Mark Robuck, Joseph Wilkerson, and Yiyi Zhang

The Boeing Company, Philadelphia, Pennsylvania

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Daniel Vonderwell

Rolls-Royce Corporation, Indianapolis, Indiana

Design Study of Propulsion and Drive Systems for

the Large Civil TiltRotor (LCTR2) Rotorcraft

NASA/TM—2013-218102

December 2013
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NASA Center for AeroSpace Information

 7115 Standard Drive

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Mark Robuck, Joseph Wilkerson, and Yiyi Zhang

The Boeing Company, Philadelphia, Pennsylvania

Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Daniel Vonderwell

Rolls-Royce Corporation, Indianapolis, Indiana

Design Study of Propulsion and Drive Systems for

the Large Civil TiltRotor (LCTR2) Rotorcraft

NASA/TM—2013-218102

December 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

67th Annual Forum and Technology Display (FORUM 67)

sponsored by the American Helicopter Society International (AHS)

Virginia Beach, Virginia, May 3–5, 2011
Acknowledgments

We would like to thank Susan Gorton and management of the Subsonic Rotary Wing vehicle project within the NASA Aeronautics 

Research Mission Directorate, Fundamental Aeronautics Program for supporting this research. In addition, special thanks and 

appreciation to propulsion and drive system staff at NASA Glenn Research Center for guidance and support through the entire 

effort. Technical assistance and support from many individuals at Boeing and Rolls Royce are also acknowledged, including, 

Robert Maciolek, Lawrence Basil, Paul Rudisaile, and Carl Nordstrom (Rolls Royce).

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only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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. 

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
ABSTRACT1

Boeing, Rolls Royce, and NASA have worked together to complete a parametric sizing study for NASA’s 

Large Civil Tilt Rotor (LCTR2) concept 2nd iteration. Vehicle gross weight and fuel usage were evaluated as 

propulsion and drive system characteristics were varied to maximize the benefit of reduced rotor tip speed during 

cruise conditions. The study examined different combinations of engine and gearbox variability to achieve rotor 

cruise tip speed reductions down to 54% of the hover tip speed. Previous NASA studies identified that a 54% 

rotor speed reduction in cruise minimizes vehicle gross weight and fuel burn. The LCTR2 was the study baseline 

for initial sizing. This study included rotor tip speed ratios (cruise to hover) of 100%, 77% and 54% at different 

combinations of engine RPM and gearbox speed reductions, which were analyzed to achieve the lightest overall 

vehicle gross weight (GW) at the chosen rotor tip speed ratio. 

Different engine and gearbox technology levels are applied ranging from commercial off-the-shelf (COTS) 

engines and gearbox technology to entry-in-service (EIS) dates of 2025 and 2035 to assess the benefits of 

advanced technology on vehicle gross weight and fuel burn. Interim results were previously reported1

. This 

technical paper extends that work and summarizes the final study results including additional engine and drive 

system study accomplishments. New vehicle sizing data is presented for engine performance at a single operating 

speed with a multispeed drive system. Modeling details for LCTR2 vehicle sizing and subject engine and drive 

sub-systems are presented as well. This study was conducted in support of NASA’s Fundamental Aeronautics 

Program, Subsonic Rotary Wing Project. 

	

	
	 

Design Study of Propulsion and Drive Systems for 

the Large Civil TiltRotor (LCTR2) Rotorcraft 

Mark Robuck, Joseph Wilkerson, and Yiyi Zhang 

The Boeing Company 

Philadelphia, Pennsylvania 19142 

Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Daniel Vonderwell 

Rolls-Royce Corporation 

Indianapolis, Indiana 46206 
INTRODUCTION 

Rotorcraft propulsion systems generally operate within 

a narrow range of rotor tip speeds. However, tiltrotor aircraft 

are able to utilize a wider range of rotor cruise tip speed 

(Vtip). For example, the V22 operates at a higher RPM 

(103.8%) for hover operations and at a lower RPM (84%) 

for airplane cruise conditions. This study was conducted to 

identify and evaluate propulsion system concepts to achieve 

a rotor cruise Vtip of approximately 54% of the hover Vtip 

for a large civil tiltrotor air vehicle. It also investigates the 

most advantageous speed variation strategies and 

technologies for the integrated engine and drive system. The 

evaluation was performed for the NASA Large Civil 

Tiltrotor (LCTR2) configuration, shown in Figure 1, resizing 

the aircraft according to the impact of component weights, 

engine and rotor performance, and mission fuel. .

The NASA LCTR2 payload is 90 passengers, weighing 

19,800 pounds with baggage. Vehicle characteristics include 

a takeoff gross weight of 107,700 pounds, with 65 foot 

diameter rotors near the wing tips. The LCTR2 design rotor 

tip speed is 650 fps during takeoff / hover to maintain high 

rotor efficiency and to manage noise levels during takeoff 

and hover. The vehicle then decreases to a 350 fps rotor Vtip 

for cruise, or 54% of the hover RPM. The four engine 

arrangement was selected for hover OEI conditions, with 

transmission technology comparable to Advanced Rotorcraft 

Transmission Program (ART I & II, AATD Program2

).

This paper summarizes many of the efforts and 

accomplishments by Boeing and Rolls-Royce engineers 

under NASA NNA06BC41C Task Order 10 entitled, 

“Engine/Gearbox Assessment for 50% Variable Rotor Tip 

Speed”. The study analyzed operation at full rotor tip speed 

and at partial tip speeds of 77% and 54% for climb and 

cruise segments of the mission profile, which dominate fuel 

usage. The cruise condition is 310 knots, at 25,000 ft altitude 

for a range of 1000 nm. 

While an overview of the project and results were 

presented in Reference 1, the current paper provides 

additional details of the analysis methodology, notional 

propulsion and drive system configurations, and additional 

vehicle sizing data for an EIS 2035 engine configuration 

focused on optimized engine performance near 100% 

(engine) speed with a fixed-geometry variable-speed power 

turbine (VSPT). 

The primary goal of this study is to identify the best 

engine and drive system concepts and technology to achieve 

a 54% rotor cruise Vtip variation, and desired operating 

economics, for the LCTR2 rotor disk loading, fuselage size,

and mission profile. Project tasks include an evaluation of 

LCTR2 vehicle sizing and performance characteristics,

development of sizing methodology, generation of engine 

data for COTS and advanced technology engines (EIS 2025 

and 2035), development of the drive system configurations 

and performance, analysis of prop-rotor performance, 

assessment of advanced technologies and operational 

scenarios conducted at 54% rotor cruise Vtip, and 

identification of technology challenges and needs for the 

overall system.

Figure 1: Conceptual view of LCTR2. 

TECHNICAL APPROACH 

Three engine and drive system technology 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. These configurations 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 are all 

presented in this report.

Rotor speed variability of 100% to 54% was achieved 

with two methods that were investigated as a part of this 

study: changing gear ratios in the output/transmission drive 

train and/or using highly variable output speed gas turbine 

engines. An additional engine was added to the study to 

evaluate the value of the highly variable output speed 

approach using fixed geometry. This fixed-geometry VSPT 

was optimized for operation over a large output speed range 

with 2035 technology. Table 1 contains the combinations of 

engine and drivetrain options that were evaluated in this 

study. 

	

	
	 	
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. 

An Excel based aircraft performance and sizing tool was 

constructed to include many of the performance and sizing 

procedures from VASCOMP3

. This format allowed Boeing 

to perform “Concept Evaluation” analysis for the LCTR2 air 

vehicle. The sizing tool used basic aircraft parameters from 

the LCTR2 vehicle design. such as fuselage size and wing 

characteristics, and incorporated scalable engine 

performance models, prop-rotor performance models, and 

drive system characteristics. Fuselage characteristics 

remained constant while wing parameters and all 

performance models (engine, drive and rotors) provided 

scalable input and output for the vehicle sizing model. The 

sizing tool is described in the Vehicle Sizing section of this 

paper which is preceded by descriptive sections on engine, 

prop-rotor and drive system models.

ENGINE MODELS 

Team-mate Rolls-Royce provided tabulated engine data 

for different advanced technology engines at each of the 

specified engine operating speeds. Four engine technologies 

at three operating RPMs gave a total of twelve combinations 

of engine data. Each set of data covered power available, 

fuel flow and residual thrust over an operating range of 

Mach number and altitude. 

