Designs and Technology Requirements for Civil Heavy Lift Rotorcraft

Designs and Technology Requirements for Civil Heavy Lift Rotorcraft

Wayne Johnson, Gloria K. Yamauchi, Michael E. Watts National Aeronautics and Space Administration Ames Research Center, Moffett Field, CA and Langley Research Center, Hampton, VA Ab stract The NASA Heavy Lift Rotorcraft Systems Investigation examined in depth several rotorcraft configurations for large civil transport, designed to meet the technology goals of the NASA Vehicle Systems Program. The investigation identified the Large Civ il Tiltrotor as the configuration with the best potential to meet the technology goals. The design presented was economically competitive, with the potential for substantial impact on the air transportation system. The keys to achieving a competitive aircr aft were low drag airframe and low disk loading rotors; structural weight reduction, for both airframe and rotors; drive system weight reduction; improved engine efficiency; low maintenance design; and manufacturing cost comparable to fixed - wing aircraft. Risk reduction plans were developed to provide the strategic direction to support a heavy - lift rotorcraft development. The following high risk areas were identified for heavy lift rotorcraft: high torque, light weight drive system; high performance, struct urally efficient rotor/wing system; low noise aircraft; and super - integrated vehicle management system.

flights in 2017 as candidates for RIA. By removing 10% of .

INTRODUCTION the flights from the primary runways, Ref. 1 projects 79% The Rotorcraft (RC) Sector was established in January 2004 less delay in 2 017, roughly equivalent to a cost avoidance of as one of six vehicle sectors within the Vehicle Systems $181B per year. Alternatively, replacing the removed short - Program (VSP) of th e NASA Aeronautics Research Mission haul flights with medium - and long - haul flights would Directorate. The principal aim of the RC Sector is to increase system capacity by 152 billion revenue passenger improve public mobility and access to air transportation.

miles, which translates into added ser vices to the public in The technology goals of the Sector originated from industry addition to substantial revenue for the airlines. Reference 3 studies and workshops during 2001 - 2004 that focu sed on a describes three RIA configurations analyzed by the new class of vehicles known as Runway Independent rotorcraft industry: the quad tiltrotor (Bell Helicopter), the Aircraft (RIA). References 1 - 2 showed that RIA can relieve reverse velocity rotor concept (Sikorsky), and the tiltro tor runway and terminal area congestion by replacing small (Boeing). The studies identified the benefits of advanced aircraft and short - haul flights that use primary runways. The technology and the resulting effects on operating cost. In primary runways would then be used exclusively for larger summary, Refs. 1 - 3 provide justification for the aircraft and medium/long - haul flights. RIA would operate overwhelming positive impact that RIA can have on the from stub runways and/or helicopter landing pads. This national air space.

operational concept would increase the capacity of the air Using the RI A studies as motivation, the RC Sector is transportation system. The increased capacity co uld then be focusing on enabling technology for a notional civil VTOL used to increase throughput or reduce delay throughout the transport capable of carrying 120 passengers at a cruise system. Reference 1 conservatively estimates 10.2% of speed of 350 knots at 30,000 ft altitude with a range of 1200 nm (without refueling). This heavy - li ft transport will be .

Presente d at the AHS Vertical Lift Aircraft Design "neighborly" quiet when operating near communities, Conference, San Francisco, California, January 18 – 20, economically competitive with a Boeing 737 aircraft, and 2006.

will exploit available airspace and ground space (excluding expositions on autorotation and one - engine - inoperative primary runways). Specific 15 - year technology goals for the requirements for heavy lift. Also under contract were notion al transport are shown in Table 1. These extreme Pennsylv ania State University (blade and wing structural mission and technology goals were established by the RC design, airfoil design), and University of Maryland and Sector to push the state - of - the - art in rotorcraft technology. Georgia Institute of Technology (assessments of slowed - For comparison, the Mi - 26, the largest helicopter in the rotor compound configurations, including reaction drive).

world today, has a maximum speed of 160 knots with a An independent review group comprised of five non - service ceiling of approximately 15,000 ft and a range of government senior rotorcraft technologists with extensive 435 nm. The NASA Heavy Lift Rotorcraft Systems design experience in the rotorcraft industry, U. S. Army, Investigation, the focus of this paper, is the first step toward and academia provided feedback on the process and content attaining the RC Sector goals. of the investigation.

The objective of the investigation was to select a heavy lift This paper presents the results of the NASA Heav y Lift rotorcraft system that has the best chance of meeting the Rotorcraft Systems Investigation. It describes the approach goals of Table 1 while being economically competitive. The used for developing the designs for the tiltrotor, tandem - first four goals of Table 1 were given highest priority. The rotor compound, and the advancing blade concept deliverables of the investigation were a candidate configurations. Completed designs are presented together configuration for a large civil VTOL transport, and a with trade studies to quan tify the impact of technology and description of the research and development required for examine alternate missions. The configurations are then risk reduction. A NASA - led team of rotorcraft technologists ranked in terms of ability to meet the RC Sector mission analyzed three notional vehicle configurations suggested by and goals. Finally, high risk areas for the selected the rotorcraft industry: a tiltrotor, a tandem - rotor compound, configuration are identified and plans to mitigate the r isks and an advancing blade concept configuration. These are presented.

configurations were deemed, as a first cut, to be technically promising. In contrast to the RIA configuration study of DESIGN APPROACH AND ANALYSIS TOOLS Ref. 3, the present investigation assesses all the candidate The approach taken was to design large VTOL transports configurations against the same RC Sector mission and that are economically competitive with today's regional jet technology goals and provides detailed analysis in multiple airliners, and meet the RC Sector mission and goals. The technology areas. In approximately 12 months, the team principal cost drivers are we ight and power. Advances in performed extensive engineering analysis including aircraft structural efficiency, aerodynamic efficiency, control design, perf ormance optimization, blade and rotor concepts, propulsion concepts, dynamics solutions, and aerodynamics, airframe aerodynamics, loads and stability prediction capability should allow substantial reductions in analysis, blade structural design, external noise, one - engine empty weight, power, and fuel. Low power is ensured by inoperative requirements, handling qualities, and cost lo w rotor disk loading and low aircraft drag. Light weight at drivers. The team was divided into subgroups representing large size requires advanced technology. The heavy lift aeromechanics, acoustics, propulsion, structures, handling rotorcraft designs required tasks covering aircraft design, qualities, and cost. This approach was highly successful in performance optimization, aerodynamics analysis (airfoil, attacking this complex design problem. Team members blade, airframe, rotor , aircraft), structural design (airframe, included Ames Research Center (primary responsibility for wing, blade), rotor loads and stability analysis, assessment developing co ncepts), Glenn Research Center (engine and of propulsion, noise, and handling qualities, one - engine propulsion), and Langley Research Center (acoustics and inoperative review, and cost estimation. The intent of the structures). The Advanced Design Team of the U. S. Army investigation team was to perform the se analysis tasks in as Aeroflight - dynamics Directorate assisted with system much detail and as much depth as possible during the 12 - design. The U. S. Army provided additional assis tance in month period, in order to inform and support the aeromechanics (RDECOM/ AFDD), engine and propulsion recommendations for risk reduction activities.

(ARL), structures and materials (ARL/VTD, AMCOM/ The code RC performed the sizing of the rotorcraft, and the AATD). Contracts were established with Bell Helicopter, comprehensive analys is CAMRAD II was used for Boeing, and Sikorsky Aircraft to provide feedback on the performance optimization, and loads and stability NASA designs and risk reduction p lans in addition to calculations. The sizing code incorporated significant conducting limited sizing, design, and analysis of some of weight savings (relative to current technology scaled to the concepts being investigated. Bell and Sikorsky prepared large size) as a result of structure, drive train, and engine techno logy. Cost models were developed, and used to An assessment of engine and drive train technology was estimate the purchase price and direct operating cost of the made in order to define and substantiate the sizing code heavy lift rotorcraft designs. The sizing code was used to models. The engine model represented what could be perform sensitivity analyses, first to optimize the aircraft obtained from (or required of) modern technology engines.

(variations including disk lo ading, tip speed, and number of Drive train concepts were developed for the heavy lift engines); and then to quantify the influence of advanced rotorcraft designs.

technology.

