NASA Heavy Lift Rotorcraft Systems Investigation — 20060023335 (2005)

section operating conditions (lift coefficient, Mach collective was varied to trim the rotor to a specified

CAMRAD II was used to define the rotor blade since the axial flow environment is periodic). The section operating conditions (lift coefficient, Mach collective was varied to trim the rotor to a specified number, and Reynolds number) for hover, cruise, thrust, but bending and torsion deflections of the and maneuver points. The Eppler incompressible blade were ignored.

airfoil code was used to obtain initial contours, given the airfoil geometric constraints and flow The OVERFLOW-D calculations were used to physics issues. Then the MSES transonic airfoil determine the maximum allowable thickness of the code (Euler solver plus boundary layer analysis, proprotor blades. Structural considerations imply a ref. 6) was used with an optimization routine to large thickness ratio, while minimizing produce airfoil contours that satisfied the compressibility power losses implies small constraints. A basic philosophy of the design thickness ratio. Using current technology airfoils approach was to relax pitching moment constraints scaled in thickness, it was determined that a tip for individual sections, and instead design for small thickness ratio of 8% or less is needed, in order to integrated torsion moment over the entire blade. avoid excessive wave drag. With an 8% thick tip, a range of root thicknesses from 20 to 26% gave Figure 28 shows the airfoils designed for the LCTR nearly the same propulsive efficiency.

blade: inboard, mid-span, and tip sections. The objective was to reduce the section drag relative to The LCTR airfoils were designed for maximum current technology airfoils at hover and cruise performance at the hover and cruise operating conditions, while maintaining or increasing conditions. The LCTR tip airfoil is 9.0% thick.

maximum lift capability for maneuver conditions. When the LCTR airfoils were used in the Figure 29 shows the envelope (for variations of OVERFLOW-D calculations, there were no shocks twist) of cruise propulsive efficiency and hover at the tip. However, shocks did form at the root, figure of merit with the current technology (SOA) extending out to 0.14R with a peak Mach number of and LCTR airfoils. The state-of-the-art curve 1.10 (fig. 31(a)). This shock corresponded to the corresponds to the envelope of figure 12. The point of maximum Mach number on the spinner.

airfoils designed specifically for the LCTR produce Hence the spinner diameter was reduced in the improvements of about 2% in cruise and 1% in vicinity of the rotor plane, which reduced the peak hover. Mach number to 1.01 (fig. 31(b)). The result was a 2% improvement in propulsive efficiency.

All of the performance calculations from the comprehensive analysis, to which the sizing code The ROT3DC Navier-Stokes flow solver was used performance analysis was calibrated, were to calculate the LCTR airframe aerodynamics, in performed using current technology airfoils (with particular the hover download. ROT3DC solves Reynolds number corrections implemented in incompressible Navier-Stokes equations on a CAMRAD II). Figure 29 therefore illustrates the Cartesian grid, with the rotors modelled by time- performance improvements that can be expected by averaged momentum source disks. Figure 32 shows introducing specially designed airfoils. the wing and body pressures from a download calculation for high wing and low wing LCTR Aerodynamics configurations. The download with the low wing was about 8%, compared to 10% for the high wing; The Navier-Stokes flow solver OVERFLOW-D was half of the reduction comes from the loading on the used to analyze the transonic flow about the tiltrotor wing and half from the loading on the fuselage.

and spinner in cruise flight, with the objective of optimizing rotor performance and minimizing Noise Assessment interference drag. Figure 30 shows the grid system.

Overset, body fitted, curvilinear grids were nested The NASA technology goal is to reduce the aircraft inside a series of Cartesian background grids. noise by 80%, hence a 14 EPNdB reduction below Approximately 8 million grid points were used (to the SOA level. Achieving this goal will require a model one blade and one-quarter of the centerbody, combination of design for low noise, active control for noise source reduction, and flight operational frequencies from 45 to 56 Hz. Therefore, any noise procedures to reduce the noise impact on the below 45 Hz is ignored by the EPNL metric. These community. represent difficult issues to resolve and may require the FAA to reassess, and possibly change, An initial acoustic assessment of the heavy lift certification requirements for heavy lift rotorcraft.

vehicles was performed. This investigation involved examination of each concept vehicle; identification The Rotorcraft Sector community noise goals are of potentially important interior and exterior noise based on the EPNL metric and on the microphone sources; assessment of the possible impact of these locations used in FAA Noise Certification.

sources on the community and passengers; However, there are two important differences assessment of proposed heavy lift concepts using a between the FAA Certification noise limits and the state-of-the-art prediction capability called the Rotorcraft Sector goals. First, the Rotorcraft Sector Comprehensive Analytical Rotorcraft Model for community noise goal does not require use of the Acoustics (CARMA); and development of risk FAA flight profiles; low noise flight profiles can be reduction plans to define and prioritize key research developed and used. Second, the FAA Certification efforts required to accomplish the Rotorcraft Sector requirements are based on noise limits at specific noise goals. microphones, whereas the Sector goals are based on EPNL averaged across microphones and flight For noise certification, the Federal Aviation conditions. The noise reduction goal can be Administration (FAA) specifies a descent condition, achieved by any combination of the low noise flight a level flight condition, and a take-off condition at operations and vehicle/rotor/engine design. Since which certain maximum noise criteria must be met. the goal is based on the EPNL metric, only noise in Although FAA Noise Certification is not part of the the audible range is considered. The goal metric was NASA Rotorcraft Sector goals, eventually any chosen prior to knowledge of low frequency noise commercial heavy lift vehicle will have to meet characteristics of the current heavy lift designs. The these requirements. Some of the issues that will current heavy lift rotor designs all have blade arise are associated with the sheer size of the passage frequencies below the lower limit of the vehicle. To clarify this, consider that the level flight audible range (which is approximately 20Hz), and condition for the FAA Noise Certification requires much lower than typically found on most an over-flight at an altitude of 120 m, regardless of helicopters. There are likely to be additional noise the size of the vehicle. Certification of a issues at these low frequencies that will need to be hypothetical heavy lift vehicle, with rotors twice the addressed. In addition, the appropriateness of the size of a conventional vehicle, would be equivalent NASA goal metrics must be reassessed.

to requiring a conventional vehicle to be certified at an altitude of 60 m (i.e., half the altitude). Because Infrasonic and very low frequency noise can induce of the size difference, spherical spreading alone vibrations in structures. Community acceptance of suggests that the noise would be approximately 6 dB noises that rattle windows and dishes tends to be higher for the heavy lift vehicle. low. This low acceptance will impact the operations of aircraft by airport operators. Low frequency Another issue is that the heavy lift rotorcraft under noises at very high levels can also have negative consideration are expected to have high levels of physiological and psychological (psycho-acoustic) noise at low frequencies. The present heavy lift effects on humans (refs. 11 and 12). These effects rotorcraft designs have fundamental blade passage will most likely be transient in nature for frequencies in the range of 9 to 12 Hz. FAA communities as vehicles pass nearby. Nonetheless, Certification noise limits and the NASA Rotorcraft these effects will have an influence on community Sector goals both are cast in terms of the Effective acceptance of the vehicles. Psycho-acoustic effects Perceived Noise Level (EPNL) noise metric. The will be more prominent for passengers and crew EPNL metric is computed using one-third octave because of long exposure times; these effects must bands, with the lowest band considered having a be quantified further.

center frequency of 50 Hz; this lowest band spans Under the same atmospheric conditions, low by averaging the levels at the three microphones for frequency noise propagates over longer distances a given condition, then averaging these three new than higher frequency noise because of different values directly over the three flight conditions. The rates of atmospheric attenuation for different NASA Rotorcraft sector goal is to reduce the frequencies. Therefore it is expected that heavy lift aircraft noise by 80%, hence a 14 EPNdB reduction vehicles will be heard or felt from much longer below the SOA level. Achieving this community distances. goal will require a combination of design for low noise, active control for noise source reduction, and The primary noise prediction tool applied to flight operational procedures.

quantify the external noise during the assessment task has been CARMA (ref. 8). CARMA is a The infrasonic range is typically defined as sound at collection of stand-alone analyses that are linked by frequencies below approximately 20 Hz; this is interface codes that facilitate information transfer. below the normal human audible range. High levels The CARMA system currently accomplishes a of infrasonic noise at frequencies of around 8 to 12 sequence of tasks that includes (1) comprehensive Hz create acoustic resonances in certain human analysis of a rotorcraft vehicle, including high body cavities and can vibrate internal organs. This resolution analysis of the rotor blade motion and noise is felt rather than heard until the levels aerodynamics; (2) noise prediction on a hemisphere become very high. At levels of approximately 130 near the vehicle (excluding atmospheric effects); dB, the infrasonic noise reaches audible levels and and (3) propagation of the noise from all approaches the threshold of pain. To avoid the hemispheres on a vehicle to distant observers acoustic range where the noise is felt, energy in (including atmospheric effects). frequencies less than approximately 20 Hz should be avoided. This places a constraint on the rotor CARMA was used to show how a heavy lift design.

rotorcraft can be examined in the context of the FAA Noise Certification requirements and in the This investigation has identified the importance of context of the NASA Rotorcraft Sector goals. For the effects of low frequency, high amplitude noise this sample assessment, three flight conditions are on humans and structures. This is a sensitive considered: level flight at 63 knots, descent at subject, economically, politically and legally. A 80 knots on a 6 deg glide slope, and take-off at lower limit in the design process on the blade 80 knots in a 6 deg climb. Microphones and passage frequency, at say 20 Hz, may be necessary.

flightpaths used were as specified in the FAA The blade passage frequency is 9.3, 10.8, and 11.3 Noise Certification guidelines. The certification Hz, for the current LCTR, LCTC, and LABC requirement is applied to the maximum level seen at designs respectively (all with a 650 ft/sec hover tip any of the three microphones. speed). Increasing the hover tip speed would increase the noise levels, and so probably higher Figure 33(a) shows the LCTR results compared to disk loading (smaller rotor diameter) and a larger the FAA Noise Certification requirements. The solid number of blades will be required. New or revised lines are the FAA noise limits. The LCTR level metrics will likely be required for heavy lift flight noise is higher than the FAA limit, but the rotorcraft noise assessment and certification. The descent and take-off conditions are lower than the NASA technology goal will also need to be revised limits. These calculations were restricted to 1/3- to fully address environmental impact of the noise octave bands (predicted tone noise converted to 1/3 spectrum for heavy lift rotorcraft.

octave), and broadband and other noise sources (such as engine and airframe) were not included; Yaw Control and EPNdB effectively filters out information below the 50 Hz 1/3-octave band. Figure 33(b) shows the Both the tiltrotor and the tandem compound utilize same information as in figure 33(a), but it was rotor inplane forces produced by cyclic pitch for low processed to match the NASA Rotorcraft Sector speed yaw control. The designs considered had goal metrics. The NASA goal metrics are computed hingeless rotors. Figure 34 shows the thrust vector tilt per degree of rotor cyclic as a function of blade c) mission fuel (lb) flap frequency from 1/rev (zero effective hinge d) purchase price ($M) offset) to 4/rev. The calculations were performed for e) direct operating cost DOC+I (cents/ASM) a flapping rotor in hover, with a Lock number of 8.0 and solidity of 0.075. Inflow gradients caused by The fuel weights in table 11 include the reserves.

hub moments, which can substantially change the Sensitivity studies were conducted using the sizing behavior at low thrust, were also included. With code, to optimize the designs by examining gimballed or low-hinge-offset flapping blades, variations in disk loading and number of blades. The cyclic control tilts the tip-path plane, which in turn influence of hover tip speed and cruise tip speed tilts the rotor thrust vector. Differential cyclic were also examined, considering in particular the control on the two rotors produces opposing in- noise requirements.

plane hub forces, and hence a yaw moment on the aircraft. With a hingeless rotor, the flapping Figure 36 shows the influence of disk loading and produced by cyclic pitch is reduced, but inplane hub number of blades on the LCTR design. The forces are still produced by cyclic. For 1/rev flap optimum disk loading is about 10 lb/ft . Indicated frequency, the response is 1 deg of thrust vector tilt on the figure is the disk loading above which the per degree of cyclic (fig. 34), and as the flap OEI requirement determines engine size, and below frequency increases, the thrust vector tilt per cyclic which cruise determines engine size. The optimum is reduced. At high thrust (levels required for yaw disk loading occurs where the OEI and cruise power control), the thrust vector tilt approaches 0.5 deg per requirements balance. The sizing code shows the degree of cyclic. metrics decreasing as the number of blades increases. However, with four blades the blade Figure 35 shows the hub moment, which for large aspect ratio is already reasonably large, and there is flap frequency approaches the value of the some concern that deicing system weights might aerodynamic hub moment produced by cyclic. The reverse the trend with blade number. Figure 37 variation of hub moment with collective (mean shows the influence of hover tip speed and number thrust) is a result of the influence of the wake (a lift of blades on the LCTR design. These design deficiency function).

parameters are expected to be influenced by the noise criteria. Figure 38 shows the influence of the Hence yaw moment capability with hingeless rotors number of engines on the LCTR design. The is reduced compared to a gimballed or low-hinge- increase in metrics with 2 or 3 engines reflects the offset articulated rotor, but at most by a factor of 2.

influence of the OEI requirement.

Yaw control requires a corresponding increase in cyclic control authority. With a hingeless rotor Figure 39 shows the influence of cruise tip speed on however, there is a substantial hub moment the LCTR design. For the 1-speed transmission, the associated with this yaw control. A hovering turn rotor-to-engine speed ratio is fixed, so as cruise tip might in fact be a critical case for the blade design.

speed is reduced so is the engine speed, with For a tiltrotor, introducing nacelle tilt may reduce eventually a significant increase in SFC. This is the these hub moments.

conventional approach for tiltrotor design, which leads to an optimum at a cruise tip speed of about SENSITIVITY STUDIES 85% of hover tip speed, as found here. For the 2-speed transmission results in figure 39, the engine Design Optimization is operated at optimum speed, regardless of the rotor speed, and it is assumed here that there is no Table 19 compares the baseline designs for the transmission weight or efficiency penalty from the LCTR, LCTC, and LABC, in terms of the following 2-speed capability. Rotor propulsive efficiency metrics: increases as the cruise speed is decreased, hence the optimum is at a cruise tip speed of about 350 ft/sec.

a) aircraft mission gross weight (lb) b) installed engine power (hp) Figure 40 addresses the question of a weight penalty where M is the aircraft (not isolated rotor) hover for the 2-speed transmission. The results in figure of merit, and L/D is the aircraft cruise lift-to- figure 40 for no weight penalty are the same as the drag ratio. In these equations, δ is the pressure ratio, 2-speed results in figure 39. Shown in figure 40 are θ the temperature ratio, L the engine lapse rate, and the metrics for a 20% transmission weight penalty. N the number of engines. Then e Figure 40 also shows a break-even drive system weight, established by increasing the transmission f = M / (L/D) eff weight until the operating cost equaled that for a

f = ( 1.33(N -1) ) / ( 0.9N ) = 1.1083

cruise tip speed of 85% hover tip speed (the OEI e e conventional approach). The difference between the

f = ( δ L( θ ) ) / ( δ L( θ ) ) = 1.3414

two curves for drive system weight in figure 40 is eng OEI OEI CR CR the penalty that can exist with the 2-speed 2 2 2

W/A = 2 ρ V [ f f f ] = 2948 f

transmission remaining a cost-effective choice.

OEI CR eff OEI eng eff with numerical values given for the OEI and cruise Figure 41 shows the influence of disk loading and conditions considered here. For example, M = 0.64 number of blades on the LCTC design. The and L/D = 11.1 (LCTR) gives a disk loading of optimum disk loading is about 15 lb/ft . As with the W/A = 9.8 lb/ft .

LCTR, the optimum disk loading occurs where the OEI and cruise power requirements balance.

Technology Payoff Figure 42 shows the influence of hover tip speed and number of blades on the LCTC design. These The impact and payoff of advanced technology were design parameters are expected to be influenced by quantified using the sizing code. For this purpose, the noise criteria. Figure 43 shows the influence of the technology factors were changed from values the number of engines on the LCTC design.

representing advanced technology to values representing current technology. The technology For the LABC, the sizing code models the rotor factors for weights are given in table 7, and the performance in terms of the rotor L/D, which is engine model is described in table 8. Table 20 and obtained from the comprehensive analysis figures 44–46 show the percentage increase calculations of performance. Disk loading for the (a negative value is good) in the five metrics, caused LABC was therefore determined by iterating by removal of various aspects of the advanced between RC and CAMRAD II, not by simply technology from the design assumptions. Results are varying disk loading in RC as for the LCTR and given for the LCTR and LCTC, but not for the LCTC (figs. 36 and 41).

LABC because of the level of maturity of the sizing code for that configuration. Figure 44(a) shows the Disk Loading impact of rotor blade and hub weight reduction, and figure 44(b) shows the impact of all structural Assuming a balance of OEI and cruise power weight reductions (blade, hub, fuselage, and wing).

requirements for the design permits a simple Individually the hub weight, fuselage weight, and estimate of the disk loading. The engine power wing weight had small influence; collectively they available as a function of lapse rate, and the power contribute the significant influence shown in required for OEI and cruise are: figure 44(b). Figure 44(c) shows the impact drive system weight reduction. Figures 45(a) and 45(b) engine: P = P δ L( θ ) show the impact of drag reductions, and figure 45(c) shows the impact of all aerodynamics (drag, rotor OEI: 1.33(N -1)P = 0.9P e OEI H figure of merit and cruise efficiency, and 1/2 download). Individually the hover figure of merit,

hover: P = W ( DL/2 ρ ) / M

H OEI cruise efficiency (propulsive efficiency for LCTR, cruise: N P = WV / L/D rotor drag for LCTC), and download had small e CR influence; collectively they contribute the significant influence shown in figure 45(c). Figure Designs were developed for the tiltrotor (LCTR), 46(a) shows the major impact engine technology has advancing blade concept (LABC), and slowed rotor on the designs. Increases in vibration treatment compound with tandem main rotors (LCTC). In weight and acoustic treatment weight were also addition, a slowed rotor compound with a single examined, and found to have a small impact on the main rotor was examined, for both shaft drive metrics. (LCSC) and reaction drive (LRDC).

A conservative design approach, based on past The comprehensive analysis was used to optimize aircraft design experience, would increase the the rotor performance. Table 22 shows the aircraft estimated power required (and hence fuel burned) characteristics assumed for the performance by 25%, and increase the estimated empty weight by optimization, and the results of the optimization for 15%, for a fixed payload and performance the three primary configurations. All cases used requirement. The penalty for imposing these weight current technology rotor airfoils with Reynolds and power contingencies is shown in figure 46(b). number corrections for the drag (and stall delay for Since the need for large contingencies is attributed the tiltrotor). For the slowed-rotor compound to lack of accuracy of current design and analysis configurations, rotor/rotor and rotor/wing tools, figure 46(b) shows the economic payoff interference were included in the comprehensive possible by improving these tools. analysis model (otherwise the tandem and single main rotor configurations would have identical Figure 47 shows the costs for the LCTR with and performance). The comprehensive analysis results without the cost technology factors, and figure 48 were used to estimate aircraft L/D = WV/P as a presents the corresponding DOC+I breakdown. function of flight speed (fig. 49). The relative These results emphasize and quantify the efficiency of the configurations was the same for all importance of controlling the maintenance costs for conditions. For the slowed rotor compound, the heavy lift rotorcraft. rotor/rotor interference resulted in a small reduction in aircraft L/D for the tandem configuration Alternate Missions compared to the single main rotor. The differences between powered (optimum thrust) and autorotating The three configurations were also sized for an operation were also examined, with the latter giving alternate mission, composed of three 400 nm slightly worse performance. Also note that the segments (takeoff, climb, cruise at 30k, descent, and LABC efficiency improves as the altitude decreases landing; with one reserve segment), instead of a (the higher density making it easier for the rotor to single 1200 nm segment. Table 21 compares the generate the required lift), while the efficiency aircraft designed for the baseline and alternate improves as altitude increases for the other mission. The additional climb and descent time in configurations (which use the fixed wing for lift in the 3x400 mission resulted in heavier aircraft cruise). Significant improvements in the calculated carrying more fuel. The relative efficiency of the L/D are possible using specially designed airfoils, three configurations remained unchanged. particularly for the LABC, which will also benefit from further optimization of the planform.