Power available is tabulated at the takeoff max rated 

power (MRP), intermediate rated power (IRP), and max 

continuous power (MCP) for each of the twelve 

combinations of engine data. Referred fuel flow collapsed 

well versus referred horsepower for all power settings and 

altitudes, and was modeled in Visual Basic (VB) as part of 

the Excel sizing program. Figure 2 shows sample data 

supplied by Rolls-Royce for the 2035 variable-geometry 

VSPT engine (PD 647).

Figure 2: 2035 EIS Variable Geometry Engine Power 

Available at 54% Speed and 77% (sample data supplied 

by Rolls-Royce) 

The baseline COTS (2015) engine is a current 

technology core of the appropriate flow size with a free 

power turbine driving the rotorcraft transmissions. The 

engine design and cycle performance are representative of 

commercial production technology generally available in the 

industry today. Engine component matching is optimized to 

provide good performance and high levels of efficiency over 

a broad power and speed regime, see Figure 3.

	

	
	 
Figure 3: Representative Image of the EIS 2015 Engine. 

The 2015 engine configuration is an axial core with a 

conventional compressor and cooled turbine, along with a 

free power turbine. The turbine in this turboshaft application 

is only driving a power output shaft and will therefore be 

referred to it as a power turbine (PT), consistent with the 

helicopter world. The compressor has variable geometry 

stators to allow satisfactory operation at off-design speeds. 

The power turbine matching was optimized to provide good 

efficiency between 80% and 100% RPM. As such, the 

engine is well suited for a variable speed transmission/rotor 

system with operation down to a 77% shift point. When 

coupled with a fixed transmission gear ratio, there is an 

appreciable drop in performance at PT speeds below 77%, 

resulting in non-optimal performance at 54% PT speed due 

to the wide variation in power turbine inlet incidence angle, 

which occurs at significantly reduced power turbine speeds. 

The 2025 engine utilizes COTS engine architecture with 

future technology insertion to improve performance and 

provide greater efficiency at reduced output speeds. It

incorporates cooling and materials improvements to allow an 

increase in cycle temperatures based on projected 

technology maturation over the next 10 years. This engine 

also features the addition of variable geometry in the power 

turbine. With the fixed geometry COTS engine, incidence 

angle changes in the power turbine due to the speed 

difference between 100% takeoff and the 54% cruise 

operating rotor speed leads to efficiency losses. For the 

2025 engine, turbine geometry is varied to accommodate 

wide variations in incidence resulting in appreciable 

improvement in specific fuel consumption (SFC) at the 

cruise condition, which is a major driver in mission fuel 

consumption. 

The 2025 engine variable turbine control system and 

mechanism does result in an increase in power plant system 

weight, which is accounted for in the aircraft studies. The 

2025 engine data were supplied to Boeing in tabular form, 

with scaling factors to allow performance, weight, and 

envelopes established across a broad power range. 

Two versions of the advanced technology 2035 VSPT 

engines were constructed for this study, one with a variable 

geometry turbine shown in Figure 4, the other with a fixed 

geometry turbine. These engines are based on a new, higher 

technology core with a high cycle pressure ratio, improved 

engine component efficiencies, and increases in turbine inlet 

temperature representative of technologies expected for the 

2035 timeframe. The aggressive overall pressure ratio 

(OPR) target of the 2035 engine resulted in a departure from 

the architecture employed in the 2015 and 2025 engines. 

Figure 4: Representative Image of the EIS 2035 Engine. 

To provide good operability and part power efficiency, 

the Rolls-Royce PD647 2035 variable geometry engine is a 

three-shaft design with Intermediate Pressure (IP) and High 

Pressure (HP) spools. The IP compressor is an all-axial 

configuration, while the HP compressor is an axialcentrifugal unit that has an appreciable efficiency benefit 

over an all-axial design given the low exit corrected flow 

rates produced by the high OPR cycle. Both the HP and IP 

turbines make full use of the advanced materials and cooling 

technologies based on projected technology maturation for 

this time period. The advanced power turbine was an 

uncooled variable geometry that provided substantially 

improved power available and reduced fuel flow at reduced 

operating RPM along with significantly reduced envelope 

and weight. The engine also embodies advanced controls 

and diagnostic technologies. 

The Versatile Affordable Advanced Turbine Engine 

(VAATE) technologies reflected in the PD647 provided a 

significant weight reduction relative to the 2015 and the 

2025 engines, as shown in Table 2. But the variable 

geometry power turbine feature that provided the excellent 

performance also carried a weight penalty. 

TABLE 2: ENGINE DRY WEIGHTS 

Engine

Installed 

SHP

(MRP/SLS)

Engine 

Dry

Weight

HP/Weight 

Ratio

2015

(COTS) 8100 HP 1356 lb 5.97

2025

(PD646) 8088 HP 1556 lb 5.20

2035

Variable Geom. 

VSPT (PD647)

8088 HP 1020 lb 7.93

2035

Fixed Geom.

VSPT (PD628)

8086 HP 807 lb 10.0

	

	
	 
NASA Glenn Research Center wanted to include in this 

assessment the application of 2035 VAATE technologies for 

a fixed geometry variable speed power turbine (VSPT). The 

core would be the same as the previous advanced and high 

performance EIS 2035 engine. For a typical aircraft mission, 

such an engine design would have a 3 stage power turbine, 

optimized for operation around 90 to 100% rpm and limited 

capability outside this range (much like the COTS engine). 

But due to recent VSPT research efforts4

, Rolls-Royce 

generated performance data for this engine assuming VSPT 

technology, optimized around 90% rpm. 

The fixed geometry VSPT design includes an extra 

power turbine stage which was used in the overall design to 

improve performance and operability over the variable speed 

range with only minimal additional weight and complexity. 

This PD628 engine was rated at essentially the same max 

power at 100% RPM and sea level standard (SLS) 

conditions as the previous three engines. As shown in Table 

2, the PD628 engine with its fixed geometry VSPT weighs 

213 lb less than the 2035 engine with a variable geometry 

VSPT – a substantial 20% weight reduction. And it was 40% 

lighter than the 2015 COTS engine. 

Engine power and fuel flow at reduced engine RPM was 

a primary focus of the study. Figure 2 showed an example 

of MRP and MCP shaft horsepower (SHP) for the 2035 

PD647 engine at 77% and 54% RPM. In this case, available 

SHP actually increased at reduced RPM, relative to 100% 

RPM, for this advanced engine with a variable geometry 

power turbine, an opposite trend from normal engine 

performance at reduced RPMs. 

Engine data from each case was reformatted for use 

with the aircraft sizing tool. Representative fuel flow and 

residual thrust curves were derived. Fuel flow at the power 

required was obtained from polynomial curve fits to referred 

fuel flow versus referred power.

Figure 5 summarizes the relative fuel flow versus SHP 

of the 2015, 2025 and the 2035 variable geometry VSPT 

engine, compared to each other at the three cruise RPMs 

considered in this study. 

Fuel flow characteristics are seen to change 

significantly with engine RPM, and with engine technology. 

The 2015 engine fuel flow increased some at 77% RPM, but 

dramatically increased at 54% RPM. While the 2025 engine 

was previously shown to be heavier, its fuel flow 

characteristics actually improved at reduced RPM. Lastly, 

the 2035 engine with its variable speed power turbine 

(VSPT) has substantially lower fuel flow than either of the 

other two engines at all RPMs and decreases some at lower 

RPM. And it had a higher HP/lb, shown in Table 2. 

54% Engine RPM

Shaft Horsepower

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

77% Engine RPM

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

100% Engine RPM

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

54% Engine RPM

Shaft Horsepower

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

54% Engine RPM

Shaft Horsepower

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

77% Engine RPM

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

100% Engine RPM

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

100% Engine RPM

Absolute Fuel Flow, lb/hr

2015 Engine

2025 Engine

2035 Engine

Figure 5: Comparison of Three Engines’ Fuel Flow 

versus Horsepower.

PROP-ROTOR PERFORMANCE MODELS 

 NASA performed extensive studies to refine the design 

of the LCTR2 rotor system, including aeroelastic, 

performance and dynamic analyses. The reference LCTR2 

rotor is a four-bladed, 65 ft diameter rotor, with an overall 

taper ratio of 0.70, a bi-linear blade twist of -38°/-30°, and a

thrust-weighted solidity of 0.133.