Blade structural loads calculations were used to design rotor The code RC (Ref. 4) was the principal rotorcraft sizing and blade sections; and the resulting blade structural and in ertial performance analysis tool for this investigation. RC was properties were used to repeat the loads calculations. This developed by the Advanced Design Team of the U. S. Army structural design required an assessment of advanced Aeroflightdynamics Directorate, RDECOM. Designer materials and application of innovative design and inputs to RC include design strategy (engine sizing, rotor optimization techniques, in order to achieve a low weight at sizing, etc.), rotorcraft parameters (drag coefficients, tail large size. A similar ap proach was used for the structural volume ratio, etc.), and requirements and constraints (take - design of the wing sections. The resulting wing structural off, pa yload, range, etc.). RC finds the aircraft that satisfies and inertial properties were used to develop NASTRAN the designer inputs, then produces the rotorcraft description, finite element models of the airframe. The NASTRAN and conducts the performance analysis. modes were used in CAMRAD II to calculate stability (parti cularly tiltrotor whirl flutter), linearized matrices for Technology in the sizing code is introduced in terms of handling qualities analysis, and vibration.

technology factors and performance model s. Weights (at the group weight level of detail) are estimated from statistical The handling qualities of the aircraft were assessed, and the equations. These equations are calibrated to current results used to guide the choice of configuration parameters technology level by comparing with existing aircraft. for the sizing code. Exposition s on autorotation and one - Technology factors are then applied to represent the impact engine inoperative requirements for heavy lift rotorcraft of advanc ed technology. In this approach, technology is a were developed independently by Bell Helicopter and change from the statistical equation, attributed to a new Sikorsky Aircraft, considering requirements and design configuration or concept, new materials, new design implications. One - engine inoperative requirements were methods, new operating procedures, etc. There are defined fo r use in the sizing code.

technology factors for blade and hub weight, vi bration The rotor performance model in the RC sizing code was treatment, drive system weight, and fuselage, wing, and tail calibrated using the performance calculated by CAMRAD weight. Technology also influences performance, in II, and the sizing task repeated. An estimate of the drag of particular rotor hover and cruise efficiency, hub drag, and the airframe was used to define the aerodynamic model for the engine weight and performance.

the sizing code and the comprehensive analysis. Based on CAMRAD II is an aeromechanical analysis o f helicopters aerodynamic environment calculations from CAMRAD II, and rotorcraft that incorporates a combination of advanced rotor blade airfoils were designed using the code MSES technologies, including multibody dynamics, nonlinear (Ref. 6). Airfoil decks were constructed for the new airfoils, finite elements, and rotorcraft aerodynamics (Ref. 5). The and used in the performance cal culations. The contours of trim task finds the equilibrium solution (constant or these airfoils were used in the blade structural design. A periodic) for a steady state operating condition, and similar approach was used for aerodynamic design of wing produces the solution for performance, loads, and vibration. airfoils. The three - dimensional Navier - Stokes analysis The flutter task linearizes the equations about the trim OVERFLOW - D was used to calculate the flow about the solution, and produces the stability results. The tiltro tor proprotor and pylon/nacelle. In addition, low aerodynamic model includes a wake analysis to calc ulate fidelity CFD calculations using the Rot3DC code (Ref. 7) the rotor nonuniform induced - velocities, using rigid, were performed of the entire tiltrotor flow field, including prescribed or free wake geometry. CAMRAD II has cruise drag and hover download calculations.

undergone extensive correlation with performance and Making use of the comprehensive analysis mo del, the loads measurements on helicopters, tiltrotors, and other aircraft noise was assessed using the CARMA system (Ref.

rotorcraft configurations. Complete ae roelastic models were 8), and the results used to guide the choice of configuration developed for each of the configurations considered in this parameters for the sizing code. An assessment was made of investigation.

the relative contributions of aircraft configuration parameters, rotor active control, and flight procedures by advanced technology. Note in particular the importance towards the acoustics goals. of maintenance costs.

The result of this process was three heavy lift rotorcraft For the same mission, a VTOL aircraft will have higher designs supported by substantial in - depth engineering gross weight and higher installed power than a CTOL analyses, and guidance and focus for the development of the aircraft. In addition, there are complexity factors in the risk reduction plans. VTOL model, including number of rotors and number of blades. Thus there is still a cost of VTOL capability in the cost model, even when the maintenance and flyaw ay price COST MODELS technology factors are used.

Cost models were developed for VTOL and CTOL aircraft, based on statistical information for current operations. The CONFIGURATIONS cost metrics considered were flyaway cost (purchase price, in 2005 US dollars) and direct operating cost plus interest, Three aircraft configurations were the primary subject of D OC+I (in 2005 US cents/ASM). The components of the Heavy Lift Rotorcraft Systems Investigation: DOC+I were maintenance (airframe, engine, rotor and 1) Large Civil Tiltrotor (LCTR) drive), flight crew, fuel and oil, depreciation, insurance, and 2) Large Civil Tandem Compound (LCTC) finance cost.

3) Large Advancing Blade C oncept (LABC) A principal source for the cost models was Ref. 9 and its These configurations were selected by industry as the most unpublished extension s. The parametric estimate of flyaway promising candidates for the civil mission. The cost was based on data for 120 helicopters and 2 tiltrotors, conventional two - rotor tiltrotor configuration was with the U. S. multi - engine turbine helicopters covering a considered, since a quad tiltrotor would not present as much weight range from the Bell 206L to the CH - 53E. The of a challenge in terms of rotor size. A low rotor speed was parametric equation gave flyaway cost from empt y weight used for the tiltrotor in cruise, to improve the proprotor and installed power ($/lb nearly just a function of W /P), E propulsive efficiency. The LCTC and LABC use edgewise and the number of rotors and number of blades. The rotors in cruise, hence the rotor rotation must be slowed as parametric estimate of maintenance cost was based on civil the flight speed increases, to keep the adva ncing tip Mach operations; the result was a function of weight empty and number reasonable. The LCTC is a slowed - rotor compound: installed power. Fli ght crew costs were proportional to it has a wing and auxiliary propulsion for cruise, so the block hours. Depreciation, insurance, and finance cost were rotors are operated in an unloaded condition. The LABC all proportional to flyaway cost.

uses stiff coaxial main rotors capable of carrying significant The CTOL cost model was based on the economics of U. S. roll mome nt, hence generating lift on the rotor advancing airline operations. side in forward flight. The LABC requires auxiliary propulsion at high speeds, but has no wing.

In order to compare VTOL and CTOL costs, the two c ost models were applied to a Boeing 737 - 700 at a stage length The slowed - rotor compound considered had shaft - driven of 500 miles. For the 737 in the VTOL cost model, the tandem main rotors. Single main rotor and coaxial main minimum complexity was used (one rotor and one blade), ro tors are alternate configurations. The number and and an installed power trend was used to get an equivalent arrangement of the main rotors affects performance through turboshaft power. The cost s are substantially higher with the rotor/rotor and rotor/wing interference; and affects the VTOL model. With these results it is possible to establish aircraft size because of antitorque and transmission layout cost technology factors: issues. An alternative to shaf t drive is a reaction drive configuration, typically using jets at the blade tips. The Maintenance tech factor = 0.9/9.8 = 0.092 reaction drive is used in hover; in cruise the rotor is Flyaway price tech factor = 48.0/83.6 = 0.57 operated in autorotation. With reaction drive the Insurance, depreciation, and finance co sts are driven by transmission weight is greatly reduced, but the rotor cruise flyaway price. Baseline cost estimates for the heavy lift pe rformance is compromised by the need for thick blades, rotorcraft designs were obtained using the above cost and the hover performance is poor because of high energy technology factors. A significant part of the differences losses entailed in delivering the air to the blade tips.

between VTOL and CTOL costs must be the very different A major objective of the Rotorcraft Sector programs is to operations th at produced the cost data used to develop the examine the potential of acti ve control as enabling models. The remaining differences in cost must be attacked technology for heavy lift, based on weight reduction and/or implies a disk loading on the order of W/A = 10 lb/ ft . The solution of dynamic or aerodynamic problems. In particular, actual disk loadings of the designs were determined based attention is being given to on - blade control, including on minimum aircraft weight, power, and cost. For this trailing edge flaps, leading edge droop, active twist, an d heavy lift rotorcraft investigation, the target airframe and 2/3 active flow control. The present investigations contributed wing drag was D/q = 1.6(W/1000) . This drag level is to identifying what problems (loads, vibration, stability, higher than current turboprop aircraft, although about 35% noise, gust response, etc.) must be attacked using active lower than is customary in the helicopter industry. So good control. aerodynamic design practice should be sufficient to achieve the target for airframe drag. For concepts with edgewise rotors in cruise, hub drag must be adde d to the airframe and MISSION AND DESIGN CONDITIONS wing drag of the aircraft. For this investigation, the target Based on the Rotorcraft Sector notiona l vehicle capabilities 2/3 hub drag was D/q = 0.4(W/1000) , which is less than half and technology goals, a civil mission was defined. This of current hub drag levels. Achieving this hub drag level investigation is not intended to specify the market, but will require advanced technology, certainly fairings bu t rather to identify enabling technology for civil applications possibly also active flow control.