To explore the influence of the design condition on the comparative performance of the heavy lift Spreadsheets were developed to produce designs rotorcraft configurations, the performance incorporating the comprehensive analysis perfor- optimization and aircraft sizing were performed for mance (fig. 49), component weight calculations, and the following alternate design cruise conditions: mission analysis. The results are given in tables 23 to 27. For the tiltrotor, advancing blade, and a) 30k/ISA and 350 knots (baseline) compound configurations respectively, the disk b) 20k/ISA and 350 knots loading was 10, 25, and 15 lb/ft ; the drag was 2/3 c) 20k/ISA and 250 knots D/q/(W/1000) = 1.5, 1.3, 1.9; the thrust-weighted d) 10k/ISA and 250 knots C / σ = 0.14, 0.10, 0.14. The performance W o e) 5k/ISA+20 C and 250 knots calculations shown in figure 49 used a disk loading 2 2 of 20 lb/ft for the LABC, but 25 lb/ft produced cost, efficiency, and productivity (tables 23–27). All lower weight and cost. For all configurations the of the configurations would benefit from improved hover tip speed was 650 ft/sec; wing loading was 80 airfoils. A detailed analysis of the reaction drive is lb/ft ; power turbine efficiency was 83% needed for a more accurate estimate of the weight (considered a conservative value); body weight was and power compared to shaft drive systems. A more fixed at the values from the baseline designs accurate estimate of the LABC weight and power (table 11); and fuel price for DOC+I calculations requires a better analysis of rotor and hub, specially was $5.00/gal. Note that the spreadsheet design designed airfoils, and further optimization of the process is not as sophisticated as the RC sizing planform.

code, so the results in table 23 are slightly different from those in tables 10 and 21.

ASSESSMENT OF CONFIGURATIONS For the LCTR, the hover figure of merit and cruise propulsive efficiency were set to the values from the For the NASA civil mission, the LCTR had the best comprehensive analysis, for 350 ft/sec tip speed and cruise efficiency, hence the lowest weight and optimized twist (from table 22). The wing efficiency lowest cost. The LCTR is the configuration with the was estimated to be 1.00. The transmission weight most promise to meet the NASA technology goals.

fraction was fixed at the value from the baseline design (table 11).

The LCTC had good cruise efficiency, but less than the tiltrotor, and higher development risk than the For the LABC, the hover figure of merit and cruise tiltrotor. Single main rotor and tandem rotor rotor effective L/D were set to the values from the configurations were comparable in efficiency and comprehensive analysis, for optimized M , at risk. Even if reaction drive produced the smallest collective, twist, and lift offset (table 22). The slowed-rotor compound rotorcraft, the high installed propeller propulsive efficiency was set to 0.90. The power compromises efficiency, and the reaction transmission weight fraction was fixed at the value drive system has higher noise and substantially from the baseline design (table 11).

increased risk.

For the slowed-rotor compounds, the hover figure of The LABC had lower cruise efficiency than the merit and rotor drag D/qA were set to the values tiltrotor for the NASA civil mission, and higher from the comprehensive analysis, for optimized M , at development risk than the tiltrotor.

collective, and twist (table 22). In addition, the wing induced efficiency and the fraction of weight carried The LCTR design presented was economically by the wing were set to the values from the competitive with comparable fixed wing aircraft, comprehensive analysis. The propeller propulsive with the potential for substantial impact on the air efficiency was set to 0.90. It was assumed that 40% transportation system. The keys to achieving a of the tandem transmission weight was the propeller competitive aircraft are: low drag airframe and low gearbox (which is the same for all the compound disk loading rotors; structural weight reduction, for configurations); and that the rotor transmission both airframe and rotors; drive system weight weight for the single main rotor and reaction drive reduction; improved engine efficiency; low were respectively 67% and 20% of the tandem. Thus maintenance design; and manufacturing cost the transmission weight fraction for the tandem, comparable to CTOL aircraft.

single rotor, and for reaction drive cases was fixed at respectively 100%, 80%, and 50% of the value Thus the LCTR design demonstrated the potential from the baseline design (table 11). The reaction for achieving the Rotorcraft Sector goals of tables 1 drive efficiency used for the hover performance was and 2. With a disk loading of 10 lb/ft compared to 50%.

the state-of-the-art value of 20 lb/ft , the 40% increase in hover efficiency was attained.

For all of the design operating conditions Considering the OEI hover power (power from 3 out considered, the tiltrotor configuration had the best of 4 engines), the power loading was W/P = 6.0. At The risk reduction plans parallel a prototype the cruise conditions, the aircraft lift-to-drag ratio development program, and feed design solutions at was L/D = 14.5 (table 16), exceeding the 44% high technology readiness levels (TRL) to the improvement goal. The airframe drag was estimated prototype program. Although a prototype is not part 2/3 to be D/q = 1.5/(W/1000) (table 13). The weight of the NASA plan, it serves to provide schedule pull technology factors (table 7) led to about 22% for the required tasks. The risk reduction program reduction in gross weight (table 20), from a 30% elements are technology readiness level reduction in empty weight, which was consistent benchmarks, program tasks, and strategic direction.

with the goal of a 25% reduction in empty weight Technology readiness level benchmarks are major excluding engines. The design had an empty weight milestones demonstrating significant increases in fraction of 0.65 (table 11), or 0.62 excluding TRL, principally by integrated tests of hardware.

engines, so technology countered the growth in There are two primary thrusts: early technology empty weight fraction with aircraft size and speed. leads to advanced technology and supports a The calculated noise was 9.3 EPNdB below possible year 7 prototype; advanced technology certification requirements (fig. 33), compared to the leads to a year 10 prototype.

goal of 14 EPNdB, with active control and flight operations available to obtain the full reduction as Tasks were identified for each of the four risk areas.

well as deal with low frequency noise. The tasks were organized by discipline (propulsion, structures, aeromechanics, acoustics, and handling qualities) although all elements are connected because of the interdisciplinary nature of rotorcraft RISK REDUCTION FOR HEAVY LIFT problems. Detailed task descriptions have been ROTORCRAFT developed, including schedules showing ground test, wind tunnel test, flight test, and decision milestones, The NASA Heavy Lift Rotorcraft Systems together with connections to the TRL benchmarks.

Investigation was a focused and coordinated The tasks constitute the work required to achieve the analytical effort to select the best configuration for TRL benchmarks. The strategic direction provides meeting the Rotorcraft Sector vehicle technology guidance for selecting highest priority activities, goals. During the course of the investigation, high aimed at the highest risk areas of heavy lift risk areas were identified. The definition of high rotorcraft development.

risk is one or both of the following: capability or attribute unavailable today, so it is necessary to The risk reduction plan has the following TRL assume advanced technology will be available in the benchmarks.

future in order for the aircraft to achieve the technology goals; or cost prevents the vehicle from a) Full-scale propulsion system ground test: being economically competitive, so the payoff of complete transmission and engine arrangement; advanced technology is essential to achieving the metrics are weight, cost, and noise.

goals. The following were identified as high risk areas for heavy lift rotorcraft: b) Full-scale structure ground test: rotor blade and hub, airframe components, wing components; a) High torque, lightweight drive system.

metrics are weight, cost, and interior noise.

b) High performance, structurally efficient rotor/ wing system.

c) Large-scale rotor system wind tunnel test: c) Low noise aircraft.

dynamically scaled rotor and hub; metrics are d) Super-integrated vehicle management system.

performance, loads, vibration, control, and noise.

Plans were then developed to mitigate the above risks. The risk reduction plans provide the strategic d) Flight simulation test: including elastic airframe direction to support a heavy-lift rotorcraft and load control system; metrics are handling development.

qualities, control, and noise.

e) Noise and control flight test: existing aircraft, established, and the control method selected (IBC, with rotor active control and low-noise or on-blade, or airframe).

operations; metrics are noise, vibration, control, handling qualities. Design guidelines for noise must be established (such as minimum blade-passage frequency), based f) Integrated large-scale wind tunnel test: utilize on physiological and psychological effects on propulsion, structure, rotor systems from earlier human response. Low noise concepts and benchmarks; metrics are system integration.

approaches must be identified (design features, active control, flight operations). Certification and Concept Development community noise impact requirements must be defined. The requirements and certification The first year of the risk reduction program is also approach for handling qualities must be established, devoted to further refinement of the rotorcraft including one-engine inoperative.

concept. Decisions must be made to narrow the focus of the program elements, since parallel The technology contributions to reduction of research lines cannot be afforded. It is essential to purchase price and maintenance cost must be focus the work on areas of high payoff for heavy lift identified. A public benefit model and reality-based rotorcraft. This concept development should be cost model should be developed.

conducted by several companies, for the selected heavy lift rotorcraft configuration (which for the NASA mission is the LCTR). The first phase would STRATEGIC DIRECTION be completed within 12 months. A second phase would be completed in year 5, for the advanced The strategic direction provides guidance for technology thrust. A number of LCTR configuration selecting highest priority activities, aimed at the features require decisions: the wing (high or low), four highest risk areas of heavy lift rotorcraft nacelle (engine tilt or not), tail, and number of development. Note that there are some important engines.

and difficult tasks that are not high risk, including rotor aerodynamic design and optimization, rotor For the propulsion system, it is first necessary to and wing airfoil design, airframe aerodynamics, and establish what engine technology is available from airframe structures.

the engine manufacturer, in particular the possibility for maintaining good SFC over a wide engine rpm High Torque, Lightweight Drive System range (by design point, or variable geometry, or a multi-speed output shaft gearbox). The required Innovative design is required for low drive system engine development must be defined. Then the weight. Large size implies high torque and high transmission requirements can be defined, including weight fraction, hence drive system weight the system configuration.

reduction is essential for an efficient and economical aircraft. The focus must be on design Structural design concepts must be identified for concept, advanced-technology components, and lightweight, heavy lift rotorcraft, specifically for the materials.

blade and hub, for the airframe, and for the wing. A blade and hub concept must be identified that first is Low maintenance is required for low operating cost.

a solution to the stability issues (whirl flutter for the Low maintenance must be a primary design LCTR), then is a solution for the strength and requirement, even ahead of weight and performance.

weight requirements. The requirement for active control of loads must be established, and the High flight speed requires, or at least benefits from, approach defined (flight condition limiting, active a variable speed propulsion system design. First it is load control, others). The requirement for control of necessary to establish the speed range available noise, vibration, gust, and performance must be from advanced engine technology, and to define the in the rotor and airframe. Hence an expanded engine required for the heavy lift rotorcraft concept. integration of the vehicle management system is required: a flight control system for good handling High Performance, Structurally Efficient qualities and gust response, active control of Rotor/Wing System vibration and noise, and rotor load limiting and active control. The focus must be on load limiting Innovative rotor and wing design is required, and system integration.

probably with unconventional dynamics. Large size implies high weight fraction, high speed introduces stability issues, and good rotor system performance CONCLUSION is essential for an efficient and economical aircraft.

The focus must be on integrated rotor/wing The NASA Heavy Lift Rotorcraft Systems performance and dynamic behavior.

Investigation examined in depth several rotorcraft configurations for large civil transport, designed to Structural efficiency is required for low rotor and meet the technology goals of the NASA Vehicle hub and wing weight. The focus must be on design Systems Program. Design and analysis tools were concepts for durability and damage tolerance.

applied to define three configurations: Large Civil Tiltrotor (LCTR), Large Civil Tandem Compound Low maintenance is required for low operating cost.

(LCTC), and Large Advancing Blade Concept Low maintenance must be a primary design (LABC).

requirement, even ahead of weight and performance.

For the NASA civil mission, the Large Civil Low Noise Aircraft Tiltrotor had the best cruise efficiency, hence the lowest weight and lowest cost. Thus the LCTR is New approaches are required to meet the challenge the configuration with the best potential to meet the of low noise. Large size implies low frequency NASA technology goals. The design presented was noise and expanded acoustic footprint. An economically competitive, with the potential for understanding of heavy lift vehicle acoustic substantial impact on the air transportation system.

phenomena (low frequency and relative distance to While fixed wing aircraft for this mission exist, the community) is required, including psychoacoustics investigation showed only the potential for a high for low frequency. New rotor design guidelines and speed, heavy lift rotorcraft. The keys to achieving a annoyance metrics must be developed. The focus competitive aircraft were low drag airframe and low must be on a combination of rotor design, active disk loading rotors; structural weight reduction, for control, and flight operations.

both airframe and rotors; drive system weight reduction; improved engine efficiency; low Super-Integrated Vehicle Management maintenance design; and manufacturing cost System comparable to fixed-wing aircraft.

Broad spectrum active control is required for an Risk reduction plans were developed to provide the effective heavy lift rotorcraft. Large size implies a strategic direction to support a heavy-lift rotorcraft significant influence of low frequency airframe development. The following high risk areas were elastic modes on flight dynamics. Active control is identified for heavy lift rotorcraft: high torque, required to achieve the goals of low rotor-induced lightweight drive system; high performance, vibration and noise. Safe operation in one-engine structurally efficient rotor/wing system; low noise inoperative conditions is essential for civil aircraft; and super-integrated vehicle management rotorcraft. Rotor load limiting and active control are system.

needed for full utilization of the structural capability 9. Harris, F. D.; and Scully, M. P.: Rotorcraft Cost REFERENCES Too Much. Journal of the American Helicopter Society, Vol. 43, No. 1, January 1. Johnson, J.; Stouffer, V.; Long, D.; and Gribko, 1998.

J.: Evaluation of the National Throughput Benefits of the Civil Tiltrotor. NASA/CR– 10. Harris, F. D.: An Econonic Model of U.S.

2001-211055, September 2001. Airline Operating Expenses. NASA/CR– 2005-213476, December 2005.

2. Stouffer, V.; Johnson, J.; and Gribko, J.: Civil Tiltrotor Feasibility Study for the New 11. Kryter, K. D: The Effects of Noise on Man.

York and Washington Terminal Areas. (Second Edition), Academic Press, Inc., NASA/CR–2001-210659, January 2001. New York, NY, 1985.

3. Smith, D. E.; Wilkerson, J.; Montoro, G. J.; 12. Burns, W.: Noise and Man (Second Edition), Coy, J.; and Zuk, J.: Technology J. B. Lippincott Company, 1973.

Development for Runway Independent Aircraft. Proceedings of the American Helicopter Society 59th Annual Forum, Phoenix, AZ, May 2003.

4. Preston, J.; and Peyran, R.: Linking a Solid- Modeling Capability with a Conceptual Rotorcraft Sizing Code. American Helicopter Society Vertical Lift Aircraft Design Conference, San Francisco, CA, January 2000.

5. Johnson, W.: Rotorcraft Aeromechanics Applications of a Comprehensive Analysis.

HeliJapan 98: AHS International Meeting on Advanced Rotorcraft Technology and Disaster Relief, Gifu, Japan, April 1998.

6. Drela, M.: Newton Solution of Coupled Viscous/Inviscid Multielement Airfoil Flows. AIAA Paper No. 90-1470, 21st Fluid Dynamics, Plasma Dynamics, and Lasers Conference, Seattle, WA, June 1990.

7. Rajagopalan, R. G.: A Procedure for Rotor Performance Flowfield and Interference: A Perspective. AIAA Paper No. 2000-0116, 38th Aerospace Sciences Meeting, Reno, NV, January 2000.

8. Boyd, Jr., D. D.; Burley, C. L.; and Conner, D.

A.: Acoustic Predictions of Manned and Unmanned Rotorcraft Using the Comprehensive Analytical Rotorcraft Model for Acoustics (CARMA) Code System. AHS International Specialists' Meeting on Unmanned Rotorcraft Design, Control, and Testing, Chandler, AZ, January 2005.

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 efficiency, 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 Noise and Vibration 77dBA & 0.05g TABLE 2. ROTORCRAFT SECTOR GOALS, OBJECTIVES, TECHNICAL CHALLENGES, AND APPROACH.

GOALS OBJECTIVES TECHNICAL CHALLENGES APPROACHES Hover Efficiency Increase overall hover Reduce rotor power with Design large, lightweight (W/P = 6, SOA = 4.3) efficiency by 40% for large minimal weight impact and aeroelastically stable rotor rotorcraft acceptable rotor dynamics systems Efficient Cruise Reduce drag by 44% at a Reduce drag without adversely Investigate novel rotor (L/D = 12, SOA = 6.5) cruise Mach of 0.60 affecting performance, noise, and configurations enabling vibration high-lift and high-speed cruise w/low noise and vibration Optimize airframe propulsion integration to reduce interference drag Empty Weight Fraction Reduce rotor weight by 25% Reduce rotor weight fraction as Develop lightweight drive (excluding engines, vehicle size doubles systems and transmission Reduce subsystem by 25% We/W = 0.41, SOA = 0.55) concepts (drive system, high-lift, etc.) Reduce subsystem empty weight fraction as vehicle size doubles Reduce airframe structural weight by 25% Reduce Community Noise Reduce rotorcraft system Reduce RC systems source noise Develop and validate source (SOA – 14EPN dB) noise by 80% without degrading performance noise prediction and propagation capabilities Reduce noise through flight capabilities and operational Develop and validate system procedures that are safe & noise prediction capabilities certifiable with acceptable Develop and demonstrate community impact low noise operations and capabilities with acceptable handling qualities TABLE 3. COMPARISON OF CTOL AND VTOL COST MODEL RESULTS (BOTH FOR BOEING 737 AIRCRAFT).

737-700 run using CTOL cost model VTOL cost model Flyaway Cost, $ (US1999) $ 48.0 M $ 83.6 M Total Direct Operating Cost + Interest 6.8 18.9 (Cents/ASM, US 1999) Maintenance 0.9 9.8 Airframe 0.6 4.9 Engine 0.3 2.9 Rotor & Drive System n/a 2.0 Flight Crew Salary & Expense 0.9 0.9 Fuel & Oil 0.9 0.9 Depreciation 2.4 4.1 Insurance Cost 0.2 0.3 Finance Cost 1.6 2.8 TABLE 4. 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 5. 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 6. 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 antisymmetric 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 7. 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 8. ENGINE TECHNOLOGY USED FOR AIRCRAFT SIZING.

Current Advanced technology 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 ratio lb/hr 0.781 0.781 Fuel flow at 30k ISA ratio lb/hr 0.334 0.334 TABLE 9. ADVANCED TECHNOLOGY ESTIMATES.

LCTR LCTC LABC Tiltrotor Tandem Advancing blade compound 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 — Optimum Disk loading, W/A lb/ft 10 15 25 Maximum M 0.70 0.80 0.85 at Cruise tip speed ft/sec 350 205 255 TABLE 10. HEAVY LIFT ROTORCRAFT DESIGNS.

