NASA/CR–2010-216384

Technology Assessment for Large Vertical-Lift 

Transport Tiltrotors

Peter J. Germanowski, Brandon L. Stille, and Michael P. Strauss

Sikorsky Aircraft Corporation

Stratford, Connecticut

May 2010
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NASA/CR–2010-216384

Technology Assessment for Large Vertical-Lift 

Transport Tiltrotors

Peter J. Germanowski, Brandon L. Stille, and Michael P. Strauss

Sikorsky Aircraft Corporation

Stratford, Connecticut

May 2010

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000
Available from:

NASA Center for AeroSpace Information National Technical Information Service

7115 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161

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

Acknowledgments

This report was written and compiled by the following authors:

Peter Germanowski Technical Lead

Brandon Stille Lead Designer

Michael Strauss Attributes Manager

With technical contributions from:

Steve Behnfeldt Program Manager

Andre Lewis Systems Engineer

S. Jon Davis Aerodynamics

Robert Moffitt Aerodynamics

Stephen Gatto Affordability Analysis

Gary Heiden Affordability Analysis

Greg Kiviat Affordability Analysis

Bryan Baskin Rotors & Blade Design

Ryan Craft Rotors & Blade Design

Todd Walker Rotors & Blade Design

Kris Hayden Rotors & Blade Design

Mikel Brigley Dynamics

Bryan Mayrides Dynamics

Jinlu Wang Dynamics

Cody Fegely Handling Qualities

Melvin Garelick Handling Qualities

Erez Eller Loads & Criteria

Sean Ison Mass Properties

Robert Sartorelli Propulsion

Joan Pham Reliability Analysis

Akif Ersahin Structures

William Townsend Structures

Bryan Holasek System Engineering

Thomas Sbabo Transmission Design

Todd Garcia Transmission Design

Bruce Plasse Transmission Design
iii 

Table of Contents 

List of Figures .................................................................................................................................iv 

List of Tables ..................................................................................................................................vi 

Nomenclature ................................................................................................................................ vii 

Abstract ............................................................................................................................................ 1 

1 Executive Summary ................................................................................................................... 2 

2 VDTR Historical Summary ....................................................................................................... 5 

2.1 Technology Needs for High-Speed Rotorcraft (’90–’91) ................................................ 7 

2.2 Wind Tunnel Test of a VDTR Model (’94) ..................................................................... 7 

2.3 VMS Simulation of a Variable-Diameter Tiltrotor (’97) ................................................. 8 

2.4 Variable-Diameter Retraction-Mechanism Risk Reduction and Trade Studies 

 (’98–’99) .......................................................................................................................... 9 

2.5 Variable-Diameter Tiltrotor Technology (’99–’00) ......................................................... 9 

2.6 Variable-Geometry Advanced Rotor Technology (VGART) Phase 1 (’99–’00) ............ 9 

2.7 Revolutionary Concepts in Aeronautics (REVCON) Phase 1 (’00–’01) ....................... 10 

2.8 Variable-Geometry Advanced Rotor Technology (VGART) Phase 2 (’00) ................. 11 

2.9 VDTR Measured Acoustic and Aerodynamic Data Compared with the V-22 (’02) ..... 11 

3 Design and Synthesis ............................................................................................................... 12 

3.1 Synthesis Approach ........................................................................................................ 13 

3.2 Design Space Exploration and Optimization Strategy ................................................... 15 

3.3 Design Descriptions ....................................................................................................... 16 

3.3.1 Main-Rotor System ............................................................................................ 20 

3.3.2 Drive System ..................................................................................................... 21 

3.3.3 Flight-Control System ....................................................................................... 22 

3.3.4 Propulsion System ............................................................................................. 22 

3.3.5 Wings and Winglets ........................................................................................... 23 

4 Attributes and Aircraft Performance ....................................................................................... 25 

4.1 Attributes Summary ....................................................................................................... 25 

4.1.1 Air Vehicle Aerodynamics ................................................................................ 25 

4.1.2 Rotor Aerodynamics .......................................................................................... 26 

4.1.3 Wing Aerodynamics .......................................................................................... 30 

4.1.4 Aeroelastic-Stability Analysis ........................................................................... 31 

4.1.5 Handling Qualities ............................................................................................. 40 

4.1.6 Mass Properties .................................................................................................. 41 

4.2 Aircraft Performance ...................................................................................................... 43 

4.2.1 Downwash/Outwash Characteristics ................................................................. 43 

4.2.2 Hover Performance ............................................................................................ 43 

4.2.3 Short-Takeoff Capability ................................................................................... 46 

4.2.4 Speed Capability ................................................................................................ 49 

4.2.5 Mission-Performance Capability ....................................................................... 54 

4.3 Configuration Evaluation ............................................................................................... 58

5 Alternate Designs ..................................................................................................................... 61 

5.1 Fixed-Wing Variant Description .................................................................................... 61 

5.2 Shipboard-Suitable VDTR Description ......................................................................... 62 

6 Conclusions .............................................................................................................................. 64 

7 References ................................................................................................................................ 66 

8 Bibliography ............................................................................................................................ 68 
iv 

List of Figures 

Figure 1. Sikorsky VDTR wind tunnel model ...................................................................................... 8 

Figure 2. NASA Vertical Motion Simulator ......................................................................................... 9 

Figure 3. General arrangements of CTR, VDTR, and VSTR ............................................................. 12 

Figure 4. LVLT synthesis process ...................................................................................................... 14 

Figure 5. SDGW comparison.............................................................................................................. 16 

Figure 6. WE comparison ................................................................................................................... 16 

Figure 7. Installed MRP at SLS comparison ...................................................................................... 17 

Figure 8. Transmission-rating comparison ......................................................................................... 17 

Figure 9. Internal-fuel-tank volume comparison ................................................................................ 17 

Figure 10. Long-range-tank volume comparison .............................................................................. 17 

Figure 11. Rotor hover-diameter comparison ..................................................................................... 18 

Figure 12. Rotor cruise-diameter comparison .................................................................................... 18 

Figure 13. Rotor hover-tip-speed comparison .................................................................................... 18 

Figure 14. Rotor cruise-tip-speed comparison .................................................................................... 18 

Figure 15. Hover diskloading comparison .......................................................................................... 19 

Figure 16. Rotor solidity comparison ................................................................................................. 19 

Figure 17. Hover-efficiency comparison ............................................................................................ 19 

Figure 18. Cruise-efficiency comparison............................................................................................ 19 

Figure 19. Development-cost comparison .......................................................................................... 20 

Figure 20. Unit-flyaway-price comparison ......................................................................................... 20 

Figure 21. Equivalent flat-plate drag comparison .............................................................................. 25 

Figure 22. Hover-cruise efficiency-design trends ............................................................................... 28 

Figure 23. Design FM comparison ..................................................................................................... 28 

Figure 24. Design ηP comparison ....................................................................................................... 28 

Figure 25. Cruise-efficiency performance checks .............................................................................. 29 

Figure 26. Hover-performance checks ................................................................................................ 30 

Figure 27. Scaled airfoil Mach contours ............................................................................................. 31 

Figure 28. Designed airfoil Mach countours ...................................................................................... 31 

Figure 29. Wing-beam-Mode Frequency Correlation ........................................................................ 34 

Figure 30. Wing-beam-mode damping correlation ............................................................................. 34

Figure 31. Whirl-flutter-stability sensitivity to delta-3 coupling ........................................................ 37

Figure 32. Whirl-flutter-stability sensitivity to wing stiffness ........................................................... 38 

Figure 33. Wing-beam-mode damping vs. airspeed | 70% reduction in wing frequencies ................ 39 

Figure 34. Wing-weight comparison .................................................................................................. 41 

Figure 35. Rotor-weight comparison .................................................................................................. 41 

Figure 36. Body-weight comparison .................................................................................................. 42 

Figure 37. Fuel-system-weight comparison........................................................................................ 42 

Figure 38. Engine-weight comparison ................................................................................................ 42 

Figure 39. Transmission-weight comparison...................................................................................... 42 

Figure 40. Flight-controls comparison ................................................................................................ 43 

Figure 41. Engine/XMSN/rotor-weight comparison .......................................................................... 43 

Figure 42. Tiltrotor-outwash comparison ........................................................................................... 44 
v 

List of Figures (cont.) 

Figure 43. Outwash-force comparison................................................................................................ 44 

Figure 44. Outwash-moment comparison ........................................................................................... 44 

Figure 45. ISA WAT-capability comparison ...................................................................................... 45 

Figure 46. 95 ̊F WAT-capability comparison..................................................................................... 45 

Figure 47. STO ground-roll-capability comparison | SL/hot .............................................................. 46 

Figure 48. STO total-distance-capability comparison | SL/hot .......................................................... 47 

Figure 49. STO ground-roll-capability comparison | Hi/hot .............................................................. 47 

Figure 50. STO total-distance-capability comparison | Hi/hot ........................................................... 48 

Figure 51. STO ground-roll | MTOW | SL/hot ................................................................................... 48 

Figure 52. STO total distance | MTOW | SL/hot ................................................................................ 48 

Figure 53. STO ground roll | MTOW | Hi/hot .................................................................................... 49 

Figure 54. STO total distance | MTOW | Hi/hot ................................................................................. 49 

Figure 55. VBE capability comparison | ISA ....................................................................................... 50 

Figure 56. VBE capability comparison | 95°F ...................................................................................... 50 

Figure 57. VBR | capability comparison | ISA ..................................................................................... 51 

Figure 58. VBR capability comparison | 95°F ..................................................................................... 51 

Figure 59. VMCP | capability comparison | ISA ................................................................................... 52 

Figure 60. VMCP | capability comparison | 95°F .................................................................................. 52 

Figure 61. VBE | 30k/ISA | capability comparison .............................................................................. 53 

Figure 62. VBE | 4k/95 | capability comparison ................................................................................... 53 

Figure 63. VBR | 30k/ISA | capability comparison .............................................................................. 53 

Figure 64. VBR | 4k/95 | capability comparison .................................................................................. 53 

Figure 65. VMCP | 30k/ISA | capability comparison ............................................................................ 54 

Figure 66. VMCP | 4k/95 | capability comparison ................................................................................ 54 

Figure 67. Hi/hot ROA mission | payload-radius comparison ............................................................ 55 

Figure 68. SL/hot ROA mission | payload-radius comparison ........................................................... 55 

Figure 69. External-load mission | payload-radius comparison.......................................................... 56 

Figure 70. Long-range insertion mission | payload-radius comparison .............................................. 56 

Figure 71. Low-speed high-endurance mission | payload-radius comparison .................................... 57 

Figure 72. High-speed high-endurance mission | payload-radius comparison ................................... 57 

Figure 73. Long-range high-payload mission | payload-range comparison ....................................... 58 

Figure 74. Response surface methodology used for configuration evaluation ................................... 59 

Figure 75. VBG results | average value .............................................................................................. 60 

Figure 76. VBG results | 1st place likelihood ..................................................................................... 60 

Figure 77. VBG results | 2nd place likelihood .................................................................................... 60 

Figure 78. VBG results | 3rd place likelihood .................................................................................... 60 
vi 

List of Tables 

Table 1. Measured and Predicted Wing-Nacelle Frequencies ............................................................ 32 

Table 2. Comparison of Frequencies of the Clamped WRATS Blade ............................................... 32 

Table 3. Nonrotating Rotor Frequencies with Flexbeam and Rotor Control | Gimbal Locked .......... 33 

Table 4. Nonrotating Rotor Frequencies with Flexbeam and Rotor Control | Free Gimbal ............... 33 

Table 5. V-22 Collective Frequencies | RCAS Model vs. EMD Report ............................................ 35 

Table 6. V-22 Cyclic Frequencies | RCAS Model vs. EMD Report .................................................. 35 

Table 7. Whirl Flutter Speeds for a 70% Reduction in Wing Frequencies ........................................ 39 

Table 8. Handling Qualities Summary ............................................................................................... 41 

Table 9. Attribute Ranking ................................................................................................................. 58 

Table 10. VBG Response Surface Linear Multivariate Regression Coefficients ............................... 59 
vii 

Nomenclature 

2-D two-dimensional 

ANSYS Finite Element Analysis Software 

ARC active rotor control 

ARL Army Research Lab 

BVI blade–vortex interaction 

C risk consequence 

CAMRAD Comprehensive Analytical Model of Rotorcraft Aerodynamics & Dynamics 

CATIA Computer Aided Three-Dimensional Interactive Application 

CCHAP Circulation Coupled Hover Analysis Program 

CF centrifugal force 

cg center of gravity 

CFD computational fluid dynamics 

CDI Continuum Dynamics Incorporated 

CHARM Comprehensive Hierarchical Aeromechanics Rotorcraft Model 

CTD Critical Technology Demonstrator 

CTR conventional tiltrotor 

DARPA Defense Advanced Research Projects Agency 

DOD Department of Defense 

EHPIC Evaluation of Hover Performance using Influence Coefficients 

FAA Federal Aviation Administration 

FBW fly-by-wire 

FCS flight-control system 

FDTR fixed-diameter tiltrotor 

FEM finite element method 

FM figure of merit 

FTR future transport rotorcraft 

FWV Fixed-Wing Variant 

GDATS General Dynamics Advanced Technology Systems, Inc. 

GenHel General Helicopter Flight Dynamics Simulation program 

GVT Ground Vibration Test 

GW gross weight 

HERO Helicopter Rotor Optimization 

HOGE hover out of ground effect 

IPS inlet particle separator 

IR infrared 

ISA International Standard Atmosphere 

L risk likelihood 

L/D lift/drag ratio 

LOS line of sight 

LSAF Lifting Surface Aerodynamics & Performance Analysis of Rotors in Axial Flight 

LVLT large vertical-lift transport 

MCP maximum continuous power 
viii 

Nomenclature (cont.) 

MEA multielement rotors 

MGB main-rotor gearbox 

MRP Maximum Rated Power 

MSES Multi-element Airfoil Design/Analysis Software 

MTOW maximum takeoff gross weight 

NASA National Aeronautics and Space Administration 

NASTRAN NASA Structural Analysis Program 

OEI one engine inoperative 

PASTA Proprotor Aeroelastic STability Analysis 

psf pounds per square foot 

QFD Quality Function Deployment 

RDM Rotorcraft Design Model 

RCAS Rotorcraft Comprehensive Analysis System 

RDS-21 Rotorcraft Drive System for the 21st Century 

REVCON Revolutionary Concepts in Aeronautics 

ROA radius of action 

rpm revolutions per minute 

SAC Sikorsky Aircraft Corporation 

SDD system design and development 

SDGW structural design gross weight 

SFC engine-specific fuel consumption 

SHCT Short Haul Civil Tiltrotor 

SL sea level 

SLS sea-level standard 

STO short takeoff 

STOL Short Takeoff and Landing 

t/c thickness/chord 

TDT Transonic Dynamics Tunnel 

TOC total operating cost 

UMARC University of Maryland Advanced Rotorcraft Code 

USAF United States Air Force 

UTC United Technologies Corporation 

VBG Value-Based Gage 

VDIVE dive speed 

VDTR variable-diameter tiltrotor 

VDTR-S shipboard-suitable VDTR 

VGART Variable-Geometry Advanced-Rotor Technology 

VH maximum level cruise speed 

VMS Vertical Motion Simulator 

VROC vertical rate of climb 

VSTR variable-speed tiltrotor 

VTOL vertical takeoff and landing 
ix 

Nomenclature (cont.) 

WAT weight-altitude-temperature 

WE Weight Empty 

WRATS Wing and Rotor Aeroelastic Test System 

XMSN Transmission 

°F degrees Fahrenheit 

CD drag coefficient 

Cf skin-friction coefficient 

CL lift coefficient 

CT thrust coefficient 

NR percent rotor rotational speed 

VBE velocity for best endurance 

VBR velocity for 99% best specific range 

VMCP velocity for engine maximum continuous power or transmission torque 

ηP rotor propulsive efficiency factor 

σ rotor solidity (blade area over disk area) 

1 

Abstract 

The technical community has identified rotor efficiency as a critical enabling technology for large 

vertical-lift transport (LVLT) rotorcraft. The size and performance of LVLT aircraft will be far 

beyond current aircraft capabilities, enabling a transformational change in cargo transport effectiveness. Two candidate approaches for achieving high efficiency were considered for LVLT 

applications: a variable-diameter tiltrotor (VDTR) and a variable-speed tiltrotor (VSTR); the 

former utilizes variable-rotor geometry and the latter utilizes variable-rotor speed. Conceptual 

aircraft designs were synthesized for the VDTR and VSTR and compared to a conventional 

tiltrotor (CTR). The aircraft were optimized to a common objective function and bounded by a set 

of physical- and requirements-driven constraints. The resulting aircraft were compared for weight, 

size, performance, handling qualities, and other attributes. These comparisons established a 

measure of the relative merits of the variable-diameter and -speed rotor systems as enabling 

technologies for LVLT capability. 
2 

1 Executive Summary 

A system-level analysis of three tiltrotor aircraft configurations was conducted to compare the relative merits of the aircraft for a common set of large-vertical-lift transport LVLT mission requirements. The goal of this comparison was to quantify the relative value of various system-level technologies, including variable geometries and advanced drive systems. The analysis included a 

conventional tiltrotor (CTR) as the baseline and two alternate configurations examined as a means 

of increasing hover and cruise efficiency: a variable-diameter tiltrotor (VDTR) and variable-speed 

tiltrotor (VSTR). The VDTR employed the ability to reduce its rotor diameter by 23% during airplane mode, and the VSTR employed a two-speed transmission allowing its rotor to operate 

between 100% and 60% rotor rotational speed (NR) during airplane mode. Upon completion of the 

synthesis for each aircraft, the attributes of the aircraft, including size, weight, and performance, 

were compared. The relative strengths and weaknesses of the designs were quantified as an indication of the merits of the underlying technology for each concept. 