of heavy lift rotorcraft. Table 2 describes the mission, a nd Table 3 describes the payload and fuselage. Note in The weight technology factors used for the three baseline particular the OEI requirement in Table 2: at takeoff rotorcraft designs are summarized in Table 5. In the RC o conditions (5k ISA+20 C) the contingency power of the weight equations, the blade and hub weight technology remaining engines (133% OEI MCP) must be greater than were actually characterized by the blade flap frequ ency; the 90% hover out - of - ground - effect power re quired (the factor equivalent multiplicative factors are given in Table 5. The of 90% accounting non - zero speed and some altitude loss baseline technology for the present designs was hingeless during the takeoff).

rotors. Advanced technology rotors have light blades, hence the actual blade flap frequencies are high. Weight reduction For maximum utilization, the aircraft must have a wide obtaine d from technology was specified by a reduced range of capabilities. Although the aircraft were designed to equivalent flap frequency in the weight equations, reflecting the mission defined in Table 2, hence with very little hover new design concepts for the blades and hub. In these terms, time, efficient hover and low speed capability is essential to the flap frequency was reduced by the factor 0.91 relative the RIA operational concept. This is reflected in the current technology, resulting in the multiplicative factors requirement for essentially OEI hover capability. The given in Table 5. In addition, the weight equations used had resulting designs optimize at balanced cruise and OEI hover a factor of 1.18 for tiltrotor blades compared to helicopter power, so the cruise speed of 350 knots can be viewed as a blades, based on calibration with current technology. The fallout of the OEI requirement. Reasonable downwash and drive system weights for the baseline aircraft were outwash from the rotors hovering in ground effect is calculated using the technology factor given in Table 5, required for effective utilization. For example, a downwash without any penalty for using a two - speed transmission of 20 lb/ft would produce an outwa sh with a peak velocity design.

of over 90 knots. As a result of these considerations, high disk loading aircraft (such as tiltwings) were not among the A scaled engine model was used by the sizing code. The configurations considered here.

current and advanced engine technology is characterized in Table 6. This mo del and technology were defined for Critical design conditions appropriate for civil heavy lift engines with SLS MCP greater than 5000 hp.

rotorcraft operatio ns were defined for calculation of performance, loads, and stability. Table 4 summarizes these The definition of the technology level in the sizing code aeromechanics analysis conditions.

also involves performance and aerodynamics. For the rotor, the design blade loading C / σ was prescribed, based on an W TECHNOLOGY FACTORS AND DESIGN as sessment of what advanced technology could provide.

PARAMETERS Rotor induced and profile power in the sizing code were Meeting the technology goals of the NASA Rotorcraft calibrated to the results of the comprehensive analysis Sector requires high speed, high altitude, and long range for calculations. Thus the sizing code performance represented productivity. The heavy lift rotorcraft must have low disk a rotor with optimum twist, taper , cruise tip speed, etc.

loading for good hover efficiency, and low drag for efficient However, current technology airfoils were used in the cruise. The target for improvement in hover efficiency comprehensive analysis optimization. Some further improvement in aircraft performance can thus be expected blade weight, while a smaller rotor would require more from the use of advanced technology airfoils, especially if power hence more engine and fuel weight). Table 9 specifical ly designed for these aircraft. Airframe drag was compares the component weights of the three designs. The specified as described above. Current technology values empty weight fraction is about 65%. The fixed weight is were used for hover download. Some further improvement compa rable to current commercial jet aircraft. Table 10 in aircraft performance might be obtained from download shows the cruise drag buildup. The drag of the LCTR is reduction. comparable to good turboprop aerodynamic design. The LCTC adds the drag of the hub (less than current The statistical weight equat ions used in the design code technology levels), and the LABC does not have t he drag of incorporate an influence of aircraft size, based on historical the wing. This LABC design was produced by the sizing trends. For rotorcraft designed to fixed disk loading, tip code using a rotor cruise L/D that was higher than that e speed, blade loading (solidity), and number of blades, these predicted by the comprehensive analysis.

equations imply that rotor blade, rotor hub, and drive system weight scale with gross weight to the 1.26, 1.39, and The aircraft cruise L/D=WV/P (based on cruise power, 1.12 power, respectively. So for an increase in gross weight including losses, at design gross weight ) was the principal by a factor of 2.0, the rotor blade, rotor hub, and drive efficiency metric. For the mission considered, the LCTR system weight increase by factors of 2.4, 2.6, and 2.2; and had the best cruise efficiency, hence the smallest design the a ircraft structural and drive system weight therefore gross weight and the smallest installed power (Table 8).

increases by about a factor of 2.2. In order to maintain Next in efficiency is the LCTC, and after that the LABC.

aircraft empty weight fraction as size increases, the design Figure 4 shows the flyaway cost and DOC+I for the three approach must be changed, which conventionally has heavy lift rotorcraft configurations, and Figure 5 presents resulted in an increase in disk lo ading with size.

the DOC+I breakdown for the 1200 nm design mission.

Basic parameters of the rotorcraft were chosen for the three These figures include the Boeing 737 costs for comparison.

heavy lift configurations based on an assessment of current The block hours per year value was ba sed on Southwest and future technology (Table 7). The rotor blade loading Airlines operations. The difference in dead time between (C / σ , based on gross weight and thrust - weighted solidity) the VTOL and 737 reflected the difference in operations.

W wa s chosen considering low speed maneuverability For the VTOL costs, the aircraft parameters (empty weight, requirements. The C / σ values in Table 7 correspond to installed power, number of rotors and number of blades) W about an 8% improvement in maximum lift capability, and the mi ssion parameters (fuel weight, block time and compared to current technology. A relatively low hover tip block speed for a specified range) were obtained from the speed was used, reflecting the importance o f the noise goal. RC code.

The cruise tip speed was chosen to optimize the The VTOL cost model is driven by gross weight and power, performance. To be conservative, hover download values so the LCTR has the lowest cost, followed by the LCTC consistent with current technology were used. A low wing and then the LABC. At the design stage length, the LCTR loading was chosen, for good low speed maneuverability cost is about 20% higher than that of a current 737. That is and wide conversion spe ed range. The same blade loading the cost of VTOL capability. The LCTR is more and wing loading design values were used for both tiltrotor economical than the 737 for stage lengths below about 200 and slowed - rotor compound configurations.

miles.

SUMMARY OF DESIGNS LARGE CIVIL TILTROTOR (LCTR) The heavy lift rotorcraft designs are summarized in Table 8.

The configuration of the Large Civil Tilt Rotor (LCTR) is Three - views of the aircraft are shown in Figures 1 – 3. Recall shown in Figure 1. The aircraft had two tilting rotors at the that for these designs the blade loading, hover tip speed, and wing tips, a low wing, non - tilting engines, and a horizontal wing loading were specified, based on assessments of the tail. A quad tiltrotor (two wings and four rotors) would have technology. Cruise tip speed was optimized based on cruise smaller rotors, but increased complexity and increased efficiency. The disk loading was optimized, based on aerodynamic interference. The conventional two - rotor airc raft weight, power, and cost. Basically the optimum disk tiltrotor configuration was considered here, which allowed loading produces a balance in power requirement between more exploration of the implications of large size on the cruise and OEI hover. Cruise efficiency defines the power rotor system design. A low wing was adapted for better available, then the disk loading is chosen that uses that structural load path s between wing, airframe, and landing power in hover (a larger ro tor would increase the rotor and gear. The horizontal tail was sized by trim requirements at the design conditions. These results are for current rather than stability, because the rotors can be used for technology rotor airfoils. Figures 8 and 9 show the hover flight dynamics stabilization as well as control. A vertical and cruise performance of the main rotor.

tail is not shown, but could be added if n eeded for yaw trim.

