 100 percent rpm 

exhibited in the EIS 2025 engine results. 

Figure 6.—Representative image of the EIS 2035 engine. 

Figure 7.—2035 engine performance with wide speed range 

power turbine optimized for 77 percent rpm. 

2035 Engine Technology

Shaft Horsepower

Absolute Fuel Flow, lb/hr

100% Engine RPM

77% Engine RPM

54% Engine RPM
NASA/TM—2010-216908 7

Engine Weight 

The engines were scaled to meet maximum power required over the mission (which generally was for 

the hover condition). The addition of the wide-speed range power turbine incurred a significant weight 

increase for the 2025 and 2035 engines; preliminary estimates put the additional dry weight at 200 and 

150 lb, respectively, per unscaled engine. Engine weights at the unscaled power level are shown in 

Table 2. 

Drive System 

The NASA LCTR2 vehicle configuration parameters and mission specifics were used to develop 

configuration data and concepts for the integrated engine and drive systems used in the study. As a 

Tiltrotor vehicle, the LCTR2 general arrangement is similar to the V22 Osprey drive system. The LCTR2 

configuration has evolved to a high wing, tilting nacelle aircraft like the V22 in many respects except with 

4 engines, 2 engines at each nacelle. The nominal drive system architecture (using helical gears for idlers 

and bull gear) was devised after consideration of many alternatives and a block diagram is shown in 

Figure 8. This diagram represents a (single speed ratio) direct drive configuration (rotor speed variations 

are achieved through engine speed variation). 

Speed reduction capability for the drive systems were based on compound planetary systems. The 

speed is changed by restraining an element of the planetary system, either a ring gear or carrier with a 

(multiple disk) clutch, causing the gear ratios to change. Figure 9 shows a schematic arrangement 

“Configuration B” that proved favorable for weight and operating characteristics to the extent examined 

in this study. The speed change modules are incorporated into the drive system as shown in the block 

diagram of Figure 10. 

TABLE 2.—UNSCALED ENGINE WEIGHTS 

Engine Installed SHP 

(Maximum rated power at sea level, static) 

Engine dry weight, 

lb 

COTS 8100 1,356 

2025 8088 1,556 (1,356+200) 

2035 8088 1,020 (870+150) 

Figure 8.—Drive system block diagram for the reference LCTR2 vehicle configuration. 
NASA/TM—2010-216908 8

Figure 9.—Speed changing compound planetary schematic. 

Figure 10.—Drive system block diagram with speed changing modules. 

191 rpm 191 rpm

3.1 : 1

3 : 1

2.5 : 1

3.1 : 1

3 : 1

2.5 : 1

1.54 :1 8,000 rpm 1.54 :1 

5,194 rpm 5,194 rpm

MID-WING 

GEARBOX FOR 

ACCY POWER

TILT AXIS GEARBOX

GEN

PUMP

2.4 : 1

or

4.4 : 1 

15,000 rpm 15,000 rpm

2.4 : 1

or

4.4 : 1 

15,000 rpm 15,000 rpm

PROP ROTOR GEARBOX

3.5:1 3.5:1

Same engine configuration 

as shown on the left, but 

out-of-plane 

Same engine configuration 

as shown on the right, but 

out-of-plane 

Speed 

Changing 

Module
NASA/TM—2010-216908 9

From these configurations, estimates were made for weights and power losses for all combinations of 

engines and gearbox speed changes to meet the desired rotor cruise tip speeds. To ease comparisons, 

weight data is given assuming 7,500 hp at the lowest operating speed. This weight information is 

subsequently scaled to the appropriate power within the vehicle sizing program. Weight trends for current 

production and applying a weight technology factor of 0.8 (20 percent weight reduction) for COTS 

technology is given in Table 3, along with the xcruise speed percentage power loss. For the EIS 2025 

technology, it was assumed the weight technology factor would be 0.75 (25 percent weight reduction 

from current production) and a 5 percent reduction in cruise power losses from COTS levels. For the 

EIS 2035, the weight technology factor was estimated to be 0.70 (30 percent weight reduction from 

current production) and cruise power losses would be 10 percent less than COTS levels. The resulting 

weights and percentage cruise power losses are given in Table 4. It can be seen that reducing rotor speeds 

increases the drive system weight, due to the increase in maximum torque or mission time at high torque 

levels. For multispeed gearboxes, increasing the amount of speed reduction achieved through the gearbox 

also increases the drive system weight. 