Sikorsky was selected to conduct this analysis in large part because of the company’s previous 

VDTR development efforts, spanning approximately 10 years during the 1990s. These efforts 

included preliminary and detailed design of both the aircraft and rotor-system components as well 

as proof-of-concept testing of the system-level benefits and risk-reduction testing of critical subsystem components. Significant milestones included wind tunnel tests of the VDTR rotor aerodynamics, dynamics, and acoustics; full-scale ground tests of the blade retention and retraction 

mechanism; and simulations of the VDTR one engine inoperative (OEI) takeoff and landing capability using the NASA Vertical Motion Simulator (VMS). This historical background formed the 

foundation for the current design and analysis of the VDTR and provided lessons learned, which 

influenced the design of multiple sub-systems and the overall aircraft integration. 

Aircraft Synthesis: Aircraft synthesis was performed using the Sikorsky Aircraft synthesis tool, 

the Rotorcraft Design Model (RDM), embedded within the Phoenix Integration ModelCenter®

design environment. The aircraft were synthesized as “clean-sheet” designs, optimized to a 

common objective function by perturbation of common design variables and subjected to a set of 

physical- and requirements-driven constraints. 

Design Description: The aircraft were designed to offer high cruise efficiency while maintaining 

significant hover capability. The tiltrotors utilize nongimbaled, flexbeam-style proprotors mounted 

on pivoting nacelles. The engine layout includes four engines, configured with two engines 

mounted in each nacelle. The engines remain horizontal in all modes of flight; only the transmissions and rotors rotate. The aircraft include wingtips that extend beyond the engine nacelles. The 

wingtips rotate independently of the nacelles to minimize vertical drag in hover and parasite drag 

during transition and level flight, increasing the overall wing-span efficiency. VDTR rotor geometry is controlled by a hydraulic reel mechanism embedded in the blade feathering axis. VSTR rotor 

revolutions-per-minute (rpm) variation is enabled by the addition of a shifting gearbox to the facegear transmission-system layout common to all three aircraft. The composite fuselage is pressurized and has an integral rear ramp for rapid loading and unloading without interference from 

engine exhaust. The landing-gear configuration comprises three two-wheel main-gear bogeys on 

each side and a two-wheel truck for the nose gear. 
3 

Rotor Aerodynamics: Rotor aerodynamic design and performance estimation was performed 

using the Evaluation of Hover Performance using Influence Coefficients (EHPIC) tool. Validation 

of EHPIC was performed by correlating with V-22 experimental data. Figure-of-merit (FM) rotor 

propulsive efficiency factor (FM-ηP) boundaries for all three aircraft were defined that progress 

from an ideal hover twist to an ideal propeller twist. The unique boundaries for each configuration 

were used to select the optimum blade-twist distributions that balanced hover and cruise efficiency. EHPIC was then used to generate comprehensive rotor performance data, which were then 

incorporated into the aircraft synthesis tool RDM. Sikorsky rotor performance estimates compared 

very favorably against existing data and other codes, such as Comprehensive Hierarchical Aeromechanics Rotorcraft Model (CHARM) and Lifting Surface Aerodynamics & Performance Analysis of Rotors in Axial Flight (LSAF). A Circulation Coupled Hover Analysis Program (CCHAP) 

model, correlated to V-22 data, was used to derive the vertical drag of the three aircraft. 

Aeroelastic Stability Analysis: Aeroelastic stability analysis was performed using the aeroelastic 

tool, Rotorcraft Comprehensive Analysis System (RCAS). Validation of RCAS for whirl-flutter 

stability was performed by correlating with V-22 test data from the Wing and Rotor Aeroelastic 

Test System (WRATS) test rig at the NASA Langley Transonic Dynamics Tunnel. RCAS was 

then used to perform stability analysis of the tiltrotor configurations. Initially, a gimbal-style hub 

was selected and analyzed; however, whirl-flutter stability analysis indicated a stiffer hub to be 

more desirable, leading to the design of a flexbeam-style rotor with greatly improved whirl-flutter 

stability. The final designs for all three aircraft were predicted to be stable throughout the required 

operating range. 

Handling Qualities: The handling-qualities analysis was performed using the Sikorsky General 

Helicopter Flight Dynamics Simulation (GenHel) program. A correlated XV-15 tiltrotor aircraft 

GenHel simulation was created and used to build a generic GenHel tiltrotor simulation, including 

variable diameter rotor and variable incidence outboard wing tips. The generic simulation was 

used for handling qualities and performance analysis of the three tiltrotor aircraft designs. Trim 

characteristics of the three configurations were calculated and used to derive schedules of mast 

angle, flap position, and wing-tip incidence in helicopter, transition, and airplane modes at airspeeds from hover to greater than 300 knots. Helicopter-mode hover flight dynamics and airplanemode longitudinal static stability and longitudinal flight dynamics were examined to evaluate 

handling qualities and verify operation of the simulations. 

Mass Properties: The weight-prediction methodology employed for the tiltrotors was based on a 

combination of statistical, semianalytical, and analytical methods. These methods were mathematically modeled in RDM to permit weight trending as a function of the primary aircraft design 

parameters. Parametric weight estimates were generally correlated to V-22 weight data, with technology or requirements factors applied where appropriate. Requirements factors were applied to 

the VDTR rotor system and the VSTR drive system to account for their unique design features. 

Rotor and Blade Design: A trade study focusing on weight minimization and whirl-flutter stability margin determined a gimbaled rotor solution to be unviable for the LVLT tiltrotor application. 

As a result, a nongimbaled, flexbeam-style rotor was adopted for all three tiltrotors. The flexbeam 

hub was mated to a blade pitch bearing couple, allowing blade pitch independent of flexbeam 

motion. The four light-weight, stiff rotor blades were designed using a mix of high-modulus and 
4 

high-strength composites to provide blades that met the dynamics requirements while still enabling 

operation in the austere environment of military aircraft. The VDTR blades include a telescoping 

spar section mated with a hydraulically actuated strap-and-reel retraction mechanism embedded in 

the feathering axis of the blade. The placement of the retraction mechanism in the feathering axis 

reduces the adverse reliability issues associated with past VDTR retraction-mechanism concepts 

and offers improved strap life. The strap design assumed a current state-of-the-art technology fiber 

mix as a conservative estimate and includes ample margin for wear and aging. 

Transmissions Design: A state-of-the-art split-torque-style transmission design was found to be 

too heavy for the tiltrotor application, so a face-gear-style transmission was designed that yielded 

significant weight savings. The design was matured using knowledge gained from the Rotorcraft 

Drive System for the 21st Century (RDS-21) program and incorporating several technology 

enhancements currently in development. Additional effort was invested for designing a separate 

gearbox module based on an existing Sikorsky patent for the VSTR configuration. 

 

Wing Design: A supercritical airfoil section was designed for the wings of the aircraft to balance 

the requirements of wing stiffness for dynamics and minimal drag for cruise efficiency. A NASA 

18%-thick supercritical airfoil section was scaled to the required thickness and then modified using 

the Multi-element Airfoil Design/Analysis Software (MSES) analysis code to optimize the airfoilsection performance at the LVLT design conditions. The airfoil modification was done by using 

the MSES inverse-airfoil design method to eliminate local supersonic flow on the airfoil upper surface. The resulting airfoil-section aerodynamic characteristics were used to calculate wing aerodynamic characteristics for the GenHel simulations and for implementation into RDM for aircraft 

synthesis. The wing planform and twist geometry of each concept was optimized for operation in 

the unique rotor-wash environment of the three rotor types. Fuel is carried in the composite wings, 

and the wings offer fuel volume in excess of the volume dictated by the mission requirements. 

Affordability Analysis: Parametric cost estimates were prepared for the three aircraft using the 

commercially available PRICE-H/HL suite and total-operating-cost (TOC) models from PRICE 

Systems LLC, calibrated to actual cost data. Unit-flyaway-price, system-design-and-development 

(SDD), and critical-technology-demonstrator (CTD) costs were estimated. 

Alternate Aircraft Analysis: In addition to the three “clean-sheet” tiltrotor aircraft, two alternate 

configuration aircraft were synthesized as further points of comparison: a shipboard-suitable 

VDTR concept, arranged in a canard configuration with shortened fuselage and tandem cockpit; 

and a fixed-wing variant. 

The remainder of this report discusses, in detail, the design synthesis and analysis process used to 

develop the three tiltrotor platforms considered for this comparison. Furthermore, the resulting aircraft attributes and performance capabilities are also compared in order to quantify the relative 

merits of the variable-diameter and -speed technologies compared to the baseline CTR. Additionally, a more detailed description of the preceding development work conducted by Sikorsky is presented, as well as a general description of the anticipated performance improvements attributable 

to the variable-diameter rotor technology. Finally, the two alternate configurations are also 

presented. 
5 

2 VDTR Historical Summary 

The VDTR configuration balances hover performance with cruise efficiency. Power loading is 

directly determined by the hover requirement that sets the engine power, transmission ratings, and 

rotor diameter. The engine and engine-performance characteristics are configuration-independent. 

While the power and drive system are determined by hover, the cruise efficiency is a function of 

wing geometry, fuselage drag, and propeller efficiency. The fuselage geometry and subsequent 

drag are essentially configuration-independent. The wing geometry of the VDTR is unique in that 

it can be tailored for the cruise rotor diameter, allowing for a slightly smaller wingspan and optimization of the wingtip extensions. The ability to tailor the wing geometry for the VDTR combined with the overall improvement in propeller efficiency afforded by the increased blade 

loadings for the VDTR in cruise improve the overall fuel efficiency of the aircraft. 

The variable-diameter concept was first explored as an application for single-main-rotor helicopter 

platforms. As early as 1972, research and design activities were being conducted to explore the 

benefits of a retracting rotor system. For compound, single-main-rotor helicopters, variablediameter rotors offer the potential to increase the aircraft speed capability by mitigating blade 

aeroelastic responses at high speed, minimizing the drag of the exposed blades and decreasing the 

stowed size of the rotor. For tiltrotor platforms, the variable-diameter rotor provides a more 

optimal solution for both the hover and cruise condition; increasing the rotor diameter in hover 

decreases the blade loading leading to improved efficiency, and retracting the rotor during cruise 

increases the blade loading, providing increased propeller efficiency and lift/drag ratio (L/D). The 

work of Fradenburgh, Scott, and Wang (refs. 1 through 8 provide detailed comparisons of VDTR 

and conventional tiltrotor aircraft). 

The decreased blade loadings in hover and increased blade loadings in cruise afforded by the ability to vary the rotor geometry offer more optimal solutions for the tiltrotor concept because in each 

state the rotor better matches the operational requirements placed upon it. The hover condition 

requires that the rotor be capable of generating enough thrust to lift the gross weight of the vehicle. 

The cruise condition requires that the rotor generate significantly less thrust as the total thrust need 

only balance the drag of the aircraft in trim, level flight. The wing carries the aircraft weight during cruise operations. If the rotor is sized to meet the hover requirement, then it is likely to be significantly oversized for generating the required cruise thrust. Similarly a rotor sized for the cruise 

condition will not provide nearly enough thrust to meet the requirements of the hover condition. A 

compromised rotor with a fixed geometry will not be ideally suited for either condition; the rotor 

size and blade loadings will not be adequate for hover or cruise, leading to compromised efficiency 

in both spectrums. 

The variable-diameter concept alleviates some of the need to compromise the rotor design in order 

to meet the thrust requirements of the hover and cruise condition. The rotor operates at its maximum diameter in hover, typically extending over the main body of the fuselage, to minimize the 

hover disk loading, and then contracts during the transition to cruise flight. This contraction 

increases the blade loading during cruise, primarily by decreasing the rotor tip speed and blade 

area, leading to an overall increase in rotor propulsive efficiency. This increase in efficiency is 

generated without the need to significantly reduce rotor rpm by either slowing the power turbine 
6 

speed or reducing the main gearbox ratio.The main power turbine can operate within the 80% to 

100% NR range, allowing engine-specific fuel consumption (SFC) to remain near its peak during 

cruise. 

Typically, conventional tiltrotor aircraft operate with very high disk loadings in hover, where 18 to 

24 pounds per square foot (psf) is normal. These disk loadings are twice the loadings seen for conventional, single-main-rotor helicopters. Disk loadings in the 18 to 24 psf range are accompanied 

by excessive downwash and outwash velocities as well as poor autorotative characteristics. The 

VDTR mitigates some of these concerns. All other design attributes being equal, a VDTR typically 

offers the designer the ability to include a larger rotor diameter in hover than could be accommodated by a conventional tiltrotor. This larger diameter contributes to reduced disk loadings, lower 

downwash velocities, and improved autorotative capability. 

The VDTR offers benefits not only in hover operations, but also in cruise operations. Typically 

conventional tiltrotor rotors have a large diameter because of the requirements of the hover spectrum. These rotors then operate in cruise as well, and partly because of their large diameters, are 

susceptible to propeller whirl-flutter instability. In fact, this instability often limits the forwardflight speed for tiltrotor aircraft. The VDTR has the potential to offer higher forward-flight-speed 

capabilities because the reduced rotor diameter is less susceptible to the whirl-flutter instability. 

The variable-diameter concept can potentially affect the tiltrotor component designs in several 

areas. If the rotor and drive system are sized in hover, then the ability to reduce the disk loading in 

hover for a given blade loading may be advantageous. Reducing the disk loading will likely reduce 

the required installed power, and thereby overall aircraft weight for the hover condition. However, 

the capability to reduce the disk loading is not freely available. The VDTR mechanism increases 

the overall weight and complexity of the rotor blades and hub. This increased weight may offset 

the savings yielded by the decreased disk loading in hover. The VDTR concept also affects the aircraft wing design in two ways. Since the rotor can expand and contract, the rotor blades can overlap the fuselage in hover. Then as the rotors transition to forward cruise the blades can retract to 

provide adequate fuselage clearance. Typically, the overall wingspan from the fuselage to the 

nacelles can be reduced for a VDTR design compared to a similar conventional tiltrotor aircraft. 

Furthermore, the VDTR concept potentially opens the design to higher cruise speeds than those 

seen for current tiltrotor platforms. The ability to increase the maximum airspeed capability, as 

well as the cruise-speed capability, influences the aircraft wing design by reducing the required 

wing incidence and wing area. Reduced incidence will likely benefit aircraft-induced drag, and 

reduced wing area will reduce the overall wing weight. Aircraft gross weight may trend lower as 

airspeed increases for a VDTR concept. 

With regard to operational suitability, the VDTR has the advantage of a large, lightly loaded rotor 

for hover operations with the added ability to reduce the rotor size for cruise and ground operations. The variable-rotor geometry provides the capability to minimize the aircraft operational 

footprint during parking and taxi operations. Minimizing this footprint allows the aircraft to be 

parked and to taxi in confined areas. 

While the benefits of a variable-diameter rotor for compound, single-main-rotor platforms should 

not be ignored, tiltrotor platforms experience a much greater potential benefit. During the late 
7 

1980s, 1990s, and the early part of the 21st century, Sikorsky Aircraft Corporation conducted significant research and development activities to mature the variable-diameter tiltrotor concept. 

These activities were conducted in partnership with the National Aeronautics and Space Administration (NASA), the United States Army, the United States Air Force (USAF), and the Defense 

Advanced Research Projects Agency (DARPA). The various activities included preliminary 

design, detailed design, model-scale experiments, and full-scale testing of the VDTR concept and 

its subcomponents. Each of the programs added to the fundamental understanding of the VDTR 

concept, and they progressively matured the design. Several of the more significant efforts are 

described in greater detail in the following sections. 

2.1 Technology Needs for High-Speed Rotorcraft (’90–’91) 

References 5 and 7 describe an analytical study conducted to identify rotorcraft concepts best 

capable of combining a cruise speed of 350 to 450 knots with helicopter-like low-speed attributes, 

and to define the technology advancements needed to make them viable for full-scale development 

by the year 2000. A systematic approach was used to compare the relative attributes and required 

gross weights for a wide range of concepts, resulting in a downselect to the most promising 

concept/mission pairs. For the transport missions, tiltwing and VDTR concepts were found to be 

superior. For a military scout/attack role, the VDTR was best, although a shrouded-rotor concept 

could provide a highly agile, low observable alternative if its weight-empty fraction could be 

reduced. A design speed of 375 to 425 knots was found to be the maximum desirable for transport 

missions, with higher speed producing rapidly diminishing benefits in productivity. The key 

technologies that required advancement to make the tiltwing and VDTR concepts viable were in 

the areas of wing and proprotor aerodynamics, efficient structural design, and flight controls. 