LARGE CIVIL TANDEM COMPOUND (LCTC) Table 11 gives the aircraft characteristics. Performance, The configuration of the La rge Civil Tandem Compound loads, and stability calculations were performed for the (LCTC) is shown in Figure 2. The aircraft had two main conditions defined in Table 4. For helicopter mode loads rotors in tandem configuration, a high wing, pusher calculations, lateral flapping was trimmed to zero using propellers for cruise propulsion, and a horizontal tail. The lateral cyclic. Symmetric trim was used for cruise length of the fuselage follows from the specification of the performance and helicopter mode loads calculations (trim p ayload, and the disk loading was optimized to balance the aircraft lift, drag, and pitching moment). For cruise stability cruise and hover power. As a result there was no overlap of calculations, the rotor was trimmed to conditions known to the rotors. The horizontal tail was sized by trim simulate extremes of whirl flutt er behavior: the rotor requirements rather than stability.

trimmed to zero power; or the rotor trimmed for aircraft drag equilibrium up to maximum power, and then trimmed Table 11 gives the aircraft characteristics. Perform ance, to constant power.

loads, and stability calculations were performed for the conditions defined in Table 4. The comprehensive analysis A hingeless rotor hub was used. To reduce mean blade modelled the auxiliary propulsion as forces applied to the bending loads, the hub incorporated 6 deg precone and airframe. Rotor/rotor and rotor/wing interference were 0.002R torque offset. For blade stability, the chordwise accounted for using t he vortex wake model.

center of gravity offset was constrained to be no farther than 5% chord aft of the quarter chord. Excessive coning can In hover and low speed flight, standard tandem helicopter significantly reduce hover figure of merit. So a tip mass of controls, plus aircraft pitch and roll attitude, could be used 1.5 slug was p laced on each blade at 95%R, in order to to trim this aircraft. At moderate speeds, the pitch angle reduce coning and thereby improve hover performance (an could be fixed and the propeller thrust trimmed instea d.

increase in hover figure of merit of about 2% was Even at low speeds, the lateral stick would be connected to produced). Figure 6 shows the calculated blade frequencies, the ailerons, and the longitudinal stick to the elevator. For at collective pitch angles representative of heli copter mode the 80 knot load factor sweep (to obtain blade loads), the and cruise. At helicopter mode tip speeds, the lag frequency mean propeller thrust was fixed at the aircraft drag value, was above 2/rev and the torsion frequency above 12/rev.

and the pilo t's controls plus aircraft pitch and roll attitude With these dynamic characteristics, no stability issues were were used to trim the aircraft (with pilot's collective, observed, either blade or whirl flutter.

longitudinal cyclic, lateral cyclic, and pedal connected to mean rotor collective, differential collective, ailerons, and The blade twist and taper were varied to optimize the rotor differential propeller thr ust respectively). In addition, for hover and cruise performance. The hover condition was o flapping was trimmed to zero (for load control) using rotor 5k ISA+20 C, 650 ft/sec tip speed, C / σ = 0.1557. The T cyclic pitch; thus there were 10 trim variables for the load cruise condition was 350 knots, 30k ISA, 350 ft/sec tip factor sweep.

speed, trim aircraft drag. The twist distribution had two linear segments, inboard (0.0R to 0.5R) and outboard (0.5R In cruise the aircraft was trimmed using lateral stick to the to 1.0R). The comprehensive analysis did not have a ailerons, longitudinal stic k to the elevator, pedal to collocation poi nt at 0.5R, so a transition from inboard slope differential propeller thrust; plus propeller thrust, and to outboard slope was not modelled. The taper model aircraft pitch and roll angles. Front and rear rotor collective considered was constant thrust - weighted solidity (constant pitch angles were set to values optimized for cruise 75%R chord). Figure 7 presents the results for twist performance (optimized rotor thrust). In addition, ro tor optimization, showing the typical hover - cruis e compromise.

flapping was trimmed to zero (for load control) using rotor The result was an optimum twist of – 32 deg inboard and – longitudinal and lateral cyclic; thus there were 10 trim 30 deg outboard; and an optimum taper of 0.8 (tip/root variables for cruise.

chord).

A hingeless rotor hub was used. Figure 10 shows the The rotor performance from the sizing code and the calculated blade frequencies, at a collective pitch angle of comprehensive analysis are compared in Table 12. The RC 10 deg. At helicopter mode tip speeds, the lag frequency model was adjust ed to match the CAMRAD II performance was above 6/rev and the torsion frequency about 7.5/rev.

With these dynamic characteristics, no stability issues were Ducted propellers on stub wings might be a better observed, either in hover or in high advance ratio forward configuration for the auxiliary propulsion.

flight.

Table 11 gives the aircraft characteristics. Performance, The blade twis t and taper were varied to optimize the rotor loads, and stability calculation s were performed for the for hover and cruise performance. The hover condition was conditions defined in Table 4. The comprehensive analysis o 5k ISA+20 C, 650 ft/sec tip speed, C / σ = 0.1491. The modelled the auxiliary propulsion as forces applied to the T cruise condition was 350 knots, 30k ISA, 205 ft/sec tip airframe. Rotor/rotor interference was accounted for using speed, 138764 lb gross weight. The twist distribution had the vortex wake model.

two linear segments, inboard (0.0R to 0.5R) and outboard In hover and low speed flight, s tandard coaxial helicopter (0.5R to 1.0R). The taper model considered was constant controls, plus aircraft pitch and roll attitude, were used to thrust - weight ed solidity (constant 75%R chord). Figure 11 trim the aircraft. At moderate speeds, the pitch angle was presents the results for twist optimization, showing the fixed and the propeller thrust trimmed instead. Even at low hover - cruise compromise. For each value of outboard twist, speeds, the pedal was connected to the rudder, a nd the the inboard twist values are 3, 0, – 3, and – 6 deg. The result longitudinal stick to the elevator. In addition, differential was an optimum twist of 0 deg inb oard and – 12 deg hub moment was trimmed to zero (for load control) using outboard; and an optimum taper of 0.8 (tip/root chord).

differential cyclic; thus there were 8 trim variables for low Collective pitch of the front and rear rotors was varied to speed flight.

find the optimum rotor thrust for high speed cruise flight.

In cruise the aircraft was trimmed using lateral stick to rotor For an untwisted rotor, the best aircraft performance would lateral cyclic, longitudinal stick to the elevator, pedal to the be obtained with zero collective (no lift, no induced power, rudder; plus propeller thrust, and aircraft pitch and roll minimum profile power). With negative outboard twist, for angles. Lift offset (rotor differential roll moment) was improved hover performance, the optimum collective was – trimmed to a specified value using differential lateral cyclic.

2 deg, which resulted in the rotors carrying about 10% of Ro tor collective pitch angles were set to values optimized the aircraft lift (the rot or thrust variation with collective was for cruise performance (optimized rotor angle of attack). In negative at this high advance ratio). This optimum occurred addition, rotor pitch moment was trimmed to zero (for load with a small, positive shaft power to the rotors. With the control) using rotor longitudinal cyclic; thus there were 9 rotor in autorotation (achieved using an aft tilt of the rotor) trim variables for cruise.

the rotor thrust was large, hence the total rotor drag larger A hingeless rotor hub was used. Figure 14 shows the and the aircraft L/D somewhat smaller.

calculated blade frequencies, at collective pitch angle of 0 The rotor advancing tip Mach number was varied, and the deg. At helicopter mode tip speeds, the flap frequency was optimum cruise performance was found at M = 0.80 (for at about 3/rev, the lag frequency about 9/rev, and the torsion the airfoils used). Further reductions in rotor rotational frequency above 15/rev. With these dynamic characteristics, speed did not im prove the aircraft L/D.

no stability issues were observed, either in hover or in high The rotor performance from the sizing code and the advance ratio forward flight.

comprehensive analysis are compared in Table 12. The RC The blade twist and taper were varied to optimize the rotor model was adjusted to match the CAMRAD II performance for hover and cruise performance. The hover condition was at the design conditions. These results are for current o 5k ISA+20 C, 650 ft/sec tip speed, 160636 lb gross weight.

technology roto r airfoils. Figures 12 and 13 show The cruise condition was 350 knots, 30k ISA, 255 ft/sec tip respectively the hover performance of the main rotor and speed, 160636 lb gross weight. The twist distribution had the aircraft cruise performance. The rotor performance in two linear segments, inboard (0.0R to 0.5R) and outboard cruise is presented in terms of aircraft L/D=WV/P, (0.5R to 1.0R). The resu lt was an optimum twist (equivalent calculated without accessory or other losses, and using a twist rate from root to tip) of 2 deg inboard and – 12 deg propeller efficiency of 0.86 (from the sizing code).

outboard. The chord distribution also consisted of two segments with linear variation, inboard and outboard of LARGE ADVANCING BLADE CONCEPT (LABC ) 0.5R. The result was an optimum taper ratio (equiva lent tip The configuration of the Large Advancing Blade Concept chord to root chord ratio) of 2 inboard and 1/3 outboard.