LCTR LCTC LABC Tiltrotor Tandem compound Advancing blade Mission gross weight (lb) 123562 138764 160636 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 – aspect 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 11. 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 Structure 5545 9206 0 Fairings 1111 1442 0 Fittings 1349 1349 0 Surface Controls 797 0 0 Primary Thruster 13714 11494 24572 Blades 6968 5863 11204 Hub & Hinge 6498 5631 13369 Fairing / Spinner 248 0 0 Tail / Aux Thrust 594 2870 4135 Body Group 7072 10194 11596 Landing Gear Group 3228 3625 4197 Nacelle 2422 1091 2411 Nacelle (Engine Support) 495 678 1055 External Cowling 296 413 1355 Pylon Support Structure 1631 0 0 Air Induction 270 397 617 Propulsion 18373 24928 29021 Engine installation 4540 6446 9284 Engine 3698 5179 7630 Exhaust System 683 956 1409 Accessories 160 311 245 Fuel System 556 957 887 Drive System 13277 17525 18850 Gearbox & Rotor Shaft 11880 16602 18397 Gearboxes 10336 14444 16340 Rotor Shaft(s) 1544 2158 2058 Flight Controls 4927 4628 5238 Rotary Wing Controls 1211 3242 3805 Non-Boosted 144 234 215 Boost Mechanism 452 398 478 Boosted 614 2609 3112 Other Scaled Weight 7716 6273 8621 USEFUL LOAD 42860 47684 56645 Crew 1110 1110 1110 Fixed Useful Load 100 100 100 Fluids 240 240 240 Fuel 15010 19834 28795 TABLE 12. FIXED EMPTY WEIGHT SUMMARY (SAME FOR ALL CONFIGURATIONS).

FIXED WEIGHT 13583 Flight Controls 210 Cockpit Controls 75 Electronics and Sensors 135 Equipment 13373 Aux Power Unit (APU) 600 Instruments 150 Utility Hyd & Pneumatics 150 Electrical System 1500 Avionics (MEP) 800 Furnishings & Equipment 8,588 Air Conditioning 1,200 Ice Protection (fixed) 100 Aircraft Handling 100 Load Handling 185 Furnishings & Equipment 8588 Seats (crew and passenger) 2656 Galleys 515 Lavatory (3) 495 Floor Covering 375 Panels, Doors, Partitions 3872 Emergency Equipment 674 Portable Oxygen 174 Fire Extinguishers 375 Escape Provisions 125 TABLE 13. 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/1000) 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 14. COST PARAMETERS.

VTOL CTOL (B737) Block hour per year hr 3750 3750 Non-flight time per trip min 12 45 Flight speed knots 350 460 Number of seats 120 137 Fuel price $/gal 5.00 5.00 TABLE 15. HEAVY LIFT ROTORCRAFT CHARACTERISTICS.

LCTR LCTC LABC Design gross weight (lb) 123562 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 Solidity (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 0.333 Blade twist (deg), inboard of 50%R –32 0 0 Blade twist (deg), outboard of 50%R –30 –12 –10 Lock number 12.1 13.0 19.1 Single blade weight (lb), from blade structural design 745 646 1080 Total blade weight (lb), all rotors 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 16. 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 248 256 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 17. 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 23803 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 18. 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 170476 159068 159049 C / σ 0.0952 0.0955 1.2468 1.2467 T Rotor shaft power 32497 39613 0 -220 Induced power 30482 37545 11668 Profile power 2015 2068 11910 Ind+pro power 14251 23577 κ 1.186 1.454 3.414 ind c 0.0076 0.0078 0.0278 do Figure of merit 0.791 0.652 Drag 13268 22156 Rotor L/D 12.0 7.2 e L/D = WV/P 7.2 5.0 TABLE 19. 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 80.0 110.5 DOC (cents/ASM) 13.3 17.2 23.9 TABLE 20. 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 12 All structural weight 21 22 18 27 21 Drive system weight 23 24 21 29 23 Airframe drag (+25%) 6 14 14 12 12 All aerodynamics 10 20 21 17 16 Engine technology 23 28 70 28 46 Weight and power contingency 13 25 25 22 19 LCTC Blade weight 9 9 8 11 9 All structural weight 20 20 18 24 20 Drive system weight 29 28 26 35 39 Airframe drag (+25%) 5 10 13 8 10 Hub drag 9 22 24 17 20 All aerodynamics 17 38 39 31 33 Engine technology 48 63 124 60 87 Weight and power contingency 13 25 25 22 20 TABLE 21. DESIGNS FOR BASELINE (1200 nm SEGMENT) AND ALTERNATE (THREE 400 nm SEGMENTS) MISSION.

LCTR LCTC LABC Baseline Alternate Baseline Alternate Baseline Alternate Gross weight (lb) 123562 138318 138764 155540 160636 189690 Engine power (hp) 4x6914 4x7754 4x9684 4x10819 4x14267 4x16566 Mission fuel (lb) 13624 19725 17902 24894 26008 37021 Purchase price ($M) 61.9 69.5 80.0 89.5 110.5 130.0 DOC (cents/ASM) 13.3 17.7 17.2 21.9 23.9 32.0 Rotor diameter (ft) 88.7 93.8 76.7 81.2 90.4 98.3 Disk loading (lb/ft ) 10 10 15 15 25 25 o C / σ (T-wt, 5k ISA+20 C) 0.141 0.141 0.141 0.141 0.090 0.090 W Number of blades per rotor 4 4 4 4 5 5 Chord (75%R, ft) 3.06 3.24 3.98 4.22 4.89 5.31 Wing loading (lb/ft ) 80 80 80 80 Drag, D/q (ft ) 37.3 40.7 50.3 55.0 38.1 42.23 TABLE 22. ROTORCRAFT CHARACTERISTICS FOR PERFORMANCE OPTIMIZATION AT ALTERNATE DESIGN CONDITIONS.

(a) Parameters independent of design conditions.

LCTR LCTC LABC Disk loading, lb/ft 10 15 20 2/3 Total cruise drag, D/q / (W/1000) 1.5 1.9 1.3 o Hover C / σ (5k ISA+20 C) 0.141 0.141 0.100 W Number of blades 4 6 6 Solidity (thrust weighted) 0.0881 0.1321 0.1239 Blade taper 0.8 0.8 0.333 Hover tip speed, ft/sec 650 650 650 Cruise tip speed, ft/sec 350 Wing loading, lb/ft 80 80 (b) Parameters dependent on design speed and altitude.

Design speed, knots 250 250 250 350 350 o Design altitude, ft 5k+20 C 10k 20k 20k 30k LCTR Optimum Blade twist, deg inboard –46 –46 –46 –32 –32 outboard –34 –34 –34 –30 –30 LCTC Optimum Maximum M 0.60 0.65 0.65 0.85 0.85 at Blade twist, deg inboard 0 0 0 0 0 outboard –12 –12 –12 –12 –12 LABC Optimum Collective, deg 4 4 6 1 0 Maximum M 0.80 0.80 0.80 0.85 0.85 at Blade twist, deg inboard 0 0 0 0 0 outboard –10 –10 –10 –10 –10 Lift offset 0.4R 0.4R 0.4R 0.25R 0.2R TABLE 23. HEAVY LIFT ROTORCRAFT DESIGNS FOR CRUISE AT 350 KNOTS, 30,000 FT ALTITUDE.

LCTR LABC LCTC LCSC LRDC Configuration Tiltrotor Advancing Tandem Single Single comp., blade compound compound reaction drive Mission gross weight (lb) 121640 158484 140233 139088 146934 Engine size (hp) 4x6923 4x18142 4x11094 4x12551 4x19809 Rotor diameter (ft) 88.00 89.84 77.15 108.66 111.68 Disk loading W/A (lb/ft ) 10 25 15 15 15 o C / σ (geom, 5k ISA+20 C) 0.1328 0.0675 0.1328 0.1328 0.1328 W o C / σ (T-wt, 5k ISA+20 C) 0.1406 0.0900 0.1406 0.1406 0.1406 W Wing loading (lb/ft ) 80 80 80 80 Wing span (ft) 104.25 145.03 144.44 148.46 Hover figure of merit 0.786 0.698 0.708 0.708 0.700 Cruise performance Propulsive efficiency 0.811 0.900 0.900 0.900 0.900 Drag D/q (ft ) 36.82 38.07 51.28 51.01 52.91 Wing induced drag (ft ) 18.00 11.16 10.05 10.62 Oswald efficiency 1.00 0.80 0.82 0.82 Wing lift (%W) 100% 85% 82% 82% Rotor drag (ft ) 23.84 22.35 25.96 Rotor L/D 7.62 Cruise power (hp) 11268 31874 15980 15446 16572 Weights Weight empty (lb) 79870 95172 92753 91715 96308 Operating weight empty (lb) 81320 96622 94203 93165 97758 Payload (lb) 26400 26400 26400 26400 26400 Fuel weight (lb) 13920 35462 19629 19524 22776 Rotor group weight (lb) 13432 13811 11655 13567 16301 Blade weight (lb) 6831 7481 5941 6383 8285 Hub weight (lb) 6357 6330 5714 7184 8016 Propulsion group weight (lb) 18049 31300 26050 23354 23464 Engine dry weight (lb) 3703 9703 5933 6713 10594 Transmission weight (lb) 13018 18543 17711 14053 9279 Flyaway cost (2005) $66.6M $131.8M $94.3M $93.8M $118.4M DOC+I (¢/asm) (2005) 13.66 29.76 18.91 18.90 22.77 Aircraft L/D=WV/P 11.60 5.34 9.43 9.67 9.52 Productivity, V*PL/(W-PL) 97 70 81 82 77 TABLE 24. HEAVY LIFT ROTORCRAFT DESIGNS FOR CRUISE AT 350 KNOTS, 20,000 FT ALTITUDE.

LCTR LABC LCTC LCSC LRDC Configuration Tiltrotor Advancing Tandem Single Single comp., blade compound compound reaction drive Mission GW (lb) 130087 159669 156468 155490 169013 Engine size (hp) 4x7404 4x16496 4x12378 4x14031 4x22785 Rotor diameter (ft) 91.00 90.18 81.49 114.88 119.78 Disk loading W/A (lb/ft ) 10 25 15 15 15 o C / σ (geom, 5k ISA+20 C) 0.1328 0.0750 0.1328 0.1328 0.1328 W o C / σ (T-wt, 5k ISA+20 C) 0.1406 0.1000 0.1406 0.1406 0.1406 W Wing loading (lb/ft ) 80 80 80 80 Wing span (ft) 107.25 153.20 152.72 159.22 Hover figure of merit 0.786 0.698 0.708 0.708 0.700 Cruise performance Propulsive efficiency 0.797 0.900 0.900 0.900 0.900 Drag D/q (ft ) 38.51 38.26 55.17 54.94 58.08 Wing induced drag (ft ) 9.59 6.43 5.68 6.17 Oswald efficiency 1.00 0.80 0.82 0.82 Wing lift (%W) 100% 87% 83% 83% Rotor drag (ft ) 23.37 21.66 25.92 Rotor L/D 9.97 Cruise power (hp) 14326 29206 22411 21702 23781 Weights Weight empty (lb) 84292 97170 101210 100295 108081 Operating weight empty (lb) 85742 98620 102660 101745 109531 Payload (lb) 26400 26400 26400 26400 26400 Fuel weight (lb) 17945 34649 27408 27345 33083 Rotor group weight (lb) 14675 16548 13475 15736 19624 Blade weight (lb) 7433 9340 6819 7343 9880 Hub weight (lb) 6981 7208 6657 8392 9744 Propulsion group weight (lb) 19302 30375 29066 26108 26990 Engine dry weight (lb) 3960 8823 6620 7504 12186 Transmission weight (lb) 13919 18681 19762 15711 10673 Flyaway cost (2005) $71.1M $125.8M $104.9M $104.5M $135.7M DOC+I (¢/asm) (2005) 15.99 28.84 23.56 23.60 29.27 Aircraft L/D=WV/P 9.75 5.87 7.50 7.70 7.63 Productivity, V*PL/(W-PL) 89 69 71 72 65 TABLE 25. HEAVY LIFT ROTORCRAFT DESIGNS FOR CRUISE AT 250 KNOTS, 20,000 FT ALTITUDE.

LCTR LABC LCTC LCSC LRDC Configuration Tiltrotor Advancing Tandem Single Single comp., blade compound compound reaction drive Mission GW (lb) 123474 155346 138116 139076 151623 Engine size (hp) 4x6948 4x16050 4x10926 4x12550 4x20441 Rotor diameter (ft) 88.66 88.95 76.56 108.65 113.45 Disk loading W/A (lb/ft ) 10 25 15 15 15 o C / σ (geom, 5k ISA+20 C) 0.1328 0.0750 0.1328 0.1328 0.1328 W o C / σ (T-wt, 5k ISA+20 C) 0.1406 0.1000 0.1406 0.1406 0.1406 W Wing loading (lb/ft ) 80 80 80 80 Wing span (ft) 104.91 143.94 144.43 150.81 Hover figure of merit 0.795 0.698 0.708 0.708 0.700 Cruise performance Propulsive efficiency 0.825 0.900 0.900 0.900 0.900 Drag D/q (ft ) 37.19 37.57 50.77 51.00 54.03 Wing induced drag (ft ) 34.69 20.09 20.46 22.30 Oswald efficiency 1.00 0.81 0.82 0.82 Wing lift (%W) 100% 84% 85% 85% Rotor drag (ft ) 19.98 19.84 23.75 Rotor L/D 9.29 Cruise power (hp) 7537 17867 8731 8776 9620 Weights Weight empty (lb) 80775 95010 91656 91708 98795 Operating weight empty (lb) 82225 96460 93106 93158 100245 Payload (lb) 26400 26400 26400 26400 26400 Fuel weight (lb) 14848 32486 18610 19518 24978 Rotor group weight (lb) 13700 15890 11422 13565 16993 Blade weight (lb) 6961 8963 5828 6382 8619 Hub weight (lb) 6491 6927 5594 7183 8375 Propulsion group weight (lb) 18268 29552 25657 23352 24213 Engine dry weight (lb) 3716 8584 5844 6712 10932 Transmission weight (lb) 13212 18176 17444 14052 9575 Flyaway cost (2005) $67.1M $122.5M $93.0M $93.8M $122.0M DOC+I (¢/asm) (2005) 16.96 32.29 22.02 22.68 28.97 Aircraft L/D=WV/P 12.57 6.67 12.14 12.16 12.09 Productivity, V*PL/(W-PL) 68 51 59 59 53 TABLE 26. HEAVY LIFT ROTORCRAFT DESIGNS FOR CRUISE AT 250 KNOTS, 10,000 FT ALTITUDE.

LCTR LABC LCTC LCSC LRDC Configuration Tiltrotor Advancing Tandem Single Single comp., blade compound compound reaction drive Mission GW (lb) 130681 160246 153435 155225 178628 Engine size (hp) 4x7353 4x16556 4x12138 4x14007 4x24081 Rotor diameter (ft) 91.21 90.34 80.70 114.79 123.14 Disk loading W/A (lb/ft ) 10 25 15 15 15 o C / σ (geom, 5k ISA+20 C) 0.1328 0.0750 0.1328 0.1328 0.1328 W o C / σ (T-wt, 5k ISA+20 C) 0.1406 0.1000 0.1406 0.1406 0.1406 W Wing loading (lb/ft ) 80 80 80 80 Wing span (ft) 107.46 151.71 152.59 163.69 Hover figure of merit 0.795 0.698 0.708 0.708 0.700 Cruise performance Propulsive efficiency 0.815 0.900 0.900 0.900 0.900 Drag D/q (ft ) 38.63 38.35 54.45 54.88 60.26 Wing induced drag (ft ) 19.38 12.04 11.89 13.68 Oswald efficiency 1.00 0.80 0.82 0.82 Wing lift (%W) 100% 85% 85% 85% Rotor drag (ft ) 20.36 19.45 24.65 Rotor L/D 12.14 Cruise power (hp) 8519 16363 11571 11486 13134 Weights Weight empty (lb) 84548 97459 99625 100156 113254 Operating weight empty (lb) 85998 98909 101075 101606 114704 Payload (lb) 26400 26400 26400 26400 26400 Fuel weight (lb) 18282 34937 25959 27219 37524 Rotor group weight (lb) 14763 16637 13130 15700 21117 Blade weight (lb) 7476 9390 6653 7327 10592 Hub weight (lb) 7025 7246 6477 8373 10525 Propulsion group weight (lb) 19335 30485 28502 26064 28525 Engine dry weight (lb) 3933 8855 6492 7491 12879 Transmission weight (lb) 13983 18749 19379 15684 11280 Flyaway cost (2005) $70.9M $126.3M $102.9M $104.3M $143.4M DOC+I (¢/asm) (2005) 19.10 33.94 26.80 27.73 37.71 Aircraft L/D=WV/P 11.77 7.51 10.17 10.37 10.43 Productivity, V*PL/(W-PL) 63 49 52 51 43 TABLE 27. HEAVY LIFT ROTORCRAFT DESIGNS FOR CRUISE AT 250 KNOTS, O 5000 FT ISA+20 C.

LCTR LABC LCTC LCSC LRDC Configuration Tiltrotor Advancing Tandem Single Single comp., blade compound compound reaction drive Mission GW (lb) 136168 171348 163345 166719 201915 Engine size (hp) 4x7662 4x17703 4x12922 4x15044 4x27221 Rotor diameter (ft) 93.11 93.42 83.26 118.96 130.92 Disk loading W/A (lb/ft ) 10 25 15 15 15 o C / σ (geom, 5k ISA+20 C) 0.1328 0.0750 0.1328 0.1328 0.1328 W o C / σ (T-wt, 5k ISA+20 C) 0.1406 0.1000 0.1406 0.1406 0.1406 W Wing loading (lb/ft ) 80 80 80 80 Wing span (ft) 109.36 156.53 158.14 174.03 Hover figure of merit 0.795 0.698 0.708 0.708 0.700 Cruise performance Propulsive efficiency 0.810 0.900 0.900 0.900 0.900 Drag D/q (ft ) 39.70 40.10 56.78 57.55 65.39 Wing induced drag (ft ) 17.06 10.82 10.77 13.05 Oswald efficiency 1.00 0.80 0.82 0.82 Wing lift (%W) 100% 85% 85% 85% Rotor drag (ft ) 21.13 20.78 27.60 Rotor L/D 12.68 Cruise power (hp) 9146 17336 12866 12923 15377 Weights Weight empty (lb) 87434 103049 104812 106215 125891 Operating weight empty (lb) 88884 104499 106262 107665 127341 Payload (lb) 26400 26400 26400 26400 26400 Fuel weight (lb) 20884 40449 30683 32654 48173 Rotor group weight (lb) 15586 18372 14266 17266 24843 Blade weight (lb) 7873 10383 7198 8016 12357 Hub weight (lb) 7440 7989 7068 9250 12487 Propulsion group weight (lb) 20147 32597 30343 27993 32244 Engine dry weight (lb) 4098 9468 6911 8046 14558 Transmission weight (lb) 14570 20048 20630 16845 12751 Flyaway cost (2005) $73.8M $134.9M $109.3M $111.9M $162.0M DOC+I (¢/asm) (2005) 20.72 37.65 29.89 31.30 45.17 Aircraft L/D=WV/P 11.42 7.58 9.74 9.90 10.07 Productivity, V*PL/(W-PL) 60 46 48 47 38 Rotorcraft Description and Performance Output Sizing and Performance Performance Design Sizing Performance Sizing Input File Input File Analysis Analysis Tool Rotorcraft Characteristics Changes Made to Rotorcraft in File (RC Format) Solid-Modeling Tool Do Not Incorporate Issues Accounted pdate Inputs Review Sizing and Performance U For in Sizing and Performance Designer Analysis Tool (Currently) C hange Dimensions Format Translator Update Shape Rotorcraft Characteristics File (Pro/E Format) Solid Model Solid-Modeling Solid-Modeling Template File Output Tool Review Changes Due to Layout Record of Changes in Rotorcraft Characteristics File Figure 3. Simplified conceptual design process.