Specific attention was needed regarding proprotor-blade-geometry optimization, refinement of the 

geared flap-pitch control system, expansion of the speed/descent envelope, and the structural and 

aerodynamic trade-offs of wing thickness and forward sweep. 

Of the 16 high-speed VTOL concepts studies, five of these concepts were deemed to have desirable high-speed vertical-takeoff-and-landing (VTOL) attributes. Preliminary design efforts were 

undertaken to explore the physical characteristics, performance, and sensitivity to technology 

assumptions for the most promising concepts. The study defined levels of new technology that 

were required to enable low-risk development and determined that weights, structural design, flyby-wire, drag reduction, airfoil design, and propulsion-system design were critical to the success of 

these concepts. 

2.2 Wind Tunnel Test of a VDTR Model (’94) 

A wind tunnel test of a 1/6th scale VDTR was conducted as a joint effort between NASA Ames 

Research Center and Sikorsky Aircraft (ref. 9). The objective was to evaluate the aeroelastic and 

performance characteristics of the VDTR in conversion, hover, and cruise. The rotor diameter and 

nacelle angle of the model were remotely changed to represent tiltrotor operating conditions. Data 

were presented showing the propulsive force required in conversion, blade loads, angle-of-attack 

stability and simulated gust response, and hover and cruise performance. This test represented the 

first wind tunnel test of a variable-diameter rotor applied to a tiltrotor concept. The results 
8 

confirmed some of the potential advantages of the VDTR and established the variable-diameter 

rotor as a viable candidate for an advanced tiltrotor. Current predictions for the LVLT VDTR 

concept indicate that predictions for hover figure-of-merit and cruise efficiency correlate well with 

the data generated by this test effort. 

This wind tunnel test successfully demonstrated the feasibility of the VTDR for tiltrotor aircraft. A 

wide range of test points were taken in hover, conversion, and cruise modes. The concept was 

shown to have numerous advantages over conventional tiltrotors, such as reduced hover downwash 

with lower disk loading and significantly reduced longitudinal gust response in cruise. 

In the conversion regime, a high propulsive force was demonstrated for sustained flight with 

acceptable blade loads. The VDTR demonstrated excellent gust-response capabilities. The horizontal gust response correlated well with the predictions, revealing only half the response to turbulence relative to the conventional tiltrotor. Figure 1 shows the Sikorsky VDTR wind tunnel 

model. 

Figure 1. Sikorsky VDTR wind tunnel model. 

2.3 VMS Simulation of a Variable-Diameter Tiltrotor (’97) 

A joint NASA/Sikorsky simulation was performed in the NASA-Ames Vertical Motion Simulator 

to evaluate the performance and handling qualities of a VDTR compared to a fixed-diameter tiltrotor (FDTR) during normal and emergency terminal-area operations (ref. 10). Pilots from the 

Federal Aviation Administration (FAA), NASA, academia, and industry evaluated both the VDTR 

and FDTR during OEI takeoffs and approaches and all-engine inoperative landings. The simulation math model was based on NASA’s preliminary-design version of the Generic Tiltrotor Simulation code. Both fixed- and variable-diameter rotor versions exhibited Level One handling qualities during normal maneuvers with both engines operating. In general, the VDTR had a greater 

power margin and performance in helicopter mode resulting from a lower disk loading requiring 

20% to 25% less power at similar thrust levels. This power margin was an advantage during oneand all-engine inoperative procedures. This simulation study quantified the performance merits of 

the VDTR concept and demonstrated that the VDTR offers enhanced performance and terminalarea safety for future civil tiltrotor transports. The simulation was also beneficial in identifying 

areas for further investigation for the VDTR and basic civil tiltrotor concepts. Figure 2 shows the 

NASA VMS Simulator. 
9 

Figure 2. NASA Vertical Motion Simulator. 

2.4 Variable-Diameter Retraction-Mechanism Risk Reduction and Trade Studies 

(’98–’99) 

Mechanical trade studies and a risk-reduction plan were developed for the VDTR retraction 

mechanism. The viability of the VDTR had been demonstrated under NASA’s Short Haul Civil 

Tiltrotor (SHCT) program through model scale tests and piloted simulation studies. These riskreduction plans and trade studies addressed the next challenge for the VDTR development: determining the feasibility of VDTR applied to full-scale flight vehicles. This report clarified VDTR 

technical and mechanical issues and presented a risk-reduction plan for large-scale retraction 

mechanism testing and wind tunnel testing. 

2.5 Variable-Diameter Tiltrotor Technology (’99–’00) 

The Variable-Diameter Tiltrotor Technology program addressed the fundamental risk issues of the 

VDTR concept through a combination of large-scale component testing. This cooperative 

DARPA/Sikorsky program was established to explore several fundamental issues surrounding 

VDTR, including the feasibility of a large-scale retraction mechanism under conditions 

representative of a realistic flight environment, including a centrifugal force field and vibration and 

the aeroelastic stability of a VDTR in high-speed airplane-mode flight. Accelerated ground testing 

of the critical mechanisms of a large-scale VDTR retraction mechanism was performed under 

realistic environmental conditions. Development of a VDTR aeroelastic wind tunnel model was 

initiated that was aimed at validating the aeroelastic stability of a VDTR in the high-speed airplane 

mode. Congressional funding cuts forced termination of this program for convenience prior to 

completion of all planned risk-reduction activities. 

2.6 Variable-Geometry Advanced Rotor Technology (VGART) Phase 1 (’99–’00) 

The Variable-Geometry Advanced-Rotor Technology (VGART) Phase I study evaluated the application of active rotor control systems to both a helicopter and a VDTR, sized to meet the futuretransport-rotorcraft (FTR) requirements. This study was performed by a consortium comprising 

Sikorsky Aircraft Corporation (SAC) and General Dynamics Advanced Technology Systems, Inc. 
10 

(GDATS). The study concluded that a trailing-edge flap driven by a hydraulic lapfit actuator provided the lightest weight, smallest, most affordable single-main-rotor helicopter solution for FTR. 

Similarly, a hydraulically powered individual-blade control system provided the smallest, lightest 

weight, most affordable VDTR solution for FTR. Both aircraft were sized for the same FTR missions and requirements, except, naturally, for speed. The VDTR solution had a maximum cruise 

speed of 295 knots, while that of the helicopter was 175 knots. 

In arriving at recommended design solutions, a thorough life-cycle cost analysis was conducted, 

comparing numerous active control approaches and actuator concepts. The chosen concepts 

represented the solutions with the lowest life-cycle cost. The study also assessed technical, scaling, 

reliability, maintainability, manufacturing, produceability, and affordability issues. The results 

indicated that all major hardware scalability, produceability, accessibility, and implementation 

issues could be resolved though risk-reduction tests. [Examples of recommended risk-reduction 

tests included wind tunnel tests to verify flap effectiveness, actuator bench tests under load, 

hydraulic power-supply-development tests, and manufacturing-risk-reduction tests. 

The project also addressed the existence of potential interactions between the aircraft flight control 

system and the active control system, and found these interactions to be negligible. Further investigation of this issue, using higher-fidelity simulations, was recommended in order to resolve any 

potential software implementation issues. 

2.7 Revolutionary Concepts in Aeronautics (REVCON) Phase 1 (’00–’01) 

The Revolutionary Concept Study “Development and Flight Test of a Variable Diameter Tilt 

Rotor Experimental Aircraft” was conducted in 2000–2001 as part of the REVCON Phase 1 Cooperative Agreement between NASA and Sikorsky Aircraft Corporation. The REVCON Phase 1 

study evaluated the system benefits and flight feasibility for a VDTR experimental aircraft, and 

defined an appropriate REVCON flight-test program. Teamed with Sikorsky in this eight-month 

effort were NASA Dryden Flight Research Facility, NASA Ames Research Center, and NASA 

Langley Research Center, with NASA Glenn Research Center joining the team towards the end of 

Phase 1. A systems benefits analysis was performed utilizing the V-22 as the baseline configuration to help expedite the study because a significant database was available on the aircraft. In 

addition, a significant database already existed on VDTR technology, resident at NASA centers, as 

a result of some eight years of prior investigations on this concept as part of the Short Haul Civil 

Tilt Rotor Program. This team’s systems evaluation confirmed significant potential performance 

and safety benefits for VDTR compared to conventional tiltrotor designs. 

VDTR safety benefits were quantified as a result of a joint NASA/Sikorsky piloted simulation 

study conducted in the Ames VMS facility (ref. 10). The system analysis also determined that the 

extra power margin inherent with the VDTR gave it a large advantage during emergency situations 

with one engine inoperative. The Category A-style takeoff procedures provided a framework to 

quantify this benefit. Overall, the VDTR had a power margin and performance capability such that 

the pilots were able to comfortably continue the takeoff if the engine failed after the takeoff decision point. Analyses of various other emergency scenarios indicate that the VDTR significantly 

reduced takeoff distances for VTOL takeoffs for rejected takeoffs compared to conventional 
11 

tiltrotors. Furthermore, VDTR improved safety during all-engines-out maneuvers because of 

decreased disk loadings and high rotor inertia compared to conventional tiltrotors. 

2.8 Variable-Geometry Advanced Rotor Technology (VGART) Phase 2 (’00) 

The VGART program was established to determine the feasibility of critical technologies and 

components required for high-risk rotor concepts needed to provide substantial improvements in 

the aerodynamic performance, external noise, and vibration characteristics of future rotorcraft. The 

tasks conducted under VGART Phase 2 developed and demonstrated technologies for active rotor 

control (ARC), high lift, multielement rotors (MEA), and VDTRs. ARC tasks included simulating 

a coupled active rotor-control–flight-control system to identify any unfavorable interactions, 

demonstrating wireless rotating to fixed-frame information transfer, studying the benefits of an 

on-blade active flap at model scale, and adapting and applying a closed-loop active controller as 

part of a cooperative study of an active twist rotor. MEA tasks included a full-scale, twodimensional (2-D) aerodynamic evaluation, a design trade study of application of active and 

passive MEA concepts, and a wind tunnel experiment on a Mach-scale MEA model rotor. 

Aeroelastic stability of the VDTR in forward flight (propeller mode) was experimentally 

investigated to verify the predicted increase in the speed of the whirl-flutter boundary. 

2.9 VDTR Measured Acoustic and Aerodynamic Data Compared with the V-22 (’02) 

A wind tunnel test was conducted using the NASA Langley 14- by 22-Foot Subsonic Wind Tunnel 

to evaluate the acoustic and aerodynamic benefits of the Sikorsky Variable Diameter Tilt Rotor. 

This test is discussed in reference 8. The results were evaluated against a baseline currenttechnology V-22 rotor. The test evaluated four VDTR configurations, each representing a potential 

noise-reduction technology compared to conventional proprotors. In addition, baseline data were 

obtained for a V-22 rotor for comparison with the measured VDTR acoustic and aerodynamic 

results. The aim of the program was to demonstrate a 6-dBA reduction in blade–vortex interaction 

(BVI) source noise through design, relative to the current-technology V-22. This reduction was 

part of the total SHCT external-noise-reduction goal of 12 dBA. The noise-reduction goals were to 

assist in overcoming the primary obstacle to the ready acceptance of tiltrotors for commercial 

carrying service. Sikorsky’s VDTR succeeded in achieving the SHCT external-noise-source 

reduction goal. 
12 

3 Design and Synthesis 

The LVLT aircraft were optimized and designed around a common set of requirements with the 

intent of exposing the nuances that differentiate the configurations. Where common designs could 

be applied to all of the configurations, a single design was applied to all three configurations. 

Where a specific design effort could differentiate between the configurations, these areas were the 

primary focus of effort. 

The tiltrotors utilize nongimbaled flexbeam-style proprotors mounted on pivoting nacelles. The 

engine layout includes four engines, configured with two engines mounted in each nacelle, which 

remain horizontal in all modes of flight (only the transmissions and rotors rotate). The aircraft 

include wingtips that extend beyond the engine nacelles. The wingtips rotate independently of the 

nacelles to minimize vertical drag in hover and parasite drag during transition and level flight, 

increasing the overall wingspan efficiency. VDTR rotor geometry is controlled by a hydraulic reel 

mechanism embedded in the blade feathering axis. VSTR rotor rpm variation is enabled by the 

addition of a shifting gearbox to the common face-gear transmission-system layout. The composite 

fuselage is pressurized and has an integral rear ramp for rapid loading and unloading without interference from engine exhaust. The landing-gear configuration comprises three two-wheel main-gear 

bogeys on each side and a two-wheel truck for the nose gear. Figure 3 shows the general arrangements of the CTR, VDTR, and VSTR. 

CTR VDTR VSTR

Figure 3. General arrangements of CTR, VDTR, and VSTR. 
13 

The aircraft were synthesized as “clean-sheet” designs, optimized to a common-objective function 

by perturbation of common design variables and subjected to a set of physical- and requirementsdriven constraints. This approach provided a means of fairly comparing the emergent performance 

and attributes (physical size, weight, and geometry) of the aircraft necessary to meet the design 

requirements, based on the unique characteristics of each configuration. 

3.1 Synthesis Approach 

Aircraft sizing was performed using the Sikorsky aircraft synthesis tool, Rotorcraft Design Model 

(RDM). RDM is a comprehensive aircraft design and analysis code developed for new conceptual 

design, derivative design, and competitive analysis. RDM is a physics-based model coupled with 

empirical corrections derived from flight-test data and/or higher-fidelity analysis. Aircraft sizing 

and performance are based on a longitudinal force trim, parametric weight estimation correlated to 

existing aircraft, scalable engine power and fuel flow-lapse rates, and flexible mission definition. 

RDM is primarily executed in one of two modes: design/sizing mode and mission-analysis mode. 

In design/sizing mode, RDM calculates the aircraft weights, geometry (rotor diameter, rotor solidity, wing area, wing chord, etc.), and propulsion-system attributes (installed engine power and 

SFC) required to perform a user-prescribed design mission profile. The design/sizing mode can be 

run in one of two fashions: for a given payload, the structural design gross weight is determined; 

or, for a given gross weight, the resulting payload is found. In mission-analysis mode a sized aircraft must be provided with fully defined weights, geometry, and propulsion-system characteristics. RDM segregates a mission profile into discrete elements at required altitudes, temperatures, 

speeds, and durations or distances. Forward-flight segments may be at specified speeds or coded to 

either provide speeds such as maximum range or maximum endurance or match a gearbox design 

rating or engine power setting. Fuel burn-off is then tracked per mission segment. Weight and drag 

changes can be applied per segment to account for payload unloading, ordinance expenditure, or 

external lift pickup. 

LVLT synthesis was accomplished using Phoenix Integration’s ModelCenter®. ModelCenter®

enables multiple instances of RDM to be graphically integrated into a seamless, automated engineering process. A highly coupled design environment was constructed to design the optimal aircraft while balancing the requirements of each mission. This environment included one instance of 

RDM running a composite sizing mission in design/sizing mode and other instances, each running 

one of the LVLT mission profiles in mission-analysis mode. The synthesis process, illustrated in 

Figure 4, was as follows: 

1. Starting values were assumed for aircraft weights and the fuel capacity. RDM was executed in 

design/sizing mode using a defined structural design gross weight (SDGW). The empty weight 

was calculated from this SDGW based on a detailed set of parametric-weight correlations and 

fixed-weight allocations. Rotor diameter was determined based on a design disk loading at a 

given midmission gross weight and rotor solidity was determined based on a design CT/σ

(thrust coefficient/rotor solidity) at the same midmission gross weight and ambient conditions. 

Wing area was determined based on a design wing loading at SDGW, and wingspan was 

determined based on a design-wing aspect ratio. 
14 

 

ModelCenter® Design Environment

Initialization:

SDGW

Critical Weights

Fuel Capacity

Disk Loading

Wing Loading

Wing AR

Weights

Geometry

Power & Torque

MISSION ANALYSIS

Mission 1 Mission 2

Mission 3 Mission 4

Mission 5

Performance

Checks

Mission 6

Mission 7

Mission 8

Driving

Missions

Critical Weights

Rotor D / σ

Engine Power

Transmission Torque

MTOWSDGW

Fuel Capacity

Max. Fuel

Capacity

Converged?

Sizing NO YES Exit

RDM Design/Sizing

RDM Mission Analysis

Start

Figure 4. LVLT synthesis process. 

2. All required missions were executed using RDM in mission-analysis mode using the weight, 

power, and geometry information generated by RDM in its first run in design/sizing mode. 

The baseline fuel required of each mission was compared with the maximum value extracted 

as the updated value of the fuel capacity. The takeoff gross weight of each mission was 

compared with the maximum value extracted as the updated value of SDGW. A critical 

midmission hover-out-of-ground-effect (HOGE) weight was also extracted. 

3. RDM was executed in design/sizing mode again and the updated values of SDGW, 

midmission gross weight, and fuel capacity were used to determine a new aircraft empty 

weight, engine power required, transmission rating, rotor geometry, and wing geometry. 