 Using the NASA defined LCTR2 rotor blades, the 

current study re-evaluated rotor performance with standard 

Boeing tools within the scope of the project. The geometric 

twist distribution of the LCTR2 blade was maintained for the 

350 fps cruise Vtip rotor and Boeing’s predicted rotor 

performance was very close to that from NASA.

Boeing continued to use NASA airfoil data, the LCTR2 

radial distribution of airfoils, blade planform and rotor 

solidity to define rotor designs for the two additional rotors 

with cruise Vtip of 500 fps and 650 fps. They were given 

modified twist distributions to align the local blade chordline 

with the oncoming flow at the nominal design cruise 

condition of 310 ktas at their respective cruise tip speeds. In 

accordance with the statement of work, no blade 

	

	
	 
optimization was performed to further refine the resulting 

twist distributions for the cruise condition or to balance the 

design for hover performance. 

NASA supplied ‘C81’ format airfoil data for the 

LCTR2 rotor design, which Boeing converted to a format 

required for the Boeing B08 rotor performance analysis. 

The airfoil tables were installed as library files available to 

the B08 program. Boeing applied the NASA blade airfoil 

performance characteristics and definition of relative chord 

throughout this study, and the LCTR2 rotor solidity was also 

preserved. Absolute chord lengths changed with the rotor 

radius as a result of resizing the aircraft. 

A comparison of the twist distributions for the three 

rotor designs are shown in Figure 6, and are compared to the 

helical inflow angles for each rotor, operating at 310 ktas. 

The NASA bi-linear twist for the LCTR2 rotor with the 350 

fps cruise Vtip closely agrees with the helical inflow angle 

(twist atan (μ/x)) with a bi-linear twist (-38°/-30°). 

The blade design for the rotor with 500 fps cruise Vtip 

rotor had a bi-linear twist (-50°/-34°) to closely match the 

helical inflow distribution at 310 ktas. It retained the LCTR2 

solidity, reference blade planform and airfoil distribution. 

A bi-linear twist distribution proved to be inadequate to 

properly align the blade for the 650 fps cruise Vtip and a trilinear twist was used instead. Blade twist for the 650 fps 

rotor cruise Vtip was (-63°/-42°/-33°) for good cruise 

efficiency.

-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)

-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)

at

/x)

Figure 6: Comparison of Rotor Blade Twist 

Distributions 

Rotor performance in hover and cruise for the objective 

rotor cruise Vtip was evaluated with Boeing’s B08 blade 

element/momentum theory program for static and axial 

flight proprotor performance. The method incorporates the 

effect of tip loss associated with a finite number of blades 

through Prandtl’s tip loss correction. Tip compressibility 

relief associated with three-dimensional flow effects near the 

tips is treated using the Lenard correction. The B08 program 

was considered sufficient for this trade study, since proprotor cruise efficiency is dominated by blade profile drag 

with relatively low induced drag, and cruise fuel is the 

dominant portion of mission fuel. 

Rotor hover performance for each rotor design Vtip was 

modeled as tables of Figure of Merit (FM) versus the hover 

thrust coefficient (CT) at the LCTR2 hover Vtip of 650 fps 

and takeoff condition. Calculated installed hover 

performance for 500 fps and 350 fps designs are shown in 

Figure 7 at the LCTR2 takeoff condition of 5,000’,

ISA+20°C, all at 650 fps hover Vtip. The estimates of 

isolated hover performance were reduced by 4% thrust to 

account for installation effects. 

Rotor cruise performance was modeled as tabulated 

cruise propulsive efficiency (η) versus advance ratio (μ) and 

thrust coefficient (CT), for each rotor design cruise Vtip.

The rotor solidity (σ) matches the NASA LCTR2 design – a

result of preserving the LCTR2 values of CT/σ, disc loading, 

and hover Vtip. An essential element of the sizing model 

was to capture differences in rotor cruise propulsive 

efficiency for each rotor’s design cruise Vtip. Maps of rotor 

cruise efficiency from the B08 analysis are presented in 

Figure 8.

Blade Loading CT/ TW

Twist for 500 fps Cruise Tip Speed

Isolated

Installed

T

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Twist for 500 fps Cruise Tip Speed

Isolated

Installed

Hover

Figure of

Merit

FM = CT /(√2CP)

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Hover

Figure of

Merit

2CP

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Twist for 350 fps Cruise Tip Speed 

Isolated

Installed

3/2

Hover

Figure of

Merit

FM = CT /(√2CP)

Hover

Figure of

Merit

2CP

3/2

Blade Loading CT/ TW

Twist for 500 fps Cruise Tip Speed

Isolated

Installed

T

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Twist for 500 fps Cruise Tip Speed

Isolated

Installed

Hover

Figure of

Merit

FM = CT /(√2CP)

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Hover

Figure of

Merit

2CP

0.40

0.45

0.50

0.55

0.60

0.65

0.70

0.75

0.80

0.85

0.90

0.00 0.04 0.08 0.12 0.16 0.20

Twist for 350 fps Cruise Tip Speed 

Isolated

Installed

3/2

Hover

Figure of

Merit

FM = CT /(√2CP)

Hover

Figure of

Merit

2CP

3/2

Figure 7: Rotor Hover Figure of Merit; 500 and 350 fps 

Cruise Tip Speed Design

	

	
	 
Figure 8: Rotor Propulsive Efficiency for 350 and 500 

fps Cruise Tip Speed Designs 

DRIVE SYSTEM MODELS 

As a tiltrotor vehicle, the LCTR2 drive system 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 

four engines, 2 engines at each nacelle. The notional 

baseline drive system for this study consists of 5 

transmissions – A left hand (LH) and right hand (RH) 

Proprotor Gearbox (PRGB, borrowing V22 nomenclature), 

LH and RH Tilt Axis Gearboxes (TAGB) and a Mid-Wing 

Gearbox (MWGB) for cabin accessory power. 

The PRGB transmissions are power-combining 

transmissions which collect power from the 2 engines (per 

nacelle) and deliver power to the rotor system. The PRGB 

transmissions are located near the rotor system to minimize 

the weight of the heavy rotor shaft. The TAGB 

transmissions are located on the nacelle tilting axis which is 

assumed to be aft of the wing rear spar similar to the V22. 

For operational scenarios where all the rotor speed reduction 

is accomplished with engine speed variation (like the V-22),

a single ratio transmission is required, see Figure 9.

191 rpm 191 rpm

2.4 : 1 3.1 : 1

Configuration #1 

 Fixed Ratio 100% Speed

3 : 1 

2.5 : 1

1.54 : 1

8,000 RPM

3.1 :1 2.4 : 1

3:1

2.5 : 1

1.54 : 1

PROP ROTOR GEARBOX

MID-WING

GEARBOX FOR 

ACCY POWER

TILT AXIS GEARBOX

GEN

PUMP

5,194 rpm

3.5:1 3.5:1

5,194 rpm

15,000 rpm 15,000 rpm

Same engine configuration 

as shown on the left, but 

out-of-plane 

15,000 rpm 15,000 rpm

Same engine configuration 

as shown on the right, but 

out-of-plane 

Figure 9: LCTR2 Single Speed (Direct Drive) Drive 

System Schematic Diagram with Helical Gears 

To satisfy the reduced rotor Vtip in cruise segments of 

the LCTR2 mission, a variable or multi-speed configuration 

is needed. This configuration is shown in Figure 10. Speed 

changing modules are located at the input stage of the PRGB 

transmissions for all configurations in this study. This 

requires 4 speed changing modules, one at each engine input 

shaft. This configuration is potentially the lightest weight 

and most flexible configuration for speed changing events. 

There are additional benefits with this location in that the 

modules would be accessible and repairable since they can 

be configured as a ‘line replaceable unit’

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 10: LCTR2 2-Speed Drive System Schematic 

Diagram with Helical Idler gears 

Characteristics of the notional drive systems are: 

Speed changing gearboxes are located in the high 

speed portion of the drive train to minimize weight 

impacts for those devices. Engine input speed is 

based on a maximum of 15,000 RPM for all engines. 

A Helical Idler geartrain is used to transfer power 

from engines to Bull Gear, Planetary Systems and 

Rotor Shaft. 

	

	
	 
Output Planetary System reduction ratios are 

moderate to low to allow for a rotor shaft that 

extends through the gearbox and is supported by a 

bearing in the base of the Proprotor Gearbox, similar 

to the V22. 

