Mark Robuck and Joseph Wilkerson

The Boeing Company, Philadelphia, Pennsylvania

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Yiyi Zhang and Bob Maciolek

The Boeing Company, Philadelphia, Pennsylvania

Study and Sub-System Optimization of Propulsion 

and Drive Systems for the Large Civil TiltRotor 

(LCTR2) Rotorcraft

NASA/TM—2013-218103

December 2013
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Mark Robuck and Joseph Wilkerson

The Boeing Company, Philadelphia, Pennsylvania

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Yiyi Zhang and Bob Maciolek

The Boeing Company, Philadelphia, Pennsylvania

Study and Sub-System Optimization of Propulsion 

and Drive Systems for the Large Civil TiltRotor 

(LCTR2) Rotorcraft

NASA/TM—2013-218103

December 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

69th Annual Forum and Technology Display (AHS FORUM 69)

sponsored by the American Helicopter Society

Phoenix, Arizona, May 21–23, 2013
Acknowledgments

The authors would like to thank all contributors and acknowledge the following Boeing, Rolls-Royce Corporation and NASA 

experts for their technical contributions to this assessment of advanced technology on civil transport rotorcraft. Overall Management: 

Ram Janakiram, Mission Analysis and Vehicle Sizing: Joseph Wilkerson, Drive System Technology: Mark Robuck, Yiyi Zhang

Prop-Rotor Performance and Aerodynamics: Robert Maciolek Propulsion Technology: Lawrence Basil and Rolls-Royce team 

members Carl Nordstrom, Dan Vonderwell, Adam Ford, and Craig Heathco Weights Evaluation and Fixed Equipment Assessment: 

Paul Rudisaile. In addition we would like to acknowledge Susan Gorton as the NASA Subsonic Rotary Wing (SRW) lead for 

fi nancial support to complete the project, Christopher Snyder of NASA Glenn Research Center, the NASA project Contract Offi ce 

Technical Representative, for his guidance throughout the project, and Jeffrey Sinsay of U.S. Army Aerofl ightdynamics Directorate 

at NASA Ames Research Center (ARC) for putting together the LCTR2 package. Additional helpful insights, considerations, and 

comments were provided by Wayne Johnson and Wally Acree of NASA ARC.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
ABSTRACT 

In a series of study tasks conducted as a part of NASA’s Fundamental Aeronautics Program, Rotary Wing Project, Boeing 

and Rolls-Royce explored propulsion, drive, and rotor system options for the NASA Large Civil Tilt Rotor (LCTR2) 

concept vehicle. The original objective of this study was to identify engine and drive system configurations to reduce rotor 

tip speed during cruise conditions and quantify the associated benefits. Previous NASA studies concluded that reducing 

rotor speed (from 650 fps hover tip speed) during cruise would reduce vehicle gross weight and fuel burn. Initially, rotor 

cruise speed ratios of 54% of the hover tip speed were of most interest during operation at cruise air speed of 310 ktas. 

Interim results were previously reported1

 for cruise tip speed ratios of 100%, 77%, and 54% of the hover tip speed using 

engine and/or gearbox features to achieve the reduction. Technology levels from commercial off-the-shelf (COTS), 

through entry-in-service (EIS) dates of 2025 and 2035 were considered to assess the benefits of advanced technology on 

vehicle gross weight and fuel burn. This technical paper presents the final study results in terms of vehicle sizing and fuel 

burn as well as Operational and Support (O&S) costs. New vehicle sizing at rotor tip speed reduced to 65% of hover is 

presented for engine performance with an EIS 2035 fixed geometry variable speed power turbine. LCTR2 is also 

evaluated for missions range cases of 400, 600, 800, 1000, and 1200 nautical miles and cruise air speeds of 310, 350 and 

375 ktas. 

INTRODUCTION 

To explore the benefits and possibilities of tiltrotors for 

commercial operations, NASA contracted Boeing to 

evaluate propulsion system concepts for the NASA Large 

Civil Tilt Rotor (LCTR2) concept vehicle shown in Figure 1. 

Vehicle characteristics include a takeoff gross weight (GW) 

weight of 107,700 lb, with 65 foot diameter rotors near the 

wing tips. The payload for NASA’s LCTR2 is 19,800 lb, 

which includes 90 passengers and baggage. The propulsion 

system is primarily contained in the two nacelles with two 

engines per nacelle. The nacelles tilt forward to cruise after a 

vertical take-off or hover. The LCTR2 design rotor tip 

speed (Vtip) is 650 fps during takeoff / hover to maintain 

high rotor efficiency and to manage noise levels. The 

vehicle rotor speed then decreases to a 350 fps rotor tip 

speed for cruise, or 54% of the hover RPM. 