4. Steps 2 and 3 were repeated until the weights and fuel capacity were converged upon. 

Two primary missions emerged as the aircraft sizing drivers. The first was a radius-of-action 

(ROA) mission with a midmission HOGE, which sets the minimum fuel capacity and whose midmission hover criterion dictates the main-rotor transmission rating and the installed engine power 

required. The HOGE takeoff criterion for this mission dictates the vertical takeoff weight of the 

aircraft. The second mission was a long-range transport mission that drove the maximum takeoff 

gross weight (MTOW) of the aircraft and in turn the structural design gross weight (SDGW). 
15 

3.2 Design Space Exploration and Optimization Strategy 

The design space exploration and optimization approach was to present the best CTR, best VDTR, 

and best VSTR for the given LVLT requirements, optimizing each aircraft independent of the 

other two so that objective comparisons could be made on the merits of each configuration. The 

three configurations were sized to the same objective function and minimum aircraft empty weight 

(to minimize logistic footprint and cost), and subjected to performance and several physical, geometric, or operational constraints. Thus, all three aircraft emerged with comparable performance, 

but different attributes (physical size, weight, and geometry) necessary to meet the LVLT requirements indicative of each configuration. 

A single performance-based constraint requiring a minimum threshold for cruise efficiency, 

derived from the LVLT requirements, coupled with the minimum empty-weight objective function, was chosen to yield the “fastest, lightest, cheapest” aircraft possible. 

The independent design variables chosen for optimization were rotor disk loading, wing loading, 

and wing aspect ratio. These design variables are inextricably linked to the physical/geometric 

constraints present in the design space. These constraints include: 

1. Minimum wing inboard aspect ratio: Minimum wing root chord is dictated by wing structural 

requirements. While higher wing aspect ratios are desirable because they decrease induced 

drag and result in reduced fuel consumption, high wing aspect ratios correspond to lower wing 

root chords for a given span. Therefore a balance between cruise efficiency and adequate 

structural stiffness was required. 

2. Maximum rotor-blade aspect ratio/minimum rotor-blade chord: Maximum rotor-blade aspect 

ratio was dictated by structural requirements, dynamic requirements, and the physical 

constraints of the VDTR retraction mechanism. Lower rotor disk loadings were desirable for 

all configurations because installed engine power required is driven by the HOGE condition, 

and specifically for VDTR because the VDTR suffers a figure-of-merit penalty due to the 

elliptically shaped inboard section of the rotor. However, as disk loading decreases, for a given 

hover design CT/σ, blade aspect ratio becomes higher. A maximum blade-aspect-ratio limit 

was imposed on the design to bound the disk-loading trade space. Further, the VDTR blade 

was constrained to have a minimum allowable geometric chord based on the packaging volume 

needed for the retraction mechanism. 

3. Maximum rotor disk loading: A maximum rotor disk loading for the midmission hover 

condition was selected to maintain outwash velocities less than or equal to those currently 

accepted within the Department of Defense (DOD) inventory. 

4. Maximum operating-wing lift coefficient (CL): The wing loading was constrained to yield the 

maximum allowable operating-wing CL. The CL corresponding to the maximum lift/drag 

coefficient (CL/CD) was extracted for each Mach number to derive the maximum wing CL as a 

function of Mach number. For any given Mach number, the maximum CL, with a margin of 

20%, was selected as an upper bound to the capability of the wing section. Higher wing 
16 

loadings were desirable because they corresponded to less wing area for a given vehicle 

weight, directly reducing wing weight and overall vehicle empty weight. However, for a given 

span, higher wing loadings resulted in increasingly high operating-wing CL. 

A final imposed constraint was the relationship between hover rotor diameter and wingspan. While 

the wingspan, aspect ratio, and wing loading should be set to achieve the most efficient wing 

possible, the hover disk loading should be set to achieve the most efficient hovering rotor possible. 

Inboard (nacelle-to-nacelle) wingspan and rotor diameter are coupled by a physical constraint. For 

CTR and VSTR, that constraint is the fuselage width plus one foot of margin on either side; and 

for VDTR the constraint is that the two rotors are not permitted to intermesh (intermeshing rotors 

pose the concerns of asymmetrical dynamics and increased fuselage downloading as more of the 

rotor wake contraction is impeded by the fuselage). 

3.3 Design Descriptions 

Figure 5 through Figure 20 graphically compare the three NASA LVLT tiltrotor concepts. Key 

parameters and weights are shown, with the CTR considered the nominal 100% point and the 

values for the other aircraft shown as relative to the CTR. The VDTR rotor-retraction mechanism 

results in a 9% larger hover diameter (translating to 15% lower disk loading, 15% less rotor 

solidity required, and 3% greater overall hover efficiency) and a 16% lower cruise diameter 

relative to the CTR. The VDTR cruise tip speed is 13% less than the CTR and only 2% greater 

than the VSTR (a result of both reducing the rotor diameter and slowing the rotor NR). These 

benefits come at the expense of 3% greater weight empty and 2% greater required fuel capacity 

relative to the CTR. Development cost is 11% greater than the CTR for the VDTR and 5% greater 

than the CTR for the VSTR, with the difference due primarily to the level of risk inherent in each 

configuration. Unit flyaway price trends similarly with risk, but more so with vehicle weight, with 

the VDTR being 9% greater than the CTR and the VSTR being 7% greater than the CTR. Recall 

that all three aircraft were optimized to equivalent cruise efficiency. 

100 102 100

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

SDGW

Aircaft Comparison

Structural Design Gross Weight

100 103 100

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

WE

Aircaft Comparison

Empty Weight

Figure 5. SDGW comparison. Figure 6. Weight empty (WE) comparison. 
17 

100 95 101

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Installed Engine Power

Aircaft Comparison

Installed Engine Power, MRP SLS

100 96 101

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

XMSN Rating

Aircaft Comparison

Main Rotor Gearbox Rating

Figure 7. Installed maximum rated power (MRP) 

at sea-level standard (SLS) comparison. 

Figure 8. Transmission rating comparison. 

100 102 100

0

50

100

150

200

250

300

CTR VDTR VSTR

Internal Tank Fuel Capacity

Aircaft Comparison

Internal Tank Fuel Capacity

100 102 100

0

50

100

150

200

250

300

CTR VDTR VSTR

Long Range Tank Fuel Capacity

Aircaft Comparison

Long Range Tank Fuel Capacity

Figure 9. Internal-fuel-tank volume comparison. Figure 10. Long-range-tank volume comparison.
18 

100

109

96

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Main Rotor Diameter (Hover)

Aircaft Comparison

Main Rotor Diameter (Hover)

100

84

96

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Main Rotor Diameter (Cruise)

Aircaft Comparison

Main Rotor Diameter (Cruise)

Figure 11. Rotor hover-diameter comparison. Figure 12. Rotor cruise-diameter comparison. 

100 100 100

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Main Rotor Tip Speed (Hover)

Aircaft Comparison

Main Rotor Tip Speed (Hover)

100

77 75

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Main Rotor Tip Speed (Cruise)

Aircaft Comparison

Main Rotor Tip Speed (Cruise)

Figure 13. Rotor hover-tip-speed comparison. Figure 14. Rotor cruise-tip-speed comparison.
19 

100

85

108

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Hover Disk Loading

Aircaft Comparison

Hover Disk Loading

100

85

108

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Rotor Solidity

Aircaft Comparison

Rotor Solidity

Figure 15. Hover disk-loading comparison. Figure 16. Rotor solidity comparison. 

100 103 102

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Hover Efficiency

Aircaft Comparison

Hover Efficiency

100 100 100

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Cruise Efficiency

Aircaft Comparison

Cruise Efficiency

Figure 17. Hover-efficiency comparison. Figure 18. Cruise-efficiency comparison. 
20 

100

111 105

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Development Cost (SD&D)

Aircaft Comparison

Development Cost (SD&D)

100

109 107

0

25

50

75

100

125

150

175

200

CTR VDTR VSTR

Unit Flyaway Price

Aircaft Comparison

Unit Flyaway Price

Figure 19. Development-cost comparison. Figure 20. Unit-flyaway-price comparison. 

The three LVLT aircraft configurations are similar in their overall architecture and are described in 

the following section collectively. Unique features of the configurations are noted and described. 

The LVLT aircraft all follow the lines of classic tiltrotor configurations. A large rotor is mounted 

on each wingtip, providing axial propeller thrust in the cruise configuration, and is rotated around 

a tilt axis to provide vertical-lift thrust in the hover configuration. 

3.3.1 Main-Rotor System 

All three aircraft employ a composite, flexbeam, nongimbaled rotor hub design as an alternative to 

a classic, gimbaled design. In concept, the composite flexbeam is mated with a pitch-bearing pair 

and allows for flapping motion while maintaining a stiff-in-plane rotor, an advantageous arrangement for mitigating propeller-whirl-flutter instability. The pitch-bearing pair isolates the flexbeam 

from the extreme collective pitch-range requirements required for high-speed tiltrotors. 

The VDTR retraction mechanism is separated from the physical hub geometry, allowing for a 

simpler rotor hub, improved retraction-system life, and blade fold. The VSTR rotor design mirrors 

the work for the other configurations. During cruise the rotor rpm and centrifugal forces (CF) are 

significantly reduced. Maintaining rotor stability at these low rpms and reduced CF is a challenge 

for the VSTR rotor design, requiring the rotor of the VSTR to be designed with significantly 

increased hub and blade stiffness to maintain blade stability. The variable-speed aspect of the 

VSTR also affects the rotor design in more subtle ways such as blade-twist geometry and rotorblade dynamics and tuning for the large-rpm variation. 

The CTR and VSTR blades are traditional composite blades utilizing a graphite-blade spar, 

fiberglass skins, and titanium/nickel abrasion strip. Cross-sections were developed in Computer 

Aided Three-Dimensional Interactive Application (CATIA) and then iterated through ANSYS and 

Rotorcraft Comprehensive Analysis System (RCAS) to develop a Southwell plot that met dynamic 

stability requirements and resulted in a rotor with acceptable aerodynamic performance and strain 

levels. 
21 

The VDTR blade-design effort had two primary focuses: developing and refining a retraction 

system to govern the reduction in rotor diameter and then integrating that system into an 

acceptable rotor design. The retraction-drive research and development was assisted by the largescale testing of VDTR systems completed by Sikorsky in the late 1990s, and the resulting LVLT 

VDTR system is believed to be an improvement over those efforts. 

A trade study was conducted to select a blade retraction and extension system. While leaning 

heavily on past Sikorsky efforts, new configurations not previously studied were also considered. 

Previous Sikorsky efforts concentrated on two broad concepts: a ball screw embedded in the blade 

to serve as a linear actuator, and a fibrous strap that retains the blade and is wound around a 

retraction drive in the hub. Each concept was considered in detail and traded against one another. 

The strap-driven concepts studied react the blade centrifugal force with a woven strap or cable and 

then retract the blade by winding the strap around a drum at or near the center of the rotor. A 

primary advantage of these concepts is reduced weight. A second advantage is increased ballistic 

tolerance. A woven strap would be several inches wide, and the likelihood of a round of 

ammunition cutting through a critical cross-section is lower than through a ball screw of smaller 

diameter. Furthermore, a damaged strap can still be wound onto a drum, while the precise 

tolerance nut and ball bearings are unlikely to travel over a damaged section of a ball screw. Given 

the advantages of the strap and reel retraction mechanism compared to ball-screw actuators, a strap 

was selected as the primary candidate for design of the retraction system. 

Several power sources are compatible with each of these ideas, including electrical, hydraulic, or 

direct drive from the transmission. The final selection of a concept and power source was a 

balance between weight, reliability, ease of integration, and redundancy, along with the 

requirement to maintain synchronous retraction and extension of the four rotor blades to maintain 

the rotor balance. 

3.3.2 Drive System 

The drive systems for the three aircraft configurations (CTR, VDTR, and VSTR) were developed 

around a fixed-engine configuration baseline that allows the rotor and primary gearbox to rotate 

for hover, but keeps the engines horizontal. This configuration eliminates hot exhaust-gas 

impingement on tarmac and ship-deck surfaces, but requires a unique gearbox and engine 

arrangement. 

The drive systems of the three LVLT configurations are similar in that each has two turboshaft 

engines that lie horizontally, aft of the gearboxes in the nacelle. The engines are spaced to provide 

an adequate mounting footprint for the main gearbox, which must rotate with the rotor. Each 

engine feeds into a stationary nose gearbox, the redirected output of which feeds into the main 

module section of the main-rotor gearbox (MGB) and supplies power to the rotor though a collector face-gear mesh and conventional planetary assembly. Structurally, in addition to containing the 

drive gears, the gearbox must support all of the flight loads from the rotor and distribute them to 

the airframe. This support is achieved with a titanium main-rotor shaft, or mast, and an advanced 

aluminum housings and bearing system. 
22 

All three LVLT configurations employ a cross-shaft system to synchronize the proprotors and to 

provide protection against an engine failure. The VSTR gearbox requires an additional gearbox to 

provide the shifting mechanism to enable low rotor speeds for high-speed flight. The shifting gearbox is a two-speed mechanism using a low-power locking clutch (ref. 5). 

3.3.3 Flight-Control System 

The flight-control systems (FCSs) for the LVLT aircraft are digital, dual-triple-redundant fly-bywire (FBW) flight-control systems incorporating advanced control laws and autopilot capability. 

The LVLT aircraft employ FBW in the FCS to capitalize on significant savings in system weight, 

complexity, and maintenance cost, as well as improvements in safety and capability over traditional mechanical flight-control systems. The FBW FCS enables tailored handling qualities and 

precision approach modes using the coupled flight director. 

With FBW, the maintenance burden is significantly reduced because of the elimination of various 

components (e.g., trim servos, mixer assemblies, etc.) that require servicing, rigging checks, and 

other maintenance actions. The FBW electronic components are highly reliable and redundant. 

3.3.4 Propulsion System 

The propulsion systems of the three LVLT aircraft are based on four engines mounted in pairs in 

each wingtip nacelle. The engines are mounted in the fixed frame, as opposed to the tilting frame 

like the V-22 and XV-15 configurations. By not rotating the engines with the rotors, some weight 

savings was realized in the engines themselves as the engine bearings and lubrication systems were 

not required to operate over such a large variation in orientation. Operational considerations, however, favored the fixed-engine arrangement. 

The engines selected for the aircraft are high-efficiency turboshaft engines designed to operate as 

efficiently as possible over a large range of rpms. These engines extract the maximum amount of 

energy from the exhaust but also have very high exhaust temperatures. Without cooling air mixing 

and significant infrared (IR) suppression efforts, these temperatures could pose a significant threat 

to the deck surfaces of ships and the asphalt surfaces of tarmacs, and could make operations from 

unprepared surfaces impossible. By keeping the engines fixed in the wing frame, the exhaust is 

ejected horizontally into the rotor downwash, providing ambient air mixing and preventing the 

direct impingement of exhaust gases on any ground surfaces or personnel below the rotors. 

3.3.4.1 Inlets and Exhaust 

To maximize the forward-flight performance cruise efficiency, the minimal inlets required for 

cruise flight face forward on the lower edges of the nacelle below the wing. The inlets were 

designed for minimal losses in the high-speed airflow environment. 

On the lower surfaces of the nacelles, there is one larger hover inlet for each engine. The air 

passing through the hover inlets is “filtered” using an inlet particle separator (IPS) with no moving 

parts and very high separation efficiency. During hover and low-speed flight, the cruise inlets are 

blocked off, forcing all of the inlet flow through the IPS. During cruise flight, the positive pressure 

in the inlet duct prevents additional flow from coming through the IPS. 
23 

The exhaust systems of the aircraft employ a simple IR suppressor system. Since the engines are 

fixed horizontally during all flight regimes, the IR suppressor is not required to mix all of the 

exhaust flow down to safe operating temperatures for ground operations, but simply needs to 

provide the best line-of-sight (LOS) blockage to the engines while minimizing the performance 

impact on the engines. 

3.3.4.2 Fuel System 

The fuel system of the LVLT aircraft is similar to other fixed-wing transport aircraft. The primary 

fuel cells are located between the spars of the main wing in typical “wet-wing” fashion. The fuel 

cells in the wings are capable of carrying the fuel required for all the LVLT missions, allowing the 

cabin to be clear of auxiliary fuel systems for these missions. For longer-range missions, auxiliary 

tanks could be mounted in the cabin and plumbed into the aircraft fuel system, allowing for the 

vehicle management system to manage the appropriate fuel distribution for the given mission and 

cruise condition. There are no center wing tanks, but wing box tanks could be added if required. 

All tanks are cross-linked, allowing the vehicle management system to manage the fuel 

distribution for balance and mission requirements. As the aerodynamic and structural requirements 

for the wing drove the minimum wing thickness ratio, there is ample fuel volume available within 

the wings for each of the configurations. 

3.3.5 Wings and Winglets 

The wings of the LVLT aircraft are key components in the LVLT system design. The wing design 

completed a full iteration, including the design of a customized, supercritical airfoil, optimized 

wing planform, and a finite-element-method (FEM) analysis of the wing structure. As in other tiltrotor configurations such as the V-22 and XV-15, the LVLT wing is thicker than large fixed-wing 

transport designs. The wings of the V-22 and XV-15 aircraft are 23% thick, whereas the wings of 

the LVLT aircraft are more aggressive. The thinner LVLT wing sections improve cruise efficiency 

but complicate the aeroelastic concerns of the design. 

The structure of the LVLT wings is very similar to that of typical fixed-wing designs with a main 

spar forward and a secondary spar aft. The spars are joined by structural wing skins that carry the 

wing torsion moments created by the sweep and nacelle position. 