A Mid Wing gearbox is required to provide auxiliary 

power for control systems and cabin environmental 

and electrical requirements. 

Potential location for the over-running clutch is after 

the speed changing gearbox so that a failure in the 

engine or speed changing gearbox can be isolated 

from the remaining functional propulsion system. 

The LCTR2 four-engine configuration may appear to be 

more complex than a two-engine tiltrotor configuration but 

the four-engine configuration has some distinct advantages. 

In the event of an engine failure, the one engine inoperative 

(OEI) power available from the remaining engines is only

marginally less than with ‘all engines operating’ (AEO) and 

the power transfer through the wing shafting is assumed to 

be less in this study. This results in a lighter weight wing 

shaft system. There are also perceived benefits in the speed 

changing mechanisms, even though there are more speed 

changing boxes needed. With this distributed system, it may 

be easier to implement a (modified) sequential shifting 

strategy similar to the method described in NASA Report 

TM 2007-2148425

. 

The concepts study for multi-speed systems was not 

exhaustive, but the scope was sufficient to support the 

integration and optimization for an LCTR2 scaled aircraft. 

Criteria used to evaluate potential multi-speed transmissions 

in this study include the following: 

The desired speed shifting range is 54%, which 

corresponds to the rotor tip speed reduction from 

650 fps to 350 fps. Additional reduction ranges of 

70% and 77% were defined to provide a mid-range 

data point in the study at 500 fps rotor speed. In this 

report the ratio (factor) between low and high speed 

reduction ranges will be referred to as the “speed 

change ratio”, which is the 54% or 77% goals noted 

above. 

Overall reduction ratios for the speed changing unit 

must be kept low to reduce the weight in the 

remainder of the drive system components. For 

example, it is preferable to have a speed changing 

module that varies between a ratio of 2 and 4 than a 

module that varies between 4 and 8. This is 

particularly true with the series of helical idler gears 

that are located in the Proprotor Gearbox, since a 

high reduction ratio speed changing module would 

present a larger torque to this train and each gear 

weight would increase. 

To meet the above criteria, the speed changing 

mechanisms considered in this study were based on 

compound planetary systems that can be enabled with one 

control input. Either a ring gear or carrier is restrained by 

an active (multiple disk) clutch, causing the gear ratios to 

change. Figure 11 shows a schematic arrangement 

‘Configuration B’ that proved favorable for weight and 

operating characteristics. This configuration was practical 

for a large ratio change while maintaining a lower overall 

reduction ratio. Planet speeds were considered reasonable 

and this configuration worked well with the full LCTR2 

drive system as shown in previous diagrams. 

15,000 RPM

INPUT SHAFT

SPRAG CLUTCH

RING GEAR #1

FRICTION CLUTCH

CARRIER

PLANET

OUTPUT SHAFT

Figure 11: Speed Changing Planetary Schematic. 

Similar arrangements were suggested in the NASA 

sponsored study described in report CR-2002-2115646

. 

Relevant concepts were also discussed in CR-2002-2115637

and in TM-2008-2152768. Boeing has recent experience in 

this area from the A160 program where a 2 speed main rotor 

transmission is currently in limited production. 

Drive System analysis included evaluation of drive 

system losses for the configurations used in the sizing study 

as noted above. The drive system power losses were 

evaluated for the cruise rotor speed condition for each 

configuration, since cruise segments dominated the defined 

mission, and differences for hover conditions were 

considered in the study. Power loss was calculated using 

methods based on test experience gathered from previous 

programs. This method assigns a loss factor per mesh based 

on the type of gearing with an adjustment factor for gear 

speed. The loss factor includes windage, bearing friction, 

seals and other losses. Power loss for the high speed (helical 

idler) portion of the rotor gearbox was studied in greater 

depth since it is an area of significant power losses for the 

V22 drive system. Information was extrapolated from a 

NASA technical memorandum9

. Table 3 and Table 4

summarize the weight and power losses of various 

configuration combinations. 

	

	
	 
Table 3: COTS Drive System Weights 

and Cruise Power Losses 

Configuration

Speed, %

Engine, %

Drive, %

Rotor RPM

Tip Speed, fps

Trend Weight, lbs, 

Current Production

COTS Weight, lbs, 

Technology Factor 

0.8

COTS Power Loss 

at Cruise Speed, %

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, %

Engine, %

Drive, %

Rotor RPM

2025 Weight, lbs,

Technology Factor 

0.75

2035 Weight, lbs,

Technology Factor 

0.70

2025 Power Loss, %. 

Technology Factor 

0.95

2035 Power Loss, %. 

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

Configuration B consists of a Sun Gear # 1 as the input 

and a Sun Gear # 2 as the output. Speed changing is 

accomplished by holding either Ring Gear # 1 or the Carrier 

stationary with clutches while the other rotates freely. In this 

case, a (spring apply, hydraulic pressure release) friction 

clutch is used to stop and hold the ring gear during hover 

while a sprag clutch is used to hold the carrier stationary for 

cruise condition. Figure 12 and Figure 13 show recent 

progress for the speed changer concept sketches. 

Figure 12: Speed Changer Gearbox: Isometric View 

Figure 13: Speed Changer Gearbox: Section View 

Models have been generated for most of the major 

components and supporting analysis done to size gears, 

bearings and clutch elements. Details regarding this drive 

systems configuration will be reported in later works. 

VEHICLE SIZING APPROACH 

NASA provided reference values for the LCTR2 aircraft 

dimensions, empty weight, mission fuel, and empty 

weight/gross weight ratios (EW/GW), rotor performance and 

mission performance. This data provided the basis for all 

drag and performance calculations. 

The NASA LCTR2 fuselage size was used throughout 

the study. The component weights were scaled up or down 

with size relative to the Boeing weight estimate for the 

baseline LCTR2 design, using VASCOMP parametric 

weight relationships. Structural weights were based on 2025 

technology throughout this study to avoid confusing the 

results by introducing another variable. 

Boeing aircraft drag was primarily based on the LCTR2 

reference data, scaling the wing profile drag with wing area 

and using a slightly modified efficiency for induced drag. 

Rotor diameter, wing span and area, and horizontal tail area 

changed with GW, maintaining the LCTR2 disk loading and 

wing loading. 

Model assumptions relative to the LCTR2 configuration 

design are listed below: 

LCTR2 wing loading, sweep, aspect ratio and taper 

ratio were preserved. Wing area depended on GW. 

LCTR2 wing tip extensions and span were 

preserved. 

The LCTR2 rotor tip clearance from the fuselage 

side to inboard rotor tip was preserved. 

LCTR2 fuselage size was preserved, including 

diameter, length, wetted area, and tail moment arm. 

	

	
	 
LCTR2 horizontal tail volume coefficient was 

preserved. Horizontal tail area depended on wing 

area and MAC (mean aerodynamic chord). 

LCTR2 rotor hover Ct/σ, disc loading, and number 

of blades were preserved. Solidity was therefore 

preserved. Rotor diameter depended on GW. 

Model assumptions relative to the LCTR2 performance 

are listed below: 

Retained NASA hover download/Thrust 

Retained NASA fuel flow conservatism of 5% 

Equivalent flat plate area (fe) was scaled from the 

NASA fe of 34.18 sq.ft. according to the area of the 

wing and tail surfaces. Fuselage fe was retained for 

the NASA LCTR2 fuselage. A VASCOMP value 

was used for the Oswald induced drag factor. 

Transmission was sized to Hover out of Ground 

Effects or Cruise torque (cruise for low Vtip). 

Retained NASA mission profile, fixed equipment, 

and payload weight for 90 passengers with baggage. 

HP available for climb and cruise was limited by 

transmission rating at cruise RPM. 

The LCTR 4-engine arrangement was preserved. 

The one-engine-inoperative (OEI) performance was 

preserved (90% of HOGE SHP is obtained with an 

assumed 20% contingency power for 4 engines, 

when engines are sized to HOGE at the design GW). 

The LCTR2 limit load factor of 3.0 was preserved at 

the design takeoff GW. 

Aircraft size and performance was evaluated with an 

Excel performance and sizing tool, This excel tool modeled 

most of the VASCOMP performance and sizing procedures 

in a format that allowed Boeing to perform “Concept 

Evaluation” analysis for the LCTR2 air vehicle. Boeing 

generally uses the VASCOMP sizing program10 to evaluate 

aircraft size and performance for tiltrotor type aircraft. 