Figure 1: Conceptual view of LCTR2

Previous summaries of study methods and results were 

presented at AHS conferences (references 1 & 2). The 

summaries included an evaluation of LCTR2 vehicle sizing 

and performance characteristics over a range of propulsion 

system variations. Three engine and drive system technology 

	

	
 

Study and Sub-System Optimization of Propulsion and Drive 

Systems for the Large Civil TiltRotor (LCTR2) Rotorcraft 

Mark Robuck and Joseph Wilkerson 

The Boeing Company 

Philadelphia, Pennsylvania 19142 

Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Yiyi Zhang and Bob Maciolek 

The Boeing Company 

Philadelphia, Pennsylvania 19142 
levels were studied in this effort: commercial off the shelf 

(2015 / COTS), and technology levels expected for 2025 

entry into service (EIS), and 2035 EIS. A primary goal of 

this study was to identify favorable engine and drive system 

concepts to achieve a 54% rotor cruise tip speed variation.

The summary also reported on the development of sizing 

methodology, generation of engine data for COTS and 

advanced technology engines (EIS 2025 and 2035), 

development of the drive system concept architecture and 

characteristics, analysis of prop-rotor performance, 

assessment of advanced technologies, as well as 

identification of technology challenges and needs for the 

overall system. 

Rotor speed variability of 100%, 77% and 54% was 

achieved with two methods: changing gear ratios in the 

output/transmission drive train and/or using highly variable 

output speed gas turbine engines. Table 1 contains the 

combinations of engine and drivetrain options that were 

previously reported. 

TABLE 1: ROTOR CRUISE TIP SPEED 

Engine 

Technology (for 

all combinations)

Rotor 

Cruise 

Tip 

Speed, 

(%)

Engine 

Cruise 

RPM

(%)

Drive 

System 

Cruise 

RPM, %

COTS 2015 

Engine

650 fps, 

(100%) 100% 100%

EIS 2025 Engine 

1*

500 fps, 

(77%)

100% 77% 

(2-speed)

77% 100%

EIS 2035 Variable 

Geometry Engine

1*

&

EIS 2035 Fixed 

Geometry Engine 

2*

350 fps, 

(54%)

100% 54% 

(2-speed)

77% 70% 

(2-speed)

54% 100%

1* refers to variable geometry ‘Variable Speed’ power 

turbine technology 

2* refers to fixed geometry ‘Variable Speed’ power turbine 

technology. 

The conclusions reached from the initial study, given 

the limitations of the methodology and the constraints of the 

configuration, were that the lightest vehicle weights were 

produced from advanced engines with rotors operating near 

500 fps tip speed. The lightest GW design was 91,923 lb, for 

the 500 fps cruise tip speed with a single-speed transmission 

and 77% engine RPM with the 2035 fixed geometry (FG) 

variable-speed power turbine (VSPT). The second lightest 

GW design was slightly heavier at 91,989 lb. It also occurs 

at the 500 fps cruise tip speed, but with a 2-speed 

transmission and 100% engine RPM. Vehicle GW for the 

2035 engine with variable geometry (VG) VSPT was 1540 

lb heavier than the 2035 FG-VSPT engine at Vtip= 500 fps. 

The 350 fps rotor tip speed cases sized very close to each 

other, between 93,900 lb and 94,900 lb GW. 

The most dramatic effects on vehicle sizing were 

obtained with advanced engine technology, which resulted 

in reduced fuel burn, and the most favorable operating 

condition was near 500 fps tip speed with cruise speed of 

310 ktas. Generally, the study results were insensitive to the 

method of speed reduction, nearly the same sizing results 

were obtained whether the speed reduction was achieved 

with reduced engine speed or with two speed transmissions. 

To further understand the sensitivities to tip speed, engine 

and gearbox speed reduction method, and mission 

parameters, additional analysis was performed, 

The current paper reports results from additional recent 

tasks accomplished by Boeing under contract to NASA.

Using only the EIS 2035 fixed-geometry variable speed 

power turbine (FG-VSPT) as the baseline engine, tasks 

included the following: 

Vehicle sizing for a new 65% rotor cruise tip speed 

case. 

Vehicle sizing for missions range cases of 400, 600, 

800, 1000, and 1200 nautical miles and cruise air 

speeds of 310, 350 and 375 ktas. 

Operating and Support (O&S) cost information for 

new sizing conditions 

Additional details of the analysis methodology, notional 

propulsion, rotor and drive system configurations, and 

vehicle sizing data are reported as well. 

BACKGROUND 

Previous Results 

Configurations and technology levels of Table 1 were 

evaluated to find the propulsion and drive system 

configuration that results in minimum vehicle weight and 

fuel burn for the three technology levels evaluated. 

Operational variables affecting that balance include engine 

speed reduction fraction, drive system speed reduction 

fraction, technology factors, efficiencies, and configuration 

variables (fuel quantity, vehicle size). 

Mission characteristics of range, cruise speed, and 

altitude were constrained to the original NASA design. 

Climb and cruise segments drove the fuel consumption in 

this study, which had a major effect on rotorcraft sized for 

long-range such as the LCTR2. Results of the sizing studies, 

engine and drive system configuration data, and study 

methodologies were presented previously, and the sizing 

comparison for the FG-VSPT study at 310 ktas airspeed are 

shown in Figure 2 and Figure 3 in this report. 