The wings are equipped with both flaps and ailerons for flight control in hover and forward-flight 

modes. The sizing efforts for all three configurations have assumed simple flap and aileron 

geometries with the flaps and ailerons sized for 25% of wing chord. This assumption provided a 

good balance between reducing download and the remaining structural effectiveness of the wing 

spar and torque tube. 

3.3.5.1 Wing-Planform Optimization 

The LVLT aircraft sizing optimization was completed with the wing-planform variables allowed to 

float. The optimizer was allowed to adjust aspect ratio, incidence, and wing loading. For the CTR 

and VSTR, the inboard wing (between the rotors) is determined by the rotor radius and fuselage 

width. The space between the rotors was set at 25 ft, which includes space for the fuselage and 

additional clearance between the rotor tips and the fuselage. For the VDTR in hover, the rotors 

were allowed to overlap the fuselage but not intermesh. As a result, VDTR rotors with a 77% 
24 

reduction ratio allow the inboard wing to be smaller than the other configurations while still 

allowing adequate clearance between the rotor tips and the fuselage in the cruise configuration. As 

RDM manipulates the sizing variables (such as disk loading and wing loading), the weight effects 

were calculated and allowed to influence the overall design. 

The overall wing planform (wing plus wing tips) was determined by wing loading and total aspect 

ratio. Wing loading was one of the independent optimization variables but was subject to the constraint that the CL at the midpoint of the heaviest cruise segment of the LVLT mission be set to be 

at least 20% less than the maximum CL of the airfoil at the cruise speed of the segment. This constraint prevented the optimizer from progressively increasing the wing loading, limiting wing size 

and weight. 

3.3.5.2 Winglet Design 

Articulated winglets proved to be the optimal solution to balancing weight and performance across 

the entire flight envelope. The wingtip configuration is also ideal for short takeoff performance 

optimization as the wingtips can be tailored for the desired flight performance. The final planform 

of the winglets for each configuration was determined by the wing planform optimization, and the 

span of the winglet was determined by the difference between the overall wingspan and the 

inboard wingspan. Because there is no significant mass outboard of the winglet, the airfoil of the 

winglet is much thinner than the inboard-wing airfoil. 

The winglets are actuated to operate in the range of approximately +10° to –90° independently of 

the nacelle position. During hover operations the winglets are vertical, and during transition the 

winglets are actuated so that the winglet operates in a low-drag lifting position. During high-speed 

cruise flight, the winglets are fixed for efficient cruise. 
25 

4 Attributes and Aircraft Performance 

The baseline CTR, VDTR, and VSTR LVLT aircraft performance was modeled using Sikorsky’s 

Rotorcraft Design Model (RDM) using rotor aerodynamic performance data generated by the 

EPHIC (Evaluation of Hover Performance using Influence Coefficients) program. 

As illustrated in Figure 4, the LVLT hover requirement sized the rotor diameter, solidity, engine 

power, and gearbox power and torque limits. The inboard span of the wing was dictated by the 

rotor diameter to provide adequate fuselage clearance in cruise for CTR and VSTR and to provide 

adequate rotor-to-rotor clearance for the VDTR in hover. Wing area was set to provide reasonable 

cruise CL for the aircraft based on the fallout cruise speed capability and LVLT cruise-mission 

weights. This area was added to both the outboard wingtips, and chord length was added over the 

inboard wingspan. The sensitivity of the design based on aspect ratio, wing loading, and disk loading was considered; the final geometry offering minimum empty weight was ultimately chosen. 

4.1 Attributes Summary 

4.1.1 Air Vehicle Aerodynamics 

The total equivalent flat-plate drag areas at zero-degrees angle of attack of the CTR, VDTR, and 

VSTR were estimated using a methodology based upon wind tunnel drag measurements, empirical 

drag data, and computational fluid dynamics (CFD)-based analytical tools. Figure 21 illustrates the 

CTR, VDTR, and VSTR equivalent flat-plate drag area versus structural-design gross weight with 

historical aircraft for comparison. The constant lines represent levels of skin-friction coefficient, 

Cf, which is the constant of proportionality between the equivalent flat-plate drag area, fe, and 

gross weight (GW)2/3. Eq. Flat Plate Area (sqft)

Takeoff Gross Weight (lb)

Aircraft Comparison

Parasite Drag Trend

Rotorcraft

Fixed Wing

CTR/VDTR/VSTR

Figure 21. Equivalent flat-plate drag comparison. 
26 

The vertical drag of the LVLT aircraft was calculated using a Circulation Coupled Hover Analysis 

Program (CCHAP) file created with aircraft design parameters based on the CTR configuration, 

including fuselage and rotor-blade properties (planform and airfoil data). The CCHAP input file 

represented one-half of the fuselage, including one wing and one rotor. A symmetry plane was 

assumed at butt-line 0.0. Rotor thrust was assumed to be one-half the total thrust in hover. 

Assuming a rotor thrust and wing CD, the vertical drag for the CTR wing under the rotor was 

calculated, resulting in a vertical drag similar to that of existing tiltrotor aircraft. A vertical-drag 

increment was applied to represent the rotor-wake-fountain contribution to download, based on 

V-22 data. 

4.1.2 Rotor Aerodynamics 

This study employed a method to quickly identify optimized blade geometries for any specified 

tiltrotor mission. A modified version of the Evaluation of Hover Performance using Influence 

Coefficients/Helicopter Rotor Optimization (EHPIC/HERO) analysis was developed by 

Continuum Dynamics Incorporated (CDI), as discussed in reference 11. A unique aspect of the 

approach is the ability to address highly nonlinear distributions of blade chord, twist, and sweep. 

The EHPIC/HERO analysis combines a highly efficient and accurate free-wake performanceprediction model with a comprehensive design-optimization algorithm to create an effective tool 

for design optimization of rotorcraft in hover and axial flight. Furthermore, this study incorporated 

Sikorsky proprietary code for optimizing blade geometry to simultaneously meet two separate 

design conditions. For proprotors, these conditions are the hover and cruise design points. 

The development of this analysis was rooted in the study of "open-loop" performance prediction 

for hovering rotors. Early efforts to develop free-wake hover models using time-domain 

calculations were hampered by long computation time and poor (or nonexistent) convergence due 

to the inherent instability of the wake beneath a hovering rotor. Reference 12 describes the 

development of the baseline EHPIC free-wake relaxation scheme using influence coefficients that 

avoid these convergence problems. An initial guess for the blade loads and wake geometry is 

adjusted in a succession of solution steps using influence coefficients until the physically correct, 

self-preserving wake geometry is obtained. The unsteady wake beneath a hovering rotor is 

unstable about this solution, but average performance characteristics can be obtained using this 

wake geometry. This fact has been substantiated through numerous correlation studies with hoverperformance measurements, some of which are cited in references 12 and 13. In EHPIC, rotorperformance characteristics are obtained by coupling the free-wake analysis to a vortex lattice 

treatment of the blade similar to that described in reference 14. This analysis can be extended to 

axial flight (i.e., helicopters in vertical climb or tiltrotors in cruise flight) simply through the 

addition of an appropriate uniform free-stream velocity. 

The single-point optimization found in the general release of EHPIC/HERO optimizes blade 

geometry for hover or cruise, but not both simultaneously. Single-point optimization is valuable 

for understanding blade-geometry traits that improve performance in hover or in axial flight, but 

cannot be used to address the design compromises necessary for rotors that need to operate 

efficiently in several different flight conditions. With dual-point optimization, blade geometries are 

found that meet any prescribed weighting combination of the hover and cruise design points. 

Further, sweeping the design-goal weighting factor quickly produces a family of optimized blade 
27 

geometries that extend from pure hover to pure cruise. Reference 11 provides correlation 

comparisons of the EHPIC/HERO results with V-22 test data. 

The desire to obtain the minimum possible rotor diameter during cruise was found to limit the 

combined hover/cruise performance of the VDTR. Although reducing blade diameter to limit 

excessive cruise-blade-profile power is fundamental to the VDTR concept, excessive diameter 

reduction actually decreased blade efficiency. This phenomenon occurred because excessive 

retraction overly constrained other important rotor design variables. Specifically, maximizing 

blade retraction constrains blade twist across 85% of the hover radius and 74% of the cruise radius. 

Because the large region of linear twist required for the retractable portion of the blade inhibited 

the nonlinear twist needed to approximate desirable distributions, overall blade efficiency was 

penalized rather than improved. 

The solution to this conflict was to limit the cruise-blade retraction to 77% of the hover radius. 

This limitation was used with an NR reduction to further increase the cruise-blade loading and 

decrease the cruise-tip Mach number. A secondary VDTR efficiency improvement was obtained 

by thinning the blade outer airfoil sections. Previous 66% radius cruise-retraction designs required 

a larger over-sliding region than the current 77% scheme. Relieving the need for thicker airfoils on 

a portion of the outer radius was beneficial in both hover and cruise. 

The rotor-blade twist distribution for each aircraft was selected from a FM-ηP boundary that 

progressed from an ideal hover twist to an ideal propeller twist. Twist distributions for boundary 

points between the two extremes were then interpolated from the two end-point patterns. Figure 22 

shows the relative FM-ηP boundaries for all three aircraft, normalized to the peak CTR FM and the 

peak CTR ηP. Figure 23 and Figure 24 show the relative design FM and ηP of the three aircraft. 

Note that the VDTR achieves 4% greater propulsive efficiency than the CTR, while losing only 

3% FM. 

Estimated rotor efficiencies were judged against similar projections using codes other than EHPIC. 

These calculations added confidence when the checks align, and served as a flag for more in-depth 

scrutiny when not aligned. To that end, limited cruise- and hover-efficiency comparisons were 

made between EHPIC predictions and those calculated with two additional rotor aerodynamic 

codes. Specifically the CDI CHARM code was exercised to estimate comparative hover FM and 

cruise ηP values for the CTR configuration, and both CHARM and LSAF were used to check the 

VDTR calculations. These additional evaluation codes were selected as having mature evaluation 

capabilities. 

Figure 25 shows good EHPIC ηP validation against both CHARM and LSAF for the VDTR and 

between EHPIC and CHARM for the CTR. The high degree of agreement was not unexpected, 

since accurate propulsive-efficiency estimation for most tiltrotor designs is not difficult because of 

the relatively low importance of induced power loading for propellers operating at high advance 

ratios. Similarly, ηP prediction requires only that the codes properly resolve spanwise angle-ofattack and Mach number distributions, while referencing common 2-D airfoil tables. 
28 

90

91

92

93

94

95

96

97

98

99

100

101

102

103

104

105

95 96 97 98 99 100 101 102 103 104 105

Relative Rotor Hover Efficiency (FM)

Relative Rotor Cruise Efficiency (ηP)

Aircraft Comparison

FM-η P Performance Boundary Comparison

CTR

VDTR

VSTR

Figure 22. Hover-cruise-efficiency design trends. 

100

97

101

95

96

97

98

99

100

101

102

103

104

105

CTR VDTR VSTR

Relative Design FM

Aircraft Comparison

Relative Design FM

100

104

103

95

96

97

98

99

100

101

102

103

104

105

CTR VDTR VSTR

Relative Design 

ηP

Aircraft Comparison

Relative Design η P

Figure 23. Design FM comparison. Figure 24. Design ηP comparison. 

 
29 

100.00 100.00 100.12 100.00 99.64

90

92

94

96

98

100

102

104

106

108

110

VDTR

EHPIC

VDTR

LSAF

VDTR

CHARM

CTR

EHPIC

CTR

CHARM

Propulsive Eficiency EPHIC = 100

Aerodynamics Methodology Validation

Propulsive Efficiency Comparison

Figure 25. Cruise-efficiency-performance checks. 

Figure 26 presents the corresponding hover-efficiency comparisons. In this case the agreement for 

the VDTR was again good. However, significant hover discrepancies were shown for the CTR 

rotor geometry. This result was unexpected since the VDTR case was deemed more challenging 

because of unavoidable twist and airfoil t/c constraints on the over-sliding portion. Based on the 

observed VDTR FM prediction consistency, only the CTR EHPIC and CHARM cases were further 

examined. 

The figure shows predicted EHPIC and CHARM FM values for the Sikorsky CTR blade design. 

Again, both calculations were carried out at the CTR design CT/σ and a common tip Mach 

number. It is noted that the large FM drop predicted with CHARM was limited to calculations at 

high blade loadings. For these conditions, CHARM predicted large tip angles of attack with local 

drag divergence, caused by released CHARM CTR tip-vortex filaments that stayed in the rotor 

plane and impacted the following blade. 

The red bar in Figure 26 shows that the adverse CHARM CTR response to the calculated tip 

environment can be lessened. Substituting a 10% thick SC36210 tip airfoil for the base CTR 7% 

tip airfoil increases the drag divergence boundary, removing about half of the FM penalty. 

Additionally, a modest change to the CHARM sweep model, bringing it into alignment with the 

EHPIC formulation, moves the CHARM FM up to 0.80 at the CTR design condition. Further 

CHARM FM improvements are not likely, given the high tip angles of attack seen by the code. 
30 

100.00 99.75 99.13

100.00

91.15

95.64

96.97

90

92

94

96

98

100

102

104

106

108

110

VDTR

EHPIC

VDTR

CHARM

VDTR

LSAF

CTR

EHPIC

CTR

CHARM

CTR

CHARM

+t/c

CTR

CHARM

+t/c + 

sweep

Hover Eficiency EPHIC = 100

Aerodynamics Methodology Validation

Hover Efficiency Comparison

Figure 26. Hover-performance checks. 

4.1.3 Wing Aerodynamics 

Wing planform selection was based on the need to meet specified mission performance while 

minimizing vehicle empty weight. Because of the interplay between wing structural weight and 

wing aerodynamic performance, precise wing planform closure required several individual 

structural and aerodynamic optimization loops. Wing structural optimization preceded the 

equivalent wing aerodynamic optimization. This approach was undertaken since wing geometric 

pitch washout, as well as wing airfoil shape, offered practical knobs for fine tuning aerodynamic 

performance for a given wing planform. 

Wing-performance tables were constructed for the isolated wing-prop geometry. These tables 

included separate induced and profile power terms for a variety of off-design wing flight 

conditions. After correction for wing-body and wing-nacelle installation, the tabular data were then 

used to produce an Oswald efficiency model for the installed wing-body–nacelle combination. The 

Oswald efficiency model was incorporated into RDM for the aircraft synthesis. 

Wing-airfoil-section aerodynamic design was performed using the MSES CFD program. Initially, 

an airfoil section was generated by scaling a NASA 18% thick supercritical airfoil to the root 

thickness required by the wing structural analysis. At the airfoil section design condition, a 

supersonic region and shock formed on the upper surface. The initial airfoil was then modified 

using the MSES mixed (viscous) inverse procedure with the objectives of eliminating wave drag at 

the design condition and improving the airfoil-section drag at lift coefficients greater than the 

design target by delaying the onset of supersonic flow on the upper surface. A shock-free 

chordwise pressure distribution was prescribed on the airfoil upper surface, and the inverse method 

was allowed to modify the mean camber line while maintaining the initial design thickness 
31 

distribution. This modification produced the desired pressures on the upper surface while allowing 

the pressures on the lower surface to change. The Mach contours of the scaled airfoil and the 

designed airfoil are shown in Figure 27 and Figure 28, respectively. 

At the design condition, the resulting designed airfoil had the same drag as the initial airfoil 

because the reduction in wave drag had been offset by an increase in viscous drag. At higher lift 

coefficients, the designed airfoil had lower drag because the development of supersonic flow was 

delayed relative to the initial airfoil. The improved performance of the designed airfoil for higher 

lift coefficients, however, was the result of increased camber, which caused more negative 

(unfavorable) values of pitching moment. 

Supersonic region 

terminates with weak 

shock at x/c = 0.16

No supersonic flow on 

upper surface

Figure 27. Scaled airfoil Mach contours. Figure 28. Designed airfoil Mach contours.

4.1.4 Aeroelastic-Stability Analysis 

A literature review was conducted on the state-of-the-art analytical methods and tools available for 

aeroelastic stability predictions of tiltrotor aircraft. Theoretical and experimental studies of 

proprotor and tiltrotor whirl flutter were surveyed (refs. 5 and 15 through 18). The influence of 

design variables such as wing design (refs. 5 and 18), active stability augmentation (ref. 19), and 

variable geometry rotors (refs. 13, 20, and 21) on whirl-flutter stability was reviewed. Tiltrotor 

aeroelastic stability modeling methods and capabilities of aeroelastic analysis codes such as 

RCAS, Comprehensive Analytical Model of Rotorcraft Aerodynamics & Dynamics (CAMRAD), 

DYMORE, University of Maryland Advanced Rotorcraft Code (UMARC), and Proprotor 

Aeroelastic Stability Analysis (PASTA) were investigated (refs. 8, 9, 14, 20, 22, and 23). Based on 

the literature review and Sikorsky experience, RCAS was selected as the primary aeroelastic 

analysis tool. 

RCAS validation was twofold. First, RCAS was validated for accuracy of whirl-flutter stability 

predictions though correlation with test data from the Wing and Rotor Aeroelastic Testing System 

(WRATS) and the corresponding DYMORE model data (Test 543 of the V-22 1/5 scale model 

conducted in the NASA Langley Transonic Dynamics Tunnel (TDT) in 2000). Second, the process 
32 

of modeling a detailed tiltrotor in RCAS was validated by comparing V-22 model results to analytical results from tools that were used to design the V-22 rotor. 