However, the work to be performed in this study required 

evaluation at different combinations of engine RPM and 

drive system RPM, which are not independently modeled in 

VASCOMP. The alternative spreadsheet approach was 

chosen to utilize data from various sources and formats 

without the expense of modifying standard Boeing tools and 

engine decks to achieve the same result. 

Mission performance was evaluated with standard 

performance equations for hover, climb, and cruise at 

specific airspeed and altitude. Rotor cruise performance was 

modeled with table lookup routines of cruise efficiency 

versus rotor thrust coefficient and advance ratio, similar to 

VASCOMP. Rolls-Royce engine shaft horsepower available 

data was tabulated at MRP, IRP and MCP versus altitude 

and Mach number (all climb and cruise flight segments were 

at ISA conditions). 

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 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 gross 

weight 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, or cruise airspeed, or to evaluate other 

mission ranges. The LCTR2 sizing mission profile is 

described in Figure 14. 

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 

nm

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

30 minute reserve fuel at Vbr, 10,000’/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 14: Mission Profile

VEHICLE SIZING RESULTS 

All LCTR2 sizing cases were run with the NASA 

choice of a 650 fps rotor hover tip speed. The rotor cruise tip 

speed of 500 fps was obtained by either; 

A 2-speed gearbox (77%) with the engine at 

100%RPM. 

A single-speed gearbox with the engine at 70%

RPM. 

	

	
	 

The rotor cruise tip speed of 350 fps was obtained by either; 

A 2-speed gearbox (54%) with the engine at 100% RPM 

A 2-speed gearbox (70%) with the engine at 77% RPM 

A single-speed gearbox with the engine at 54% RPM. 

Sizing With The 2015 Engine 

Sizing results for the COTS engine study matrix are 

presented as bar graphs in Figure 15 and Figure 16. The 

three cases at 350 fps Vtip (54% of hover RPM) examines

the effect of engine RPM reduction versus drive system 

RPM reduction. The engine was sized by hover power 

required, except for the case with the engine operating at 

54% RPM, pointing out the need for an engine design with 

improved cruise performance at this low cruise RPM.

The gross weight at 54% engine RPM (350 fps rotor 

cruise Vtip) was driven up by an 11% increase in required 

fuel relative to the 100% engine RPM case. Notably, the 

engine was sized by cruise power required at this 54% 

engine RPM, and it required more installed SHP than either 

the 100% or the 77% engine RPM cases. The least takeoff 

GW for the 350 fps rotor cruise Vtip was at the intermediate 

condition of 77% engine RPM, although that did not 

demonstrate much improvement from the 100% engine 

RPM. 

While the 650 fps rotor cruise Vtip had the second 

highest GW, it was in fact no worse than the result for the 

350 fps case with a two-speed gearbox, even though the 

helical tip speed in cruise was 840 fps (M 0.82) at 650 fps 

Vtip and cruise airspeed. Installed SHP was still determined 

by the hover condition for this case, with a simple singlespeed transmission. However, not surprisingly, it had the 

lowest rotor cruise efficiency and therefore required more 

mission fuel than most other cases.

The minimum GW solution occurred at the 500 fps 

cruise Vtip, not for the objective 350 fps cruise Vtip. Both 

of the 500 fps cruise Vtip cases resulted in lighter overall 

GW than the other four cases. The 500 fps rotor cruise Vtip 

had nearly the same rotor cruise propulsive efficiency as the 

350 fps cases (0.839 vs. 0.845), and it gave the lightest EW 

and GW. The higher rotor tip speed (500 fps) had a 9%

lower drive system weight than the 350 fps cases, reflecting 

about 30% less output torque required at the higher rotor tip 

speed. The lightest GW solution was for the 500 fps rotor 

cruise Vtip with a light weight single-speed drive system and 

the engine operating at 77% RPM, which did not carry the 

fuel flow penalty of engine operation at 54% RPM. It’s GW

was 2,600 lb lighter than the 350 fps rotor cruise Vtip case at 

100% engine RPM. 

In general, the following may be concluded from the 

study with the COTS engine. 

Gross Weight variation was less than expected at the 

different rotor cruise tip speeds. 

Sizing the engine SHP to hover resulted in smaller 

engines than the NASA LCTR2 (different criteria). 

Boeing transmission weights and rotor weights were 

generally higher than NASA LCTR2. 

Sensitivity to design cruise airspeed was found to have 

as much effect on GW as rotor cruise tip speed. 

Two-speed transmissions were a more efficient means 

of obtaining the 350 fps rotor Vtip than reducing the 

engine RPM. 

Reduced Engine RPM was equally as efficient as a 2-

speed transmission for the 500 fps cruise Vtip. 

The 500 fps rotor Vtip resulted in lower GW than the 

350 fps rotor Vtip, suggesting that the optimum rotor 

cruise tip speed may lie near 500 fps for a 310 ktas 

cruise airspeed, when engine fuel flow and drive 

system weights are properly accounted for. 

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

350 350 350 500 500 650

Gross Weight, lb,

Empty Weight, lb

Gross Weight

Empty Weight

Cruise Tip Speed, fps

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2015 Engine Technology

NASA LCTR2 SDGW

64.6% 65.2% 65.1% 64.1% 64.6% 63.6%

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

350 350 350 500 500 650

Gross Weight, lb,

Empty Weight, lb

Gross Weight

Empty Weight

Cruise Tip Speed, fps

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2015 Engine Technology

NASA LCTR2 SDGW

64.6% 65.2% 65.1% 64.1% 64.6% 63.6%

Figure 15: 2015 Engine - Effect of Rotor Tip Speed and 

Engine/Drive System RPM on GW 

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Installed SHP

Engine Weight

NASA LCTR2 Installed max SHP =7500 per engine

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2015 Engine Technology

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Installed SHP

Engine Weight

NASA LCTR2 Installed max SHP =7500 per engine

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2015 Engine Technology

Figure 16: 2015 Engine - Installed SHP and Weight 

	

	
	 
Sizing With The 2025 Engine 

The concept for the 2025 engine was to accept a 

relatively small increase in engine weight to gain a large 

expected benefit from more efficient fuel burn. The 2025 dry 

engine weighed more than the 2015 engine (Table 2), but its 

performance was a major improvement over the COTS 

engine for operations at reduced RPM. It displayed increased 

MCP power available and lower SFC at reduced RPM, 

whereas the COTS engine lost significant MCP power and 

suffered increased SFC at 54% RPM, typical of current 

engines, refer back to Figure 5. 

Sizing results for LCTR2 with the 2025 EIS engine 

(PD646_11751) also reflect 2025 drive system technology 

that reduced drive system weight and had lower drive system 

losses. Sizing results, summarized in Figure 17, show the 

2025 engine resulted in an overall 2% to 7% increase in 

aircraft gross weight, relative to the 2015 engine. 

64.6% 65.2% 65.1% 64.1% 64.6% 63.6%

NASA LCTR2 SDGW

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

310 KTAS Cruise Airspeed, 25,000 ft

2025 Engine Technology

Gross Weight

Empty Weight

64.6% 65.2% 65.1% 64.1% 64.6% 63.6%

NASA LCTR2 SDGW

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

310 KTAS Cruise Airspeed, 25,000 ft

2025 Engine Technology

Gross Weight

Empty Weight

Figure 17: 2025 EIS Engine - Effect of Rotor Tip Speed 

and Engine/Drive System RPM on GW 

The engine was sized by hover for all cases except for 

the 650 fps cruise Vtip case. Gross weight at 100% engine 

RPM is nearly the same for 650 fps cruise Vtip and for 350 

fps cruise Vtip, reflecting the 2025 engine’s significantly 

higher fuel flow at 100% engine RPM shown in Figure 17. 

The GW trend at 350 fps rotor cruise Vtip was very 

different than the COTS engine. GW from the COTS engine 

cases increased with reduced engine RPM, but GW for the 

2025 engine actually decreased with reduced engine RPM, 

owing to the significant fuel efficiency from the 2025 

engine’s variable geometry power turbine. The lowest GW 

solution came from the 500 fps Vtip rotor, with improved 

2025 engine fuel efficiency at 77% engine RPM, coupled 

with a light weight single-speed transmission. 