1990 Helicopter Technology 1990 Helicopter Technology 2/3 2/3 D/q = 2.5 (W/1000) D/q = 2.5 (W/1000) 100 100 1975 Helicopter Technology 1975 Helicopter Technology Figure 1. Runway Independent Aircraft (RIA) 2/3 2/3 D/q = 5.0 (W/1000) D/q = 5.0 (W/1000) industry concepts from 2002: Sikorsky Reverse Velocity Rotor (top), Boeing Tiltrotor (center), Bell Quad Tiltrotor (bottom).

Jet Transport Jet Transport 10 10 2/3 2/3 D/q = 0.75 (W/1000) D/q = 0.75 (W/1000) Drag D/q, sqft Drag D/q, sqft Propeller Transport Propeller Transport 2/3 2/3 D/q = 1.4 (W/1000) D/q = 1.4 (W/1000) Heavy Lift Rotorcraft Fuselage Heavy Lift Rotorcraft Fuselage propulsion system sizing 2/3 2/3 noise D/q = 1.6 (W/1000) D/q = 1.6 (W/1000) 1 1 airframe aerodynamics RC 1,000 10,000 100,000 1,000,000 1,000 10,000 100,000 1,000,000 handling qualities Gross Weight, lb Gross Weight, lb performance opt Figure 4. Aircraft drag trends (courtesy F. D. Harris).

CAMRAD II rotor airfoils blade aerodynamics loads, stability 100 100 Possible Hub Technology Possible Hub Technology 2/3 2/3 D/q = 0.85 (W/1000) D/q = 0.85 (W/1000) CAMRAD II Typical Hub Technology Typical Hub Technology hub concept 2/3 2/3 D/q = 1.20 (W/1000) D/q = 1.20 (W/1000) blade structural design materials 10 10 noise Drag D/q, sqft Drag D/q, sqft control concept Heavy Lift Rotorcraft Hub Heavy Lift Rotorcraft Hub vibration flight profiles 2/3 2/3 D/q = 0.40(W/1000) D/q = 0.40(W/1000) Faired Hubs Faired Hubs weight, performance, drag 1 1 resolve differences with RC 1,000 10,000 100,000 1,000,000 1,000 10,000 100,000 1,000,000 Gross Weight, lb Gross Weight, lb Figure 5. Helicopter hub drag trends (courtesy Figure 2. Outline of iterative design process.

F. D. Harris).

Figure 6. Three-view of Large Civil Tiltrotor (LCTR).

Figure 7. Three-view of Large Civil Tandem Compound (LCTC).

Figure 8. Three-view of Large Advancing Blade Concept (LABC) Figure 9. Flyaway price (2005 USD) and DOC+I Figure 10. Cost elements compared for heavy lift (2005 cents/ASM) comparisons for baseline designs.

rotorcraft and B737 (1200 nm, 120 passengers, including technology factors for rotorcraft costs).

40 40 40 40 30 30 30 30 more flap than lag more flap than lag more lag than flap more lag than flap per rev per rev 20 20 20 20 operating operating airframe sym airframe sym frequency (Hz) frequency (Hz) frequency (Hz) frequency (Hz) airframe antisym airframe antisym 10 10 10 10 0 0 0 0 40 80 120 160 40 80 120 160 40 80 120 160 40 80 120 160 rotor speed (rpm) rotor speed (rpm) rotor speed (rpm) rotor speed (rpm) Figure 11. LCTR blade and airframe frequencies. Collective = 0 deg (left figure, appropriate for 140 rpm operation) and collective = 60 deg (right figure, 75 rpm operation).

0.85 0.85 0.82 0.82 0.82 inboard twist = -28 inboard twist = -28 inboard twist = -28 inboard twist = -30 inboard twist = -30 inboard twist = -30 0.81 0.81 0.81 0.80 0.80 inboard twist = -32 inboard twist = -32 inboard twist = -32 inboard twist = -34 inboard twist = -34 0.80 0.80 0.80 inboard twist = -36 inboard twist = -36 0.75 0.75 inboard twist = -38 inboard twist = -38 0.79 0.79 0.79 5k+20 5k+20 figure of merit figure of merit inboard twist = -40 inboard twist = -40 0.70 0.70 RC code RC code cruise propulsive efficiency cruise propulsive efficiency cruise propulsive efficiency outboard twist = -30 outboard twist = -30 0.78 0.78 0.78 0.77 0.78 0.79 0.80 0.77 0.78 0.79 0.80 0.77 0.78 0.79 0.80 optimum (-32/-30) optimum (-32/-30) 0.65 0.65 hover figure of merit hover figure of merit hover figure of merit 0.08 0.12 0.16 0.20 0.08 0.12 0.16 0.20 CT/s CT/s Figure 12. LCTR twist optimization. Figure 13. LCTR rotor hover performance.

0.90 0.90 0.75 0.75 outboard twist = -15 deg outboard twist = -15 deg 0.73 0.73 outboard twist = -12 deg outboard twist = -12 deg 0.70 0.70 outboard twist = -9 deg outboard twist = -9 deg 0.71 0.71 outboard twist = -6 deg outboard twist = -6 deg outboard twist = -3 deg outboard twist = -3 deg SLS SLS Figure of merit Figure of merit 0.69 0.69 outboard twist = 0 deg outboard twist = 0 deg 0.50 0.50 30000 ft 30000 ft inboard twist = 0 inboard twist = 0 RC code RC code 0.67 0.67 optimum (0/-12) optimum (0/-12) propulsive efficiency propulsive efficiency 9.9 10.1 10.3 10.5 9.9 10.1 10.3 10.5 0.30 0.30 aircraft L/D=WV/P aircraft L/D=WV/P 150 250 350 450 150 250 350 450 Figure 16. LCTC twist optimization (inboard twist = speed (knots) speed (knots) 3, 0, –3,–6 deg).

16 16 12 12 0.85 0.85 8 8 5k+20 5k+20 0.80 0.80 RC code RC code 4 4 aircraft L/D=WV/P aircraft L/D=WV/P 0.75 0.75 0 0 figure of merit figure of merit 150 250 350 450 150 250 350 450 0.70 0.70 speed (knots) speed (knots) 0.65 0.65 Figure 14. LCTR rotor cruise performance.

0.08 0.12 0.16 0.20 0.08 0.12 0.16 0.20 CT/s CT/s Figure 17. LCTC rotor hover performance.

more flap than lag more flap than lag more lag than flap more lag than flap per rev per rev operating operating 30 30 14 14 airframe sym airframe sym airframe antisym airframe antisym 12 12 10 10 20 20 8 8 frequency (Hz) frequency (Hz) 6 6 30000 ft 30000 ft 4 RC code 4 RC code aircraft L/D=WV/P aircraft L/D=WV/P 10 10 2 2 0 0 150 250 350 450 150 250 350 450 speed, knots speed, knots 0 0 Figure 18. LCTC aircraft cruise performance.

40 80 120 160 200 40 80 120 160 200 rotor speed (rpm) rotor speed (rpm) Figure 15. LCTC blade and airframe frequencies (collective = 10).

0.80 0.80 40 40 0.75 0.75 0.70 0.70 more flap than lag more flap than lag 30 30 more lag than flap more lag than flap figure of merit figure of merit per rev per rev 0.65 0.65 5k/ISA+20 5k/ISA+20 operating operating RC code RC code airframe modes airframe modes 0.60 0.60 0.06 0.10 0.14 0.18 0.22 0.06 0.10 0.14 0.18 0.22 20 20 CT/s CT/s frequency (Hz) frequency (Hz) Figure 20. LABC rotor hover performance.

30000 ft 30000 ft 10 10 10 rotor L/D 10 rotor L/D rotor L/D, RC code rotor L/D, RC code 8 8 6 6 4 4 aircraft L/D=WV/P aircraft L/D=WV/P 0 0 2 2 40 80 120 160 40 80 120 160 0 0 rotor speed (rpm) rotor speed (rpm) 150 250 350 450 150 250 350 450 flight speed, knots flight speed, knots Figure 19. LABC blade and airframe frequencies (collective = 0).

Figure 21. LABC aircraft cruise performance.

2.2E-03 2.9E-03 1 2.0E-03 2.5E-03 r/R=0.50 mid span r/R=0.25 root 1.5E-03 2.0E-03 1.0E-03 1.5E-03 5.0E-04 1.0E-03 Min (-1800) Max (2200) Min (-1300) 0.0E+00 5.0E-04 Max (2800) 0.5 -5.0E-04 0.0E+00 0.5 -1.0E-03 -5.0E-04 -1.4E-03 -1.0E-03 -1.5E-03 -1.8E-03 0 1 2 3 0 1 2 3 2.9E-03 3.0E-03 1 1 2.5E-03 2.5E-03 r/R=1.0 tip 2.0E-03 r/R=0.75 2.0E-03 1.5E-03 1.5E-03 1.0E-03 1.0E-03 5.0E-04 5.0E-04 0.0E+00 0.0E+00 Max (2900) 0.5 Max (2900) 0.5 Min (-2000) -5.0E-04 -5.0E-04 Min (-2300) -1.0E-03 -1.0E-03 -1.5E-03 -1.5E-03 -2.0E-03 0 0 0 1 2 3 0 1 2 3 Figure 22. LCTR blade section design; normal strain (microstrain).

6.0E+06 8.0E+06 5.0E+06 1 r/R=0.50 mid span 7.0E+06 4.0E+06 6.0E+06 3.0E+06 r/R=0.25 root 5.0E+06 2.0E+06 4.0E+06 1.0E+06 Max (6.9E+6) 3.0E+06 0.0E+00 Max (9.2E+6) 0.5 Min (-4.2E+6) Min (-5.9E+6) 2.0E+06 0.5 -1.0E+06 1.0E+06 -2.0E+06 0.0E+00 -3.0E+06 -1.0E+06 -4.0E+06 -2.0E+06 -3.0E+06 -4.0E+06 -5.0E+06 0 1 2 3 0 1 2 3 8.0E+06 r/R=1.0 tip 8.0E+06 6.0E+06 6.0E+06 r/R=0.75 4.0E+06 4.0E+06 2.0E+06 2.0E+06 0.0E+00 0.0E+00 0.5 Max (9.2E+6) 0.5 -2.0E+06 Max (8.7E+6) -2.0E+06 -4.0E+06 -4.0E+06 Min (-6.7E+6) -6.0E+06 Min (-4.9E+6) 0 1 2 3 0 1 2 3 Figure 23. LCTR blade section design; normal stress (lb/ft ).

0.25 0.15 0.05 -0.05 -0.15 -0.25 Figure 26. LCTC NASTRAN model (non-structural 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 masses not shown).

Figure 24. LCTR wing section (24% thick).

Figure 27. LABC NASTRAN model (non-structural Figure 25. LCTR NASTRAN model (non-structural masses not shown).

masses not shown).

(a) Root section: AFDD CTR1544, c = –0.160, m t/c = 15.3%.

Figure 30. OVERFLOW grid for LCTR rotor and (b) Mid-span section: AFDD CTR4475, c = 0.027, m spinner analysis.

t/c = 11.3%.

Figure 28. LCTR airfoil sections.

(a) Simple spinner geometry.

0.84 0.84 0.82 0.82 0.80 0.80 SOA airfoils SOA airfoils 0.78 0.78 LCTR airfoils LCTR airfoils cruise propulsive efficiency cruise propulsive efficiency 0.76 0.76 0.76 0.77 0.78 0.79 0.80 0.76 0.77 0.78 0.79 0.80 (b) Improved spinner geometry.

hover figure of merit hover figure of merit Figure 31. Regions of supersonic flow (red area) on Figure 29. Influence of airfoils on LCTR hover and LCTR in cruise.

cruise performance.

EPNdB Weight (lb/100000) (a) Descent, takeoff, and level flight.

(a) Low wing.

EPNdB Weight (lb/100000) (b) Average.

(b) High wing.

Figure 33. Assessment of LCTR noise: calculated Figure 32. Hover download calculation for LCTR: rotational noise relative current certification wing and body pressures, comparing high wing and requirements and technology goals.

low wing configurations.

1.2 1.2 0.3 0.3 1.0 1.0 0.2 0.2 0.8 0.8 0.2 0.2 0.6 0.6 0.1 0.1 0.4 0.4 collective = 2 collective = 2 collective = 2 collective = 2 collective = 5 collective = 5 collective = 5 collective = 5 0.1 0.1 thrust tilt/cyclic magnitude thrust tilt/cyclic magnitude 0.2 0.2 collective = 10 collective = 10 collective = 10 collective = 10 hub moment/cyclic magnitude hub moment/cyclic magnitude collective = 15 collective = 15 collective = 15 collective = 15 0.0 0.0 0.0 0.0 1 2 3 4 1 2 3 4 1 2 3 4 1 2 3 4 flap frequency (per-rev) flap frequency (per-rev) flap frequency (per-rev) flap frequency (per-rev) Figure 34. Rotor thrust vector tilt with cyclic control Figure 35. Hub moment produced by cyclic control (deg/deg).

(C / σ per rad).

M 140 9.5 16.5 3 bladed 4 bladed 5 bladed 6 bladed V OEI sized cr 15.5 8.5 3 blades 14.5 7.5 Gross weight, lb (*1000) Fuel burned, lb (*1000) Engine size, shp (*1000) 13.5 6 blades 115 6.5 13 15.5 110 16 15.5 14.5 14.5 14 90 13.5 Cost, cent/ASM Rotor diameter, ft 13.5 Drive system weight, lb (*1000) 12.5 70 12.5 8 10 12 14 16 8 10 12 14 16 8 10 12 14 16 Disk loading, psf Disk loading, psf Disk loading, psf Figure 36. Influence of disk loading and number of blades on the LCTR design (hover tip speed 650 ft/sec, cruise tip speed 350 ft/sec, 4 engines).

260 16 30 3 bladed V OEI sized cr 220 25 4 bladed 5 bladed 6 bladed 180 20 3 blades Fuel burned, lb (*1000) Gross weight, lb (*1000) 8 Engine size, shp (*1000) 6 blades 100 6 10 40 130 30 25 20 Cost, cent/ASM Rotor diameter, ft Drive system weight, lb (*1000) 10 10 400 500 600 700 800 400 500 600 700 800 400 500 600 700 800 Hover tip speed, fps Hover tip speed, fps Hover tip speed, fps Figure 37. Influence of hover tip speed and number of blades on the LCTR design (cruise tip speed 350 ft/sec, disk loading 10 lb/ft , 4 engines).

170 30 19 160 25 150 20 140 15 Fuel burned, lb (*1000) Gross weight, lb (*1000) 10 Engine size, shp (*1000) 5 13 24 105 22 Cost, cent/ASM Rotor diameter, ft Drive system weight, lb (*1000) 12 12 2 3 4 5 6 2 3 4 5 6 2 3 4 5 6 Number of engines Number of engines Number of engines Figure 38. Influence of number of engines on the LCTR design (hover tip speed 650 ft/sec, cruise tip speed 350 ft/sec, disk loading 10 lb/ft , 4 blades; for 6-engine case there is no cross-shafting, so OEI condition is loss of one engine on each side).

150 10 19 OEI sized 1-speed 1 speed 2-speed 140 17 8 16 Fuel burned, lb (*1000) Engine size, shp (*1000) Gross weight, lb (*1000) V OEI sized cr 2 speed 120 6 13 18 98 18 tip speed 17 96 ratio = 0.85 Cost, cent/ASM Rotor diameter, ft Drive system weight, lb (*1000) 13 12 300 400 500 600 700 300 400 500 600 700 300 400 500 600 700 Cruise tip speed, fps Cruise tip speed, fps Cruise tip speed, fps Figure 39. Influence of cruise tip speed on the LCTR design (hover tip speed 650 ft/sec, disk loading 10 lb/ft , 4 engines, 4 blades), for 1-speed transmission (conventional tiltrotor approach, variable engine rpm) and for 2-speed transmission (optimum engine rpm, with no transmission penalty).

140 9 17 constant cost no W-penalty 20% W-penalty 135 16 130 15 125 14 Gross weight, lb (*1000) Engine size, shp (*1000) Fuel burned, lb (*1000) 6 13 19 94 tip speed ratio = 0.85 92 15 Cost, cent/ASM 14 Rotor diameter, ft Drive system weight, lb (*1000) 12 13 300 400 500 600 700 300 400 500 600 700 300 400 500 600 700 Cruise tip speed, fps Cruise tip speed, fps Cruise tip speed, fps Figure 40. Break-even weight assessment for 2-speed drive system. Influence of cruise tip speed on the LCTR design with 2-speed transmission: with no drive system weight penalty, with 20% transmission weight penalty, and with transmission weight adjusted for constant operating cost (hover tip speed 650 ft/sec, disk loading 10 lb/ft , 4 engines, 4 blades).

155 12 21 11.5 4 bladed 20 V cr OEI sized 5 bladed 6 bladed 10.5 4 blades 9.5 Fuel burned, lb (*1000) Gross weight, lb (*1000) Engine size, shp (*1000) 6 blades 8.5 17 20 105 19 19 18 18 Cost, cent/ASM Rotor diameter, ft 17 17 Drive system weight, lb (*1000) 16 16 10 12 14 16 18 20 10 12 14 16 18 20 10 12 14 16 18 20 Disk loading, psf Disk loading, psf Disk loading, psf Figure 41. Influence of disk loading and number of blades on the LCTC design (hover tip speed 650 ft/sec, cruise tip speed 205 ft/sec, 4 engines).

200 18.5 28 4 bladed 16.5 5 bladed 24 OEI sized 6 bladed 14.5 4 blades 12.5 Fuel burned, lb (*1000) 10.5 Gross weight, lb (*1000) Engine size, shp (*1000) 6 blades 120 8.5 16 30 95 30 25 25 20 20 Cost, cent/ASM Rotor diameter, ft Drive system weight, lb (*1000) 15 15 500 600 700 800 500 600 700 800 500 600 700 800 Hover tip speed, fps Hover tip speed, fps Hover tip speed, fps Figure 42. Influence of hover tip speed and number of blades on the LCTC design (cruise tip speed 205 ft/sec, disk loading 15 lb/ft , 4 engines).

160 35 22 155 21 150 20 145 19 140 18 Fuel burned, lb (*1000) Gross weight, lb (*1000) Engine size, shp (*1000) 5 17 19.5 82 18.5 Cost, cent/ASM Rotor diameter, ft 17.5 Drive system weight, lb (*1000) 17 16 2 3 4 5 6 2 3 4 5 6 2 3 4 5 6 Number of engines Number of engines Number of engines Figure 43. Influence of number of engines on the LCTC design (hover tip speed 650 ft/sec, cruise tip speed 205 ft/sec, disk loading 15 lb/ft , 4 blades).

LCTR LCTC 11 11 9 9 percent increase percent increase Fuel DOC Fuel DOC Price Price Power Power Weight Weight (a) Rotor blade and hub weight (advanced technology factors 0.79 and 0.96 respectively).

21 22 20 20 18 18 percent increase percent increase Fuel DOC Fuel DOC Price Price eight Power Power Weight W (b) All structural weight.

29 29 percent increase percent increase DOC Fuel DOC Fuel Price Price Power Power Weight Weight (c) Drive system weight (advanced technology factor 0.67).

Figure 44. Impact of technology on the LCTR and LCTC designs: percentage increase caused by changing the technology from advanced to current level.

LCTR LCTC +0.25 D/q +0.25 D/q -0.25 D/q -0.25 D/q 14 14 14 12 12 12 12 6 6 percent change percent change -3 -3 -3 -3 -3 -3 -7 -7 -11 -11 (a) Airframe drag ( ± 25% relative baseline).

percent change -4 -4 -4 -7 k=0.85 -11 k=0.20 2/3 (b) Hub drag, D/q = k(W/1000) (k = 0.85 (SOA) and 0.20, relative baseline k = 0.44).

17 17 percent increase percent increase Fuel DOC Fuel DOC Price Price Power Power Weight Weight (c) All aerodynamics.