TABLE 3.—COTS DRIVE SYSTEM WEIGHTS AND CRUISE POWER LOSSES 

Configuration Speed, 

percent 

Engine, 

percent 

Drive, 

percent 

Rotor, 

rpm 

Tip 

speed, 

fps 

Trend weight, 

pounds, 

Current 

production 

COTS weight, 

pounds, 

Technology 

factor 0.8 

COTS power 

loss at cruise 

speed, 

percent 

1 100 100 100 191.0 650.0 11,236 8989 4.10 

2B 100 100 100 191.0 650.0 11,758 9406 4.70 

2B 77 100 77 147.1 500.5 11,758 9406 4.35 

1 77 77 100 147.1 500.5 11,236 8989 3.85 

2B 77 77 100 147.1 500.5 11,872 9497 4.35 

3B 54 100 54 103.1 351.0 12,086 9669 3.90 

1 54 54 100 103.1 351.0 11,236 8989 3.40 

2B 53.9 77 70 102.9 350.4 11,872 9497 3.80 

TABLE 4.—2025 AND 2035 DRIVE SYSTEM WEIGHTS AND CRUISE POWER LOSSES 

Configuration Speed, 

percent 

Engine, 

percent 

Drive, 

percent 

Rotor, 

rpm 

2025 weight, 

pounds, 

Technology 

factor 0.75 

2035 weight, 

pounds, 

Technology 

factor 0.70 

2025 power 

loss, percent 

Technology 

factor 0.95 

2035 power 

loss, percent 

Technology 

factor 0.90 

1 100 100 100 191.0 8427 7866 3.90 3.69 

2B 100 100 100 191.0 8819 8231 4.47 4.23 

2B 77 100 77 147.1 8819 8231 4.13 3.92 

1 77 77 100 147.1 8427 7866 3.66 3.47 

2B 77 77 100 147.1 8904 8310 4.13 3.92 

3B 54 100 54 103.1 9065 8460 3.71 3.51 

1 54 54 100 103.1 8427 7866 3.23 3.06 

2B 53.9 77 70 102.9 8904 8310 3.61 3.42 

 
NASA/TM—2010-216908 10

Rotor Performance 

Although NASA has performed extensive studies to refine the LCTR2 rotor system design, Boeing 

applied NASA descriptions of the LCTR rotor to perform an independent, in-house estimate of rotor 

hover and cruise performance. The NASA geometric twist distribution was maintained for the 350 fps 

cruise tip speed, but was modified for the two additional cruise tip speeds (500 and 650 fps) with the goal 

of locally aligning the blade with oncoming flow at 300 knots for vehicle cruise. Resource constraints 

permitted no blade optimization to further refine the resulting blade twist distributions for cruise or to 

balance the design for hover performance. A comparison of rotor blade twist distributions is shown in 

Figure 11 and compared to the helical inflow angles for each rotor, operating at 310 knots. Boeing applied 

a bi-linear twist distribution for the 500 fps tip speed, similar to the NASA twist parameterization. A 

bi-linear twist distribution proved to be inadequate to properly align the blade for the 650 fps cruise tip 

speed and a tri-linear twist was used instead. Maps of rotor cruise efficiency from the Boeing B08 

(Refs. 6 and 7) analysis are presented Figures 12 to 14. 

Figure 11.—Comparison of rotor blade twist distributions. 

-10

-5

0

5

10

15

20

25

30

35

40

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

Dimensionless Blade Radius, r/R

Local Blade Twist Angle, deg

NASA LCTR2 Twist (350 fps Cruise Tip Speed)

Boeing Twist for 500 fps Cruise Tip Speed

Boeing Twist for 650 fps Cruise Tip Speed

Dashed Lines: Radial distribution 

of helical inflow angle at cruise 

flight speed = arctan (/x)
NASA/TM—2010-216908 11

Figure 12.—Rotor cruise propulsive efficiency, 650 fps cruise tip speed design. 