Installed SHP and engine weight for the 2025 engine are 

shown in Figure 18 with a trend similar to the GW trend. 

Engine weights are considerably more than for the 2015 

engine, due to the double effect of more installed SHP due to 

the higher GW and the lower HP/weight ratio of the 2025 

engines. 

350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

Installed SHP

Engine Weight

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP,

Engine Weight, lb

310 KTAS Cruise Airspeed, 25,000 ft

2025 Engine Technology

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

Installed SHP

Engine Weight

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP,

Engine Weight, lb

310 KTAS Cruise Airspeed, 25,000 ft

2025 Engine Technology

Cruise Tip Speed, fps

Figure 18: 2025 EIS Engine - Installed SHP and Weight

Sizing With The 2035 Variable Geometry VSPT Engine 

The LCTR2 was resized again for the advanced RollsRoyce 2035 EIS variable speed power engine (PD647-

11772) with VAATE technology and its variable geometry 

power turbine. These cases applied the new engine 

performance and weight, and the estimated weight and 

efficiency for a 2035 drive system. Aircraft structural 

weights remained based on 2025 technology and the same 

three rotor designs were used. Results are shown in Figure 

19 and Figure 20. 

Fuel flow of the 2035 engine was significantly less than 

either the COTS engine or the 2025 engine at all operating 

RPMs. And the 2035 engine was significantly lighter; 

weighing 25% less than the COTS engine (per shp), and 

34% less than the 2025 engine (per shp). The combination of 

reduced mission fuel and reduced engine weight had a 

dominant effect on LCTR sizing. Gross weight results from 

the 2015 engine (Figure 15) ranged from 105,700 to 110,600 

lb, and results from the heavier 2025 engine (Figure 17) 

ranged from 108,000 to 114,700 lb. But the double benefit of 

reduced fuel and reduced engine weight for the 2035 

variable geometry VSPT engine substantially reduced 

aircraft GW for all cases. GW for the lighter, more fuel 

efficient 2035 engine (Figure 19) ranged from 93,500 to 

97,500 lb, a remarkable 14% average reduction in GW. 

Overall, the 2035 engine fuel flow was much less 

sensitive to operating RPM than either of the previous 

engines, resulting in very little variation in GW across the 

combinations of engine and drive system RPM. Once again, 

the lowest GW was for the 500 fps rotor Vtip with a 77% 

engine RPM and the lighter weight single-speed drive 

system. That was closely followed by the 350 fps rotor Vtip 

with a 54% engine RPM and a single-speed drive system.

	

	
	 	
Figure 20 displays installed SHP and the weight of one 

engine for the six combinations of rotor cruise Vtip, engine 

and drive system RPM. As with the other engine 

technologies, the trend of installed SHP follows the GW. 

NASA LCTR2 SDGW

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Variable Geometry Power Turbine

Gross Weight

Empty Weight

NASA LCTR2 SDGW

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Variable Geometry Power Turbine

Gross Weight

Empty Weight

Figure 19: 2035 Variable Geometry Engine - Effect of 

Rotor Tip Speed and Engine/Drive System RPM on GW 

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

NASA LCTR2 Installed max SHP =7500 per engine

Installed SHP

Engine Weight

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Variable Geometry Power Turbine

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

NASA LCTR2 Installed max SHP =7500 per engine

Installed SHP

Engine Weight

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Variable Geometry Power Turbine

Figure 20: 2035 Variable Geometry Engine - Installed 

SHP and Weight 

Sizing With the 2035 Fixed Geometry VSPT Engine 

The fourth and final engine explored in this study was

the Rolls-Royce’s PD628, a 2035 EIS engine with a fixed 

geometry VSPT, previously discussed in this paper. This 

was the lightest engine of the four, with a HP/lb ratio of 10, 

while retaining very good specific fuel consumption. 

Figure 21 compares the ratio of MCP cruise power 

available at 25,000 ft/ISA/Mach 0.5 to the Max Rated Power 

(takeoff) at SLS, across the spectrum of operating RPM for 

all four engines. Clearly the 2035 EIS engine with a variable 

geometry power turbine excelled in this measure, achieving 

more power at reduced RPM than it did at 100% RPM. And 

its excellent fuel consumption resulted in considerably 

lighter GW, evidenced in Figure 19. But the substantial 

cruise power available over-shot the LCTR2 need for cruise 

power. So, good as it is, it may not provide the ideal hovercruise power match. 

The PD628 engine with a fixed geometry VSPT 

provided about 12% more cruise power than the COTS 2015 

engine at the very low 54% RPM, but far less than the 2035 

engine with the variable geometry power turbine. Still, the 

PD628 turned out to be the best overall engine for LCTR2, 

providing sufficient cruise power with a very light engine 

(HP/lb = 10). 

0.40

0.45

0.50

0.55

0.60

0.65

0.70

50% 60% 70% 80% 90% 100% 110%

Engine RPM, %

Ratio of: M

CP 

@ 25,000', Mach 0.5 

Max Rated Power, SLS 

2035 Engine, Std PT

2035 Engine, Variable Geom PT

2025 Engine

2015 Engine

Relative HP Available At 25,000 ft, Mach 0.5

0.40

0.45

0.50

0.55

0.60

0.65

0.70

50% 60% 70% 80% 90% 100% 110%

Engine RPM, %

Ratio of: M

CP 

@ 25,000', Mach 0.5 

Max Rated Power, SLS 

2035 Engine, Std PT

2035 Engine, Variable Geom PT

2025 Engine

2015 Engine

Relative HP Available At 25,000 ft, Mach 0.5

Figure 21: Relative MCP Cruise Power Available 

That is a significant lesson to be taken from examining 

these multiple engines, drive systems and rotor tip speeds, 

even if not a surprising one. The best engine for the aircraft 

is the one that best fits the rotorcraft’s particular hover and 

cruise requirements, while providing low fuel consumption. 

Features that provide excess cruise power that add to engine 

weight, which cannot effectively be used in flight, may not 

pay their way into the aircraft design. 

The LCTR2 was resized with the PD628 engine to 

quantify the net benefit of reduced engine weight, improved 

fuel flow, and cruise power available. The six combinations 

of rotor tip speed, engine cruise RPM, and drive system 

speed reduction were run, and are shown in Figure 22 and 

Figure 23. Key points from the PD628 engine sizing cases 

are: 

The engine was sized by the hover condition in all 

cases, showing that the PD628 engine’s MCP 

cruise power available was sufficient for the 

LCTR2. 

The 2035 advanced technology engine reduced fuel 

flow and engine weight significantly, reducing GW 

from about 106,000 lb GW for the 2015 engine 

down to about 92,000 lb for the 2035 PD628

engine. That GW difference is equivalent to another 

53 passengers. 

	

	
	 
The lightest GW design was 91,923 lb, and 

continued to be for the 500 fps cruise Vtip with a 

single-speed transmission and 77% engine RPM.

This was 1540 lb lighter than with the 2035 engine 

with variable geometry power turbine. 

The second lightest GW design was 91,989 lb, also 

for the 500 fps cruise Vtip, but with a 2-speed 

transmission and 100% engine RPM. 

The 350 fps rotor Vtip cases sized very close to 

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

exhibiting little sensitivity to the combination of 

engine RPM-drive system RPM, reflecting the 

ability of the PD628 engine to operate efficiently 

over a broad RPM range. 

Installed SHP was nearly flat at about 4500 HP per 

engine, as shown in Figure 21. The installed SHP 

was between 0.192 and 0.195 times the aircraft GW 

for all six cases. 