Figure 45. Impact of technology on the LCTR and LCTC designs: percentage increase caused by changing the technology from advanced to current level.

LCTR LCTC percent increase percent increase 28 28 Fuel DOC Fuel DOC Price Price Power Power Weight Weight (a) Engine technology (current relative advanced; note change of scale relative other figures).

25 25 25 25 22 22 increase increase percent percent Fuel DOC Fuel DOC Price Price Power Power Weight Weight (b) Penalty for weight and power contingency.

Figure 46. Impact of technology on the LCTR and LCTC designs: percentage increase caused by changing the technology from advanced to current level.

Figure 47. Effect of cost technology factors on Figure 48. Cost elements compared for LCTR flyaway price (2005 USD) and DOC+I (2005 with and without cost technology factors for lift cents/ASM) for LCTR.

rotorcraft (1200 nm, 120 passengers); in legend [x,y], x is maintenance factor and y is price factor.

5k ISA+20 and 250 kts 5k ISA+20 and 250 kts 20k ISA and 350 kts 20k ISA and 350 kts 16 16 16 16 12 12 12 12 8 8 8 8 4 4 4 4 aircraft L/D=WV/P aircraft L/D=WV/P aircraft L/D=WV/P aircraft L/D=WV/P 0 0 0 0 250 300 350 400 450 250 300 350 400 450 150 200 250 300 350 150 200 250 300 350 flight speed (kts) flight speed (kts) flight speed (kts) flight speed (kts) 10k ISA and 250 kts 10k ISA and 250 kts 30k ISA and 350 kts 30k ISA and 350 kts 16 16 16 16 12 12 12 12 8 8 8 8 4 4 4 4 aircraft L/D=WV/P aircraft L/D=WV/P aircraft L/D=WV/P aircraft L/D=WV/P 0 0 0 0 150 200 250 300 350 150 200 250 300 350 250 300 350 400 450 250 300 350 400 450 flight speed (kts) flight speed (kts) flight speed (kts) flight speed (kts) 20k ISA and 250 kts 20k ISA and 250 kts 16 16 12 12 LCTR LCTR LABC LABC 8 8 LCSC (single) LCSC (single) LCTC (tandem) LCTC (tandem) 4 4 aircraft L/D=WV/P aircraft L/D=WV/P 0 0 150 200 250 300 350 150 200 250 300 350 flight speed (kts) flight speed (kts) Figure 49. Aircraft L/D for different design conditions.

Next: APPENDIX A →

APPENDIX A

ROTORCRAFT SECTOR GOTCHA

The Rotorcraft Sector goals, objectives, technical challenges, and approaches (GOTChA) are presented in table 2, and in more detail in figures A1 and A2.

Hover Efficiency Goal

The hover efficiency goal is W/P = 6 (SOA = 4.3), where W is the gross weight (lb) and P the installed power (hp).

The SOA is derived from the V-22 and CH-53X:

V-22: W = 53000 lb, P = 12300 HP

CH-53X: W = 80000 lb, P = 18605 HP

which gives W/P = 4.3.

The objective is to increase overall hover efficiency by 40% for large rotorcraft.

The disk loading of the V-22 and CH-53X is approximately 20 lb/ft . For the 120-passenger notional vehicle, the rotor size will be larger than SOA but the weight and disk loading must be reduced from SOA levels in order to increase hover efficiency and have the low power required. Assume a target disk loading of 10 lb/ft for the notional vehicle and a figure of merit (M) unchanged from SOA levels. The disk loading is W/A, and

1/2

the definition of figure of merit gives W/P = M(2 ρ A/W) . So a factor of 2 reduction in disk loading gives

1/2

W/P = 4.3*2 = 6.

Efficient Cruise Goal

The efficient cruise goal is L/D = 12 (SOA = 6.5), where L is the aircraft lift (lb) and D is the aircraft drag (lb).

The SOA value of 6.5 is based on the approximate L/D of the V-22 airframe in airplane mode at W = 53000 lb (ref. V-22 Aerodynamic Report 901-909-6, Rev A, Sheet 70.5).

The objective is to reduce the drag by 44% at a cruise Mach number of 0.6.

From plots of airframe drag D/q (ft ) versus aircraft gross weight W (courtesy F. D. Harris),

2/3

D/q = K(W/1000) , with K = 2.5 for 1980 helicopter technology and K = 1.4 for 1980 propeller driven fixed- wing aircraft. Taking propeller driven fixed-wing drag levels as the goal, the percent reduction is 100*(2.5– 1.4)/2.5 = 44%. This reduction in drag increases the airframe L/D from 6.5 to 11.6.

Empty Weight Goal

The empty weight fraction (excluding engine) goal is W /W = 0.41 (SOA = 0.55), where W is the gross

e

weight (lb), and for this purpose

W = gross wt – fuel wt – payload wt – operating item wts – engine wt

e

or

W = airframe structure wt + airframe systems and equipment (including furnishings) wt + rotor

e

system wt + transmission/drive system wt

The rotor system is defined as the rotor blades, the control system (swash plate, linkages, etc.), and hub.

The SOA value is based on the V-22 (weights obtained from Bell-Agusta and “Pocket Guide V-22 Osprey” by the Bell-Boeing Team, March 2000):

gross weight = 53000 lb

engine weight = 2(1889) = 3778 lb

fuel + payload weight = 20000 lb

So W = 53000 – 20000 – 3778 = 29222 and W /W = 29222/53000 = 0.55. Note that the conventional

e e

definition of operating empty weight is gross weight less fuel and payload, giving OWE/GW = 33000/ 53000 = 0.62.

The objectives are (1) reduce rotor weight by 25%; (2) reduce subsystem weight (drive system, high-lift, etc.)

by 25%; (3) reduce airframe structural weight by 25%. Assuming all weight sources contributing to W are

e

reduced by 25% gives a goal of W /W = 0.75(0.55) = 0.41.

e

Community Noise Goal

The community noise goal is a reduction of 14 EPNdB below SOA level (SOA = Stage 2 certification microphones average level over 3 flight conditions).

The SOA noise level is a function of helicopter gross weight and flight condition. For a given rotorcraft, the SOA will be the average value measured by the 3 certification microphones during take-off, approach, and level flight.

The Goal assumes a reduction in the measured sound pressure level of 80%: 20*log(.2) = –14db.

The objectives are (1) reduce source noise by design and active control; (2) reduce noise impact by flight operational procedures; (1 and 2 combined) reduce rotorcraft system noise by 80%.

For rotorcraft, noise reduction will be achieved through a combination of source noise reduction and low- noise flight operational procedures. Assigning noise reduction metrics to each objective is not appropriate.

Noise reductions achieved via objective 1 or 2 separately are not necessarily additive. A reduction of rotorcraft system noise by 80% is consistent with the Goal of 14 EPNdB noise reduction.

Flight Control Goal

The flight controls and systems goal is to enable certifiable near-all-weather operations (CAT IIIC; SOA = Special VFR).

The SOA is GPS helicopter IFR. Current operations are similar to ground-based non-directional beacon, to a point in space with final approach to landing done with visual reference. This typically requires “Special VFR” operation (1000 ft ceiling and 1 mile visibility). Lower visibility requirements are attainable with appropriate equipment at airports. This requires use of the ground-based Instrument Landing System (ILS) and use of an approved helicopter IFR approach, typically flown at a constant speed of 90 knots until breakout into visual conditions. The 90 knots speed requires some visibility for the deceleration to landing.

Cat. IIIA goal for “austere” operations: IFR approach to Cat. IIIA visibility (decision height less than 100 feet and visibility of 700 feet RVR) without extensive ground infrastructure or equipage (FAA AC-120-28D).

Cat. IIIC goal at “equipped” airports: “Zero” visibility IFR operations at sites with some ground infrastructure (e.g. differential GPS) and surveyed approach/departure paths and sanitized touch-down surface.

The objective is to remove restrictions for rotorcraft operations in zero visibility and near all weather conditions (CAT IIIC).

Advanced Engine Performance Goal

The advanced engine performance goal is SFC of SOA–10%, and SHP/eng-wt of 120% SOA.

The CT7-8 (commercial version of the GE T-700) is defined as the SOA engine.

The SFC and SHP/eng-wt goals were established under IHPTET/VAATE. The SFC metric under IHPTET was SOA–40%, but since IHPTET was initiated nearly 20 years ago, the SOA has changed and therefore the metric was changed to SOA–10%.

The objectives are (1) capitalize on new engine technology to improve performance, weight, noise of rotorcraft systems; (2) reduce engine weight by 20%.

Cabin Noise and Vibration Goals

The cabin noise and vibration goal is reduction to 77dbA and 0.05g throughout the flight envelope (SOA = 88dBA and 0.15g).

The SOA interior noise levels were provided by a technical representative of a large US rotorcraft manufacturer and indicate the interior noise and vibration of a large commercial passenger helicopter with a VIP interior cabin treatment. Average interior noise levels in cruise range from 86 to 90dB(A) with an average level of 88dB(A). Hover levels averaged 79dB(A). Vibration levels are 0.15g or below.

The desired interior noise for rotorcraft is 77dB(A), which is the average interior noise of the existing fixed- wing subsonic commercial aircraft fleet as of 2003. The desired interior vibration levels are 0.05g, a 10dB reduction relative the current SOA.

The objective is passenger acceptance: noise and vibration levels comparable to current fixed-wing subsonic aircraft.

GOALS Empty Weight Fraction (excluding Hover Efficiency Efficient Cruise engine) (L/D = 12, SOA = 6.5) (Wg/P = 6, SOA = 4.3) (We/Wg = 0.41, SOA = 0.55) 1 2 3 OBJECTIVES Reduce drag by Increase overall Reduce rotor Reduce Reduce airframe 44% at a cruise hover efficiency by subsystem structural weight weight by 25% Mach # of 0.60 40% for large by 25% weight by 25% rotorcraft.

(drive system, high-lift, etc.)

01 02 03 04 05 TECHNICAL CHALLENGES Reduce drag without Reduce rotor power Reduce rotor Reduce Reduce adversely affecting with minimal weight subsystem weight fraction as airframe empty performance, noise, impact and acceptable vehicle size empty weight weight fraction and vibration.

rotor dynamics.

fraction as as vehicle size doubles.

vehicle size doubles. 02 doubles.

03 04 05 01 APPROACHES Reduce drag on Optimize Develop low drag, high- Develop lightweight Improve Design rotating lift rotor system airfoil airframe/propul. swashplateless passive and active materials and components.

integration to red section designs.

aeroelastically damage rotor systems. 04 interference drag stable structures. tolerant (boundary layer structural control, MAFC, design Integrate active airfoil design, aerodynamic flow concepts.

Design large, Improve structural etc.).

control devices for high- lightweight Determine bonding.

lift (oscillation, blowing, aeroelastically Develop aeromechanic 13 02 MAFC, etc.).

stable rotor lightweight scaling laws for systems.

drive systems large R/C.

Develop advanced and transmission variable speed drive 05 Reducing R/C Level 1 handling system concepts to concepts.

system source qualities in an better match hover and Investigate novel noise without obstacle rich 11 cruise flight. 03 rotor configs degrading environment with enabling high-lift performance low noise … ...

Reduce rotor and high-speed cabin noise and Link to Cabin Noise Tech Challenge cruise w/low noise vibration … ...

and vibration (e.g, Link to Community Noise Tech Challenge active rotors, etc.)

Link to Flight Controls Tech Challenge Figure A1. Rotorcraft Sector goals, objectives, technical challenges, and approaches (Rev 051005).

GOALS Reduce cabin noise and vibrations Advanced Engine Performance Flight Controls & systems to enable Reduce Community Noise (IHPTET/VAATE: throughout flight envelope Goal: SOA Level -14EPNdB certifiable near all weather (77dBA/0.05g ’ s, SOA = 88dBA/0.15g ’ s) SFC = SOA - 10% operations (CAT Ź IIIC, SOA: Special SOA=Stage2 cert. mics ave SHP/Wg = SOA *120%) level over 3 flight conditions VFR) 4 5 6 7 OBJECTIVES Remove restrictions for Capitalize on new Reduce Passenger Reduce noise Reduce source engine technology engine weight acceptance R/C operations in zero noise by design impact by flight to improve visibility and near all by 20%. (Noise level and active control operational performance, comparable to weather conditions procedures 06 07 weight, noise, etc. (CAT Ź IIIC).

current fixed- Reduce rotorcraft system noise by 80%.

of R/C systems. wing subsonic A/C) 08 11 TECHNICAL CHALLENGES Reduce noise Reduce rotor Reducing R/C Reduce Reduce the large Level 1 handling Ensure R/C Compact, IHPTET through flight cabin noise gearbox cabin qualities in an operations are lightweight low system source engine-out Challenges capabilities & and vibration obstacle rich noise without noise & system penalty compatible with power anti- operational 12 without vibration w/o for R/C safe environment with the evolving air icing system degrading procedures that increasing increasing low noise flight performance.

certifiable traffic not available.

are safe & weight/cost.

weight/cost.

operations. procedures. management certifiable with 13 14 Level 2 with a system.

acceptable degraded system community 11 state.

08 09 10 impact.

Develop tuned APPROACHES aircraft Develop and structures to Adaptive/Robust fully Develop and Develop and Develop new validate attenuate noise demonstrate low augmented control ATM/controller demonstrate system/source and vibration. system (includes noise operations emergency power, and operational noise &propagation automation): and capabilities concepts and auxiliary augmentation.

prediction 23 IFPC, IVHM/PHM with acceptable procedures capabilities.

Intelligent systems, handling with FAA.

Investigate Light weight power flight management qualities.

novel rotor Develop active 17 21 system (FMS). 18 sharing/ distribution configs … noise (composite cross

Develop & evaluate -

cancellation : shafting, ducting, etc.).

Link to Identify vehicle low noise rotors Structural and Evaluate and constraints of steep, Efficient (active/passive Aero (such as develop new Cruise segmented, curved, concepts) & active flutter systems Innovative propulsion decelerating RIA Approach engines suppression (electro- systems (distributed flight ops. Link to Efficient Cruise systems). expulsive, 19 28 highly redundant Approach piezo-electric, propulsion). 25 etc.). Develop source noise Develop cockpit automation Investigate Methods to automatically 22 reduction technology to enable low workload novel rotor for mechanical detect failure of power human supervision of configs … train using IVHM/PHM. components complex flight tasks in congested airspace.

Figure A2. Rotorcraft Sector goals, objectives, technical challenges, and approaches (Rev 051005).

Next: APPENDIX B →

APPENDIX B

RISK REDUCTION FOR HEAVY LIFT ROTORCRAFT

The NASA Heavy Lift Rotorcraft Systems Investigation was a focused and coordinated analytical effort to select the best configuration for meeting the Rotorcraft Sector vehicle technology goals (table 1). Three aircraft configurations were the primary subject of the Investigation: 1) Large Civil Tiltrotor (LCTR) 2) Large Civil Tandem Compound (LCTC) 3) Large Advancing Blade Concept (LABC) The conclusion of this investigation was that the LCTR is the configuration with the most promise to meet the Rotorcraft Sector technology goals. The LCTR had the best cruise efficiency, hence the lowest weight and lowest cost. The LCTC exhibited good cruise efficiency, but less than the tiltrotor, and with higher development risk than the tiltrotor. The LABC had lower cruise efficiency than the tiltrotor for the NASA civil mission.

During the course of the investigation, high risk areas were identified. The definition of high risk is one or both of the following: capability or attribute unavailable today, so it is necessary to assume advanced technology will be available in the future in order for the aircraft to achieve the technology goals; or cost prevents the vehicle from being economically competitive, so the payoff of advanced technology is essential to achieving the goals. The following high risk areas were identified for the LCTR: a) High torque, lightweight drive system.

b) High performance, structurally efficient rotor/wing system.

c) Low noise aircraft.

d) Super-integrated vehicle management system.

If the heavy lift configuration selected had an edgewise rotor (LCTC or LABC), then the high risk areas would also include low hub drag, and high advance ratio dynamics. In general, the top-level risks associated with the development of a high-speed, heavy lift rotorcraft are not configuration dependent.

Plans were developed to mitigate the above four risks. The risk reduction plans provide the strategic direction to support a heavy-lift rotorcraft development. Because of the significant overlap between the LCTR, LCTC, and LABC in terms of the risk mitigation technology, the plans are applicable to all three configurations.

There are other rotorcraft goals and missions for which the LCTC and LABC are viable concepts, hence specific tasks for these two configurations are included where appropriate.

The risk reduction plans parallel a prototype development program, and feed design solutions at high technology readiness levels (TRL) to the prototype program. Although a prototype is not part of the NASA plan, it serves to provide schedule pull for the required tasks. The risk reduction program elements are technology readiness level benchmarks, program tasks, and strategic direction. Figure B1 illustrates the program. Technology readiness level benchmarks are major milestones demonstrating significant increases in TRL, principally by integrated hardware tests. There are two primary thrusts (fig. B1): early technology leading to advanced technology and supporting a possible year 7 prototype; advanced technology leading to a year 10 prototype. The risk reduction plan has the following technology readiness level benchmarks (fig. B1).

a) Full-scale propulsion system ground test: complete transmission and engine arrangement; metrics are weight, cost, and noise.

b) Full-scale structure ground test: rotor blade and hub, airframe components, wing components; metrics are weight, cost, and interior noise.

c) Large-scale rotor system wind tunnel test: dynamically scaled rotor and hub; metrics are performance, loads, vibration, control, and noise.

d) Flight simulation test: including elastic airframe and load control system; metrics are handling qualities, control, and noise.

e) Noise and control flight test: existing aircraft, with rotor active control and low-noise operations; metrics are noise, vibration, control, and handling qualities.

f) Integrated large-scale wind tunnel test: utilize propulsion, structure, rotor systems from earlier benchmarks; metric is system integration.

The following sections describe the tasks identified for each of the four risk areas. The tasks are organized by discipline (propulsion, structures, aeromechanics, acoustics, and handling qualities) although all elements are connected because of the interdisciplinary nature of rotorcraft problems. Background information is provided for each of the discipline areas. Detailed task descriptions are presented next, including schedules showing ground test, wind tunnel test, flight test, and decision milestones, together with connections to the TRL benchmarks (figs. B2-B7). The tasks constitute the work required to achieve the TRL Benchmarks. Finally, a strategic direction providing guidance for selecting highest priority activities, aimed at the highest risk areas of heavy lift rotorcraft development, is presented.

Concept Development The first year of the risk reduction program is also devoted to further refinement of the rotorcraft concept (fig. B2). Decisions must be made to narrow the focus of the program elements, since parallel research lines can not be afforded. It is essential to focus the work on areas of high payoff for heavy lift rotorcraft. This concept development should be conducted by several companies, for the selected heavy lift rotorcraft configuration (the LCTR for the NASA mission). The first phase would be completed within 12 months. A second phase would be completed in year 5, for the advanced technology thrust. A number of LCTR configuration features require decisions: the wing (high or low), nacelle (engine tilt or not), tail configuration, and number of engines. Corresponding decisions for the LCTC are whether it has single or tandem rotors, powered or auto rotating, tail configuration, and number of engines. For the LABC, the auxiliary propulsion configuration, tail configuration, and number of engines must be selected.

For the propulsion system, it is first necessary to establish what engine technology is available, in particular the possibility for maintaining good SFC over a wide engine rpm range (by design point, or variable geometry, or a multi-speed output shaft gearbox). The required engine development must be defined; then the transmission requirements can be defined, including the system configuration.

Structural design concepts must be identified for light-weight, heavy lift rotorcraft, for the blade and hub, for the airframe, and for the wing. A blade and hub concept must be identified that first and foremost is a solution to the stability issues (whirl flutter for the LCTR, high advance ratio for the LCTC), and also a solution for the strength and weight requirements. The requirement for active control of loads must be established, and the approach defined (flight condition limiting, active load control, others). The requirement for control of noise, vibration, gust and performance, must be established, and the control method selected (IBC, or on-blade, or airframe).