Figure 13.—Rotor cruise propulsive efficiency, 500 fps cruise tip speed design. 

Cruise

Propulsive

Efficiency

PROP = CT/CP

RCRUISE = 650 ft/s 

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

(x 104)

Advance

Ratio 

0.60 0.40 0.50

0.90

0.70

0.80

0.85

24

40

20

22

28

26

30

32 34 36 Cruise 38

Propulsive

Efficiency

PROP = CT/CP

RCRUISE = 650 ft/s 

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

(x 104)

Advance

Ratio 

0.60 0.40 0.50

0.90

0.70

0.80

0.85

24

40

20

22

28

26

30

32 34 36 38

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

(x 104)

Advance

Ratio 

0.60 0.40 0.50

0.90

0.70

0.80

0.85

24

40

20

22

28

26

30

32 34 36 38

Cruise

Propulsive

Efficiency

PROP = CT/CP

RCRUISE = 500 ft/s 

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.9

1.0

1.1

1.2

30

0.5 0.7

1.3

34

38

42

46

50

54

58

62

70 66

Thrust

Coefficient CT

(x 104)

Advance

Ratio 

Cruise

Propulsive

Efficiency

PROP = CT/CP

RCRUISE = 500 ft/s 

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.9

1.0

1.1

1.2

30

0.5 0.7

1.3

34

38

42

46

50

54

58

62

70 66

Thrust

Coefficient CT

(x 104)

Advance

Ratio 
NASA/TM—2010-216908 12

Figure 14.—Rotor cruise propulsive efficiency, 350 fps cruise tip speed design. 

Results and Discussion 

The NASA LCTR2 vehicle and mission was then modeled using the estimates from Boeing for all 

weights (except engine) and performance for the vehicle, rotor and drive systems. Estimates for engine 

weights and performance were supplied by Rolls-Royce and scaled by Boeing as previously mentioned. 

Overall vehicle results were generated assuming COTS, EIS 2025, and EIS 2035 technology for both the 

engine and drive systems (while maintaining other vehicle systems technology levels fixed) to help isolate 

engine and drive system technology effects as previously mentioned. 

COTS Engine and Drive System Technology Results 

A summary of sizing results assuming COTS engine and drive system technology is given in Table 5. 

The minimum gross weights and fuel weights are very similar. For the mission requirements (speed, 

range, payload, etc.) rotor cruise tip speed of 500 fps resulted in the minimum gross weight and fuel 

weight cases; results were similar whether the speed reduction was achieved only from the engine or a 

multispeed gearbox. As shown in Figure 15, the further reduction in rotor cruise tip speed from 500 to 

350 fps results in minimal improvement in rotor propulsive efficiency, while incurring penalties in engine 

SFC and gearbox and rotor weights evident from Table 5. This may point to a deficiency to capture the 

interaction between the rotor propulsive efficiency and vehicle aerodynamics (beyond contract scope for 

this effort) that was recently reported in Reference 8. At rotor cruise tips speeds of 650 fps, lower drive 

system weight (higher rpm yields lower torque and therefore lower weight per horsepower) does not 

overcome rotor propulsive efficiency (which shows up in fuel weight). Conversely at 350 fps rotor cruise 

tip speed, achieved through the engine alone, the significant loss in engine power and efficiency at the 

low 54 percent engine rpm shown earlier resulted in cruise power requirements determining the engine 

size. This increased the engine power requirement and weight, further penalized by poorer engine specific 

fuel consumption at cruise, resulting in the highest gross and fuel weights. 