(LABC) is shown in Figure 3. The aircraft had two main rotors in coaxial configuration, pusher propellers fo r cruise propulsion, and horizontal and vertical tails for cruise trim.

Collective pitch of the rotors was varied to find the TECHNOLOGY PAYOFF optimum rotor shaft angle for high speed cruise flight. With The impact and payoff of adv anced technology were the twist used, the optimum collective was 0 deg.

quantified using the sizing code. For this purpose, the Rotor lift offset (diffe rential roll moment) was varied, and technology factors were changed from values representing the best cruise performance found for 0.2R offset advanced technology to values representing current (differential roll moment divided by gross weight and rotor technology. The technology factors for weights are given in radius). The rotor advancing tip Mach number was varied.

Table 5, a nd the engine model is described in Table 6. Table The optimum cruise performance was found at M = 0. 85 at 16 shows the percentage increase (a negative value is good) (for the airfoils used).

in the five metrics, caused by removal of various aspects of the advanced technology from the design assumptions.

The rotor performance from the sizing code and the Results are given for the LCTR and LCTC, but not for the comprehensive analysis are compared in Table 14. Figures LABC because of the level of maturity of the sizing code 15 and 16 show respectively the hover performance of the for that configuration. Table 16 shows the impact of rotor main rotor and the aircraft cruise performance. These results blade and hub weight reduction, and the impact of all a re for current technology rotor airfoils. Figure 16 also structural weight reductions (blade, hub, fuselage, and shows the cruise performance that might be achieved using wing ). Individually the hub weight, fuselage weight, and advanced airfoils, here simulated by assuming a 10% wing weight had small influence; collectively they reduction in drag and a 10% increase in critical Mach contribute the significant influence shown in Table 16.

number relative the current technol ogy airfoils (which Table 16 shows the impact of drag reductions, and the optimizes to the same twist and taper as with current impact of all aerodynamics (drag, rotor fi gure of merit and technology airfoils, but at a lift offset of 0.2%R and M = at cruise efficiency, and download). Individually the hover 0.90). The rotor performance in cruise is presented in terms figure of merit, cruise efficiency (propulsive efficiency for of aircraft L/D = WV/P, calculated without accessory or LCTR, rotor drag for LCTC), and download had small othe r losses, and using a propeller efficiency of 0.88 (from influence; collectively they contribute the significant the sizing code). For the LABC (unlike the other two influence sh own in Table 16. Table 16 shows the major designs) the RC model was not adjusted to match the impact engine technology has on the designs. Increases in CAMRAD II performance at the design conditions, because vibration treatment weight and acoustic treatment weight in order to obtain a converged design from the sizing code, were also examined, and found to have a small impact on it was necessary to assume a rotor effective L/D the metrics.

substantially larger than that obtained from the comprehensive analysis. It is anticipated that significant A conservative design approach , based on past aircraft improvements in the calculated L/D can be realized using design experience, would increase the estimated power specially designed a irfoils, thereby making the required (and hence fuel burned) by 25%, and increase the comprehensive analysis calculations closer to the estimated empty weight by 15%, for a fixed payload and performance on which the RC design was based.

performance requirement. The penalty for imposing these weight a nd power contingencies is shown in Table 16. As DESIGN OPTIMIZATION the need for large contingencies is attributed to lack of Table 15 compares the baseline designs for the LCTR, accuracy of current design and analysis tools, Table 16 LCTC, and LABC, in terms of the following metrics: shows the economic payoff possible by improving these tools.

a) aircr aft mission gross weight (lb) b) installed engine power (hp) Figure 17 shows the costs for the LCTR with and without c) mission fuel (lb) the cost technology factors, and Figure 18 presents the d) purchase price ($M) corresponding DOC+I breakdown. These results emphasize e) direct operating cost DOC+I (cents/ASM) and quantify the importance of controlling the maintenance costs for heavy lift rotorcraft.

The fuel weights in Table 9 include the reserves. Sensitivity studies were conducted using the sizing code, to optimize DESIGN REFINEMENTS the designs by examining variations in disk loading and Since the e ngine model used was intended to represent an number of blades. The influence of hover tip speed and advanced engine based largely on currently available cruise tip speed were also examined, considering in technology, it was prudent to examine the impact of higher particular the noise requirements.

SFC. Table 17 compares the baseline designs with the Thus the LCTR design demonstrated the potential for designs obtained when the engine SFC was incr eased by achieving the Rotorcraft Sector goals of Table 1. With a 10%: the mission fuel increased by about 15%, and the disk loading of 10 lb/ft compared to the state - of - the - art other metrics increased by about 5%. value of 20 lb/ft , the 40% increase in hover efficiency was attained. Considering the OEI hover power (power from 3 It is possible that achieving the noise goals will require that out of 4 engines), the power loading was W/P = 6.0. At the the blade passage frequency be kept in the audible range.

cruise conditions, the aircraft lift - to - drag ratio was L/D = Table 18 compares the baseline des igns with the designs 14.5 (Ta ble 12), exceeding the 44% improvement goal. The obtained when the blade passage frequency is raised to 20 2/3 airframe drag was estimated to be D/q = 1.5/(W/1000) Hz. For the LCTR, it was necessary to increase the number (Table 10). The weight technology factors (Table 5) led to of blades and increase the disk loading. The increased disk about 22% reduction in gross weight (Table 16), from a loading results in an increase in all metrics except gross 30% reduction in empty weight, which was consistent with weight. The sizing code model for blade and control weight the goal of a 25% reduction in empty weight excluding implies a reduction in rotor weight as the number of blades engines. The design had an empty weight fraction of 0.65 increases, resulting in a decrease in the aircraft gross (Table 9), or 0.62 excluding engines, so technology weight. For the LCTC, it was necessary to increase the countered the growth in empty weight fraction with aircraft number of blades, but di sk loading of the baseline design size and spee d. The calculated noise was 9.3 EPNdB below was already higher than that of the tiltrotor. Hence the certification requirements, compared to the goal of 14 reduction of blade weight caused by the increased number EPNdB, with active control and flight operations available of blades results in a reduction of all metrics for the LCTC.

to obtain the full reduction as well as deal with low These rotor weight trends must be confirmed, b ut the results frequency noise.

of Table 18 suggest that a blade passage frequency of 20 Hz could be an acceptable requirement.

RISK REDUCTION FOR HEAVY LI FT ROTORCRAFT ASSESSMENT OF CONFIGURATIONS The NASA Heavy Lift Rotorcraft Systems Investigation For the NASA civil mission, the Large Civil Tiltrotor was a focused and coordinated analytical effort to select the (LCTR) had the best cruise efficiency, hence the lowest best configuration for meeting the Rotorcraft Sector vehicle weight and lowest cost. The LCTR is the configuration with technology goals. During the course of the investigation, the most promise to meet the NASA technology goals.

high ris k areas were identified. The definition of high risk is The Large Civil Tandem Compound (LCTC) had good one or both of the following: capability or attribute cruise efficiency, but less than the tiltrotor, and higher unavailable today, so it is necessary to assume advanced development risk than the tiltrotor. Single main rotor and technology will be available in the future in order for the tandem rotor configurations were comparable in efficiency aircraft to achieve the techn ology goals; or cost prevents the and risk. Even if reaction drive produced the smallest vehicle from being economically competitive, so the payoff slowed - rotor compound rotorcraft, the high installed power of advanced technology is essential to achieving the goals.

compromises efficiency, and the reaction drive system has The following were identified as high risk areas for heavy higher nois e and substantially increased risk.

lift rotorcraft: The Large Advancing Blade Concept (LABC) had lower a) High torque, light weigh t drive system.

cruise efficiency than the tiltrotor for the NASA civil b) High performance, structurally efficient rotor/ wing mission .

system.

c) Low noise aircraft.