Design guidelines for noise must be established (such as minimum blade-passage frequency), based on physiological and psychological effects on human response. Low noise concepts and approaches must be identified (design features, active control, flight operations). Certification and community noise impact requirements must be defined. The requirements and certification approach for handling qualities must be established, including one-engine inoperative.

The technology contributions to reduction of purchase price and maintenance cost must be identified. A public benefit model and reality-based cost model should be developed.

Technology Background Propulsion Based on the operational requirements of all three configurations proposed for the heavy lift rotorcraft program, all require the propulsion system to vary in speed and in some cases to unload rotors and power propellers for forward flight. The speed of the rotating rotors and propellers need much larger speed changes than have been designed into the V-22 and BA-609. Therefore, the propulsion system needs to vary speed up to or greater than 50% or more from the helicopter hover mode. Speed change can be achieved by having gas turbine engines that have wide operational range or a drive system that has multiple or variable speed capability. For the designs, as currently configured, the propulsion system may require a combination of both a variable speed engine and drive system.

The propulsion system will also have to be highly reliable, maintenance free, and cost effective if the heavy lift rotorcraft is to approach fixed wing operation on a dollars per average seat mile basis. While the heavy lift rotorcraft propulsion system will handle more power than any previously built rotorcraft, the reliability and cost aspects need to be designed into the system. This means that low cost (procurement and operating) needs to be a highly valued design constraint. This may require a propulsion system that may weigh more, but will have a much longer life or mean time between overhaul. The mean time between overhaul is on the order of 25,000 flight hours for current fixed wing aircraft and the propulsion system for the heavy lift rotorcraft needs to be at least as good as that of fixed wing aircraft.

Design & Configuration Study As an initial attempt at design of the propulsion system for the three proposed concepts, the various configurations that would lead to variable propulsion system need to be evaluated. In all current rotorcraft the drive systems have been fixed ratio. Variable speed or multi-speed drive systems have only been developed to the production stage for land based systems (automobiles, farm equipment, etc.). For two of three concepts the typical load-carrying rotors need to operate in a torque-unloaded fashion during forward flight leaving this end of the high speed flight to propellers. Therefore, transition and power shifting requires careful attention.

Varying speed will be accomplished using the engine, the drive system or the combination of both. Also, the proposed propulsion systems all have four engines requiring 28,000 to nearly 52,000 hp, depending upon the aircraft. In order to keep the drive system weight down, many of the recent improvements developed under the Army’s Advanced Rotorcraft Transmission (I & II) and the Rotorcraft Drive System for the 21st Century projects must be applied, and component programs must continue to evolve to support the program goals for the heavy lift rotorcraft program. Some of these recent technologies include face gears, low-noise spiral bevel gears, split path gearing (or split torque), hybrid bearings, and other technology improvements.

For the gas turbine engine’s part of the propulsion system, a fairly flat specific fuel consumption (SFC) rating is required over a wide operating range. Smooth power through the transition from vertical to horizontal flight and back again is essential for passenger comfort and acceptance. While this program does not include an engine development task, other programs will be required to make the necessary investments to have the engine requirements in-hand when needed in the 7-year time frame. Cruise power levels required for the study aircraft may size the engines. Increased engine size may be required to achieve objective cruise power and engine life. A significant investment in engine technology development is required to achieve objective SFC levels and to raise the Technology Readiness Level.

The configuration study aspect of this program will determine the high risk technologies needing development from a low to high technology readiness level. Variable speed propulsion is a necessary technology for the heavy lift rotorcraft to be successful.

Component Development and Concept Validation Once a variable speed propulsion system configuration is chosen, the necessary design, analysis, and concept tests need to be performed. On a parallel path, the evolution of propulsion system technologies that feed into improvements of the future must be continued for future system improvements. As far as the drive system, the variable/multi-speed aspects most likely will be located close to the engine at the highest speed and lowest torque part of the system—much like the location of the over-running clutch as required on multi-engine rotorcraft. This will permit the variable/multi-speed aspect of the design to be as light as possible.

The variable speed engine capabilities need to be addressed. Previous studies have shown that power turbine efficiency can vary significantly when run at off-design speeds. Engine technologies exist that could be applied to flatten out the SFC curve and thereby deliver part of the variable speed requirement. System trade studies are required to optimize drive system speed ratios and power turbine speed ranges. The drive system may only need to have two or more speeds without variability in between.

For the drive system, the component development aspects include: analytical tools, component improvements for gears, bearings, shafts/couplings, housings, clutches and rotor brakes. Also, loss-of-lubrication, efficiency, and health and usage monitoring (diagnostics and prognostics) research and technology development needs to be continued. These technologies feed into longer life, reliability, and lower operating costs.

For gears and bearings, advanced materials are necessary to increase power-to-weight capability through increases in bending and allowable contact levels. These improvements are validated through conducting statistically significant testing that requires a great deal of test time and data to substantiate improvements.

Engineered surfaces are another technology area that holds great promise for drive system improvements.

These improvements are accomplished through paying particular attention to the surface of the mating components during operation. Some of the recent technologies that have been under consideration include super finishing and thin film coatings. Improvements at the surface level can increase performance and extend operation of components during compromised lubrication conditions.

Advanced bearing technologies will have a drastic effect on operational capabilities and possible design arrangements. Recent technologies that are currently under evaluation include hybrid (ceramic rolling elements) and wave-journal bearings. Hybrid bearings offer lighter weight (this can be a sizable savings for large bearings) and wider operating range in terms of rotational speed and temperature. Wave-journal bearings proposed for use in high bypass ratio fans for advanced civil transport can carry high loads in a small package. This permits planetary gear systems to have a reduced size (reduced weight) because of the bearing envelope being much smaller. These technologies, while showing great promise, need further investigation to validate their operation in future rotorcraft systems.

Another area of enhancement for the drive system is lubrication and performance (efficiency) improvements.

These lubrication system enhancements not only will be applied to the emergency operation of the drive system but for the normal operation investigating the windage loss reduction that could enhance drive system performance during all operating conditions. In the large power-speed requirements as proposed for the propulsion system, windage losses will be substantial and careful attention to the design of the necessary shrouding and other techniques will be useful to the efficient operation of these systems. For the emergency or loss-of-lubrication requirements for rotorcraft (30 minutes of operation in this mode), several technologies require continued research to reduce the weight and size of the system necessary to operate the rotorcraft after occurrence. One possible solution being considered for future rotorcraft is called vapor/mist lubrication. This system injects a special, once-through lubricant into strategic high-heat generation areas of the drive system to provide the necessary lubrication to permit the rotorcraft to operate and finish the mission (flight segment).

Diagnostics/prognostics are another technology area that needs further and continued research to improve detection, location, and time to failure capabilities. This technology area will improve safety (possibly lower insurance and other cost elements) and permit maintenance on a “condition based” assessment. False alarms need to be eliminated from current capabilities and missed failures are not acceptable. Current advancements such as data fusion and fuzzy logic techniques need to be further refined for advance components (such as ceramic bearings, journal bearings, housings, clutches, etc.).

Also, manufacturing technology for the large mechanical components envisioned for the heavy lift rotorcraft is necessary. While large components are made for various applications (non-aerospace), making gears, bearings, and housings to the large sizes envisioned for this program requires careful consideration. Special processing (machining, processing, heat treatment, final grinding, etc.) techniques will require development because of the tolerances required for successful mechanical system operation. For the heavy lift rotorcraft this also means investigating low cost manufacturing while improving component performance (power to weight). Technologies such as near net forging require further investigation for application to gears and other components for the drive system.

Finally, an enhanced analytical development at the basic design level for optimizing configuration layout would also be of great use for this and future programs. Currently, mechanical systems are configured conceptually by the designer then worked at the design level on a gear or bearing mesh basis through the gear system. An optimization process using a minimal amount of input information would be useful for arriving at a satisfactory drive system arrangement early in the design stage. The process would also help reduce system design cost.

Scaled Subsystem Demonstration Once the variable/multi-speed configuration is chosen for the heavy lift rotorcraft, the arrangement needs to be validated on a scaled or subsystem basis. This will require an analytical effort as well to determine scalability of the technology for the full scale demonstration to follow. Validation testing will include the entire variable/multi-speed aspects of the drive system on a subscale basis. Testing will be used to validate the configuration and to understand the operation constraints or boundaries. The testing will simulate the overall system requirements so that the tests are valid for the large full scale tests to follow.

Full Scale Propulsion System Demonstration A full scale demonstration of the propulsion system will need to be performed. At this time there is not an available testing facility in the U.S.A. that could be utilized for the complete propulsion system, and therefore a new facility will need to be designed, fabricated, and built.

The full scale propulsion system will be based on attributes, capabilities, and compromises found in the subscale tests. The full scale propulsion test will require the entire system as a ground test article. The system test is necessary to validate operation of the complete system through the entire expected flight spectrum. As mentioned earlier, the propulsion system operation should be transparent to the pilots and passengers from vertical to forward flight, and should not detract from passenger acceptance. Also, program objectives for power to weight, cost, and reliability must be validated for the propulsion system at this stage.

Structures and Materials Low-Velocity Impact and Impact Damage Tolerance For heavy lift rotorcraft, new rotorcraft structural concepts must be developed to meet strength, weight, and performance goals. The fuselage must be able to support much higher loads than previous rotorcraft. These new structural concepts may include sandwich construction with integrated stiffeners, thick skins, thick cores, and new advanced lightweight cores (e.g., X-Cor). The rotorcraft will also likely include some type of through-thickness reinforcement such as Z-Fiber. Advanced joint concepts will also be developed. All of these new structural concepts and construction techniques must be evaluated for susceptibility to impact damage and subsequent damage tolerance. In addition, the effect of cyclic loading on the impact damage must also be investigated. Damage from low-velocity impacts, or static concentrated out-of-plane loads, is a major threat to the structural integrity of this type of rotorcraft. Debris or hail, accelerated by the vehicle propulsion system, will impact the exterior surface. In addition, cargo handling and passenger usage will cause interior impacts and high out-of-plane concentrated loads. Maintenance activities will also result in both interior and exterior impact damage. Before these new structural concepts and construction techniques can be applied, the impact performance must be well understood and simplified design and analytical tools must be developed that can predict the impact damage and the corresponding post-impact performance and strength. The designers will require these types of tools to correctly size and design the structure to minimize weight and improve performance.

Delamination and Debonding Analysis Tools To date, analysis codes have not been used to certify composite rotorcraft components. Of primary concern is the need to characterize and analyze delamination and debonding as mixed-mode fracture involving crack opening (mode I), sliding shear (mode II) and tearing shear (mode III). Furthermore, although separation of energy release rates into contributions from the three fracture modes has been routinely achieved in research studies, additional work is required to provide efficient analysis tools that can be easily integrated into design practice. The virtual crack closure technique (VCCT) is widely used for computing energy release rates based on results from continuum (2D) and solid (3D) finite element analyses to supply the mode separation required when using the mixed-mode fracture criterion. However, present techniques typically require geometric non- linear finite element analyses with additional post processing routines that are currently not an integral part in most commercial codes. ABAQUS recently announced the release of a new add-on for ABAQUS 6.5 called ABAQUS-VCCT. The other large commercial finite element codes such as MSC NASTRAN or ANSYS, which are frequently used by the helicopter industry, do not offer the choice for calculating mixed mode energy release rates VCCT. Full implementation of Interlaminar Fracture Mechanics (ILFM) in design requires a continuing development effort of codes to calculate energy release rates, advancements in delamination growth criteria under mixed mode conditions, and consideration of three-dimensional geometry and spectrum loading including out-of-plane loads.

Flexible Multifunctional Composite Structures Future advanced vehicles will require flexible structures to achieve improved aerodynamic performance. For example, helicopter rotor hubs are subjected to bending and torsion loading, as well as tension. Currently, composite rotor hubs are manufactured which accommodate these loads, by varying the flexbeam thickness along the length. Studies to-date have investigated the effect of combined tension and bending loads on sub- scale tapered flexbeam components. These studies have applied a fracture-mechanics approach, using finite- element modeling along with laboratory testing to understand and predict the durability and damage tolerance of these flexbeams. This fatigue life methodology, which relies on accurate data from delamination characterization tests, has been applied to several different geometries and materials with reasonable success.

However, many issues remain to be studied. Current investigations have been limited to 2-D models, using delamination characterization data from fatigue tests in modes I and II. There are currently no standardized mode II tests (static or fatigue) and no mixed-mode fatigue test standard. In addition, the flexbeam modeling should be extended to full 3-D, which would require mode III material data. Current tests have also been limited to constant amplitude loading. The effects of a more realistic spectrum loading should be investigated.

Embedded sensors, both active and passive, have been proposed as a means of adding multi-functionality to aircraft structures. The viability of these embedded sensors in flexbeams or other highly flexible structures should be investigated to determine the effect of the embedded element on the durability of the structure and the integrity of the embedded component under fatigue loading. In addition, the effect of typical manufacturing flaws on the fatigue life of rotor and wing components will need to be evaluated. A fatigue life methodology to establish accept/reject criteria for manufacturing flaws in rotor and wing components must be developed.

Through-Thickness Reinforced, Self-Healing, Polymeric Matrix Composites One of the major limitations of laminated polymeric matrix composite structures is the potential for the formation and growth of ply interface cracking, known as delamination. This problem may be alleviated using through-thickness reinforcement, which involves the placement of continuous fibers or stitches through the laminate thickness. An example of this technology involves pultruded carbon/BMI pins inserted through the laminate thickness using an ultrasonic method. The pins (known as z-pins, due to the insertion direction) resist the opening of delaminating surfaces, and thus enhance the resistance to delamination growth.

Additionally, the energy dissipated as the z-pins are pulled out from the laminate result in a significant increase in the fracture toughness. Tests have shown that a 2% areal density of z-pins results in an increase of twenty times the mode I fracture toughness of a standard carbon/epoxy laminate. In addition, new polymeric matrix materials are being developed that incorporate microcapsules of uncured resin that release when a delamination crack grows between the layers, which reacts with a catalyst in the matrix to heal the crack. To date these self-healing polymers have only been developed in resin and filament wound composite form with room temperature cured matrices. Furthermore, a finite amount of time is required for the self-healing reaction to occur. The resistance to delamination growth supplied by the z-pins will suppress the opening displacements between the delaminating surfaces allowing the self-healing to occur. Hence, the addition of z- pins to polymer matrix composites with these self-healing matrices offers the synergy of extremely durable and damage tolerant lightweight composite structures.

Bonded Joining Technology for Primary Structure Heavy lift rotorcraft will likely utilize bonded joints in the fuselage, wing section and rotor components.

Many candidate material systems and reinforcing technologies exist for use in bonded joints, such as fiber metal laminates, z-pinning in PMC’s for enhancing damage tolerance, and self-healing polymers to enhance the durability of bonded joints. These joints will be subjected to dynamic and cyclic loads that may cause delaminations in the laminate material and debonding of the joint structure. ASTM test standards have been developed to characterize the opening mode I and mixed-mode I/II (opening and sliding shear) fracture toughness. In addition, a standard for characterizing delamination onset under mode I fatigue loading has been developed. Test standards are also under development for evaluating the delamination resistance of laminates subject to mode II and mode III loading, although further work is required to solve problems related to the test methods. In order to fully understand the delamination behavior of bonded joints under a mixed- mode cyclic loading scenario representative of a service load spectrum, the current delamination test standards need to be extended to include the characterization of delamination growth under cyclic loading.

Furthermore, the interaction between the opening, sliding shear and scissoring shear (mode III) fracture modes needs to be understood and characterized. Recent developments in failure analysis need to be applied and extended, utilizing decohesive-zone elements that emulate material damage, to predict the delamination and debonding behavior of joints. Existing analytical models for predicting delamination growth in through- the-thickness reinforced laminates also must be extended and incorporated into the finite element models.

Selectively Reinforced Metallic Structure The next generation of structural materials for aerospace application is selectively reinforced metals.

Selectively reinforced metals are the marriage of the best features of metals and fiber reinforced composite materials. Selective reinforcement (SR) is the local (intelligent) application of reinforcing material to otherwise conventional metal structure to compensate for deficiencies (such as damage tolerance, durability, stiffness and strength) associated with the base material and structure. It is equally applicable to wing, fuselage and rotor structures. Two selective reinforcement approaches are lamination and in-situ. An example of the laminate approach is the adhesive bonding of commercially available fiber metal laminates to conventional metallic fuselage/wing skin and stiffeners. In-situ SR is less mature than the lamination approach although the potential performance gains surpass what is capable via lamination. In-situ SR volumetrically replaces base metallic material with high stiffness/strength reinforcement only where needed, not in a uniform distribution as is performed in PMC materials. Typical total reinforcement volume fraction ranges between 3 and 10 percent. Studies by Alcoa indicate that potential weight reductions using advanced lamination or in-situ SR techniques, as compared to current fiber reinforced PMC materials, range from 10 to 20 percent.

Tailored Wing Structure One of the principal design challenges for high-speed tiltrotor transport aircraft is achieving acceptable proprotor aeroelastic stability margins, which can restrict the operating airspeed of the tiltrotor aircraft in the high-speed airplane mode. In a typical tiltrotor wing design, the rotor pylon pitches up as the wing bends upward in the symmetric wing beamwise bending mode. To increase the stability boundary it is necessary to minimize the ratio of the pylon pitch motion ( θ ) to the wing tip deflection (z) in the fundamental wing mode to increase the proprotor stability by reducing the rotor destabilizing forces. For a conventional tiltrotor composite wing design with structurally balanced skin laminates, the wing provides no structural pitch/bending coupling to resist the nose-up pitch due to pylon mass offsets. Unbalanced composite skins, on the other hand, can create nose-down structural twist as the wing bends upward to offset the pitch-up tendency from the pylon mass offsets. The objective is to reduce the wing weight for a given thickness ratio.

Reliability-Based Design To date, the design procedures for most aerospace vehicle structures are based on combinations of factors of safety on the loads and knockdown factors on the strength. The factors of safety and knockdown factors account for uncertainties in material properties, dimensions, loads, and other quantities that govern a design.

These procedures use deterministic analyses methods and provide no information about the risk of structural failure. A design technology referred to as reliability-based design (RBD) accounts for uncertainty by using probabilistic methods. RBD seeks a design that has a probability of failure that is less than an acceptable small value. The principal benefits of RBD are: increasing the confidence in analysis tools; reducing the design cycle time, cost and risk; and increasing the structures performance while ensuring that reliability requirements are met. In addition to performing a reliability analysis, a sensitivity analysis will be performed to determine which parameters have the greatest influence on the structures. There are several codes (e.g., Unipass, ProFES, RDCS, ROBO, and DAKOTA) that can be used to perform multidisciplinary reliability- based design.

Aeromechanics Heavy lift rotorcraft require high hover efficiency, hence low disk loading and very large rotor diameter. The large size implies high weight fraction, and high speeds introduce stability issues. Productivity requirements drive designs well beyond the limits of state-of-the-art rotorcraft, with severe demands on structural efficiency for low weight and aerodynamic efficiency for high performance. Meeting these challenges requires innovative designs (probably with unconventional dynamics), a comprehensive and highly integrated design optimization process, and advances in design/analysis tools and supporting test data. In the following sections, several key areas of aeromechanics are discussed: aerodynamic efficiency, dynamics and loads, active control, and wing design.

Aerodynamic Efficiency: Efficient Rotor and Aircraft Aerodynamic efficiency is key to economically-competitive heavy-lift rotorcraft. Existing aerodynamic databases and analysis tools are inadequate for design of such large rotorcraft. Required airframe drag levels have not been achieved by the rotorcraft industry. Edgewise rotor hub drag at the level of current helicopters would not be acceptable.