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

NASA LCTR2 SDGW

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Fixed Geometry Power Turbine

Gross Weight

Empty Weight

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

50,000

60,000

70,000

80,000

90,000

100,000

110,000

120,000

Gross Weight, lb,

Empty Weight, lb

NASA LCTR2 SDGW

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Fixed Geometry Power Turbine

Gross Weight

Empty Weight

Figure 22: 2035 Fixed Geometry VSPT Engine – Effect 

of Rotor Tip Speed and Engine/Drive System RPM on 

GW

1,000

1,500

NASA LCTR2 Installed max SHP =7500 per engine

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Fixed Geometry Power Turbine

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

Installed SHP

Engine Weight

1,000

1,500

NASA LCTR2 Installed max SHP =7500 per engine

310 KTAS Cruise Airspeed, 25,000 ft

2035 Engine, Fixed Geometry Power Turbine

2,000

2,500

3,000

3,500

4,000

4,500

5,000

5,500

6,000

Installed SHP, 

Engine Weight, lb

Cruise Tip Speed, fps 350 350 350 500 500 650

Engine % Cruise RPM

Drive System % RPM

100 77 54 100 77 100

54 70 100 77 100 100

Installed SHP

Engine Weight

Figure 23: 2035 Fixed Geometry VSPT Engine - 

Installed SHP and Weight 

Effect of Empty Weight, Fuel and Gross Weight 

The ratio of empty weight/GW and fuel weight/GW are 

important indicators of aircraft overall efficiency. Together 

they constitute between 77% and 82% of GW. The designer 

has some control over these quantities, making them 

important measures of value. The sum of these two ratios 

plus the ratio of payload/GW and fixed useful load/GW must 

add up to 1.0, i.e. the whole GW. 

The LCTR2 mission fuel is dominated by the long 1000 

nm cruise segment. A typical distribution of mission fuel is: 

7.2% for taxi, takeoff and landing segments, 6% for climb, 

75.3% for cruise and 11.5% for the 30 nm alternate 

destination and 30 min reserve fuel segments. So LCTR2 

mission fuel is dominated by the combined cruise efficiency 

of the prop-rotor, engine, and drive system. The ratio of 

total mission fuel/GW is a good measure of these combined 

efficiencies in cruise, reflecting more fuel efficient engines 

or improved prop-rotor propulsive efficiency. Table 5

summarizes this fuel/GW ratio at all combinations of proprotor cruise Vtip, engine and drive system RPM, for all four 

engines. It also shows the EW/GW ratio for all cases. For 

comparison, the NASA LCTR2 design has a fuel/GW ratio 

of 0.183, a fallout of the engine characteristics NASA used. 

Table 5: Ratio of Mission Fuel / Gross Weight 

Rotor 

Cruise 

Vtip

350 fps 500 fps 650 

fps

Engine 

Cruise 100% 77% 54% 100% 77% 100%

Dr.Sys 

Cruise 54% 70% 100% 77% 100% 100%

2015 COTS Engine Avg

Fuel / 

GW 0.154 0.156

A

0.168 0.157 0.159 0.167 0.160

EW / 

GW 0.650 0.647 0.640 0.643 0.640 0.637

2025 Advanced Technology Engine

Fuel / 

GW

B

0.169 0.156 0.153 0.171

C

0.157 0.181 0.165

EW / 

GW 0.646 0.652 0.651 0.641 0.646 0.636

2035 Variable Geometry VSPT Engine

Fuel / 

GW 0.138 0.132 0.133 0.140 0.134

D

0.148 0.138

EW / 

GW 0.645 0.646 0.643 0.638 0.639 0.633

2035 Fixed Geometry VSPT Engine

Fuel / 

GW 0.131 0.134 0.143 0.133 0.136

E

0.141 0.136

EW / 

GW 0.643 0.640 0.633 0.636 0.633 0.632

	

	
	 
The right hand column of the table shows the average 

Fuel/GW ratio of the six cases for each engine technology. 

The 2025 engine shows a slightly higher average ratio than 

the 2015 engine. Even though advanced technology gave the 

2025 engine better performance at reduced RPM, 3 of the 6 

cases were operating at 100% RPM in cruise where it lost 

performance. The two 2035 engines with VSPT technology

had similar Fuel/GW averages, much better than either the 

2015 or the 2025 engine. 

Several points in Table 5 are selected as examples. Point 

A is a particularly high Fuel/GW ratio. The 2015 engine was 

operating at its worst condition, 54% RPM, resulting in a 

high GW, over 110,000 lb. But Figure 15 showed the empty 

weight was nearly the same as the other cases at 350 fps 

Vtip, so the GW increase came from the engine’s high fuel 

demand at a low cruise RPM. The (relatively) high 

Fuel/GW ratio for point A validates the value of the metric. 

Another observation from Table 5 is that the sum of 

Fuel/GW and EW/GW fall in a narrow band, since they must 

add up to the total fraction of (Fuel+EW)/GW. When fuel 

fraction increases, the empty weight fraction generally 

decreases, and vice-versa. Interpretation requires some 

knowledge of what physical or performance attributes 

changed from one case to another, i.e. did engine fuel flow 

decrease or did engine dry weight increase? For instance, the 

Fuel/GW ratio decreased significantly at point C because the 

2025 engine was so much more efficient at that 77% RPM 

than it was at the point B 100% RPM. 

As expected, improved performance of the two 2035 

VSPT engines show much lower Fuel/GW ratios than the 

2015 or 2025 engines, at all operating RPMs. Also 

expected, the VSPT engines display higher Fuel/GW ratios 

at 650 fps Vtip (points D and E) than at lower Vtip cases, 

driven by the low rotor cruise efficiency at 650 Vtip, and 

similar to the pattern for the 2015 and 2025 engines. 

COMPARISONS AND CONCLUSIONS 

Study objectives highlighted in this paper were threefold; 

Validate the benefit of reduced rotor cruise tip speeds 

for large civil tiltrotor (LCTR2) performance. 

Assess the tradeoffs between operating at reduced 

engine RPM versus employing a 2-speed gearbox to 

achieve reduced rotor tip speeds in cruise. 

Evaluate the potential of different engine cycles and 

advanced technology to improve power available at 

reduced engine RPM and to quantify the benefit of 

improved fuel flow.

Figure 21 compares the four engines’ ability to achieve 

that objective, in terms of power. In order of their time entry 

into service (EIS) dates, the 2015 engine lost the most power 

at reduced RPM, down to half of the takeoff power at 54% 

rpm. That low speed was the only condition where the 

LCTR2 engine was sized by the cruise power required, 

rather than the hover power, and cruise only required about 

5% more power. The 77% rpm case was still sized by hover. 

The 2025 engine showed a very substantial 

improvement in cruise power available, up to 62% of the 

takeoff power. This engine was sized by LCTR2 hover 

power requirements for both the 54% and 77% rpm design 

cases. However, the 2025 engine had slightly less cruise 

power at 100% rpm than the 2015 engine, causing the engine 

to be sized by cruise at the 100% RPM design. There is 

obviously a fine line in achieving the right balance of cruise 

power available and hover power available, and that balance 

depends on the unique power requirements of the individual 

aircraft. 

 The 2035 engine with variable geometry VSPT 

achieved remarkable results. It improved the ratio of cruise 

power available at 100 % RPM, and it excelled at reduced 

RPM, providing 66% and 67% power ratios at 77% RPM 

and 54% RPM, respectively. This 2035 technology engine 

was sized by LCTR2 hover in all the sizing cases, with 

considerably more cruise power available than was needed 

for LCTR2. These results again point out the importance of 

tailoring the ratio of cruise power available to takeoff power 

available in future studies. 

Lastly, the 2035 engine with fixed geometry VSPT 

(PD628) maintained a high ratio of cruise power available at 

100% RPM (60.5%), nearly the same as the 2035 engine 

with the variable geometry VSPT. But it dropped off in 

power as RPM decreased – to about 58% at 77% RPM and 

56% at 54% RPM. These fractions of MCP cruise power 

available at reduced RPM were higher than the 2015 engine, 

but far less than either the 2025 EIS or the 2035 variable 

geometry PT engines. The advanced technology made this 

engine 20% lighter weight than the advanced PD647 and 

40% lighter than the 2015 engine, and it offered 

substantially reduced fuel flow. 

LCTR2 sizing with the 2035 PD628 engine also 

included drive system weight reductions and improved 

efficiency projections for 2035, as for the previous 2035 

PD647 engine. All sizing cases were run at a 25,000 ft 

cruise altitude and 310 ktas, matching the original NASA 

LCTR2 design. 

Average GW for five of the six LCTR2 sized cases with 

the 2015 engine was nearly equal to the NASA LCTR2 

structural design GW of 107,124 lb. The one outlier was 

the 110,570 lb GW at the 350 fps rotor cruise Vtip 

operating at 54% engine RPM. 

The cases at 500 fps rotor cruise Vtip consistently sized 

to a slightly lower GW than any of the 350 fps cases for 

all four engine technologies. Additional investigation is 

suggested for rotor cruise tip speeds between 350 and 500 

	

	
	 
fps to determine the optimum rotor cruise tip speed from 

a vehicle sizing perspective. 