A NASA Structural Analysis Program (NASTRAN) model of the WRATS Ground Vibration Test 

(GVT) configuration was developed based on the DYMORE GVT model received from the Army 

Research Lab (ARL) at NASA Langley. The geometry, element meshes, material, and section 

properties were translated from the DYMORE model into the NASTRAN model. The NASTRANpredicted modal frequencies and mode shapes at the rotor hub correlated well with the DYMORE 

and GVT measurements. A comparison of the NASTRAN, RCAS, and DYMORE analytically 

predicted frequencies with the GVT measured frequencies for both the off- and on-downstop configurations is shown in Table 1. The validated NASTRAN model was converted to the RCAS 

model for whirl-flutter analysis. 

TABLE 1. MEASURED AND PREDICTED WING-NACELLE FREQUENCIES 

Mode Prediction / 

Test Source 

Frequency (Hz)

Off-downstop On-downstop 

Wing Beam 

GVT Test 5.43 5.83 

DYMORE 5.16 N/A 

NASTRAN 5.14 5.59 

RCAS 5.28 5.68 

Wing Chord 

GVT Test 8.11 8.67 

DYMORE 8.28 N/A 

NASTRAN 8.33 8.33 

RCAS 8.30 8.31 

Wing Torsion 

GVT Test 10.54 12.02 

DYMORE 10.54 N/A 

NASTRAN 10.47 11.92 

RCAS 10.48 11.89 

RCAS predicted natural frequencies of an isolated, clamped WRATS blade were validated with 

test results. RCAS natural frequencies predictions for the clamped, nonrotating blade in a vacuum 

are compared with test results and DYMORE predictions in Table 2. 

TABLE 2. COMPARISON OF FREQUENCIES OF THE CLAMPED WRATS BLADE 

Mode Experiment (Hz) * DYMORE (Hz) * RCAS (Hz) 

1st Flap 12.29 12.08 12.02 

1st Chordwise 34.11 34.30 36.68 

2nd Flap 52.44 56.25 53.34 

1st Torsion 113.35 115.18 113.79 

* Data from ref. 9 
33 

Validation of the isolated full rotor model was also been performed with a locked gimbal. The 

purpose of running the model with locked gimbal was to obtain blade frequencies with proper 

modeling of the flexbeam and control systems. The RCAS prediction of nonrotating frequencies 

for a single blade with flexbeam and control system is compared to experimental and DYMORE 

results in Table 3. The nonrotating blade natural frequencies correspond to a case with a 10-degree 

collective angle and locked gimbal. 

TABLE 3. NONROTATING ROTOR FREQUENCIES WITH FLEXBEAM AND ROTOR 

CONTROL | GIMBAL LOCKED 

Mode Experiment (Hz) * DYMORE (Hz) * RCAS (Hz) 

1st Flap 6.60 6.83 6.91 

1st Lag 20.6 18.83 17.78 

2nd Flap 25.2 26.78 26.04 

3rd Flap 69.3 71.78 72.14 

1st Torsion 112.6 110.57 107.71 

* Data from ref. 9 

After the isolated rotor model with locked gimbal was validated, the full rotor model with 

unlocked gimbal was analyzed. The validation of RCAS-predicted frequencies with experimental 

data and DYMORE results is given in Table 4. 

 

TABLE 4. NONROTATING ROTOR FREQUENCIES WITH FLEXBEAM AND ROTOR 

CONTROL | FREE GIMBAL 

Mode Experiment (Hz) * DYMORE (Hz) * RCAS (Hz) 

Gimbal 2.0 2.33 2.02 

Cone 6.8 6.80 7.20 

1st lag 19.7 18.74 18.50 

2nd Flap 25.0 26.69 25.64 

1st Torsion 112.1 110.57 106.14 

* Data from ref. 9 

The complete semispan WRATS model, created by combining the GVT model and the rotor 

model, was used to predict the whirl-flutter stability boundary of the wind tunnel test configuration 

in airplane mode. Whirl-flutter stability analysis was performed for the off-downstop 

configuration, which has pitch-flap coupling (delta-3) of –15 degrees and a rotor speed of 742 rpm. 

The airspeed was increased from 50 knots until the whirl-flutter speed was reached. The variations 

of the predicted and measured wing-beam-mode frequency and damping with airspeed are shown 

in Figure 29 and Figure 30. 
34 

4

4.5

5

5.5

6

50 75 100 125 150

W

i

n

g

B

e

a

m

M

o

d

e

F

r

e

q

u

e

n

c

y

(

H

z

)

Airspeed (knots)

RCAS

DYMORE

Test Data

Test Curve Fit

Figure 29. Wing-beam-mode frequency correlation. 

-1

0

1

2

3

4

50 75 100 125 150

W

i

n

g

B

e

a

m

M

o

d

e

D

a

m

p

i

n

g

(

%

c

c

)

Airspeed (knots)

RCAS

DYMORE

Test Data

Test Curve Fit

Figure 30. Wing-beam-mode damping correlation. 

 
35 

The black solid circles represent the measured data points obtained from the WRATS wind tunnel 

test database. The solid black line is the curve fit of the test results, similar to the data published in 

reference 18. The blue solid line with square symbols is the DYMORE prediction (ref. 9), while 

the red solid line with diamond symbols is the RCAS prediction. The measured and predicted 

whirl-flutter speeds are 142 knots from the test, 130 knots from DYMORE, and 135 knots from 

RCAS. 

An isolated V-22 rotor model was created in RCAS, complete with representation of the yoke, 

grip, pitch bearings, control-system stiffness, and blade. The frequencies and mode shapes (both 

collective and cyclic) were calculated for a low-thrust case of 38° root collective and compared to 

those in the Bell/Boeing EMD flutter report. A root collective of 38° corresponds to about 12° of 

pitch at r/R = 0.75. Note that the elastomeric pitch bearing stiffnesses were not known, so a stiff 

spring (1x108 lb/ft) was used for the transverse flexibility in both axes of the pitch bearing. Also, 

the drive-system stiffness was not known, so the stiffness of the lag hinge was varied until the rigid 

collective lag-mode frequency matched that found in the report. A tabulated comparison of the 

normalized frequencies is shown in Table 5 and Table 6. The frequency results show reasonable 

correlation. 

TABLE 5. V-22 COLLECTIVE FREQUENCIES | RCAS MODEL VS. EMD REPORT 

Mode # Mode Description EMD Report RCAS Model

1 Edgewise 1 0.34P 0.34P 

2 Flap 1 1.17P 1.22P 

3 Flap 2 2.74P 2.81P 

4 Edgewise 2 4.37P 4.50P 

5 Torsion 1 5.33P 5.36P 

6 Flap 3 6.17P 6.42P 

7 Torsion 2 9.26P 9.36P 

TABLE 6. V-22 CYCLIC FREQUENCIES | RCAS MODEL VS. EMD REPORT 

Mode # Mode Description EMD Report RCAS Model

1 Flap 1 1.002P 1.004P 

2 Edgewise 1 1.53P 1.47P 

3 Flap 2 2.40P 2.57P 

4 Torsion 1 4.49P 4.47P 

5 Flap 3 5.36P 5.21P 

6 Edgewise 2 7.06P 6.97P 

7 Torsion 2 / Flap 4 8.67P 9.00P 

8 Flap 4 / Torsion 2 9.60P 9.31P 
36 

The stability criteria consisted of a 20% margin from the maximum level cruise speed (VH) to 

account for dive speed (VDIVE) and a 15% margin from the dive speed as a stability margin. The 

final result was a stability margin of about 40% over the cruise speed. 

The modeling approach for whirl flutter was common to all three aircraft. The fully coupled RCAS 

model consisted of a 4-bladed rotor model (including blades, torque tubes, and flexbeams) coupled 

to a modal representation of the wing, which consisted of frequencies and mode shapes at the hub 

as derived from a wing NASTRAN model. A detailed kinematic control-system model, pre-cone, 

and the dual-load-path torque tube were all included in the model. The detailed control system 

modeled the full translation of swashplate motion into blade pitch, accounting for geometric kinematic coupling. The swashplate and pitch horn were assumed to be rigid, and a dynamically stiff 

scaled V-22 stiffness was applied. The push rod was positioned (radially offset from the effective 

flapping hinge) to provide a design negative delta-3 when the top of the push rod is in the plane of 

the rotor. 

Several assumptions were made in the whirl-flutter analysis. Specifically, a semispan NASTRAN 

wing model was used, so only symmetric wing modes (the first five elastic modes) were considered, as the symmetric modes are typically most important for whirl flutter. The same wing 

model was used for all the aircraft to allow for a fair comparison. Also, no wing aerodynamic 

effects were modeled; this assumption is conservative since aerodynamic damping was not 

accounted for. However, this assumption removed the potential to analytically predict wing divergence or fixed surface flutter. A conservative critical damping was applied to the wing modes 

based on the minimum elastic fuselage damping observed in shake test results of composite fuselage structures. To prevent the rotor polar moment of inertia from coupling with the wing and 

depressing the wing-mode frequencies, a torsional hinge was added between the rotor and the wing 

in lieu of modeling a separate, complex drive train. The rotor was trimmed to zero power. 

As designed, all three aircraft satisfied the whirl-flutter stability requirements and were stable up to 

the required 140% VH speed. Two sensitivity analyses were performed on all the aircraft to gain 

confidence in the models and to better understand the design space: rotor delta-3 and wing 

stiffness. 

Rotor delta-3 coupling was controlled by offsetting the pushrod from the blade-flap axis. A negative value indicates flap up–pitch up coupling where –15° means 1° of flap induces 0.25° of blade 

pitch. For 3-bladed gimbaled rotors like the V-22, optimal delta-3 angle is near –15°, with an 

inherent trade-off between optimizing flap-lag stability, which wants more negative delta-3, and 

whirl-flutter stability, which wants less negative delta-3. 

The sensitivity of whirl-flutter stability to delta-3 for all three aircraft plus an example gimbaled 

CTR is shown in Figure 31, where VH is represented as 100, the required stability speed is 

represented as 140, and the delta-3 angle is normalized to a baseline value of 100. The gimbaled 

CTR sensitivity is consistent with the WRATS model delta-3 sensitivity (and other published 

results) and points to an optimal delta-3 between –10° and –7°. However, this optimal delta-3 still 

does not provide adequate stability to satisfy the requirements. The flexbeam-style rotors satisfied 

the whirl-flutter requirements within a reasonable band of delta-3 angles. 
37 

0

20

40

60

80

100

120

140

160

-250 -200 -150 -100 -50 0 50 100 150 200

Whirl Flutter Speed, VH = 100

% Pitch-Flap Coupling, Delta-3 (Baseline = 100)

Gimballed CTR

Flexbeam CTR

Flexbeam VDTR

Flexbeam VSTR

Rotor Pitch-Flap Coupling (Delta-3) vs. Whirl Flutter Stability

(Cruise Altitude, Trimmed to Zero Power)

Requirement

Figure 31. Whirl-flutter-stability sensitivity to delta-3 coupling.

A wing-stiffness-sensitivity study was performed to compare the aircraft and to quantify the 

effects of variable speed and variable diameter on whirl-flutter stability. The wing stiffness was 

artificially reduced by applying a uniform factor to the NASTRAN material properties, resulting in 

wing-mode-frequency reductions equal to the square root of the applied factor. The sensitivity of 

whirl-flutter stability to wing stiffness for the three aircraft plus the gimbaled CTR is shown in 

Figure 32. The same wing was used for all aircraft to allow for a fair comparison. The wing was 

initially designed for a gimbaled CTR with whirl-flutter requirements being the limiting factor in 

the design. This wing design is close to satisfying the requirements for a gimbaled CTR but falls 

about 50 knots short. Switching to the flexbeam-style rotors significantly improves stability. The 

fact that a wing stiffness that is 20% of the baseline is required to induce instability with the 

flexbeam rotors indicates that whirl flutter is most likely not the design driver for the wing 

anymore. Instead wing loads, wing divergence, or fixed surface flutter will govern the wing design 

and could result in a wing that weighs less than the current wing after detailed design iterations. 

Reducing wing frequency introduces classical whirl-flutter instability involving coalescence of the 

rotor regressive flap mode, rotor regressive lag mode, and wing-beam bending mode. Zooming in 

on the case where the wing frequencies have been reduced by 70% (10% wing stiffness) allows for 

comparison of the different technologies (variable speed and variable diameter) relative to the 

conventional tiltrotor. 
38 

0

20

40

60

80

100

120

140

160

0 10 20 30 40 50 60 70 80 90 100

Whirl Flutter Speed, VH = 100

% Wing Stiffness (Baseline = 100)

Gimballed CTR

Flexbeam CTR

Flexbeam VDTR

Flexbeam VSTR

Wing Bending Stiffness vs. Whirl Flutter Stability

Cruise Altitude, Trimmed to Zero Power

Requirement

Figure 32. Whirl-flutter-stability sensitivity to wing stiffness.

The wing-beam-mode damping versus airspeed is plotted in Figure 33 for the three flexbeam 

rotors and example gimbaled CTR rotor. The gimbaled CTR is unstable at a very low speed (as 

would be expected since the frequencies have been reduced by 70% from the design point). The 

VSTR and VDTR show a respective 22% and 28% stability improvement over the flexbeam CTR. 

Most of the improvement for the VSTR, however, comes from the fact that the flexbeam is 2x 

stiffer than the CTR and the VDTR. The same stiffened flexbeam could be used on either of those 

rotors at the expense of increased loads. 

The whirl-flutter speeds for the four aircraft including the 1x and 2x stiffness flexbeam for the 

VSTR are summarized in Table 7 for the case where the wing frequencies have been reduced by 

70%. 

The aeroelastic-stability-analysis results indicated that the flexbeam-style hub configuration offers 

significant stability advantages compared to the gimbaled hub configuration. The flexbeam-style 

rotor changes the way the blades are constrained to each other so that a perturbation of one blade 

does not strongly affect the others because the blades are not hard mounted to a gimbaled 

structure. This design reduces rotor yaw deflections which in turn reduce the destabilizing 

gyroscopic moments applied to the wing, thus improving whirl-flutter stability. Furthermore, the 

flexbeam hub design and delta-3 schedule allow for better frequency separation between rotor flap 

modes, rotor lag modes, and wing modes throughout the airspeed range because the first rotor flap 

mode is relatively constant with airspeed. 
39 

The variable-speed design (VSTR) offers a slight whirl-flutter stability improvement over a 

conventional tiltrotor (CTR) due to reduced rotor gyroscopic loads resulting from a decrease in 

rotor speed. The variable-diameter design (VDTR) offers a greater stability improvement over a 

conventional tiltrotor (CTR) due to reduced rotor gyroscopic loads resulting from decreased rotor 

speed and decreased diameter. 

-300

-250

-200

-150

-100

-50

0

50

100

150

200

250

300

0 20 40 60 80 100 120 140 160

% Wing Beam Mode Damping Ratio 

(Baseline = 100)

Whirl Flutter Speed, VH = 100

Gimballed CTR

Flexbeam CTR

Flexbeam VDTR

Flexbeam VSTR

Wing Beam Mode Damping | 70% Wing Frequency

Cruise Altitude, Trimmed to Zero Power

Requirement

Figure 33. Wing-beam-mode damping vs. airspeed | 70% reduction in wing frequencies.

TABLE 7. WHIRL-FLUTTER SPEEDS FOR A 70% REDUCTION IN WING FREQUENCIES 

Aircraft Whirl Flutter Speed (VPWF=100) 

Gimbaled CTR 22 

CTR (1x flexbeam stiffness) 62 

VSTR (1x flexbeam stiffness) 63 

VSTR (2x flexbeam stiffness) 75 

VDTR (1x flexbeam stiffness) 79 
40 

4.1.5 Handling Qualities 

Simulations of the XV-15, CTR, VDTR, and VSTR were developed from information in 

references 13 and 24, and trim and flight dynamics were analyzed with both simulations. The 

GenHel XV-15 simulation was used to evaluate longitudinal trim, trim in turns, and dynamic 

responses to control inputs, and to correlate with information in references 1, 13, and 24. A generic 

tiltrotor simulation was used to calculate trim solutions for CTR, VDTR, and VSTR in straight and 

level flight and to derive a nominal mast-angle schedule vs. flight speed. 

The simulations were constructed using the Sikorsky General Helicopter (GenHel) Flight 

Dynamics Simulation program. GenHel is a generalized, modular analytical representation of a 

total helicopter system. The basic model is a total-force, nonlinear, large-angle representation in 

six rigid-body degrees of freedom. In addition to typical GenHel helicopter methodology, a 

software module was developed to calculate wing-pylon aerodynamic forces and moments, 

including wing download due to the rotors and downwash at the horizontal tail. A nacelle module 

was written to calculate the effect of mast tilt on the location of applied rotor forces and aircraft 

mass properties. 

The GenHel simulations utilized two sets of longitudinal trimmers, one for helicopter mode and 

another for airplane mode. In helicopter mode with the rotor lifting the aircraft, pitch attitude was 

varied to trim longitudinal acceleration and collective pitch was varied to trim normal acceleration. 