Optimized efficient rotors . The rotor system must be optimized for aerodynamic efficiency in both hover and cruise while achieving maneuver capability in edgewise flight. Tiltrotors operate in axial flow at the critical design conditions, but require high twist in cruise which compromises hover efficiency. Optimum twist cannot be matched for both hover and cruise, and reduced cruise tip speed will lead to a greater difference in optimized rotors than any yet tested or flown. Three-dimensional flow effects are known to substantially improve tiltrotor lift capability in hover, but the hover performance database is extremely limited for highly twisted rotors, adding risk to design and optimization.

Conventional edgewise rotors face a severe, rapidly changing aerodynamic environment at high advance ratios, with large areas of reverse flow, which is greatly exacerbated at the high speeds required for heavy lift rotorcraft productivity. Compound rotorcraft use a wing to unload the rotor in cruise, thereby eliminating lift imbalance, but there remains the challenge of reducing high-speed drag while retaining good hover performance and adequate stability when unloaded. Coaxial rotors balance lift in cruise with counter-rotating rotors, but (roughly) half the rotor is non-lifting at high speed, adding weight and drag. Slowed-rotor concepts deliberately place much of the rotor in reverse flow, further increasing the drag penalty.

All edgewise rotors require airfoils that have high maximum lift for maneuvers, good lift/drag performance in forward flow at both low (hover) and very high (cruise) speeds, and very low drag in reverse flow. Tiltrotors must compromise blade twist distribution, therefore their airfoils must have good behavior (low drag and pitching moment) at non-optimum angles of attack. Airfoils must be developed uniquely for high-speed rotorcraft. Heavy-lift rotorcraft will extend the airfoil environment to higher Reynolds numbers than existing rotor airfoils, and their higher speeds will expand the range of Reynolds numbers over which rotor airfoils must retain good performance.

Blade planforms are constrained by structural requirements: high taper reduces weight, but may reduce performance. Hub/blade structural integration may further constrain aerodynamic optimization. For good hover performance, coning must be limited, which constrains the structural design and affects dynamics.

Local blade section lift-curve slope strongly affects stability at high speeds. Consequently, aerodynamic optimization cannot be carried out independently of rotor dynamic design.

Component drag reduction . Edgewise rotors, especially coaxials, suffer from very high hub drag. Tiltrotors experience spinner/root interference, and design solutions are complicated by the need for very large blade pitch travel.

For the LCTC and LABC, the heavy lift rotorcraft program must identify candidate passive (fairings) and active flow control methods, with detailed analysis and component tests to develop practical low drag concepts, concluding with large scale tests to demonstrate the selected concept(s). A parallel effort must be carried out for LCTR spinner drag reduction.

Aerodynamic interference. High-speed rotorcraft face significant flow-interference effects: rotor/wing interference for compounds and tiltrotors; rotor/rotor for coaxials and tandems; rotor/tail and rotor/prop for compounds, including coaxials; and rotor/inlet for compounds (and to a lesser extent, tiltrotors).

A special case of interference is wing download for compounds and tiltrotors, which is a substantial penalty.

The fuselage affects download, so a high-wing vs. a low-wing layout can, in theory, make a significant difference, but there is no comparative database to guide design decisions. Numerous methods of reducing wing download have been proposed, notably those including active aerodynamics, but few have been tested at adequate scale.

Auxiliary propulsion. Compound and coaxial rotorcraft require efficient, low noise propellers for cruise propulsion. Although propeller design is highly developed for fixed-wing aircraft, those machines do not suffer rotor/prop interference effects. Therefore, the LCTC and LABC will require thorough analysis and testing of auxiliary propeller designs, including large-scale wind-tunnel tests of a complete airframe in order to properly quantify interference effects and demonstrate good performance.

Test data and design tools. Available analysis and design tools have been proven only for conventional designs, or for advanced concepts at small scale; some proposed design features have never been tested at adequate scale. There exists a very large database for edgewise-rotor performance, loads, and stability, but little at large scale (e.g., CH-53), and nothing at both large scale and high speed. There have been no large compounds (excepting a few, marginally successful British or Soviet machines with non-representative technology). There has been only one high-speed coaxial helicopter (XH-59) and but a few high-speed compound helicopters (e.g., AH-56, RSRA), none with nearly the required performance or payload; none reached production.

The large-scale test database for tiltrotors is limited to a handful of rotors, all with similar twist distributions, and none optimal for LCTR. Only one coaxial rotor has been tested at large scale (at least in the U.S.), and only at limited speeds. No large-scale, slowed rotors have been tested at anywhere near the required speeds.

The available database against which to test design tools is consequently inadequate. In addition, rotor/airframe interactions, including rotor/wing lift sharing, have not been fully studied for the proposed concepts.

Extensive aerodynamic performance tests are, therefore, required to validate design and analysis tools. They include medium- to large-scale wind tunnel tests, with representative blade and control-system dynamics, to define and confirm aerodynamic behavior and predictive capability. At least two separate test series are warranted: a near-term series with extrapolations of existing technology (e.g., new twist distributions for tiltrotors), and a longer-term series with advanced technology (e.g., new structural concepts). Both low- and high-speed tests are required, because of the radically different aerodynamic environments. These tests should lead to an integrated, full-scale wind tunnel test utilizing propulsion, structures, and rotor systems from earlier benchmarks—the goal being demonstration of full system integration of the down-selected concept(s), ultimately leading to flight test.

Considerations for airframe aerodynamic design are covered in detail elsewhere. The issue is relevant here because of rotor/airframe interference. There must eventually be a complete airframe wind-tunnel test, including rotors, wings, tail, and auxiliary propulsion (as appropriate for the down-selected concept), to demonstrate that aerodynamic interference issues are properly understood and controlled for in the final design. Specific tests include LCTR wing download and LABC rotor/prop interference.

Dynamics and Loads: Light-weight, Strong, Stable, Low-maintenance Rotor System Reducing the weight fraction of the rotor system is crucial to economically-competitive designs, and low disk loading is essential for acceptable hover efficiency. The consequences of substantially reduced weight and low disk loading include large, relatively light rotors with novel hub and control concepts. Such rotors will have radically altered dynamic characteristics compared to current rotors. The objective is a light-weight design that is inherently stable. Emerging active load-control technology should be integrated into the design to minimize loads and hence weight.

Blade and hub concept. Tip speeds must be kept subsonic, leading to low rpm for large diameter rotors, but light-weight construction will prevent blade frequencies from scaling down with rpm. Large, light rotors will consequently exhibit higher frequencies relative to rpm than existing designs. A major objective of the heavy lift rotorcraft program is to exploit this relative frequency shift to reduce the weight penalty that would otherwise be prohibitive for very large rotors.

Instability mechanisms are different for tiltrotors (axial flow), compounds (edgewise flow), and coaxials (stiff rotors). Moreover, rotor/wing dynamic coupling creates the threat of whirl flutter for tiltrotors: although the wing can be stiffened to compensate, this adds undesirable weight and otherwise compromises the wing design. Hingeless, gimbaled, and articulated rotors have different susceptibility to instabilities, hence have different requirements for active stability augmentation (if used) and impose different stiffness requirements on tiltrotor wing design. Hingeless rotors are simpler and lighter than gimbaled or articulated rotors, and have better high-speed stability but higher loads. There is, therefore, a tradeoff between loads and stability; risk/payoff issues and demonstrated technology levels drive heavy-lift designs to hingeless rotors with loads control.

Although not needed for the concepts studied here, active stability augmentation should not be neglected. Of immediate relevance is the requirement that active controls not introduce any instabilities. Passive stability enhancements (e.g., structural tailoring) may be required to fully exploit the potential of some designs.

Coaxial rotors are also susceptible to instabilities, but their designs tend to be driven by the absolute requirement to avoid rotor/rotor tip contact. They require extremely stiff blades, structural coupling to minimize blade tip clearance, and/or active flapping control. They may also require active control for good cruise performance, which must be compatible with stability and contact avoidance.

Also desired are a minimum number of components, simple conditional monitoring, and ease of inspection and repair. Designs must be compatible with durability and damage tolerance requirements.

Active load control and automatic limiting. Rotor designs must withstand extreme dynamic loads without incurring excessive fatigue damage. Compared to traditional metal blades, composite blades have better fatigue characteristics but are less tolerant of large load excursions. Rotor load limiting, implying active control, is needed for full utilization of the structural capability of modern materials. It is here assumed that there will be no significant extension of materials technology beyond the best already qualified; therefore, the greater loads capability demanded of heavy-lift rotorcraft must be achieved by improved rotor design and control concepts.

The requirement for active control of loads can be approached two ways: by imposing limits upon the flight condition, which implies integration with the full aircraft flight-control system and compatibility with good handling qualities; or by active rotor load control, which requires monitoring the rotor state (loads, flapping, etc.) to determine appropriate adjustments to the rotor controls. Advances in automated flight-control technology may eventually eliminate the distinction between these load-control methods, but they are treated separately in the current studies.

Tiltrotors require a high collective range, which constrains control-system kinematics, which in turn strongly affects control-coupling dynamics. Coaxial rotors require dual controls within a small package to minimize hub drag. These issues complicate the application of active loads control to tiltrotors and coaxial rotors.

Concept development and demonstration. Because of its criticality to safety, a demonstration test of aeroelastic stability is mandatory for any candidate rotor concept and should be done early in the program. A small-scale wind tunnel test of a representative system is adequate. (However, “small-scale” is a relative term; a 1/5-scale model of the LCTR would easily dwarf any existing tiltrotor model.) This requires early identification of design concepts that address stability issues (whirl flutter for LCTR, high advance ratio for LCTC and LABC).

Large-scale wind tunnel tests of dynamically scaled rotors and hubs will be needed to fully evaluate down- selected concepts. Test metrics include performance, loads, vibration, control, and noise. Useful tests can overlap stability demonstrations if the flight envelope is restricted or, in the case of LCTR, the rotor system is tested in isolation.

Active load-control and automatic load-limiting are essential for the hingeless rotor concepts herein proposed.

The heavy lift rotorcraft program must define the approach—flight condition limiting or active load control— and develop the associated control concepts, including hub moment control and rotor state feedback that are all compatible with stability, handling qualities, and maneuver requirements. Active-control technology will require small-scale wind tunnel tests to demonstrate viability and refine concepts appropriate for heavy lift rotorcraft, followed by medium or large scale tests to demonstrate capability and stability. Flight tests will be required to fully demonstrate load control, including gusts response and maneuvers.

Active Control: Noise and Vibration Reduction Noise and vibration requirements cannot be met solely by design parameter selection or passive techniques.

Active control will be required.

Types of active control. Application of active controls to rotorcraft may be divided by frequency domain and by risk/payoff considerations. Loads and flapping control require low-frequency inputs, typically less than 1/rev, whereas noise and vibration reduction, performance improvements, and stability augmentation require higher frequencies, sometimes at several harmonics of 1/rev.

Active control (including 2/rev IBC) promises to improve cruise performance for slowed rotors by reducing drag in reverse flow further than can be achieved by traditional, sinusoidal cyclic controls.

Stability augmentation entails high risk, in that the consequence of failure is immediately catastrophic, as do some applications of flapping control (rotor/rotor separation for coaxials, and rotor/wing clearance for tiltrotors). None of the concepts considered here depend upon active stability augmentation or separation control. A failure of loads control may lead to rapid consumption of fatigue life, but need not endanger the aircraft, and can permit large increases in structural design allowables. Therefore, loads control should be strongly considered where it enables significant improvements in the design (a good example is hingeless rotors). Failures of active noise or vibration reduction, or of active performance improvement, have little direct impact on safety, and can therefore be freely employed. Because noise- and vibration-reduction techniques apply to all rotorcraft concepts considered in this study, they are given emphasis here.

Design of vibration-control and noise-control concepts. Design requirements for active controls vary with rotorcraft concept. Active controls may be applied to any of several areas: noise control, vibration reduction, gust-response limiting, and performance enhancement. An important consideration is the establishment of effective metrics for performance, loads, vibration, gust response, and noise, in order to perform effective design tradeoffs and rigorously evaluate competing designs.

Several means of implementing active controls have been proposed, roughly divided into rotor or airframe controls. Individual blade control (IBC) and active swash plate concepts (higher harmonic control, HHC) apply high-frequency control inputs at the blade root. A rapidly-growing technology is on-blade controls via flaperons or local active twist, which promises both lighter systems and more highly optimized (spanwise- distributed) control of the rotor. On the airframe side, active-isolation systems are already in production, but their weight penalty is high. Active noise control (anti-noise) has also been demonstrated. Active aerodynamic control, principally high-frequency flaperon inputs on tiltrotors and compounds, is yet another candidate. All of these methods face challenges when scaled to the size of heavy-lift rotorcraft.

Development and demonstration. Detailed, mutually-compatible requirements must be established to guide and evaluate active-control developments, including integration of noise and vibration control with load alleviation, flapping control, gust response control, and performance improvement, which are all consistent with handling qualities and stability requirements. A comprehensive series of wind-tunnel tests will be required, first at small scale to develop concepts, then at larger scale to demonstrate successful integration with down-selected rotor concepts. This effort includes selection of the control method(s) (IBC, on-blade, or airframe).

Large-scale wind-tunnel and flight tests of active controls are already planned and should be assiduously pursued; they include UH60-IBC and SMART rotor tests. As advanced concepts mature, flight tests on existing aircraft will demonstrate control of noise, vibration, gust response, and possibly performance, and will evaluate handling qualities. Full demonstration of loads control and stability will require integration with new rotor concepts, including rotor dynamics, which will probably dictate a dedicated demonstrator aircraft.

Wing Design for Large Tiltrotors Because whirl flutter and wing download present technology challenges for tiltrotors (and to a lesser extent, compounds), the wing design requires careful attention. Compounds can use pusher props, so whirl flutter is not an issue; coaxial concepts do not need a wing. For these and other reasons discussed below, the technology effort is focused on tiltrotor wings.

Design constraints. A tiltrotor wing must accommodate a transmission cross-shaft, and preferably include a nacelle-tilt shaft. For download reduction, the wing must also have full-span, large-chord flaps with very large deflections (up to 90 deg). The wing is tip-loaded in beam bending for hover and low-speed maneuvers, and the concentrated tip masses (engines and transmissions) drive the wing structural dynamics. Moreover, large in-plane forces generated by the rotors at high-speeds can couple with the wing modes to cause whirl flutter; high torsional stiffness is required for stability. The wing should also accommodate emerging download-reduction technology (e.g., active aerodynamics). Fixed-wing design practices are inappropriate to meet these collective requirements.

The large beam and torsional stiffnesses required for tiltrotors result in wings with unusually thick cross sections, compared with fixed-wing aircraft. Thinner wings have lower drag, but higher weight. Purpose- designed airfoils are needed to simultaneously maximize aerodynamic and structural efficiency.

The LCTR baseline design is a low-mounted wing, in contrast to current practice (e.g., V-22). The advantages over a high wing are a lighter, simpler structure to carry landing gear loads between fuselage and wing; no sponsons are needed for landing gear, hence lower drag; and a potential reduction in download (no fountain over the fuselage). Design constraints include fixed engines with tilting shafts, longer rotor shafts or extreme dihedral for fuselage clearance in hover (for cabin noise), and hingeless rotors for adequate pitch control power in hover. A serendipitous fallout of the low-wing configuration is that a hingeless rotor tends to be less susceptible to whirl flutter, so the wing need not be as torsionally stiff as would be required for a gimbaled or articulated rotor. However, a hingeless rotor will require load alleviation. For these reasons, the wing and rotor cannot be designed independently of each other.

The LCTR wing structural design is driven by a 2-g jump takeoff and VSTOL pullout requirements.

Combined with low cruise rpm, the lowest wing/nacelle frequency is greater than 2/rev. This is a very different design constraint than applies to any existing tiltrotor, so the wing structure cannot be extrapolated from current (V-22, BA-609) design practice.

Although a strength-designed wing, combined with a hingeless rotor, has no whirl-flutter issues, an advanced wing design could evolve into a low-drag, low-weight structure with inadequate torsional stiffness for stability. Therefore, flutter-alleviation measures should be considered as potential research areas.

Demonstration of whirl-mode stability is required in any event. Stability-enhancement technology includes aeroelastic tailoring of the wing (bending/torsion coupling), active flutter suppression via high-frequency rotor control inputs, and possibly passive rotor design measures (aeroelastic tailoring, planform optimization, or mass distribution).

Tests and demonstrations. Test requirements include demonstration of combined rotor/wing aeroelastic stability. Small-scale, semi-span wind-tunnel tests are adequate. Also required are download measurements (covered above in the context of aerodynamic interference).

The wing airfoil(s) will be unique and must be designed and tested independently of rotor airfoils.

Flutter-suppression concepts, if employed, must be rigorously tested. It should be possible to develop the technology at small scale, relative to heavy lift rotorcraft (1/10-scale or even smaller), but a larger scale (at least 1/5) demonstration would eventually be necessary. Any flutter-suppression technology should be flight- tested on a demonstrator before commitment to a full-up design.

Acoustics External Acoustics – Rotor Noise Sources The heavy lift rotorcraft concepts that are being proposed under the NASA Rotorcraft Sector will need to address both interior and external noise issues through design and operational procedures. The acoustic goals set by NASA are very ambitious, given that there is currently no data available for the proposed heavy lift rotorcraft concepts or for any rotorcraft vehicles of that size and class. Existing rotorcraft which are currently considered to be “heavy lift” vehicles are still significantly (20 to 50%) lighter and smaller than the proposed NASA heavy lift rotorcraft concepts. Currently, of vehicles considered to be “heavy lift” rotorcraft, all are of military origin and include the CH-53E, CH-47, and V-22. These vehicles, being of military origin, are not designed, nor are they required, to meet the strict FAA/ICAO noise certification standards. For a successful design, it is necessary for the heavy lift rotorcraft vehicles, at a minimum, to meet current noise certification standards.

Current noise certification regulations and community noise requirements are based on noise metrics designed to emphasize human audible responses to noise. For example, FAA noise certification requirements place a limit on the allowable Effective Perceived Noise Level (EPNLdB) that a helicopter may produce under given flight conditions. These metrics, being based solely on human aural response, do not capture effects of large levels of noise below approximately 50 Hz. Due to the large diameter rotors which are required to operate at low rotational frequencies, fundamental and blade passage frequencies potentially will be below approximately 20 Hz. As such, while there will be significant acoustic energy in the audible range for these heavy lift rotorcraft vehicles, there will also be significant acoustic energy in the infrasonic (below approximately 20 Hz) frequencies. It will be critical to determine and understand the physiological and psychological effects of these repetitive very low frequency and infrasonic stimuli on the human body and community before these concepts will be acceptable.

Some of the noise sources known to be very annoying to the community when they occur are blade-vortex interaction (BVI) noise, high-speed shock (HSI) noise, blade-wake interaction (BWI) noise, and blade self- noise. BVI noise can be one of the most annoying sources when it occurs. BVI noise typically occurs when the vehicle is in a mild descent, which would likely occur as the vehicle approaches an airport, flying over populated areas. When BVI noise occurs, it dominates the frequency range from 6 to 40 times the blade passage frequency (BPF). For typical helicopters flying today, this frequency range is generally in the most sensitive frequency range of human hearing. For the heavy lift rotorcraft vehicles, this frequency range spans from about 60Hz to several hundred Hertz. Active control has shown promise as a tool to mitigate BVI noise.