	

	
 	
Figure 2: 2035 FG-VSPT Engine – Effect of Rotor Tip 

Speed and Engine/Drive System RPM on GW 

Figure 3: 2035 FG- VSPT Engine -

 

Installed SHP and Weight

New Analysis 

Table 2 presents the updated matrix of rotor cruise tip 

speeds, combinations of drive system and engine rpm, and 

technology levels used to evaluate the LCTR2 overall 

vehicle size, geometry, performance, installed engine HP, 

and rotor efficiency. This table lists all conditions that were 

studied with the most recent unreported parameters of 

interest highlighted. All new sizing cases are conducted at 

EIS 2035 technology level with the FG-VSPT engine. 

Previous studies by NASA investigated rotor speed 

reduction to 54% (350fps). Boeing’s initial study results 

found that 500 fps tip speed rotor speed provided a lighter 

weight air vehicle design for the constraints and conditions 

imposed. To further articulate this study, an intermediate 

rotor speed of 422 fps (65%) was analyzed in this new work 

and is presented in this paper. Aircraft sizing results for 

airspeeds of 350 ktas, and 375 ktas. are also included to 

understand sensitivities encountered with higher operating 

speeds. Projected (relative) O&S costs are also presented to 

provide understanding of the economic effects in this design 

exploration. 

Table 2: Design Matrix of Engines, Technology and 

Cruise RPM Combinations 

TECHNICAL APPROACH 

Sizing Methodology 

Engine models, rotor system performance models, and 

drive system weight and efficiency charts were used by the 

sizing model and are covered in previous technical papers 

and a project report. The sizing methodology used in the 

latter portion of the study remained the same. Boeing used 

customized spreadsheets to evaluate the aircraft size and 

performance. This sizing tool modeled most of the 

VASCOMP (reference 3) performance and sizing 

procedures in a format that allowed Boeing to perform 

“Concept Evaluation” analysis for the LCTR2 air vehicle.

 

From this spreadsheet, aircraft weight, engine performances, 

rotor performances, mission performances, and overall 

vehicle sizing are extracted. 

Engine Model 

Engine models were provided by Rolls-Royce in 

spreadsheet format. Available shaft horsepower engine data 

was tabulated at Maximum Rated Power (MRP)

, 

Intermediate Rated Power (IRP), and Maximum Continuous 

Power (MCP) versus altitude and Mach number (all climb 

and cruise flight segments were at International Standard 

Atmosphere (ISA) conditions). This previously supplied 

engine data was extended to cover the 65% speed condition 

added to the study. The engine used in this study is a RollsRoyce engine designated PD628. This engine has Advanced 

Versatile Affordable Advanced Turbine Engines (VAATE)

advanced technology with high overall pressure ratio (>30) 

and two-spool core. The turbine was optimized for 90% 

speed operation with part speed performance down to the 

54% RPM condition. It weighs 807 lb with a reference SHP 

of 8086 HP per engine. A graphical representation of that 

engine model at 65% operating speed is shown in Figure 4. 

	

	
 
Figure 4: 2035 FG-VSPT Engine Power Available at 

65% RPM

Mission Fuel and Profile 

The installed engine power required for each LCTR2 

sizing case was scaled to the greater of hover takeoff power 

or cruise power. Engine scaling assumed specific fuel 

consumption (SFC) was preserved for the same relative 

power, altitude and Mach number. Power required for 

LCTR2 cruise performance accounted for the Rolls-Royce 

engines’ residual jet thrust. Fuel flow was obtained from 

referred fuel flow versus referred power, against Mach 

number and altitude. Mission fuel was calculated for each 

LCTR2 mission segment and summed up to total fuel 

required. Fuel was calculated at seven (7) climb altitudes, 

sequentially evaluated at the corresponding GW during 

climb, and at four (4) cruise segments. 

The NASA mission profile for the LCTR2 was used to 

size all cases. No attempt was made to find or use a more 

optimum altitude, cruise airspeed, or to evaluate other 

mission ranges. The LCTR2 sizing mission profile is 

described in Figure 5. 

 5 minute warm up at IRP power at 5,000’/ISA+20°C

 2 minute hover takeoff at 5,000’/ISA+20°C

 Climb to 25,000’ cruise altitude at MCP, ISA 

 Cruise at 25,000’/ISA, 310 ktas to a range of 1000 

nmi 

 Vertical descent (no time, no fuel, no distance)

 1 minute hover landing at 5,000’/ISA+20°C

 30 nm cruise allowance for alternate destination, Vbr 

(airspeed (velocity) for best range) at 25,000’/ISA

Taxi, 4 min

Ground check, 

1 min @ 60% 

takeoff power

Climb to 25,000’ cruise altitude @ MCP 

(25-30 min)

Takeoff & convert,

2 min

Cruise @ design airspeed 

to mission range

Alternate 

Destination 

30 nm

Transfer Altitude

(conservative

estimate of fuel)

No credit for range.