In airplane mode with the wing lifting the aircraft and the rotors tilted forward, rotor collective 

pitch was varied to trim longitudinal acceleration and pitch attitude was varied to trim normal 

acceleration. 

Flight dynamic responses to longitudinal control-stick inputs were calculated using GenHel to 

examine the simulation behavior and compare it with XV-15 flight-test data. The simulation was 

initially run in airplane mode at constant blade pitch and rotational speed, but small airspeed 

variations in response to stick inputs caused large changes in rotor thrust and torque. Therefore, a 

rotor-blade pitch controller was implemented and evaluated. 

Flight dynamic responses to lateral-control-stick and pedal inputs were calculated with GenHel to 

examine the simulation behavior and compare it with XV-15 flight-test data. The analysis included 

a comparison between simulation results and flight-test data for a lateral-stick pulse in airplane 

mode in addition to calculated responses to lateral-stick and pedal inputs in hover. 

Longitudinal trim analysis of CTR, VDTR, and VSTR was conducted by calculating the required 

power, pitch attitude, and wing lift coefficient vs. airspeed on curves of constant mast angle, flap 

deflection, and wingtip incidence. These constant curves were then examined and nominal 

schedules in level flight at constant airspeed (zero longitudinal acceleration) were defined, subject 

to constraints on power, wing lift coefficient, and pitch attitude in transition mode. The schedules 

were developed with the objectives of operating the aircraft near minimum power and near level 

pitch attitude without exceeding a wing lift coefficient of 80% (CL)MAX. Note that the maximum 

available wing lift coefficient varies with mast angle and flap deflection. Transition to airplane 
41 

mode was scheduled to occur just above an airspeed at which wing lift alone was sufficient for 

level flight, and flap retraction was also scheduled at a minimum safe airspeed. Table 8 shows a 

relative ranking of handling qualities for the three aircraft. 

TABLE 8. HANDLING-QUALITIES SUMMARY 

4.1.6 Mass Properties 

The weight-prediction methodology employed for the three tiltrotor concepts was based on a 

combination of statistical, semianalytical, and analytical methods. These methods were 

mathematically modeled in Sikorsky's Rotorcraft Design Model (RDM) computer program to 

permit weight trending as a function of primary aircraft design parameters, and to facilitate 

evaluation of the numerous point and variant designs and impact analyses required by the study. 

The parametric weight estimates were generally correlated to V-22 weight data, incorporating 

technology where appropriate. Additional factors were applied to the rotor system for the VDTR 

concept, and to the drive system for the VSTR concept to account for their unique design features. 

Figure 34 through Figure 41 show relative component weights; that is, the baseline CTR is set to 

the nominal value of 100. 

100

88 94

0

25

50

75

100

125

150

CTR VDTR VSTR

Wing Group Weight

Aircaft Comparison

Wing Group Weight

100

133

98

0

25

50

75

100

125

150

CTR VDTR VSTR

Rotor System Group Weight

Aircaft Comparison

Rotor System Group Weight

Figure 34. Wing-weight comparison. Figure 35. Rotor-weight comparison.
42 

100 101 100

0

25

50

75

100

125

150

CTR VDTR VSTR

Body Group Weight

Aircaft Comparison

Body Group Weight

100 101 100

0

25

50

75

100

125

150

CTR VDTR VSTR

Fuel System Group Weight

Aircaft Comparison

Fuel System Group Weight

Figure 36. Body-weight comparison. Figure 37. Fuel-system-weight comparison.

100 97 101

0

25

50

75

100

125

150

CTR VDTR VSTR

Engine Installation Weight

Aircaft Comparison

Engine Installation Weight

100 105 107

0

25

50

75

100

125

150

CTR VDTR VSTR

Drive System Group Weight

Aircaft Comparison

Drive System Group Weight

Figure 38. Engine-weight comparison. Figure 39. Transmission-weight comparison.
43 

100 95

104

0

25

50

75

100

125

150

CTR VDTR VSTR

Flight Controls Group Weight

Aircaft Comparison

Flight Controls Group Weight

100

116

102

0

25

50

75

100

125

150

CTR VDTR VSTR

Engine/XMSN/Rotor Weight

Aircaft Comparison

Engine/XMSN/Rotor Weight

Figure 40. Flight-controls comparison. Figure 41. Engine/transmission (XMSN)/ 

rotor-weight comparison.

4.2 Aircraft Performance 

Upon completion of the design synthesis for each aircraft, the performance of each aircraft was 

compared to the design requirements and to the relative performance of the other designs. The 

relative strengths and weaknesses of the designs were quantified as an indication of the merits of 

the underlying technology for each concept. The attributes that showed the greatest differences 

between the various technology systems are described in detail in the following sections. 

4.2.1 Downwash/Outwash Characteristics 

The expanded rotor diameter of the VDTR greatly benefits hover and low-speed operations. 

Increasing the hover diameter effectively reduces the power loading of the aircraft, allowing the 

hover-gross-weight capability of the VDTR to exceed the capabilities of the CTR and VSTR even 

though the VDTR has close to 5% less power installed on the aircraft. The large rotor diameter 

decreases the VDTR disk loading in hover, thereby decreasing the outwash velocity experienced 

by personnel and structures that are in the near vicinity of the hovering aircraft. Figure 42 provides 

a comparison of the outwash velocity for the VDTR, CTR, and VSTR. Figure 43 and Figure 44 

show the relative maximum outwash force and overturning moment exerted on a standing person. 

4.2.2 Hover Performance 

Figure 45 and Figure 46 show the International Standard Atmosphere (ISA) and constant 95°F 

weight-altitude-temperature (WAT) capabilities of the three tiltrotor aircraft. Note that while the 

VDTR has the lowest rotor figure of merit (because of its exposed inboard spar section), it has the 

greatest hover efficiency, evident in its expanded WAT envelope, because while rotor figure of 

merit is purely a measure of rotor efficiency, aircraft power loading (as manifested in disk loading) 

is a measure of overall hover efficiency. Hover efficiency is more strongly driven by disk loading 

than by rotor figure of merit. The higher disk loading of the VSTR is mitigated by its improved 
44 

rotor figure of merit mainly because VSTR benefits from the fact that the optimum rotor twist 

needed for efficient cruise performance at reduced rpm is closer to the optimum rotor twist in 

hover at 100% rpm. 

0

1

2

3

4

5

6

7

8

9

10

20 30 40 50 60 70 80 90 100 110 120

Height from Ground (ft)

Relative Outwash Velocity (CTR Max. = 100)

Aircraft Comparison

Outwash Height vs. Velocity 

CTR

VDTR

VSTR

Figure 42. Tiltrotor-outwash comparison.

100

92

112

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Outwash Force

Aircraft Comparison

Maximum Outwash Force

100 94

111

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Outwash Moment

Aircraft Comparison

Maximum Outwash Overturning Moment

Figure 43. Outwash-force comparison. Figure 44. Outwash-moment comparison.

 
45 

0

2

4

6

8

10

12

14

16

18

20

0 10 20 30 40 50 60 70 80 90 100 110

Altitude (1,000 ft)

Relative Useful Load (Max. CTR = 100)

HOGE Ceiling (0 ft/min VROC) | ISA Atmospheric Conditions

CTR

VDTR

VSTR

Figure 45. ISA WAT-capability comparison.

0

1

2

3

4

5

6

7

8

9

10

0 10 20 30 40 50 60 70 80 90 100 110

Altitude (1,000 ft)

Relative Useful Load (Max. CTR = 100)

HOGE Ceiling (0 ft/min VROC) | 95F Atmospheric Conditions

CTR

VDTR

VSTR

Figure 46. 95 ̊F WAT-capability comparison.

 

Aircraft Comparison 

Aircraft Comparison 
46 

4.2.3 Short-Takeoff Capability 

Figure 47 and Figure 48 show the short-takeoff (STO) ground-roll distance and total distance to 

clear an obstacle as a function of gross weight at SL/hot conditions. Figure 49 and Figure 50 show 

the STO ground-roll distance and total distance as a function of gross weight at high altitude/hot 

day (Hi/hot) conditions. The ordinates of Figure 47 through Figure 50 represent the aircraft weight 

at some nominal value. The maximum aircraft weight is 100. Lower weights are percentages of the 

maximum weight. For example, a nominal weight of 95 is a gross weight 5% less than the 

maximum weight. The abscissas of Figure 47 through Figure 50 represent the STO ground roll and 

distance and total distance to clear an obstacle, normalized to the CTR, respectively. 

Figure 51 and Figure 52 show the STO ground-roll distance and total distance at maximum takeoff 

gross weight (MTOW) at SL/hot conditions, normalized to the CTR, respectively. Figure 53 and 

Figure 54 show the STO ground-roll distance and total distance at MTOW at Hi/hot conditions, 

normalized to the CTR, respectively. 

Generally speaking, STO performance increases with larger wings and larger rotors. The CTR has 

the largest wing and a nominally large rotor. The VDTR has the largest rotor and nominally large 

wing. The VSTR has the smallest rotor and the smallest wing. Thus, the CTR performs nominally 

during STO operations, the VDTR has the greatest performance, and the VSTR performs worst. 

This phenomenon is clearly illustrated in Figure 47 through Figure 50, where for a given weight 

the VDTR has significantly less ground roll and less distance required to clear an obstacle. 

75

80

85

90

95

100

105

0 10 20 30 40 50 60 70 80 90 100 110

Nominal Gross Weight

Relative Ground Roll Distance (Max. CTR = 100)

Aircraft Comparison

GW vs. Ground Roll | SL/Hot

CTR

VDTR

VSTR

Figure 47. STO ground-roll-capability comparison | SL/hot.
47 

75

80

85

90

95

100

105

0 10 20 30 40 50 60 70 80 90 100 110

Nominal Gross Weight

Relative Total Dist. to Clear Obstacle (Max. CTR = 100)

Aircraft Comparison

GW vs. Total Distance to Clear Obstacle | SL/Hot

CTR

VDTR

VSTR

Figure 48. STO total-distance-capability comparison | SL/hot.

75

80

85

90

95

100

105

0 10 20 30 40 50 60 70 80 90 100 110

Nominal Gross Weight

Relative Ground Roll Distance (Max. CTR = 100)

Aircraft Comparison

GW vs. Ground Roll | Hi/Hot

CTR

VDTR

VSTR

Figure 49. STO ground-roll-capability comparison | Hi/hot.
48 

75

80

85

90

95

100

105

0 10 20 30 40 50 60 70 80 90 100 110

Nominal Gross Weight

Relative Total Dist. to Clear Obstacle (Max. CTR = 100)

Aircraft Comparison

GW vs. Total Distance to Clear Obstacle | Hi/Hot

CTR

VDTR

VSTR

Figure 50. STO total-distance-capability comparison | Hi/hot.

100

84

121

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Distance

Aircraft Comparison

Ground Roll Distance | MTOW | SL/Hot

100

94

111

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Distance

Aircraft Comparison

Distance to Clear Obstacle | MTOW | SL/Hot

Figure 51. STO ground roll | MTOW | SL/hot. Figure 52. STO total distance | MTOW | SL/hot.
49 

100 99

113

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Distance

Aircraft Comparison

Ground Roll Distance | MTOW | Hi/Hot

100 97

112

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Distance

Aircraft Comparison

Distance to Clear Obstacle | MTOW | Hi/Hot

Figure 53. STO ground roll | MTOW | Hi/hot. Figure 54. STO total distance | MTOW | Hi/hot.

4.2.4 Speed Capability 

Figure 55 through Figure 60 show the velocity for best endurance (VBE), velocity for 99% best 

specific range (VBR), and velocity for engine maximum continuous power or transmission torque 

(VMCP) speed comparison of the three tiltrotors, as a function of altitude under ISA and constant 

95 ̊F conditions. Figures 61 through 66 show the VBE, VBR, and VMCP speed capability of the three 

tiltrotor aircraft at 30k/ISA and 4k/95 conditions. 

The VDTR offers a broader spectrum of speed performance compared to the CTR and VSTR. Its 

smaller cruise rotor diameter is more efficient for high-speed flight. The VDTR transmission 

torque capability is retained because its rotor rpm is not reduced as much as that for the VSTR. 

The VSTR has two modes in which it can operate: “efficiency mode,” where the rotor is slowed 

via gearbox reduction and the engine operates at reduced rpm; and “performance mode,” where the 

rotor is slowed via engine rpm reduction only, equivalent to the rpm reduction used on the CTR 

and VDTR. In “efficiency mode,” the VSTR operates at high rotor propulsive efficiency, but 

because its transmission power available is significantly reduced by the rpm reduction, its speed 

performance is limited. In “performance mode,” the VSTR speed capability improves beyond that 

of the CTR, yet is still less than the VDTR. The performance gain of the VSTR in “performance 

mode” is at the expense of a significant propulsive efficiency penalty, as its rotor is not optimized 

for both high-speed and -rotor rpm. The VDTR, being designed for both high-speed and higherrotor rpm, enables simultaneous high-speed and high-efficiency cruise flight. 
50 

0

2

4

6

8

10

12

14

16

18

20

22

24

26

28

30

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for Best Endurance (VBE) | ISA Conditions

CTR

VDTR

VSTR

Figure 55. VBE capability comparison | ISA. 

0

1

2

3

4

5

6

7

8

9

10

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for Best Endurance (VBE) | 95°F Conditions

CTR

VDTR

VSTR

Figure 56. VBE capability comparison | 95°F. 
51 

0

2

4

6

8

10

12

14

16

18

20

22

24

26

28

30

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for 99% Best Specific Range (VBR) | ISA Conditions

CTR

VDTR

VSTR

Figure 57. VBR capability comparison | ISA. 

0

1

2

3

4

5

6

7

8

9

10

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for 99% Best Specific Range (VBR) | 95°F Conditions

CTR

VDTR

VSTR

Figure 58. VBR capability comparison | 95°F. 
52 

0

2

4

6

8

10

12

14

16

18

20

22

24

26

28

30

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for 100% Engine or XSMN MCP (VMCP) | ISA Conditions

CTR

VDTR

VSTR 60%NR

VSTR 80%NR

Figure 59. VMCP capability comparison | ISA. 

0

1

2

3

4

5

6

7

8

9

10

60 65 70 75 80 85 90 95 100 105 110

Altitude (1,000 ft)

Relative Velocity (Max. CTR = 100)

Aircraft Comparison

Velocity for 100% Engine or XMSN MCP (VMCP) | 95°F Conditions

CTR

VDTR

VSTR 60%NR

VSTR 80%NR

Figure 60. VMCP capability comparison | 95°F. 
53 

100 106 102

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Aircraft Speed

Aircraft Comparison

VBE | 30k/ISA

100

107 101

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Aircraft Speed

Aircraft Comparison

Relative VBE | 4k/95

Figure 61. VBE | 30k/ISA | capability 

comparison. 

Figure 62. VBE | 4k/95 | capability 

comparison. 

100

107 103

0

25

50

75

100

125

150

CTR VDTR VSTR

Relaive Aircraft Speed

Aircraft Comparison

Relative VBR | 30k/ISA

100 106 105

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Aircraft Speed

Aircraft Comparison

Relative VBR | 4k/95

Figure 63. VBR | 30k/ISA | capability 

comparison. 

Figure 64. VBR | 4k/95 | capability 

comparison. 
54 

100 101 99

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Aircraft Speed

Aircraft Comparison

Relative VMCP | 30k/ISA

100 101

89

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Aircraft Speed

Aircraft Comparison

Relative VMCP | 4k/95

Figure 65. VMCP | 30k/ISA | capability 

comparison. 

Figure 66. VMCP | 4k/95 | capability 

comparison. 

4.2.5 Mission-Performance Capability 

All three aircraft are capable of executing heavy-lift, long-range transport missions required of the 

LVLT mission set. While the VDTR is the heaviest of the three solutions, it also offers the 

potential for the highest cruise speed with improved efficiency over the CTR. During high-speed 

transport missions, the VDTR and VSTR have improved range capability beyond the CTR due to 

their improved cruise efficiency. During low-speed external load or rotary-wing air-to-air refueling 

missions, the VDTR and VSTR (which have smaller wings) suffer reduced range capability with 

respect to the CTR. Figure 67 through Figure 73 show several representative mission profiles for 

the LVLT mission set. In the figures, the payload-range/radius performance of the three aircraft is 

presented relative to the mission required payload and range/radius (which are normalized to 100). 
55 

0

20

40

60

80

100

120

140

160

180

200

0 50 100 150 200 250 300 350 400 450 500

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | Hi/Hot ROA Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 67. Hi/hot ROA mission | payload-radius comparison.

0

20

40

60

80

100

120

140

160

180

200

0 25 50 75 100 125 150 175 200 225 250

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | SL/Hot ROA Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 68. SL/hot ROA mission | payload-radius comparison.
56 

0

20

40

60

80

100

120

140

160

180

200

0 50 100 150 200 250 300 350 400 450 500

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | External Load Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 69. External-load mission | payload-radius comparison.

0

50

100

150

200

250

300

350

400

0 20 40 60 80 100 120 140 160 180 200

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | Long Range Insertion Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 70. Long-range insertion mission | payload-radius comparison.
57 

0

20

40

60

80

100

120

140

160

180

200

0 50 100 150 200 250 300 350 400

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | Low-Speed High Endurance Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 71. Low-speed high-endurance mission | payload-radius comparison.