 

Cruise

Propulsive

Efficiency

PROP = CT/CP

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

(x104)

Advance

Ratio 

0.8 1.0

1.3

1.4

1.5

1.6

1.7

1.8

1.2

50

60

70

80

90

100

110

120 130 140

ΩRCRUISE = 350 ft/s

Cruise

Propulsive

Efficiency

PROP = CT/CP

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

(x104)

Advance

Ratio 

0.8 1.0

1.3

1.4

1.5

1.6

1.7

1.8

1.2

50

60

70

80

90

100

110

120 130 140

ΩRCRUISE = 350 ft/s
NASA/TM—2010-216908 13

TABLE 5.—SUMMARY OF LCTR2 SIZING WITH COTS ENGINE AND DRIVE SYSTEM TECHNOLOGY 

Rotor cruise tip speed 

Engine cruise/Hover rpm 

Drive system cruise/Hover rpm 

Drive system type 

350 

100% 

54% 

2-speed 

350 

77% 

70% 

2-speed 

350 

54% 

100% 

1-speed 

500 

100% 

77% 

2-speed 

500 

77% 

100% 

1-speed 

650 

100% 

100% 

1-speed 

Gross weight, lbm 108,325 b107,882 110,571 b106,132 b105,687 108,569 

Fuel, lbm b16,710 16,882 18,628 b16,624 b16,767 18,141 

SHP 5186 5159 a5521 5168 5120 5278 

Rotor weight, lbm 9529 9477 9641 9049 9011 9261 

Drive system weight, lbm 9640 9131 8712 8296 7857 8138 

a

Engine sized by cruise requirement 

b

Minimum values for gross and fuel weight 

Figure 15.—Rotor propulsive efficiency for 350 and 500 fps cruise tip speed designs. 

2025 Engine and Drive System Technology Results 

A summary of sizing results assuming EIS 2025 engine and drive system technology is given in 

Table 6. The minimum gross weights and fuel weights are again similar. The weight penalty for the widespeed range power turbine generally resulted in heavier vehicles using more fuel than with COTS 

technology (the exception being the 350 fps rotor cruise tip speed, single-speed gearbox and engine at 

54 percent cruise to hover rpm ratio). Results favor the single-speed gearbox at 350 and 500 fps rotor 

cruise tip speed. This makes sense for the assumptions made; the vehicle has already taken a weight 

penalty to achieve more efficient engine operation at reduced rpm (wide-speed range power turbine). 

Operation at 100 percent power turbine rpm at cruise while using a two-speed gearbox resulted in an 

additional penalty in engine fuel efficiency, which makes those cases even worse than for the COTS 

technology. As seen for the COTS results, reduced rotor cruise tip speeds also increased rotor and drive 

system weights (per shaft horsepower), although principally at 350 fps. 

RCRUISE = 350 ft/s (54% rpm)

Cruise

Propulsive

Efficiency

PROP = CT/CP

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

Thrust

Coefficient CT

Advance

Ratio 

0.80.91.01.11.21.3

1.4

1.5

1.6

1.7

1.8

0.0050

0.0055

0.0060

0.0065

0.0070

0.0075

0.0090

0.0100

0.0080 0.0085

0.0095

Bi-linear twist (-38°/-30°)

310 ktas

0.0106

Extended to

Ct=0.014

0.40

0.45 Thrust

Coefficient CT

Advance

Ratio 

0.8 0.9

1.0

1.1

1.2

1.3

0.0030

0.0032

0.0034

0.5 0.6 0.7

0.0036

0.00380.0040 0.00420.0044 0.00460.0048 0.0050

Bi-linear twist (-50°/-34°)