Final approach to land

+ 30 min 

Reserve Fuel 

@ 10,000’ / ISA

310 knots

Vertical landing, 

1 min 

@ 5,000’, ISA+20C

= 1000 nm Range

Conversion from 

helicopter to 

airplane above 

1,000’ (over 

terrain)

Conversion from airplane 

to helicopter above 1,000’

(over terrain)

Taxi, 4 min

Ground check, 

1 min @ 60% 

takeoff power

Climb to 25,000’ cruise altitude @ MCP 

(25-30 min)

Takeoff & convert,

2 min

Cruise @ design airspeed 

to mission range

Alternate 

Destination 

30 nm

Transfer Altitude

(conservative

estimate of fuel)

No credit for range.

Final approach to land

+ 30 min 

Reserve Fuel 

@ 10,000’ / ISA

310 knots

Vertical landing, 

1 min 

@ 5,000’, ISA+20C

= 1000 nm Range

Conversion from 

helicopter to 

airplane above 

1,000’ (over 

terrain)

Conversion from airplane 

to helicopter above 1,000’

(over terrain)

Figure 5: Mission Profile 

Drive System 

The LCTR2 drive system configuration remains the 

same as the previous work. It has 4 engines with 2 at each 

nacelle. It consists of 5 transmissions: a left hand (LH) and 

right hand (RH) proprotor gearbox, LH and RH Tilt Axis 

Gearbox, and a mid-wing gearbox. For portions of this 

study, it is assumed that the speed reduction is achieved 

using the drive system speed reducer. To accomplish this, a 

speed changing gearbox is placed at each engine input to the 

proprotor gearbox as shown in Figure 6. A similar baseline 

configuration without two speed capability is also used for 

the analysis cases where all speed reduction is accomplished 

by the engine. 

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 

Figure 6: Drive System Schematic

Cruise Propulsive Efficiency 

Boeing constructed models of three additional rotors for 

this study based on NASA LCTR2’s rotor airfoils and blade 

platform. Twist distributions were modified for the 422 fps 

rotor speed case for airspeed of 310 ktas. Two rotor models 

were constructed to evaluate the impact of higher cruise 

airspeeds on the LCTR2 size, GW, and cost; one for 350 

 30 minute reserve fuel at Vbr, 10,000’/ISA

	

	
 
ktas cruise and the other for 375 ktas cruise. Both rotor 

designs applied the 350 fps rotor cruise tip speed, which 

corresponded to 54% RPM where existing engine data was 

available. The helical blade tip Mach number is 0.71 at 

25,000 ft, 375 ktas cruise airspeed, so this design required 

thinner airfoils over the blade radius to avoid adverse drag 

divergence. See Figure 7 for an example of blade 

distribution at 310 ktas. Maps of rotor cruise efficiency are 

illustrated in Figure 8 through Figure 10. 

-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

Boeing Twist for 650 fps Cruise Tip Speed

Boeing Twist for 500 fps Cruise Tip Speed

Boeing Twist for 422 fps Cruise Tip Speed

NASA LCTR2 Twist (350 fps Cruise Tip Speed)

Dashed Lines: Radial distribution 

of helical inflow angle at cruise 

flight speed = arctan (/x)

Figure 7: Comparison of Rotor Blade Twist Distribution 

Figure 8: Rotor Cruise Propulsive Efficiency for 422 fps 

Cruise Tip Speed Design at 310 ktas Airspeed 

 For the two additional rotors operating at increased 

cruise airspeeds of 350 ktas and 375 ktas, Boeing retained 

the NASA parameters for cruise tip speed of 350 fps and its 

relative chord distribution. Blade geometric twist was 

modified to better align local airfoil sections with helical 

inflow angle at the two higher flight speeds. During this 

process, additional consideration was given to the attendant 

increase in local blade Mach number, especially over the 

inboard portion of the rotor blade. As flight speed is raised 

from 310 ktas to 350 ktas and beyond to 375 ktas, the local 

Mach number at the blade root station (r/R = 0.10) increases 

from 0.51 to 0.58 and 0.63, respectively.

 Inspection of the properties provided by NASA for the 

LCTR2 28% thick blade root airfoil indicates that this airfoil 

cannot operate above Mach 0.60 at any angle-of-attack 

without incurring significant compressibility penalties. 

Comparison of this limit with the local Mach number 

conditions at the blade root suggests that at 350 ktas this 

airfoil will operate close to its drag divergence boundary, 

while at 375 ktas this airfoil will operate entirely beyond this 

limit and unduly penalize rotor performance at this operating 

condition. For the purpose of this study, the original NASA 

LCTR2 airfoil placement was retained for the 350 ktas rotor 

design, but was modified for the 375 ktas design by 

eliminating the 28% thick airfoil from the blade root and redistributing the remaining airfoils along the inner portion of 

the span. 

Upon re-twisting the blade to align the local airfoil 

sections with helical inflow angle, rotor cruise predictions 

were made with the Boeing B-08 rotor performance program 

at representative thrust conditions to identify the associated 

blade lift coefficient levels. From these calculations, a 

representative value of Cl = 0.30 was identified, and this 

value was used to determine the limiting outboard radial 

station at which the 18% thick LCTR2 airfoil could be 

tolerated without exceeding its performance limits. A limit 

of r/R = 0.50 was identified, and the blade thickness 

distribution of the 375 ktas rotor was tapered from 18% at 

r/R=0.225 to 12% at r/R=0.50. 