The LCTR design presented was economically competitive d) Super - integrated vehicle management system.

with comparable fixed wing aircraft, with the pot ential for substantial impact on the air transportation system. The Plans were then developed to mitigate the above risks. The keys to achieving a competitive aircraft are: low drag risk reduction plans provide the strategic dire ction to airframe and low disk loading rotors; structural weight support a heavy - lift rotorcraft development.

reduction, for both airframe and rotors; drive system weight reduction; impro ved engine efficiency; low maintenance STRATEGIC DIRECTION design; and manufacturing cost comparable to CTOL The strategic direction provides guidance for selecting aircraft.

highest priority activities, aimed at the four highest risk areas of heavy lift rotorcraft development. Note that there SUPER - INTEGRATED VEHICLE MANAGEMENT are so me important and difficult tasks that are yet not high SYSTEM risk, including rotor aerodynamic design and optimization, Broad spectrum active control is required for an effective rotor and wing airfoil design, airframe aerodynamics, and heavy lift rotorcraft. Large size implies a significant airframe structures.

influence of low frequency airframe elastic modes on flight dynamics. Active control is required to achieve the goals of HIGH TORQUE, LIGHT WEIGHT DRIVE SYSTEM low rotor - induced vibration and noise. Safe operation in Innovative design is req uired for low drive system weight.

one - engine inoperative conditions is essential for civil Large size implies high torque and high weight fraction, rotorcraft. Rotor load limiting and active control are needed hence drive system weight reduction is essential for an for full utilization of the structural capability in the rotor efficient and economical aircraft. The focus must be on and airframe. Hence an expanded integration of the vehicle design concept, advanced - technology components, and management system is required: a flight control system for materials.

good handling qualities and gust response, active control of Low maintenance is required for low operating cost. Low vibration and noise, an d rotor load limiting and active maintenance must be a primary design requirement, even control. The focus must be on load limiting and system ahead of weight and performance.

integration.

High flight speed requires, or at least benefits from, a CONCLUSION variable speed propulsion system design. Fir st it is The NASA Heavy Lift Rotorcraft Systems Investigation necessary to establish the speed range available from examined in depth several rotorcraft configurations for large advanced engine technology, and to define the engine civil transport, designed to meet the technology goals of the required for the heavy lift rotorcraft concept.

NASA Vehicle Systems Program. Design and analysis tools were applied to define three configurations: Large Civil HIGH PERFORMANCE, STRUCTURALLY EFFICIENT Tiltrotor (LCTR), Large Civil Tandem Compound (LCTC), ROTOR/WING SYSTEM and Large Advancing Blade Concept (LABC).

Innovative rotor and wing design is required, probably with unconventional dynamics. Large size implies high weight For the NASA c ivil mission, the Large Civil Tiltrotor had fraction, high speed introduces stability issues, and good the best cruise efficiency, hence the lowest weight and rotor system performance is essential for an efficient and lowest cost. Thus the LCTR is the configuration with the economical aircraft. The focus must be on integrated best potential to meet the NASA technology goals. The ro tor/wing performance and dynamic behavior.

design presented was economically competitive , with the potential for substantial impact on the air transportation Structural efficiency is required for low rotor and hub and system. While fixed wing aircraft for this mission exist, the wing weight. The focus must be on design concepts for investigation only showed the potential for a high speed, durability and damage tolerance.

heavy lift rotorcraft. The keys to achieving a competitive Low maintenance is required for low operating cost. Low aircraf t were low drag airframe and low disk loading rotors; maintenanc e must be a primary design requirement, even structural weight reduction, for both airframe and rotors; ahead of weight and performance.

drive system weight reduction; improved engine efficiency; low maintenance design; and manufacturing cost LOW NOISE AIRCRAFT comparable to fixed - wing aircraft.

New approaches are required to meet the challenge of low Ri sk reduction plans were developed to provide the noise. Large size implies low frequency noise and expanded strategic direction to support a heavy - lift rotorcraft acoustic footprint. An understanding of hea vy lift vehicle development. The following high risk areas were identified acoustic phenomena (low frequency and relative distance to for heavy lift rotorcraft: high torque, light weight drive community) is required, including psychoacoustics for low system; high performance, structur ally efficient rotor/wing frequency. New rotor design guidelines and annoyance system; low noise aircraft; and super - integrated vehicle metrics must be developed. The focus must be on a management system.

combination of rotor design, active control, and flight operations.

[8] Boyd, Jr., D.D., Burley, C.L., and Conner, D.A., ACKNOWLEDGMENTS “Acoustic Predictions of Manned and Unmanned Rotorcraft The NASA Heavy Lift Rotorcraft Systems Investigation Using the Comprehensive Analytical Rotorcraft Model for was a community effort. The participation by Bell Acoustics (CARMA) Code System,” AHS International Helicopter, Boeing, and Sikorsky Ai rcraft is gratefully Specialists' Meeting on Unmanned Rotorcraft Design, acknowledged. Contributions by The Pennsylvania State Control, and Testing, Chandler, AZ, January 2005.

University, University of Maryland, and Georgia Institute of [9] Harris, F.D., and Scully, M.P., “Rotorcraft Cost Too Technology are also much appreciated. The individuals who Muc h.” Journal of the American Helicopter Society , Vol.

contributed to the investigation are too numerous to list 43, No. 1, January 1998. (Additionally, H arris, F.D. “An here , but the authors wish to recognize the following Economic Model of U.S. Airline Operating Expenses,” dedicated individuals on the government analysis team: unpublished.)

Wally Acree, John Coy, William Decker, Robert Kufeld, Ethan Romander, Johannes van Aken (NASA Ames); Doug NOMENCLATURE Boyd, Casey Burley (NASA Langley); Preston Ma rtin, Rick A rotor disk area Peyran, John Preston, Jeff Sinsay, Hyeonsoo Yeo (U. S.

Army Aeroflightdynamics Directorate); Robert Handschuh, c mean drag coefficient for profile power do Kevin O'Brien (U. S. Army – Army Research Laboratory).

C rotor thrust coefficient, T/( ρ AV ) Finally, the guidance provided by Charles Crawford, Troy T tip Gaffey, Franklin Harri s, Robert Loewy, and Kenneth Rosen C rotor weight coefficient, W/( ρ AV ) W tip during the course of this investigation is very much D/q airframe drag divided by dynamic pressure appreciated.

L/D aircraft effective lift - to - drag ratio, W/VP (based REFERENCES on cruise power) [1] Johnson, J., Stouffer, V., Long, D., and Gribko, J., M Mach number “Evaluation of the National Throughput Benefits of the M advancing tip Mach number Civil Tiltrotor,” NASA CR 2001 - 2 11055, September 2001.

at P aircraft power [2] Stouffer, V., Johnson, J., and Gribko, J., “Civil Tiltrotor Feasibility Study for the New York and Washington R rotor radius Terminal Areas,” NASA CR 2001 - 210659, January 2001.

T rotor thru st [3] Smith, D. E., Wilkerson, J., Montoro, G. J., Coy, J., and V flight speed Zuk, J., “Technology Development for Runway Independent Aircraft,” American Helicopter Society 59th V best range flight speed br Annual Forum, Phoenix, AZ, May 2003.

V rotor tip speed tip [4] Preston, J., and Peyran, R., “Linking a Solid - Modeling W gross weight Capability with a Conceptual Rotorcraft Sizing Code,” American He licopter Society Vertical Lift Aircraft Design W empty weight E Conference, San Francisco, CA, January 2000.

W/A disk loading [5] Johnson, W., “Rotorcraft Aeromechanics Applications κ induced power factor (P /P ) ind induced ideal of a Comprehensive Analysis,” HeliJapan 98: AHS International Meeting on Advanced Rotorcraft Technology ρ air density and Disaster Relief, Gifu, Japan, April 1998.

σ rotor solidity (ratio blade area to disk area) [6] Drela, M., “Newton Solution of Coupled Viscous / Inviscid Multielement Airfoil Flows,” AIAA Paper No. 90 - 1470, June 1990.

ASM available seat miles [7] Rajagopalan, R.G. “A Procedure for Rotor Flowfield CTOL con ventional takeoff and landing and Interference: A Perspective,” AIAA Paper No. 2000 - DOC direct operating cost 0116, January 2000.