310 ktas

0.0054

Extended to 

Ct=0.0066 

RCRUISE = 500 ft/s (77% rpm) 
NASA/TM—2010-216908 14

TABLE 6.—SUMMARY OF LCTR2 SIZING WITH 2025 ENGINE AND DRIVE SYSTEM TECHNOLOGY 

Rotor cruise tip speed 

Engine cruise/Hover rpm 

Drive system cruise/Hover 

rpm 

Drive system type 

350 

100% 

54% 

2-speed 

350 

77% 

70% 

2-speed 

350 

54% 

100% 

1-speed 

500 

100% 

77% 

2-speed 

500 

77% 

100% 

1-speed 

650 

100% 

100% 

1-speed 

TOGW, lbm 115,017 b

110,350 b108,567 113,105 b107,985 114,716 

Fuel, lbm 19,459 b

17,161 b16,647 19,363 b16,994 20,832 

SHP 5462 5236 5132 5449 5179 a

5564 

Rotor weight, lbm 10023 9692 9548 9625 9188 9773 

Drive system weight, lbm 9474 8776 8146 8336 7507 8108 

a

Engine sized by cruise requirement 

b

Minimum values for gross and fuel weight

TABLE 7.—SUMMARY OF LCTR2 SIZING WITH 2035 ENGINE AND DRIVE SYSTEM TECHNOLOGY 

Rotor cruise tip speed 

Engine cruise/Hover rpm 

Drive system 

cruise/Hover rpm 

Drive system type 

350 

100% 

54% 

2-speed 

350 

77% 

70% 

2-speed 

350 

54% 

100% 

1-speed 

500 

100% 

77% 

2-speed 

500 

77% 

100% 

1-speed 

650 

100% 

100% 

1-speed 

TOGW, lbm 97,540 a95,676 a95,014 95,734 a93,463 106,397 

Fuel, lbm 13,452 a12,600 a12,633 13,404 a12,484 19,576 

SHP 4664 4572 4523 4646 4516 5121 

Rotor weight, lbm 8636 8502 8430 8163 7969 9061 

Drive system weight, lbm 7538 7098 6619 6464 5915 6902 

a

Minimum values for gross and fuel weight 

2035 Engine and Drive System Technology Results 

A summary of sizing results assuming EIS 2035 engine and drive system technology is given in 

Table 7. As before, the minimum gross weights and fuel weights are similar. The advanced 2035 engine 

and drive system technologies resulted in roughly 25 percent reduction in mission fuel and almost 

12 percent reduction in vehicle gross weight from COTS technology. The case with 500 fps rotor cruise 

tip speed and single-speed gearbox had both the minimum gross and fuel weights. It is also important to 

note that the other 500 fps case (with the two-speed gearbox) just misses being one of the three best cases 

with respect to gross weight. For this case, the engine is not operating at its optimum cruise rpm for fuel 

efficiency and includes a weight penalty of 150 lb per unscaled engine for a wide-speed range power 

turbine that it is not using effectively. It is reasonable to assume that with a lighter power turbine designed 

for optimum performance at 100 percent rpm that it would be among the best in gross weight and 

possibly fuel weight as well. Reduced rotor cruise tip speeds also increased rotor and drive system 

weights (per shaft horsepower) for this technology level as well. 
NASA/TM—2010-216908 15

Conclusions 

NASA contracted with The Boeing Company and Rolls-Royce to perform an investigation of 

different combinations of engine and gearbox variability to achieve a maximum of 50 percent rotor tip 

speed reduction from hover to cruise conditions. Advanced (EIS 2035) engine and drive system 

technology yielded an impressive 25 percent reduction in fuel weights and a 12 percent reduction in gross 

vehicle weight versus COTS technology. The 500 fps rotor cruise tip speed resulted in many of the 

minimum gross weights and fuel weights, although several of the 350 fps rotor cruise tip speed cases 

yielded similar results. For these minimum gross weights and fuel weights, results did not clearly favor 

the engine versus the two-speed gearbox approach to achieve the rotor cruise tip speed reduction. The 

reduction in rotor cruise tip speed from 500 to 350 fps resulted in increases in rotor and drive system 

weights that were not offset by improved rotor performance. A recent report (Ref. 8) has noted a 

favorable interaction between slowing the rotors and vehicle aerodynamics, but these analyses and effect 

were beyond the scope of this contract. Adding a wide-speed range power turbine to the engine, 

optimized for engine operation at 77 percent rpm, significantly improved the cases that depended on the 

engine alone to get the rotor cruise tip speed reduction to 54 percent. For other cases, the engine weight 

penalty (150 to 200 lb per unscaled engine) and minimal fuel efficiency improvement from a standard 

power turbine at 77 percent rpm (and a penalty at 100 percent engine rpm) significantly penalized those 

solutions. There are efforts underway to revisit the EIS 2035 cases using an engine with a “standard” 

power turbine. Once performed, results should give a clearer idea of the value of wide-speed range power 

turbine technology versus its weight and the multispeed gearbox approach for rotor cruise tip speed 

reduction. Finally, all these efforts are to be included in a comprehensive NASA contractor report, to help 

guide future project research efforts. 