The Boeing 350 ktas cruise airspeed rotor design had a 

tri-linear twist (-33.1°/-30.5 / -27°) to closely match the 

helical inflow distribution with a 350 fps tip speed. The 

LCTR2 solidity, reference blade planform and airfoil 

distribution were maintained. Breakpoints in the piecewise 

linear twist distribution were located at r/R = 0.45 and 0.70. 

The Boeing rotor design for 375 ktas cruise airspeed 

had a tri-linear twist (-30.8°/ -29o /-25.8°) with the LCTR2 

solidity and reference blade planform. Breakpoints in the 

piecewise linear twist distribution were located at r/R = 0.40 

and 0.70. 

Figure 9: Rotor Cruise Propulsive Efficiency for 350 

ktas Cruise Airspeed Design, 350 fps Vtip

	

	
 
Figure 10: Rotor Cruise Propulsive Efficiency for 375

ktas Cruise Airspeed Design, 350 fps Vtip 

O&S Cost 

Cost was estimated with the PRICE Estimating Suite, the 

identical PRICE model that was applied in reference 4 for 

previous civil tiltrotor analysis. Relevant output from the 

Excel sizing analysis was linked to the PRICE Estimating 

Suite and run in Phoenix Integration’s ModelCenter 

environment. The cost model assumed a fleet of 300 aircraft 

operating 2500 flight hours per year. Indirect operating costs 

were based on a service life of 20 years and a 7.5% interest 

rate, but this study focused on direct costs. 

The metric of Direct Operating Cost (DOC) per 

Available Seat-NM (DOC/ASM) is used by commercial 

passenger airlines to track the financial health of daily 

operations. The revenue side of the balance sheet is revenue 

per available seat-nmi, which is essential to the airline’s 

financial viability. 

Cash Operating Cost comprises both direct and indirect 

operating cost. The term Cash DOC refers only to the direct 

operating cost components, including fuel, oil, maintenance, 

landing fees, crew expenses, supplies and catering, flight 

crew and cabin crew salaries, as shown in Table 3. 

The ground rule utilization of 2500 flight hours per year 

actually required 2.5 flight crews and cabin crews per 

aircraft because air crews are limited to 1000 flight hours per 

year. Annual crew salaries came from Conklin & deDecker 

(“The Aircraft Cost Evaluator” http://www.conklindd.com).

They were multiplied by 2.5 crew sets and then divided by 

2500 FH/aircraft/year to express them as $/FH, per aircraft 

in the fleet. 

Mission fuel requirements came from the Excel sizing 

analysis, depending on the rotorcraft GW, cruise altitude and 

airspeed, and, as shown in this study, are greatly affected by 

advanced engine technologies. The cost of fuel and oil, flight 

crew salaries, cabin crew salaries, landing fees, crew 

expenses, and supplies and catering were added to the 

PRICE output with a Post-Price module in ModelCenter to 

arrive at Cash DOC/ASM. 

OPERATING COSTS

Direct Operating Cost (DOC)

Fuel & Oil

Maintenance (Price)

Airframe, Labor & Parts

Engine Restoration

Dynamic Systems/Life Ltd

Burden

Landing Fees

Crew Expenses

Supplies-Catering

Indirect (Fixed) Operating Cost

Flight Crew Salaries + benefits

Cabin Crew Salaries + benefits

Hanger Costs

Hull Insurance

Depreciation

Financing

Training

Computer Mgt pgm

Refurbishment

Cash 

DOC

OPERATING COSTS

Direct Operating Cost (DOC)

Fuel & Oil

Maintenance (Price)

Airframe, Labor & Parts

Engine Restoration

Dynamic Systems/Life Ltd

Burden

Landing Fees

Crew Expenses

Supplies-Catering

Indirect (Fixed) Operating Cost

Flight Crew Salaries + benefits

Cabin Crew Salaries + benefits

Hanger Costs

Hull Insurance

Depreciation

Financing

Training

Computer Mgt pgm

Refurbishment

OPERATING COSTS

Direct Operating Cost (DOC)

Fuel & Oil

Maintenance (Price)

Airframe, Labor & Parts

Engine Restoration

Dynamic Systems/Life Ltd

Burden

Landing Fees

Crew Expenses

Supplies-Catering

Indirect (Fixed) Operating Cost

Flight Crew Salaries + benefits

Cabin Crew Salaries + benefits

Hanger Costs

Hull Insurance

Depreciation

Financing

Training

Computer Mgt pgm

Refurbishment

Cash 

DOC

RESULTS 

LCTR2 Sizing 

The LCTR2 was resized with the 2035 FG-VSPT 

engine for all rotor cruise tip speeds previously evaluated 

and the additional rotor design with a 422 fps tip speed (65% 

of hover rpm) at the baseline airspeed of 310 ktas to better 

define the optimum rotor cruise tip speed. Results for 

vehicle GW are shown graphically in Figure 11 and results 

for engine power and weight are in Figure 12. Minor 

adjustments were made to some of the study parameters that 

resulted in a small effect on previously reported sizing 

results for tip speeds at 350, 500, and 650 fps. 