0

20

40

60

80

100

120

140

160

180

200

0 50 100 150 200 250 300 350 400

Relative Payload (Req. = 100)

Relative Radius (Req. = 100)

Aircraft Comparison

Payload vs. Radius | High-Speed High Endurance Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 72. High-speed high-endurance mission | payload-radius comparison.
58 

0

20

40

60

80

100

120

140

160

180

200

0 20 40 60 80 100 120 140 160 180 200

Relative Payload (Req. = 100)

Relative Range (Req. = 100)

Aircraft Comparison

Payload vs. Radius | Long-Range High-Payload Mission

CTR

VDTR

VSTR

Mission

Requirement

Max. Internal Payload Limit

Figure 73. Long-range high-payload mission | payload-range comparison. 

4.3 Configuration Evaluation 

Ten disparate attributes were chosen to evaluate the three aircraft, presented in terms of a pure 1st, 

2nd, or 3rd place ranking in Table 9. This evaluation, however, does not consider the value 

weighting associated with each attribute. Using requirements-driven weightings representative for 

the LVLT mission and usage spectrum, a resultant score can be obtained that depicts the relative 

ranking of the three concepts based on the operator’s needs. 

TABLE 9. ATTRIBUTE RANKING 
59 

The ultimate measure of the relative value of an aircraft was provided by the United Technologies 

Corporation (UTC)-proprietary Value-Based Gage (VBG) methodology (ref. 25). The VBG 

process was used to develop an objective and auditable metric or scale, which was used to 

compare the aircraft. The VBG process is based upon Quality Function Deployment (QFD) or 

“House of Quality” analysis to provide a link between the engineering characteristics (e.g., weight 

or speed) of a concept and the overall needs of the decision makers. A key feature of the gauge is 

that soft metrics (e.g., risk) can be included with hard metrics, such as weight, power consumption, 

and drag. A response-surface-approximation methodology was used to express the VBG score for 

each concept as a function of the attribute weighting factors, as depicted in Figure 74. 

10 Attributes Weighted from 0% to 100% 191 Points Total Aircraft Score

 

Figure 74. Response-surface methodology used for configuration evaluation.

The weighting factors for each attribute were systematically varied to attain a matrix of VBG sets 

valid for a variety of weighting schemes. The weighting factors were permeated to yield 191 VBG 

ranked sets. The sets included: a) all attributes weighted equally; and, b) each attribute weighted 

from 5% to 100% in steps of 5% with the other attributes equally ranked (for example, if WE was 

weighted at 10%, then all other attributes would be weighted at 9%; and, if WE was weighted at 

70%, then all other attributes would be weighted at 3%). A linear multivariate regression, shown in 

equation (1), was found to give the closest correlation to the dataset. The regression coefficients 

are given in Table 10. 

VBG = α + β1x1 + β2x2 + β3x3 + β4x4 + β5x5 + β6x6 + β7x7 + β8x8 + β9x9 + β10x10 (1)

TABLE 10. VBG RESPONSE SURFACE LINEAR MULTIVARIATE REGRESSION COEFFICIENTS

Coefficient Attribute CTR VDTR VSTR 

α --- 0.8507442427 -0.8769132653 2.2755102041 

β1 WE x1 0.0014925576 0.0087691327 -0.0129595119 

β2 MRP x2 -0.0070355404 0.0187691327 -0.0227551020 

β3 VOUTWASH x3 -0.0051196304 0.0187691327 -0.0227551020 

β4 xSTO x4 -0.0003488022 0.0187691327 -0.0227551020 

β5 VBR x5 -0.0085074424 0.0187691327 -0.0166834356 

β6 ηCRUISE x6 -0.0085074424 0.0087691327 -0.0227551020 

β7 ηHOVER x7 -0.0085074424 0.0187691327 -0.0164734857 

β8 Productivity x8 -0.0085074424 0.0130143890 -0.0127551020 

β9 Footprint x9 -0.0085074424 0.0173538275 -0.0127551020 

β10 Risk x10 -0.0025510204 0.0087691327 -0.0127551020 
60 

Figure 75 shows the average value of each aircraft (relative to the CTR) for the 191 scenarios of 

VBG weighting schemes, equivalent to weighting all attributes equally. As shown in Figure 75, if 

the CTR has a baseline value of 100, then the VDTR is 26% more valuable and the VSTR is 18% 

more valuable. Figure 76 through 78 show the likelihood of an aircraft ranking 1st, 2nd, or 3rd, 

respectively, for any of the 191 VBG weighting scenarios. 

100

126

118

0

25

50

75

100

125

150

CTR VDTR VSTR

Relative Value (CTR=100)

Aircraft Comparison

Average Value

2%

68%

31%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

CTR VDTR VSTR

Probability

Aircraft Comparison

Probability of Aircraft Ranking 1st

Figure 75. VBG results | average value. Figure 76. VBG results | 1st place likelihood.

30%

18%

51%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

CTR VDTR VSTR

Probability

Aircraft Comparison

Probability of Aircraft Ranking 2nd

69%

14% 17%

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

100%

CTR VDTR VSTR

Probability

Aircraft Comparison

Probability of Aircraft Ranking 3rd

Figure 77. VBG results | 2nd place likelihood. Figure 78. VBG results | 3rd place likelihood.
61 

5 Alternate Designs 

In addition to the three “clean-sheet” tiltrotor aircraft, two alternate-configuration aircraft were 

synthesized: a shipboard-suitable VDTR concept, arranged in a canard configuration with 

shortened fuselage and tandem cockpit, and a fixed-wing variant. These aircraft offer increased 

shipboard capability and potential cost savings, respectively, through their novel design 

characteristics. 

5.1 Fixed-Wing Variant Description 

The LVLT Fixed-Wing Variant (FWV) was synthesized using the CTR as a launching point. As 

the missions extend to the longer ranges typical of fixed-wing aircraft, the question arises as to the 

viability of the LVLT aircraft solely as a VTOL aircraft. The design team examined the paradigm 

of a mixed fleet of aircraft filling the LVLT role while simultaneously filling the role of the 

midsize fixed-wing transports. The hypothesis is that there would be a significant programmatic 

and financial benefit to producing a fixed-wing transport aircraft based on the VTOL LVLT 

fuselage. 

The fuselage of the fixed-wing LVLT would be identical to that of the VTOL aircraft, built on the 

same production line. The landing gear would be common between the fixed-wing and VTOL 

configuration since it is capable of roll-on landings at high speeds. As a result, the LVLT-FWV 

would be capable of short-takeoff-and-landing (STOL) operations from unpaved airfields. 

The fuselage-commonality concept is a key to achieving programmatic cost savings. An integrated 

production line could use common tooling and facilities for both the VTOL and fixed-wing LVLT 

variants. The cockpit and cabin systems could all be common, including environmental systems, 

cargo systems, wiring, and plumbing. The pilot interface with the aircraft (side stick versus cyclic 

stick) and some panel layout may be unique. The hydraulic arrangement of the VTOL LVLT, with 

the hydraulics contained only in the nacelle areas, could allow the fixed-wing variant to use a 

unique hydraulic system, contained only in the wing system that is independent of the fuselage. 

Electro-hydraulic and electro-mechanical actuators are used throughout the fuselage systems, 

further enhancing the production commonality. 

The wing of the fixed-wing aircraft would be completely unique but would be mounted using the 

same structure as that for the VTOL wing, requiring that only the nonstructural wing fairings be 

custom for the fuselage. The wing of the fixed-wing aircraft would be much thinner than the thick 

wing required for the VTOL aircraft and would support four wing-mounted turboprop nacelles. 

The thin wing would allow for efficient high-speed cruise flight, at speeds greater than those that a 

VTOL aircraft can achieve. 

One aspect of the wing for the fixed-wing variant that will require further study is the weight and 

balance of the fixed-wing aircraft. The VTOL configuration has its own unique balance constraints 

because the center of gravity (cg) moves during transition. In the cruise mode, the cg is well 

forward, but as the nacelles rotate during transition the cg moves aft, eventually lining up near the 

tilt axis and allowing the aircraft to hover in equilibrium without any rotor head moment or 
62 

excessive flapping. The fixed-wing variant would only be constrained by typical fixed-wing 

weight and balance limits, but would ideally be aft-swept with forward-mounted, under-slung, 

turboprop nacelles. This arrangement creates a unique balance situation and would have to be 

investigated fully to allow for the optimal positioning of the wing and engines. 

One of the greatest advantages of the fixed-wing variant is the use of engines that are common to 

the VTOL platform. History has shown that common cores can be used for turbo-shaft and 

turboprop engines. The VTOL LVLT will require new turboshaft engines, and it is foreseeable that 

these engines could be designed early in the design process for commonality or even complete 

interchangeability of the large engine sections. Prop reduction gearboxes would be required, but 

could likely be incorporated in the engine development with no detrimental performance or weight 

impact. 

The engine development will be one of the largest single-cost drivers for the LVLT program, and 

distributing this development cost across additional aircraft would appreciably reduce the cost per 

aircraft. A first-order cost analysis for the LVLT aircraft shows a cost reduction per aircraft of 4% 

to 5% for the VTOL aircraft if an equal number of additional fixed-wing fuselages and engine ship 

sets are included. This analysis is a simplified approximation, but it shows the potential benefit of 

commonality between the VTOL and fixed-wing variants. 

5.2 Shipboard-Suitable VDTR Description 

The baseline LVLT aircraft does not incorporate folding, so a shipboard suitability trade study was 

conducted. The focus of this effort was to maximize shipboard suitability, and to study the 

shipboard operation requirements, while providing an aircraft design that accentuates the strengths 

of the design with the nuances of the requirements and unique operating environment. The 

shipboard-suitable aircraft is based on VDTR and is denoted VDTR-S. 

The VDTR-S aircraft minimizes the overall aircraft length by eliminating the conventional 

empennage and replacing it with a canard for pitch control. The aerodynamics of the canard 

configuration allow the aircraft mean aerodynamic center to be much further forward than on a 

traditional wing and tail configuration. This setup enables the tilt axis to be positioned more 

favorably at the wing tips, reducing the complexity of the cross-shaft gear meshes and shortening 

the overall gearbox drive system. 

To improve the pilot visibility during shipboard operations, the cockpit system has been reconfigured to a tandem arrangement. This tandem fighter aircraft style of canopy allows the pilot to 

see the entire aircraft behind him, including the nacelles and rotors. The pilot also has better visibility around the front of the aircraft and improved downward visibility. However, this canopy 

arrangement is likely not the optimal weight solution. Another drawback of this cockpit configuration is that the crew cannot get out of their seats and move around during long missions. While 

early 1950’s large aircraft such as the B47 had tandem cockpits, modern transport-pilot crews have 

grown accustomed to working in a side-by-side collaborative environment. Fighter-pilot crews, 

however, have mastered the tandem cockpit arrangement. The pilot protocol for each arrangement 
63 

is unique, and transport-pilot crews would have to adapt to the fighter-pilot style for this cockpit to 

work effectively. 

The cockpit arrangement, although unconventional, also enables the aircraft to be loaded and 

unloaded using cargo-ramp doors at both ends of the aircraft, even with the rotors turning and the 

engines running. For tight shipboard operations where space is at a premium, the aircraft can be 

positioned for ease of loading using either ramp, opening up the loading possibilities for the ship 

crew and load master. 

The wings of the VDTR-S fold for stowage on the ships; they fold along the sides of the fuselage 

and rest vertically nestled under the nacelles with the engines in front of the cockpit. 
64 

6 Conclusions 

The technical community has identified rotor efficiency as a critical enabling technology for large 

vertical-lift transport (LVLT) rotorcraft. The size and performance of LVLT aircraft will be far 

beyond current aircraft capabilities, enabling a transformational change in cargo transport effectiveness. Several viable alternatives were identified as candidate technologies capable of meeting 

the rotor performance requirements of the LVLT mission, including rotors with widely variable tip 

speeds and rotors with variable diameters. This study quantified the capabilities and efficiencies of 

these technologies compared to a conventional tiltrotor reference design. Ultimately, three aircraft 

concepts were synthesized and the attributes of these aircraft were compared as a measurement of 

the effectiveness of the underlying rotor technology. 

Sikorsky was selected to conduct this analysis largely because of the company’s previous variablediameter-tiltrotor (VDTR) development efforts. These efforts spanned approximately 10 years 

during the 1990s. The VDTR development efforts included preliminary and detailed design of both 

the aircraft and rotor-system components as well as proof-of-concept testing of the system-level 

benefits and risk-reduction testing of the critical sub-system components. A summary of this previous work was discussed along with the relevance of these efforts to the current LVLT aircraft 

designs. 

The three aircraft were optimized and designed around a common set of requirements with the 

intent of exposing the nuances that differentiate the configurations. Where common designs could 

be applied to all of the configurations, a single design was applied to all three configurations. Subsystems such as the rotor, drive system, and wing were designed in greater detail because these 

areas had the greatest potential to affect the capabilities of the three designs. Each of these systems 

was tailored to the specific needs of the rotor system applied to a given design, affecting the 

overall size and weight of the sub-system and the overall aircraft. 

The design of the VDTR rotor was noteworthy in that the analysis brought forth an interesting 

nuance related to the VDTR synthesis for the LVLT requirements. Previous VDTR work 

emphasized the maximum rotor retraction possible relative to hover, with most designs employing 

a 34%-diameter retraction. Such large retractions, however, penalize hover performance two-fold: 

the hover twist is not optimal because of the requirement of linear twist on the retractable portion 

of the blade; and the exposed area of the elliptical inboard blade spar is greater, lowering figure of 

merit (FM). Because the LVLT mission required austere hover conditions at high mission weights, 

cruise performance was balanced by hover performance, leading to lower disk loading and a lower 

retraction ratio of 23%. 

The resulting designs for the conventional tiltrotor (CTR), VDTR, and variable-speed tiltrotor 

(VSTR) were all relatively similar in size, weight, and installed power—not necessarily a 

surprising result given that the aircraft were optimized using the same ground rules and 

requirements set. Generally, the VDTR was the heaviest of the three concepts and had the least 

installed power. The VSTR was lighter than the VDTR but had the most installed power. The CTR 

was the lightest aircraft of the three, and had more power installed than the VDTR but less power 

installed than the VSTR. 
65 

The similar size, weight, and power of the three aircraft concepts lead to similar performance and 

handling-qualities capabilities, as well as similar levels of risk and cost. Several critical 

performance attributes were evaluated. These attributes were chosen because they showed the 

greatest differentiation between the three aircraft designs. The downwash, hover, and short-takeoff 

(STO) capability analyses favored the aircraft with the largest rotor, the VDTR, since these 

attributes are heavily influenced by the power loading of the aircraft. 

The VDTR offered a broader spectrum of speed performance compared to the CTR and VSTR. Its 

smaller cruise-rotor diameter is more efficient for high-speed flight. The transmission torque 

capability of the VDTR is retained because its rotor revolutions per minute (rpm) is not reduced as 

much as with the VSTR. The VSTR has two modes in which it can operate: “efficiency mode,” 

where the rotor is slowed via gearbox reduction and the engine operates at reduced rpm; and 

“performance mode,” where the rotor is slowed via engine rpm reduction only, equivalent to the 

rpm reduction used on the CTR and VDTR. In “efficiency mode,” the VSTR operates at high 

rotor-propulsive efficiency, but because its available transmission power is significantly reduced 

by the rpm reduction, its speed performance is severely limited. In “performance mode,” the speed 

capability of the VSTR improves beyond that of the CTR, yet is still less than that of the VDTR. 

The performance gain of the VSTR in “performance mode” is at the expense of a significant 

propulsive-efficiency penalty, as its rotor is not optimized for both high-speed and high-rotor rpm. 

The VDTR, being designed for both high-speed and higher-rotor rpm, enables simultaneous highspeed and high-efficiency cruise flight. 

Similar trends were observed in the payload-range performance of the aircraft. Missions involving 

arduous hover or low-speed conditions, including external lift, favored the VDTR because of the 

large rotor-disk area. Missions that favored high cruise efficiency also favored the VDTR because 

of its small, efficient propellers. The VSTR also performed the long-range cruise missions well 

because of its high efficiency, but it required greater installed power to perform the hover and lowspeed missions. 

Analysis of the aircraft-handling qualities again yielded similar results. Maneuvers that were 

similar to those executed by a helicopter, i.e., low-speed banked turns, heading changes in hover, 

and low-speed climbs, all favored either the VDTR because of its larger rotor or the CTR because 

of its larger wing area. Finally, additional analysis was conducted to quantify the risk and cost of 

these aircraft configurations, though this part of the study is not discussed in the report. These 

analyses indicate that the VDTR faces the largest development risk, followed by the VSTR. 

Development and life-cycle costs for these aircraft are also similar. 

The comparison of the three aircraft, CTR, VDTR, and VSTR, indicates that the VDTR and VSTR 

are enabling technologies for the LVLT mission requirements, and that they can both be developed 

at reasonable levels of technical risk and development cost. Furthermore, the VDTR and VSTR 

developments are only incrementally more expensive than a similar CTR development, and add 

only incrementally greater risk. The performance benefits of the VDTR and VSTR make them 

worth the additional investment. 
66 

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