DOC+I direct operating cost plus interest ISA international standard atmosphere RIA runway independent aircraft LABC Large Advancing Blade Concept SFC specific fuel consumption LCTC Large Civil Tandem Compound SHP shaft power LCTR Large Civil Tilt Rotor SLS sea level standard MCP maximum continuous p ower SOA state of the art MRP maximum rated power VTOL vertical takeoff and landing OEI one - engine inoperative Table 1. Rotorcraft Sector capability set and technology goals.

ROTORCRAFT NOTIONAL VEHICLE 15 - YEAR CAPABILITIES Payload 120 passengers Cruise speed M = 0.60 (350 knots) at 30000 ft Cruise altitude at or above 22000 ft (icing) Range 1200 nm ROTORCRAFT SECTOR 15 - YR TECHNOLOGY GOALS Hover ef ficiency, W/P 6 Efficient Cruise, L/D 12 Empty Weight Fraction 0.41 (excluding engines) Community Noise SOA – 14 EPNdb Flight Control Automated single - pilot CAT IIIC SNI for heavy lift Advanced Engine Performance SFC = SOA*0.9, SHP/W = SOA*1.2 Cabin No ise and Vibration 77dBA & 0.05g Table 2. Civil design mission.

1200 nm range, 120 passengers Cruise at 350 knots and 30000 ft (min 22000 ft, for icing) Design mission Idle 5 min o Takeoff + 1 min Hover OGE 5k ISA+20 C [convert] Climb at V best range (0k ISA to 30k ISA, distance part of range) Cruise at 350 knots, for 1200nm range 30k ISA Reserve: 30 min + 30 nm at V 30k ISA br Descend at V (no range credit) br [convert] o 1 min Hover OGE + Landing 5k ISA+20 C Idle 5 min Design power o Hover: 95% MRP, 5k ISA+20 C Cruise: 100% MCP, 30k ISA One engine inoperative (OEI): o at 5k ISA+20 C, 133% (OEI MCP) greater than 90% (HOGE P ) req at 22k ISA, (OEI MCP) greater than (P at V ) req br 4 engines Table 3. Payload and fuselage.

Payload: 120 passengers = 26400 lb Passengers: 120 at 220 lb each (190 + 30 baggage) Flight crew: 2 at 240 lb each Cabin crew: 3 at 210 lb each Fuselage size and layout 12 first class (4x3, 38 in pitch) 108 economy class (6x18, 32 in pitch) Length = 109.61 ft, width = 12.25 ft Table 4. Critical design conditions for aeromechanics analysis.

Blade stability Thrust sweep in hover (SLS), to rotor stall Level flight speed sweep (30k ISA), to maximum power Aircraft and rotor stability Up to 350 knots at SLS, 500 knots at 30k ISA flutter speed 1.2 V = 1.2 (1.25 V ) = 1.50 V = 525 knots (FAR) dive cruise cruise flutter speed V = 1.15 (1.2 V ) = 1.38 V = 480 knots (MIL - A - 8870) L cruise cruise Tiltrotor high speed forward flight Zero and max power; sea level, 30k; symmetric and an tisymmetric modes, with drive train Ground resonance and air resonance for soft - inplane rotors Performance o Thrust sweep in hover (5k ISA+20 C), for power and figure of merit Speed sweep in high speed forward flight (30k ISA) for power and efficiency Loads (blade, hub, control), deflection, and vibration Load factor sweep at 80 knots (SLS), to 1.5g o Level flight speed sweep (5k ISA+20 C), to maximum power Nacelle angles of 80, 60 deg for tiltrotor Table 5. Weight technology factors used for aircraft sizing .

Rotor blade weight 0.79 Rotor hub weight 0.96 Drive system weight 0.67 Fuselage weight 0.88 Wing primary structure weight 0.88 Empennage weight 0.90 Table 6. Engine technology used for aircraft sizing.

current technology advanced technology SFC (SLS MCP) lb/shp - hr 0.4260 0.3243 specific power (MCP) hp/lb/sec 140.8 290.0 power/weight shp/lb 6.49 7.48 Relative SLS MRP o MRP at 5k ISA+20 C ratio shp 0.769 0.769 MCP at 30k ISA ratio shp 0.348 0.348 o Fuel flow at 5k ISA+20 C r atio lb/hr 0.781 0.781 Fuel flow at 30k ISA ratio lb/hr 0.334 0.334 Table 7. Advanced technology estimates.

LCTR LCTC LABC tiltrotor tandem compound advancing blade concept Specified o Hover C / σ , (5k ISA+20 C) 0.141 0.141 0.100 W o Hover C / σ , (4k/95 F) 0.140 0.140 0.100 W Hover download 9.7% 5.7% 5.5% Tip speed, hover ft/sec 650 650 650 Tip speed, cruise ft/sec 350 205 255 Cruise speed, 30k knots 350 350 350 2/3 2 Drag, D/q / (W/1000) ft 1.5 1.9 1.3 Wing loading lb/ft 80 80 — Opti mum Disk loading, W/A lb/ft 10 15 20 Maximum M 0.70 0.80 0.85 at Cruise tip speed ft/sec 350 205 255 Table 8. Heavy Lift Rotorcraft Designs.

LCTR LCTC LABC tiltrotor tandem compound advancing blade Mission gross weight (lb) 123562 138764 16 0636 Engines (hp) 4x6914 4x9684 4x14267 Rotor diameter (ft) 88.7 76.7 90.5 Disk loading W/A (lb/ft ) 10 15 25 o C / σ (geom, 5k ISA+20 C) 0.133 0.133 0.0675 W o C / σ (T - wt, 5k ISA+20 C) 0.141 0.141 0.090 W Hover tip speed (ft/sec) 650 650 650 Cruise tip speed (ft/sec) 350 205 255 maximum M 0.70 0.80 0.85 at Solidity 0.0881 0.1321 0.1721 Number blades per rotor 4 4 5 chord (75%R, ft) 3.06 3.98 4.89 aspect ratio 14.5 9.6 9.2 taper ratio 0.8 0.8 0.33 Drag D/q (ft ) 37.3 50.3 38.1 2/3 (D/q)/(W/1000) 1.5 1.9 1.3 Wing loading (lb/ft ) 80 80 – area (ft ) 1545 1735 – span (ft) 105 144 – asp ect ratio 7.1 12.0 – Mission, payload 120 pass 120 pass 120 pass range (nm) 1200 1200 1200 cruise altitude (ft) 30000 30000 30000 cruise speed (kt) 350 350 350 Cruise power (hp) 11904 15956 25068 Cruise L/D=WV/P 11.1 9.3 6.9 Table 9. Concept weight comparison.

LCTR LCTC LABC GROSS WEIGHT 123562 138762 160636 weight empty fraction 65.3% 65.6% 64.7% WEIGHT EMPTY 80701 91079 103991 FIXED WEIGHT 13583 13583 13583 SCALED WEIGHT 67119 77497 90408 Structure 36104 41668 47529 Wing Group 8804 11998 0 Primary Thruster 13714 11494 24572 Tail / Aux Thrust 594 2870 4135 Body Group 7072 10194 11596 Landing Gear Group 3228 3625 4197 Nacelle 2422 1091 2411 Air Induction 270 397 617 Propulsion 18373 24928 29021 Engine installation 4540 6446 9284 Fuel System 556 957 887 Drive System 13277 17525 18850 Flight Controls 4927 4628 5238 Other Scaled Weight 7716 6273 8621 USEFUL LOAD 42860 47684 56645 Crew 1110 1110 1110 Fixed Useful Lo ad 100 100 100 Fluids 240 240 240 Fuel 15010 19834 28795 Table 10. Cruise drag buildup.

LCTR LCTC LABC Wing D/q 14.10 15.84 — area 1545 1735 C .0091 .0091 D Body D/q 11.88 12.42 12.38 S 3650 3650 4290 wet C .0021 .0021 .0021 f interference 4.32 4.86 3.50 Horizontal Tail D/q 1.33 1.91 1.37 area 180 217 186 Vertical Tail D/q 1.33 Pylon D/q 10.02 9.39 9.45 Hub D/q 10.72 13.45 2/3 hub D/q / (W/100 0) 0.40 0.45 Total D/q 37.33 50.28 38.06 2/3 D/q / (W/1000) 1.50 1.88 1.29 Gross Weight 123562 138764 160636 Hover Download (%T) 9.7% 5.7% 5.5% Table 11. Heavy lift rotorcraft characteristics.