Figure 11: 2035 FG VSPT Engine: Rotor Tip Speed and 

Engine/Drive System RPM Effect on GW

Table 3. Definition of Cash DOC 

	

	
 
Figure 12: 2035 FG-VSPT Engine Installed SHP and 

Weight

As anticipated, the new sizing case at 422 fps provided 

the lightest aircraft sizing point by a narrow margin. The 422 

fps and 500 fps rotor tip speeds had a small spread of only 

648 lb between them, rather clearly showing that the 

optimum rotor cruise tip speed is in this 422 fps to 500 fps 

range. In contrast to previous results in this study, the lowest 

weight option at the 422 fps tip speed is obtained with a 2 

speed drive system used to obtain the 65% reduction, and 

engine operating at 100% speed. 

Confirming previous analysis, the 2035 FG-VSPT engine 

resulted in an average 2400 lb lower GW than the 2035 

(Variable geometry) VG-VSPT engine for all combinations 

of tip speed and engine-drive system RPM. The minimum 

GW drops down to 91,612 with the FG-VSPT and 422 fps 

tip speed. 

The closest result for 350 fps was 1912 lb heavier than 

the minimum GW case just mentioned. There was a very 

small spread of rotor cruise propulsive efficiency from 350 

fps, 422 fps, and 500 fps rotors, which is 0.841 to 0.848 at 

the 310 ktas design cruise airspeed. Cruise propulsive 

efficiency for the 650 fps case was notably lower, 0.76. 

Figure 13 graphs the propulsion system component 

weights, i.e. rotor weight, drive system weight, and total 

engine weight. The combination of rotor and drive system 

weight clearly overshadows the engine weight. The 2035 

drive system is estimated to weigh about 12.5% less than the 

2015 drive system, for a given gear reduction and power 

rating. Sizing results showed the average 2015 drive system 

weight to be about 0.41 lb/rated HP, whereas the average 

2035 drive system weighed 0.344 lb/rated HP, a significant 

weight reduction.

Figure 14 shows the variation of the fuel weight as a 

fraction of GW. The 2035 FG-VSPT engine is considerably 

lighter than either of the other engines, bringing the empty 

weight down, and it has lower fuel flow. These fuel weight 

fractions are much lower than the 2015 fuel weight fractions 

spotted on the graph. 

Figure 13: Propulsion System Component Weights for 

2035 FG-VSPT Engine

Figure 14: Mission Fuel Weight Fraction for 2035 FGVSPT Engine

Sensitivity to Increased Airspeed and Range 

Tasks were completed to explore the sensitivity of 

LCTR2 to design cruise airspeed and mission range, in 

concert with estimated operational costs. Using the best 

engine match for LCTR2, the 2035 FG-VSPT engine, three 

design airspeeds are evaluated; 

 310 ktas with the 422 fps tip speed rotor designed for 

310 ktas cruise airspeed. 

 350 ktas with the new 350 fps tip speed rotor 

designed for 350 ktas cruise airspeed. 

 375 ktas with the new 350 fps tip speed rotor 

designed for 375 ktas cruise airspeed. 

	

	
 
Aircraft Weight Growth with Design Airspeed and 

Range

The LCTR2 was resized at each design airspeed for 

mission ranges of 400 nautical miles (nmi) up to 1200 nmi, 

including estimated operating costs. The carpet plot in 

Figure 15 quantifies the growth of vehicle GW for higher 

design cruise airspeeds (more required SHP) and for longer 

range (increased mission fuel). Both trends are as expected. 

The growth of GW with design airspeed is dramatic. 

Considering the 1000 nmi mission range, GW grows from 

91,600 lb at a 310 ktas design airspeed to 110,000 lb at a 350 

ktas design airspeed, on up to over 125,000 lb at a 375 ktas 

design airspeed. This increase was driven by the added fuel 

requirement, and compounded by increased installed SHP to 

satisfy higher cruise airspeeds. Increasing mission range 

from 1000 nmi by 20% to 1200 nmi increased the takeoff 

GW by 5% to 7%. 

Figure 15: Design Gross Weight Sensitivity to Design 

Airspeed and Range

The accompanying bar chart on the left hand side of 

Figure 15 provides reference GW from three previous cases; 

the NASA LCTR2 design with 350 fps tip speed, the Boeing 

2015 design with 500 fps tip speed, and the Boeing 2035 

FG-VSPT design with 422 fps tip speed. The GW values 

displayed were the minimum GW for the selected tip speeds 

and propulsion system technology. Corresponding aircraft 

empty weight fractions (Empty Weight / GW) are shown in 

Figure 16. Higher design airspeeds require more installed 

SHP, heavier drive systems to deliver that power, as well as 

heavier rotors to provide increased thrust. All these lead to a 

higher empty weight fraction. Contrarily, at a given design 

airspeed, increased range requires more fuel, necessarily 

reducing the empty weight fraction to account for the added 

useful load (fuel). 