LCTR2 gross weight and empty weight with the EIS 

2035 PD628 engine were significantly less than any case 

with the 2015 COTS engine, and for most cases less than 

the 2035 PD647 engine with the variable geometry 

VSPT. 

Considering the engines configured at the narrow 

(engine) speed range, the 2015 COTS and the 2035 fixed 

geometry VSPT, variation in the sizing cases for each 

engine was 3% or less. This suggests that the benefits of 

reduced speed operation for the mission, operating 

conditions and vehicle configurations is a relatively small 

effect, whereas the benefits for engine technology from 

the COTS engine to the 2035 engine were dramatic. Fuel 

efficiency and engine weight differences between those 

engines resulted in a 13% reduction in vehicle size cases 

(at 500 fps Vtip) 

Examination of the results for the 2015 COTS and the 

2035 fixed geometry VSPT (PD628), at 500 fps and 350

fps rotor cruise Vtip at 100% engine RPM, showed the 

engine weight was 48% less with the PD628 and the 

resized aircraft needed 25% less fuel. 

Essentially equivalent benefits, as shown in the sizing 

results (GW), are derived from reduced engine RPM as

from a 2-speed transmission for the 500 fps Vtip for 

either the 2015 engine or the 2035 engine with fixed 

geometry VSPT. Results for the other two engines 

favored 100% engine RPM with the single speed 

transmission. 

Additional criteria such as operating ecomonics or 

development cost may also affect investment decisions 

and determine future direction for VSPT and variable 

speed drive system technologies, where they provide 

equivalent performance benefits. 

A summary comparison of GW and EW for the 2015 

and the two 2035 advanced technology engines is shown in 

Figure 24. The 2035 technology engines clearly provide 

significant reductions in gross weight, resulting from both 

their reduced fuel flow and reduced engine weight. 

The 2035 engine with variable-geometry VSPT gave the 

lowest GW solution when operating at the lowest engine 

RPM that was analyzed, because it provided plenty of power 

at lower operating RPM while maintaining good SFC. And, 

those cases benefited from a lighter single-speed gearbox. 

But the LCTR2 aircraft, with its relatively high cruise 

L/D, did not need or use the high MCP cruise power 

available from the variable-geometry VSPT at reduced 

RPM. LCTR2 suffered the weight penalty of the variable 

geometry engine, gaining little if any benefit from the 

engine’s much improved cruise performance. For contrast, 

the higher weight variable-geometry power turbine engine 

resulted in 31% higher installed engine weight and 10% 

more fuel than the 2035 with a fixed-geometry power 

turbine (for the 500 fps Vtip with 100% engine RPM and 

77% drive system RPM). 

The lighter weight 2035 advanced engine with a fixedgeometry VSPT gave the lowest aircraft gross weight of any 

engine for all six combinations of rotor cruise Vtip and 

engine and drive system RPM. 

The 500 fps Vtip cases continued to show up as the best 

rotor cruise tip speed in terms of aircraft gross weight and 

empty weight although the effects of reduced rotor speed 

operation is small compared to the effect of high efficiency 

engine technology. 

Cruise Tip Speed, fps 350 350 350

Engine % Cruise RPM

Drive System % RPM 54 70 100 77 100 100

60,000

70,000

80,000

90,000

100,000

110,000

120,000

50,000

500 500 650

Gross Weight, Empty Weight, lb

100 77 54 100 77 100 

310 KTAS Cruise Airspeed, 25,000 ft

NASA LCTR2 SDGW

Gross 

Weight

Empty 

Weight

with Std 

Power 

Turbine

2015 

COTS 

Engine

2035 

Engine. 

Fixed

Geometry

Power

Turbine 

Variable 

2035 

Engine. 

Geometry 

Power 

Turbine

Cruise Tip Speed, fps 350 350 350

Engine % Cruise RPM

Drive System % RPM 54 70 100 77 100 100

60,000

70,000

80,000

90,000

100,000

110,000

120,000

50,000

500 500 650

Gross Weight, Empty Weight, lb

100 77 54 100 77 100 

310 KTAS Cruise Airspeed, 25,000 ft

NASA LCTR2 SDGW

Gross 

Weight

Empty 

Weight

with Std 

Power 

Turbine

2015 

COTS 

Engine

2035 

Engine. 

Fixed

Geometry

Power

Turbine 

Variable 

2035 

Engine. 

Geometry 

Power 

Turbine

Figure 24: LCTR2 Gross Weight and Empty Weight 

Comparison for 2015 and 2035 Engines 

REFERENCES

1) 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.

2) Lenski, J and Valco, M “Advanced Rotorcraft 

Transmission (ART) Program-Boeing Helicopters 

Status Report” presented at the 27th Joint 

Propulsion Conference cosponsored by the AIAA, 

SAE, ASME, and ASEE, Sacramento, California, 

June 24-27, 1991 

	

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

Aircraft Sizing and Performance Computer 

Program, User's Manual", Boeing Rotorcraft, 

D210-13635-1, 2002. 

4) Ford, A., Bloxham, M., Turner, E. Clemens, E., and 

Gegg, S., “RTAPS VSPT Contract NNC10BA14B, 

Design Optimization of Incidence-Tolerant Blading 

Relevant to Large Civil Tilt-Rotor Power Turbine 

Applications,” NASA/CR-2010-217016, Nov, 

2010. 

5) Litt, Johathan S. Edwards Jason M, DeCastro, 

Jonathan A., A Sequential Shifting Algorithm for 

Variable Speed Control”, NASA Report TM 2007-

214842, June 2007 

6) Jules Kish, “Vertical Lift Drive System Concept 

Studies Variable Speed/ Two Speed 

Transmissions”, NASA CR-2002-211564, June 

2002

7) Robert Bossler, “Vertical Lift Drive System 

Concept Studies”, NASA CR-2002-211563, June 

2002

8) Stevens, Mark A., Handschuh, Robert, and 

Lewicki, David G. “Concepts for Variable/ MultiSpeed Rotorcraft Drive System”, TM-2008-

215276, September 2008 

9) Handschuh, R., and Kilmain, C., “Experimental 

Study of the Influence of Speed and Load on 

Thermal Behavior of High-Speed Helical Gear 

Trains”, NASA/TM—2005-213632, ARL–TR–

3488, July 2005 

10) 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. 

	

	
	 
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Design Study of Propulsion and Drive Systems for the Large Civil TiltRotor (LCTR2) 

Rotorcraft 

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6. AUTHOR(S)

Robuck, Mark; Wilkerson, Joseph; Zhang, Yiyi; Snyder, Christopher, A.; Vonderwell, Daniel 

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WBS 877868.02.07.03.07.01 

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This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Boeing, Rolls Royce, and NASA have worked together to complete a parametric sizing study for NASA’s Large Civil TiltRotor (LCTR2) 

concept 2nd iteration. Vehicle gross weight and fuel usage were evaluated as propulsion and drive system characteristics were varied to 

maximize the benefit of reduced rotor tip speed during cruise conditions. The study examined different combinations of engine and gearbox 

variability to achieve rotor cruise tip speed reductions down to 54 percent of the hover tip speed. Previous NASA studies identified that a 54 

percent rotor speed reduction in cruise minimizes vehicle gross weight and fuel burn. The LCTR2 was the study baseline for initial sizing. 

This study included rotor tip speed ratios (cruise to hover) of 100, 77, and 54 percent at different combinations of engine RPM and gearbox 

speed reductions, which were analyzed to achieve the lightest overall vehicle gross weight (GW) at the chosen rotor tip speed ratio.Different 

engine and gearbox technology levels are applied ranging from commercial off-the-shelf (COTS) engines and gearbox technology to entryin-service (EIS) dates of 2025 and 2035 to assess the benefits of advanced technology on vehicle gross weight and fuel burn. Interim results 

were previously reported. This technical paper extends that work and summarizes the final study results including additional engine and 

drive system study accomplishments. New vehicle sizing data is presented for engine performance at a single operating speed with a 

multispeed drive system. Modeling details for LCTR2 vehicle sizing and subject engine and drive sub-systems are presented as well. This 

study was conducted in support of NASA’s Fundamental Aeronautics Program, Subsonic Rotary Wing Project. 

15. SUBJECT TERMS

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

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