LCTR LCTC LABC Design gross weight (lb) 12 3562 138764 160636 Total cruise drag, D/q (ft ) 37.3 50.3 38.1 Disk loading, W/A (lb/ft ) 10 15 25 o Hover C / σ (T - wt, 5k ISA+20 C) 0.141 0.141 0.090 W Hover download 9.7% 5.7% 5.5% Rotor radius (ft) 44.35 38.37 45.22 Number of blades per rotor 4 4 5 So lidity (thrust weighted) 0.0881 0.1321 0.1721 Chord (75%R ft) 3.07 3.98 4.89 Maximum M 0.70 0.80 0.85 at Tip speed (ft/sec), hover 650 650 650 Tip speed (ft/sec), cruise 350 205 255 Rotor speed (rpm), hover 140 162 137 Rotor speed (rpm), cruise 75 51 54 Blade taper 0.8 0.8 2 & 1/3 Blade twist (deg), inboard of 50%R – 32 0 2 Blade twist (deg), outboard of 50%R – 30 – 12 – 12 Lock number 12.1 13.0 19.1 Single blade weight (lb), from blade structural design 745 646 1080 Total blade weight (lb), all roto rs 5960 5168 10800 data source and identification size, airframe aerodynamics: from RC code 5/13/05 4/22/05 5/12/05 blade stiffness, inertia: from structural design 6/17/05 5/13/05 7/18/05 airframe structural dynamics: from NASTRAN model 5/05 5/05 5 /05 rotor airfoils current technology airfoils, with Reynolds number correction of drag, and with stall delay for LCTR Table 12. Comparison of LCTR rotor performance from RC (sizing) and CAMRADII (comprehensive analysis).

HOVER CRUISE o 5k ISA+20 C, 650 ft/sec 350 knots, 30k ISA, 350 ft/sec V/V = 1.688, M = 0.70 tip at RC CAMRADII RC CAMRADII Thrust 136836 136751 8531 8538 C / σ 0.1557 0.1556 0.0720 0.0720 T Power 17061 17155 11283 11296 Parasite power 9163 9170 Induced power 15879 15967 248 2 56 Profile power 1182 1186 1875 1869 κ 1.186 1.193 24.323 25.125 ind c 0.0091 0.0091 0.0087 0.0087 do Figure of merit 0.785 0.780 Propulsive efficiency 0.812 0.812 L/D = WV/P 11.76 11.75 Table 13. Comparison of LCTC rotor performance from RC ( sizing) and CAMRADII (comprehensive analysis).

HOVER CRUISE o 5k ISA+20 C, 650 ft/sec 350 knots, 30k ISA, 205 ft/sec V/V = 2.882, M = 0.80 tip at RC CAMRADII RC CAMRADII Thrust 147102 146914 13608 12812 C / σ 0.1491 0.1489 0.2981 0.2806 T Power 2380 3 23513 0 16 Induced power 22520 22237 462 442 Profile power 1283 1276 2280 2348 κ 1.305 1.291 13.332 14.389 ind c 0.0088 0.0088 0.0101 0.0104 do Figure of merit 0.725 0.732 Rotor drag 2553 2583 Rotor D/q 16.5 16.7 L/D = WV/P 9.9 9.8 Table 14. Comparison of LABC rotor performance from RC (sizing) and CAMRADII (comprehensive analysis).

HOVER CRUISE o 5k ISA+20 C, 650 ft/sec 350 knots, 30k ISA, 255 ft/sec V/V = 2.319, M = 0.85 tip at RC CAMRADII RC CAMRADII Thrust 169960 170592 159068 159192 C / σ 0.0952 0.0955 1.2468 1.2478 T Rotor shaft power 32497 37961 0 - 2554 Induced power 30482 35940 11224 Profile power 2015 2022 11963 Ind+pro power 14251 23188 κ 1.186 1.390 3.221 ind c 0.0076 0.0077 0.0280 do Figure of merit 0.791 0.681 Drag 1 3268 23997 Rotor L/D 12.0 7.7 e L/D = WV/P 7.2 5.1 Table 15. Comparison of baseline designs.

LCTR LCTC LABC Gross weight (lb) 123562 138762 160636 Engine power (hp) 4x6914 4x9684 4x14267 Mission fuel (lb) 13624 17902 26008 Purchase price ($M ) 61.9 8 4.5 11 6.8 DOC (cents/ASM) 13.3 17.2 23. 6 Table 16. Impact of technology on the designs: percentage increase caused by changing the technology from advanced to current level.

weight power fuel price DOC LCTR Blade weight 12 13 11 16 13 All structural weight 21 22 18 27 22 Drive system weight 23 24 21 29 25 Airframe drag (+25%) 6 14 14 12 10 All aerodynamics 10 20 21 17 16 Engine technology 23 28 70 28 29 Weight and power contingency 13 25 25 22 19 LCTC Blade weight 9 9 8 11 1 0 All structural weight 20 20 18 24 22 Drive system weight 29 28 26 35 31 Airframe drag (+25%) 5 10 13 8 8 Hub drag 9 22 24 17 16 All aerodynamics 17 38 39 31 28 Engine technology 48 63 124 60 61 Weight and power contingency 13 25 25 22 20 Tabl e 17. Influence of increased engine SFC.

LCTR LCTC baseline increased SFC baseline increased SFC Engine SFC 100% 110% 100% 110% Gross weight (lb) 123652 128555 138764 145928 Engines (hp) 4x6914 4x7198 4x9684 4x10167 Disk Loading (lb/ft ) 10 10 15 15 Rotor diameter (ft) 88.7 90.5 76.7 78.7 Rotor solidity 0.0881 0.0881 0.1321 0.1321 Number blades per rotor 4 4 4 4 Chord (75%R, ft) 3.06 3.13 3.98 4.08 Blade aspect ratio 14.5 14.5 9.6 9.6 Rotor weight (lb) 13714 14449 11494 12285 Mission Fuel ( lb) 13624 15528 17902 20652 Increase Relative Baseline Gross weight 4% 5% Power 4% 5% Mission fuel 14% 15% Price 4% 5% DOC 5% 6% Table 18. Influence of increased rotor blade passage frequency.

LCTR LCTC baseline increased BPF baseline increased BPF Blade passage freq (Hz) 9.3 19.8 10.8 20.1 Gross weight (lb) 123652 117261 138764 129857 Engines (hp) 4x6914 4x8137 4x9684 4x9417 Disk Loading (lb/ft ) 10 14 15 16 Rotor diameter (ft) 88.7 73.0 76.7 71.9 Rotor solidity 0.0881 0 .1233 0.1321 0.1410 Number blades per rotor 4 7 4 7 Chord (75%R, ft) 3.06 2.02 3.98 2.28 Blade aspect ratio 14.5 18.1 9.6 15.8 Rotor weight (lb) 13714 9738 11494 9295 Mission Fuel (lb) 13624 14373 17902 16792 Increase Relative Baseline Gross wei ght – 5% – 6% Power 18% – 3% Mission fuel 6% – 6% Price 5% – 6% DOC 3% – 6% Figure 1. Three - view of Large Civil Tiltrotor (LCTR).

Figure 2. Three - view of Large Civil Tandem Compound (LCTC).

Figure 3. Three - view of Large Advancing Blade Concept (LABC) Figure 4. Flyaway price (2005 USD) and DOC+I (2005 cents/ASM) comparisons for baseline designs.

Figure 5. Cost elements compared for heavy lift rotorcra ft and B737 (1 , 200 nm, 120 passengers, including technology factors for rotorcraft costs).

Figure 6. LCTR blade and airframe frequencies. Collective = 0 deg (left figure, appropriate for 140 rpm operation) and collective = 60 deg (right figure, 7 5 rpm operation).

Figure 7. LCTR twist optimization. Figure 8. LCTR rotor hover performance.

Figure 11. LCTC twist optim ization.

Figure 12. LCTC rotor hover performance.

Figure 9. LCTR rotor cruise performance.

Figure 13. LCTC aircraft cruise performance.

Figure 10. LCTC blade and airframe frequencies (collective = 10).

Figure 15. LABC rotor hover performance.

Figure 14. LABC blade and airframe frequencies (collective = 0).

Figure 16. LABC aircraft cruise performance.

Figure 17. Effect of cost technology factors on flyaway price (2005 USD) an d DOC+I (2005 cents/ASM) for LCTR.

Figure 18. Cost elements compared for LCTR with and without cost technology factors (1 , 200 nm, 120 passengers)