Figure 16: Aircraft Empty Weight Sensitivity to Design 

Airspeed and Range 

Aircraft Operating Cost Variation with Design Airspeed 

and Range

Estimated values of DOC per flight hour (DOC/FH) and 

DOC/ASM are shown in Figure 17 for the same 

combinations of design airspeed and mission range shown 

above. These metrics have been normalized by PRICE 

results for the 2015 COTS engine at 100% rpm, 310 ktas and 

the 500 fps rotor tip speed. The 2035 drive system and FGVSPT engine technology results in a reduced GW for the 

310 ktas aircraft and reduced relative fuel flow/SHP. The 

relative DOC in Figure 17 for the 2035 engine and drive 

system technology shows that advanced technology can 

result in nearly 30% lower DOC/ASM and 20% lower 

DOC/FH relative to the best combination with 2015 

technology. 

DOC/FH naturally increases with aircraft GW; larger 

aircraft generally requiring more fuel per FH. But Figure 17

shows DOC/FH to be fairly flat with mission range for the 

310 ktas design, even as GW grew from about 80,000 lb at 

the 400 nmi range up to 96,000 lb for the 1200 nmi range. 

That reflects the content of DOC/FH: part fuel costs that do 

increase with GW and part fixed costs per flight hour, such 

as crew salaries and expenses (overnight stays). 

Figure 17: Relative Cost Variation with Design Airspeed 

and Range

	

	
 
Notably, DOC/FH increases significantly for design 

airspeeds of 350 ktas and 375 ktas driven by increased 

maintenance cost and fuel per FH associated with heavier, 

more powerful aircraft. DOC/FH shows more sensitivity to 

mission range at the higher cruise airspeed designs, 

presumably due to lower nmi/lb of fuel at the higher GW. 

Results for DOC/FH and DOC/ASM reveal that the 

additional speed capability comes at a price, and there is no 

cost benefit for the additional airspeed even when 

considering the costs per seat mile, which is a measure of 

productivity scaled to transporting the individual customer. 

Operating and support cost results (per flight hour) for 

310 ktas sizing cases are presented in Figure 18 and Figure 

19. The solid blue line represents the operating cost trend 

for 2 speed transmission sizing cases while the solid green 

line represents the engine based speed reduction cases. The 

dashed lines represent the GW variation for the same sizing 

cases and are proportionally scaled to the cost results. 

Trends show that there are minor differences in operating 

cost between an engine speed variant or a two speed 

transmission variant at higher cruise tip speeds of about 

500fps and above. As the cruise tip speed takes a larger 

reduction toward 350fps, the operation cost favors a two 

speed gearbox rather than the engine based speed reduction. 

The relative cost minimum seems to coincide with lowest 

weight models for the vehicle, which is between 422 fps and 

500 fps. Interestingly the cost model exhibits a greater 

variation between the two speed transmission trend and the 

engine based speed reduction trend than the GW trends 

indicate. GW variation at the minimum appears negligible 

while the operating cost advantage for a 2 speed system at 

the minimum design point is approximately 2.5%. This 

would be attributable to the greater fuel consumption for 

engine based speed reduction as indicated in Figure 14. 

Similar observations of the trends can be made for the 

operating cost per available seat mile chart of Figure 19. 

75000

80000

85000

90000

95000

100000

75%

80%

85%

90%

95%

100%

300 350 400 450 500 550 600 650 700

Aircraft Gross Weight

Relative

Direct Operating Cost(DOC per Flight Hour)

Cruise Tip Speed (ft/s)

Variation of Cruise Tip Speed on Relative Direct Operating Cost

(DOC per Flight Hour) & Gross Weight

Two Speed Transmission - DOC

Variable Speed Engine - DOC

Two Speed Transmission - GW

Variable Speed Engine - GW

Figure 18: Operating Cost per FH at Various Cruise Tip 

Speeds at 310 ktas

65000

70000

75000

80000

85000

90000

95000

100000

65%

70%

75%

80%

85%

90%

95%

100%

300 350 400 450 500 550 600 650 700

Aircraft Gross Weight

Relative

Direct Operating Cost

(DOC per Available Seat Mile)

Cruise Tip Speed (ft/s)

Variation of Cruise Tip Speed on Relative Direct Operating Cost

(DOC per Available Seat Mile) & Gross Weight

Two Speed Transmission - DOC

Variable Speed Engine - DOC

Two Speed Transmission - GW

Variable Speed Engine - GW

Figure 19: Operating Cost per ASM at Various Cruise 

Tip Speeds at 310 ktas 

CONCLUSIONS 

The 2035 FG-VSPT engine gave the lightest GW 

solution of the four engines evaluated, where the best rotor 

cruise tip speed was between 422 and 500 fps (65% and 77% 

of full rotor speed). This option had lower fuel flow and 

smaller engines (design SHP) than the 350 fps tip speed 

(54% rpm). Sizing analysis of the LCTR2 propulsion 

systems at the 422 fps rotor tip speed with the 2035 FGVSPT engine and a two-speed drive system provided the 

lightest overall vehicle GW at 91,612 lbs. The 500 fps rotor 

tip speed produced a close second, 92,012 lb GW with either 

a single-speed or a 2-speed drive system. Reduced engine 

weight and fuel consumption associated with the 2035 FGVSPT has a dramatic effect on vehicle sizing when 

compared to the COTS 2015 engine (best case) and 

represents a significant result in this study. A beneficial 

result in this study of the NASA LCTR2 design with a 350 

fps tip speed is that it has investigated practical operational 

boundaries for a tiltrotor propulsion systems, as 422 to 500 

fps tip speeds are far lower than the current V-22 cruise tip 

speed of 664 fps. 