References 

1. Johnson, W., Yamauchi, G.K., and Watts, M.E., “NASA Heavy Lift Rotorcraft Systems 

Investigation,” NASA/TP—2005-213467, September 2005. 

2. Acree, C.W., Hyeonsoo, Y., and Sinsay, J.D., “Performance Optimization of the NASA Large Civil 

Tiltrotor,” International Powered Lift Conference, London, UK, July 22–24, 2008. 

3. Wilkerson, Joseph B., and Smith, Dr. Roger L., “Aircraft System Analysis of Technology Benefits to 

Civil Transport Rotorcraft,” The Boeing Company, NASA/CR—2009-214594, June 2009. 

4. Schoen, A.H., Rosenstein, H., Stanzione, K.A., Wisniewski, J.S., “User’s Manual for 

VASCOMP II, The V/STOL Aircraft Sizing and Performance Computer Program” Prepared by the 

Boeing Vertol Company, D8-0375, 3rd revision, 1980. 

5. Wilkerson, Joseph, “VASCOMP III, The V/STOL Aircraft Sizing and Performance Computer 

Program, User’s Manual,” Boeing Rotorcraft, D210-13635-1, 2002. 

6. Liu, J., and McVeigh, M.A., “Design of Swept Blade Rotors for High-Speed Tiltrotor Application,” 

AIAA 91-3147, AIAA, AHS, and ASEE, Aircraft Design Systems and Operations Meeting, 

Baltimore, Maryland, September 1991. 

7. Liu, J., Paisley, D.J., and Hirsh, J., “Tiltrotor Aerodynamic Blade Design by Numerical Optimization 

Method,” American Helicopter Society 46th Annual Forum, Washington, D.C., May 1990. 

8. Acree, C.W., “Integration of Rotor Aerodynamic Optimization with the Conceptual Design of a Large 

Civil Tiltrotor,” AHS Aeromechanics Conference, San Francisco, CA, January 20–22, 2010. 
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Summary of the Large Civil Tiltrotor (LCTR2) Engine Gearbox Study 

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6. AUTHOR(S)

Snyder, Christopher, A.; Robuck, Mark; Wilkerson, Joseph; Nordstrom, Carl 

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WBS 877868.02.07.03.07.01 

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John H. Glenn Research Center at Lewis Field 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

In support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project, NASA is continuing to study the Large Civil Tiltrotor 

(LCTR) concept to help define/refine vehicle, system and subsystem attributes. These attributes can then be used to define performance 

requirements and identify new or advanced technologies to achieve an operational vehicle class. As part of this goal, NASA contracted with 

The Boeing Company and its subcontractor Rolls-Royce to perform an investigation of different combinations of engine and gearbox 

variability to achieve a maximum of 50 percent rotor tip speed reduction from hover to cruise conditions. Previous NASA studies identified 

the 50 percent rotor speed reduction minimized vehicle gross weight and fuel burn. The LCTR2 (LCTR-iteration 2) was the contracted study 

baseline for initial sizing. Rotor tip speed ratios (cruise to hover) of 100, 77, and 54 percent were analyzed for each combination of engine 

and gearbox speed reduction to achieve the chosen rotor tip speed ratio. Three different engine and gearbox technology levels were 

assumed; commercial off-the-shelf (COTS), entry-in-service (EIS) in 2025 and EIS in 2035. These technology levels were applied to 

determine each particular effect on vehicle gross weight and fuel burn, while other vehicle technologies were assumed constant. This report 

summarizes the work performed that is being put together into a comprehensive NASA contractor report. Some background on the LCTR 

concept and baseline vehicle will be given and then a discussion concerning the technical approach utilized. Major study assumptions and 

results will be presented and discussed. Finally conclusions will be drawn as well as suggestions provided for future efforts.

15. SUBJECT TERMS

Tiltrotor; Vehicle sizing; Gas turbine engine; Gearbox; Mission analysis 

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