The LCTR2 GW weight differences between 

configurations that used engine based speed variation vs. 

drive system speed variation were subtle considering the 

significant variations studied in this effort. As an example, 

for the 422 fps sizing cases at the 2035 technology level, 

which represents the most favorable sizing cases in the 

study, the difference between two-speed transmission and 

reduced engine speed cases (91,612 lb and 92,260 lb 

respectively) is a mere 0.7%. For the 2015 technology level, 

the difference between two-speed and reduced engine speed 

for 500 fps best sizing is 0.4%. In general the two-speed 

transmission approach becomes more favorable where the 

engine performance falls off substantially, however the 

lower rotor speed cases where engine performance is 

diminished are not the optimum (lowest GW) configurations 

in this study. 

	

	
 
The trade space examined in this study was heavily 

focused on vehicle sizing with the vehicle GW and system 

weights as the parameters of interest. A sensitivity study task 

was also conducted to evaluate weight trends and cost trends 

as mission range and speed were varied. Results are 

presented hold no surprises, the weight and cost of the 

LCTR2 vehicle rose predictably and proportionally as the 

variables of speed and range increased. It is notable 

however that operating costs favor a two speed transmission 

approach to speed reduction more distinctly than the GW 

analysis suggests. The most favorable operating costs for 

the 310 ktas cases examined were 2.5% better with the two 

speed approach as compared to engine based speed 

reduction due to differences in fuel consumption. 

REFERENCES 

1) Robuck, Mark, Wilkerson, Joseph, and Snyder, 

Christopher A., Vonderwell, Daniel, “Design Study of 

Propulsion and Drive Systems for the Large Civil 

TiltRotor (LCTR2) Rotorcraft” American Helicopter 

Society 67th Annual Forum, Virginia Beach, VA, May 

3-5th, 2011 

2) Snyder, Christopher A., Robuck, Mark, Wilkerson, 

Joseph, and Nordstrom, Carl, “Summary of the Large 

Civil TiltRotor (LCTR2) Engine Gearbox Study,” 

American Helicopter Society International Power Lift 

Conference, Philadelphia, PA, October 2010 

3) Wilkerson, Joseph, "VASCOMP III, The V/STOL 

Aircraft Sizing and Performance Computer Program, 

User's Manual", Boeing Rotorcraft, D210-13635-1, 

2002

4) Wilkerson, Joseph, Smith, Roger, “Aircraft System 

Analysis of Technology Benefits to Civil Transport 

Rotorcraft”, NASA/CR-2009- 214594 

	

	
 


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Study and Sub-System Optimization of Propulsion and Drive Systems for the Large Civil 

TiltRotor (LCTR2) Rotorcraft 

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Robuck, Mark; Wilkerson, Joseph; Snyder, Christopher, A.; Zhang, Yiyi; Maciolek, Bob 

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WBS 380046.02.03.03.03.03 

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Subject Categories: 07 and 83 

Available electronically at http://www.sti.nasa.gov 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

In a series of study tasks conducted as a part of NASA’s Fundamental Aeronautics Program, Rotary Wing Project, Boeing and Rolls-Royce 

explored propulsion, drive, and rotor system options for the NASA Large Civil TiltRotor (LCTR2) concept vehicle. The original objective 

of this study was to identify engine and drive system configurations to reduce rotor tip speed during cruise conditions and quantify the 

associated benefits. Previous NASA studies concluded that reducing rotor speed (from 650 fps hover tip speed) during cruise would reduce 

vehicle gross weight and fuel burn. Initially, rotor cruise speed ratios of 54 percent of the hover tip speed were of most interest during 

operation at cruise air speed of 310 ktas. Interim results were previously reported for cruise tip speed ratios of 100, 77, and 54 percent of the 

hover tip speed using engine and/or gearbox features to achieve the reduction. Technology levels from commercial off-the-shelf (COTS), 

through entry-in-service (EIS) dates of 2025 and 2035 were considered to assess the benefits of advanced technology on vehicle gross 

weight and fuel burn. This technical paper presents the final study results in terms of vehicle sizing and fuel burn as well as Operational and 

Support (O&S) costs. New vehicle sizing at rotor tip speed reduced to 65 percent of hover is presented for engine performance with an EIS 

2035 fixed geometry variable speed power turbine. LCTR2 is also evaluated for missions range cases of 400, 600, 800, 1000, and 1200 

nautical miles and cruise air speeds of 310, 350 and 375 ktas.

15. SUBJECT TERMS

Turboshaft; Tiltrotor; Design; Vehicle sizing; Gas turbine engine; Gearbox; Mission analysis; Operational and support costs 

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