Broadband noise sources, such as BWI and blade self-noise, typically dominate during level and mild ascent conditions. These conditions generally occur when the vehicle is departing an airport, which in most cases requires flight over populated areas. These broadband sources typically occur in a frequency range from 20 to 100 times the blade passage frequency. For conventional sized vehicles of today, this frequency range is in the mid- to upper range of human hearing, and can be of levels that dominate the spectrum in the absence of other noise sources. Because of the very low BPF associated with the heavy lift rotorcraft vehicles, these broadband noise sources will be shifted to lower frequencies and hence in the most sensitive part of human hearing.

Depending on the particular configuration of the heavy lift rotorcraft vehicle, other rotor and non-rotor noise sources could contribute significantly to the acoustic signature of the vehicle. One example of a non-rotor noise source is engine noise. Heavy lift rotorcraft vehicles will require multiple, large engines. The design, configuration, and placement of these engines could contribute significantly to both exterior and interior noise.

Since no vehicles currently exist in the proposed heavy lift rotorcraft-class of vehicles, acoustic prediction methods will be an important tool to aid in the design of a vehicle which meets noise requirements. Acoustic prediction methods are necessary and becoming more essential in assessing acoustic characteristics of designs, computationally testing the effects of design changes, and optimizing flight operations to minimize noise. In order to understand the limitations and improve prediction, the methods must be comprehensively validated with measured data for configurations which are as similar as possible to the heavy lift rotorcraft concept vehicles.

Validation of prediction methods requires that a comprehensive flight test and wind tunnel test database be available. Currently, such a database does not exist. Acquisition of such a database is a critical step which must be developed in the initial years of a heavy lift rotorcraft effort. Current “heavy lift” rotorcraft vehicles that could be used to meet these criteria are the CH-53E, the CH-47, the V-22, and the Mi-26. For a complete flight test program, a full range of flight conditions would include level flight, hover, climb, descent, and low speed mild-maneuver conditions (e.g., steady turns, mild accelerations, and mild decelerations). To adequately capture major acoustic characteristics, a full “U-shaped” microphone array is needed. This type of array consists of a linear ground array and vertically placed microphones at the two ends of the linear ground array. Such an array allows the vehicle to fly through the center of the array, capturing the noise directivity. In addition to data from the microphone arrays, human “sound jury” data is an essential component of a complete database. A human sound jury consists of volunteers who listen to vehicle sounds, whether located at a flight test site or in a sound booth. While listening to the sounds, the sound jury provides various types of feedback on how the sounds affect them. These responses can then be used to develop noise metrics to categorize human responses and to determine which noises provide the highest levels of annoyance.

With a prediction method that has been validated against adequate and applicable measured data, and with metrics that adequately mimic human responses, heavy lift rotorcraft concept vehicles can be evaluated, candidate low noise designs and controls can be computationally examined and down-selected for testing, and flight operations can be optimized to minimize noise in conjunction with other design requirements.

Psycho-acoustics Heavy lift rotorcraft will generate extremely high levels of noise in a variety of ways. Large diameter rotors require operation at very low rotational rates to keep the compressibility effects at the rotor blade tips to a minimum. Periodic loading on these blades operating at low rotational rates will generate high levels of very low frequency noise—a significant part of the sound energy will be in the infrasonic range below 20 Hz.

Though this is not in the audible range, research in other areas has shown that exposure to repetitive low frequency noise is a concern for humans. The low-frequency sound energy can also cause secondary results such as building vibration. Current Federal Aviation Administration (FAA) regulations for certification of helicopters do not adequately address these issues because of the use of certain noise metrics (e.g., EPNL) that have weightings or tone corrections which are focused on the audible range. Quantification of low frequency noise effects requires definition and development of noise metrics that take into account physiological and psychoacoustic effects. Effects such as the noise impacts on the community and the long term exposure to high levels of low-frequency noise on humans (flight attendants, crew, etc.) must also be quantified in order to assess noise reduction or noise mitigation technologies. Quantification of physiologic and psychoacoustic effects will require development and construction of ground testing facilities with new capabilities to synthesize and accurately reproduce representative acoustic environments. New computational tools and advanced models of human response to such environments will need to be developed for incorporation into system noise prediction tools.

Interior Acoustics As with exterior noise, interior noise goals for heavy lift rotorcraft concept vehicles are very aggressive.

These aggressive interior noise goals are necessary to make the vehicle cabin environment acceptable to passengers and crew. The two major paths for interior noise transmission to the cabin are the airborne path and structural path. Which paths contribute to the interior noise levels is configuration dependent.

Noise transmitted along airborne paths is typically dominated by noise from the rotor(s). For example, rotor noise impacts the fuselage and is then transmitted through the walls of the fuselage. The fuselage structure provides some measure of acoustic damping to the incoming noise; however, this damping is frequency dependent and typically very small for very low frequencies. In this case, the structure is nearly “acoustically transparent”, providing almost no inherent noise mitigation. Fortunately, the airborne noise mechanisms are configuration dependent. For example, edgewise rotors could have less noise impacting the fuselage than a tilting prop-rotor in propeller mode. Substantial research is needed to examine combinations of fuselage structure and airborne excitation and how to minimize the noise transmission through these structures. The nature of this type of interior noise is such that it requires near-full-scale testing of hardware. Active noise mitigation of low frequency airborne noise is expected to be a challenging research area and will require innovation. This is because conventional active noise reduction techniques would require very large (undesirable) amplitude structural deflections; conventional passive techniques would require placement of significant quantities of noise dampening material (heavy). Prediction of interior noise from airborne excitation will require significant integration between external noise prediction methods and interior noise analyses.

Noise transmitted along structural paths can contain noise from transmission shafts, bearings, gears, gear boxes, engines, etc. These noise sources originate in structural vibrations, which excite the air near the structure, resulting in acoustic radiation into the cabin. These noise sources are also configuration dependent.

For example, interior noise is expected to be higher if cross-shafting or gear boxes are placed in or near the fuselage cabin. Component testing to analyze these various noise sources can be conducted initially at small scale. Analysis of vibro-acoustics of these components also requires further development of prediction tools which include both structural vibration and structural acoustic analyses. To be successful and applicable, the interior noise research effort must be done in conjunction and in coordination with both exterior noise analyses and psychoacoustic analyses.

Noise Reduction/Mitigation There are numerous strategies that could be employed to reduce or mitigate the effects of exterior and interior noise from rotorcraft. Most strategies fall into one of two categories: reduction of source noise or mitigation of noise using flight operations.

Source noise reduction typically may come in the form of passive blade design or active control, both of which are designed to minimize some aspect of noise generation. Many of the recent noise reduction efforts have focused on reduction of BVI noise. Passive designs aimed at reducing noise typically have an unconventional planform (e.g., non-rectangular planform, non-square tip region, non-linear twist, sweep, etc.), which is tailored to minimize or mitigate a particular noise mechanism. Active controls have been applied at the blade root in the form of Higher Harmonic Control (HHC) and Individual Blade Control (IBC).

Both methods have been successful in demonstrating reductions in BVI noise and vibrations under certain controlled conditions; however, there is typically a penalty of increased low frequency noise associated with these methods. Distributed active controls have been applied in the form of active twist, which has been shown to reduce BVI noise and vibration under controlled conditions. Other on-blade controls such as active flaps show promise in controlling BVI noise sources. While methods of this class have shown promise in controlling BVI noise and/or vibration within certain frequency ranges for conventional sized vehicles and models, application of these techniques to heavy lift rotorcraft class vehicles must be explored to determine their ability to be cost-effective in controlling BVI noise.

Currently, there are no known effective and efficient methods of actively or passively controlling very low frequency noise without significant performance penalties. The heavy lift rotorcraft concept vehicles are expected to generate significant low frequency noise due to rotor size, low rotor rotational rates, thrust required, etc. Mitigation of the low frequency noise will require identification of innovative and revolutionary active noise control concepts. For example, since rotorcraft noise tends to be very directional in nature, using an innovative active control strategy may be possible for canceling or re-directing regions of high noise radiation away from populated areas.

Noise generation and mitigation can also be affected by flight operational procedures. For example, under NASA’s Short Haul Civil Tiltrotor (SHCT) Program, a noise reduction of 6 dBA in the BVI noise component was demonstrated by changing the descent flight profile of an XV-15 aircraft. This was necessary in order to reduce noise while maintaining safe and flyable vehicle handling qualities. Recently, many vehicle manufacturers have begun providing guidance on low noise operations for their vehicles for various flight scenarios. A similar approach must be taken with the heavy lift rotorcraft concept vehicles. However, since heavy lift rotorcraft vehicles will have very different handling qualities than those of existing vehicles, development of safe flight procedures that also mitigate noise impact on the community need to be developed. These efforts must at a minimum encompass vehicle dynamics, handling qualities, and acoustics.

Handling Qualities Configurations designed for a conceptual civil heavy-lift vertical flight transport must meet the primary mission requirements of range, speed, and payload while also satisfying regulatory and operational requirements of civil operations. Historically, most production Vertical Take-Off and Landing (VTOL) designs have been derived from military aircraft or requirements, but civil airworthiness and operations regulators have imposed requirements unique to civil operations. This section outlines some of the design considerations important to civil operations, beginning with reference to the FAA’s Rotorcraft Airworthiness Criteria. Considerations for aircraft stability and control will be noted as well as one engine inoperative criteria and instrument operations. Noise abatement considerations will be discussed. Finally, aircraft configuration control and the use of secondary surfaces will be discussed.

Civil Certification and Operation Transport Category Rotorcraft are certified for airworthiness according to FAR Part 29. The focus of Part 29 is on transport category aircraft as opposed to utility types certified under Part 27. An important distinction is the assumption that transport category aircraft will carry passengers or cargo for hire, requiring a higher degree of safety. The military never quite makes this distinction, although different handling qualities design standards have been developed for attack and scout, utility, and medium lift helicopters. Ultimately, though concerned with safety, the military is most concerned with mission accomplishment. Civil transport aircraft designs and operations must meet a safety standard first, the mission comes second. The design and operations considerations of this distinction can become significant.

Although FAR Part 29 is designated for rotorcraft, it provides the regulatory underpinnings for other VTOL designs. The FAA definition of rotorcraft includes helicopters and autogyros. Tiltrotor aircraft are considered “powered lift.” A draft airworthiness criterion, Part XX, was developed for powered lift, but it uses much of the same criteria as rotorcraft for operations near hover. The Bell-Agusta 609 will be certified to criteria largely drawn from Part 29, augmented by appropriate fixed wing criteria (Part 25) for airplane mode operations and selected powered-lift criteria for tilt-mode conversion operations. The criteria and philosophy behind Part 29, based on decades of helicopter operation, will continue to provide the underpinnings of civil VTOL certification criteria for the foreseeable future.

Most of the civil airworthiness criteria for handling qualities and operations are contained in Part 29, Subpart B—Flight. This subpart has criteria for weight and center of gravity limits, performance (including engine out performance), flight characteristics (including static and dynamic stability), ground and water handling characteristics, and vibration. Other subparts deal with strength requirements (C), design and construction (D), the power plant (E), equipment (F), and operating limitations and information (G). Requirements for one engine inoperative performance and operation, and for aircraft stability, will be discussed further below.

One Engine Inoperative Transport category aircraft must be designed and operated such that they can sustain a major system failure and continue a safe operation to either a landing or continued flight. An important failure is that of an engine, as specified in “Category A” operations. The operation and aircraft performance must be designed such that the aircraft, upon failing one engine (one engine inoperative, OEI) can either return to and stop safely on the takeoff surface or continue the takeoff and climb-out at a specified minimum rate, speed and configuration.

Given the typical rotorcraft power-required curve and installed power, this usually means the aircraft has a substantial region of height and velocity conditions (close to the ground and at low speed) that must be avoided.

Meeting the Category A OEI requirements typically requires either long clearways (including runways) or special operations. As an example of long runways, the commercial Chinook required a clearway and landing surface on the order of 2400 feet at the design gross weight to satisfy the land-back requirement of a balked takeoff. Some Category A helicopter operations satisfy the OEI requirement by backing up from the middle of a small landing zone, always keeping the land-back point in sight until they reach a safe height where altitude can be traded for airspeed for a continued takeoff and climb-out. Such operations typically use an alternate weight, lower than that certified for use with a long clearway (runway).

The Category A OEI requirements tend to drive commercial operations toward short takeoff techniques when flying with high payloads. Lower payloads are carried when pure vertical operations are required. This provides landing zone flexibility where the allowable payload is matched to the space available for takeoff and landing. The implication for conceptual design is that at least two mission alternative weights should be considered: a maximum gross weight for a defined cruise mission that will use a STOL technique and a lower mission weight that accounts for hovering operations for takeoff or landing.

Aircraft Stability The aircraft stability requirements of FAR Part 29 tend to be looser than military design standards. The FAA must set minimum required flight dynamics and stability, while the military, as a customer, can define a desired result. An aircraft that marginally meets the FAA’s minimum criteria in one area, often falls short elsewhere, so prudent design comes in from the boundaries a bit. In contrast to the FAR Part 29 requirements, the military Aeronautical Design Standard for Handling Qualities, ADS-33, has a well-constructed, modern mathematical basis, backed up by extensive simulation and flight test. The mathematical basis of ADS-33 provides for mathematical evaluation at a much earlier stage in the design process—essentially when a linear flight dynamics model first becomes available. ADS-33 can serve as a flight dynamics benchmark to at least point to areas needing further control design work or analysis for a civil design.

Inner Loop Control Dynamics An issue for large rotorcraft is the frequency overlap that can occur between aeroelastic modes, drive train modes and/or flight dynamics and control modes. While civil certification criteria does not explicitly address this (except for the prohibition of “excessive vibration under each appropriate speed and power condition”), any resonance or confluence of dynamic frequencies could have serious consequences. Most recent large/medium rotorcraft projects have encountered some form of this. Dynamic analysis during the conceptual design stage should help identify the potential for such dynamic issues. Further, the frequency overlap issue is expected to place limitations on tolerable aeroelastic modes and on advanced blade control concepts. Flight control implications will have to be dealt with as that design matures, but early identification of potential interactions might lead to alternative design solutions.

Transport Category Instrument Operations A large civil transport VTOL is expected to operate in near all-weather conditions. The NASA concept for a Runway Independent Aircraft (RIA) envisions commercial airline use of these transports. To achieve the desired airport capacity augmentation, these RIA must operate in the same weather and visibility conditions as the long haul conventional airliner fleet. Low visibility and icing are the primary challenges.

Icing protection will be needed equivalent to that employed by conventional airliners. In addition, the mission cruise altitude for the conceptual design needs to be above 22,000 feet to establish cruise flight above the icing altitude band. The consequence of a lower cruise altitude is an aircraft that cannot fly during some atmospheric conditions that conventional airplane transports are flying or that must carry extra deicing equipment which is usually complex, heavy, and has high power demands.

Operations in low visibility and turbulent winds tend to drive control designs to significant stability augmentation. Typical low visibility operations use gentle maneuvers and low acceleration to maintain a stable operating point. Without visual references for attitude and speed, the aircrew and flight control system must rely upon instrumentation to provide control around the desired operating condition and flightpath. The need for aircraft stability drives control designs to tighter control of body attitude, driving the response frequencies higher, potentially driving into the overlapping frequency issues cited above.

Although instrument flight operations are conducted without visual reference, eventually a visual reference is required, at least for the final landing touchdown. A concern to FAA operations certifiers is the maintenance of a visual sight line to the intended landing spot throughout the approach, should visual conditions permit.

High body pitch attitudes that raise the nose and instrument panel into this line of sight are frowned upon.

Passenger comfort may also produce a similar pitch attitude constraint. Helicopters, with their rotor shafts fixed relative to the body, can encounter blanking of the critical line of sight as they pitch nose-up to decelerate along the approach path. While yawing to provide a side view of the landing sight-line may be acceptable to military and some commercial operators, this is not expected to be acceptable to the envisioned RIA VTOL operation.

The design implications of this line of sight design standard may affect allowable operations with a given payload (the need for a slower approach than otherwise required), require the use of additional tail deflection, flaps or other pitch moment producers, or limit other desired operations such as noise abatement approach profiles.

Ironically, the approach sight-line design goal may not be satisfied even by use of full-aft nacelle angle in a tiltrotor aircraft. Use of full aft nacelle angle probably places the aircraft in an undesirable flight configuration for recovery from an engine failure. As seen during the studies of the NASA SHCT Program, use of aft nacelle had to be limited until after a commitment to landing was made, very late in the approach. A corollary to this is that nacelle angle movement becomes a primary longitudinal acceleration control. As such, nominal flight operations must be planned to not use the full range of the control movement, leaving some margin to pilot discretion.

The implication to aircraft designers is that not all of the static performance of an aircraft may be used to satisfy nominal requirements. Margins beyond the nominal are required.

Noise Abatement The aerophysics of rotorcraft flight produce the potential for blade vortex interaction (BVI) at positive angles of attack of the rotor disk plane. BVI has been identified as a major noise source for rotorcraft. Positive rotor angles of attack typically occur during descent and deceleration, both of which occur on approach to landing.

The rotorcraft noise problem begins with approach. Careful tailoring of the approach profile—airspeed, rate of descent, deceleration, and rotor shaft angle for variable configurations has been demonstrated to impact the production of BVI and resultant noise footprint on the ground.

The state of the art for noise abatement operations design and analysis assumes quasi-static flight conditions.

This tends to drive these operations into the same slowly changing operations needed for instrument flight operations. Some of the flight conditions desired for reduced noise can lead to unacceptably high body pitch attitudes. In general, noise abatement operations must use the same nominal range of body pitch attitudes that satisfy the approach line-of-sight goal cited with instrument operations.

Additional Aircraft Features Prompted by both noise abatement operations and control issues during approach, secondary surfaces or devices may be useful and important on a new heavy-lift VTOL. High flap angles were used during the SHCT XV-15 noise abatement flight tests to increase the drag on the aircraft, lower the body pitch attitude and increase the power setting required for descending flight. While lowering the pitch attitude should be understandable from both line of sight and comfort considerations, the increased drag and required torque may be counter-intuitive. Increased engine torque places a turbine engine in a more linear response range, good for flight control where power is being used to control flightpath angle and height. In addition, a turbine engine operating more toward its mid-range can respond better to a sudden demand for full power such as might occur in the event of an engine failure.

Additional devices to increase drag during approach may be desirable. A rapidly cleaned-up device could be valuable to rapid recovery from an engine failure during approach. Designers should look for opportunities to provide such devices, especially if they might serve a performance goal such as download reduction during hovering flight.

Summary Flight operations, performance, and handling quality considerations impact civil rotorcraft design with design constraints not immediately apparent from basic mission requirements of range, speed and payload. One engine inoperative operations requirements may define alternative mission weights. Flight stability, dynamics and controls considerations may limit allowable aeroelastic mode frequencies. Noise abatement and instrument operations may place even greater constraints on stability requirements and blade control concepts. While handling qualities and operations issues may not drive the initial design, these considerations properly belong in the design iteration loop once a basic configuration has been defined.

Risk Reduction Tasks For High Torque, Lightweight Drive System (Propulsion, Figure B3) P ROPULSION : I NNOVATIVE C ONCEPTS F OR L IGHT -W EIGHT , H IGH -T ORQUE , L OW M AINTENANCE A ND C OST Justification: Heavy lift rotorcraft design requires an innovative configuration to handle high torque at low weight, with variable speed operation capability. Propulsion system weight, cost, and efficiency depend on technology gains from advanced components. Advanced materials, processing and design can result in substantial cost reduction and improved reliability.

Design—Transmission and Engine System Configuration Study Variable speed drive system concept definition (Y1) Define mechanism components, weight penalties, operational characteristics, location for mechanism, fail-safe operation Variable speed engine capability defined (Y1) Define speed range capability, SFC penalties, identify risk reduction issues for implementation in Heavy Lift RC Choose drive system or engine speed or a combination to meet heavy lift rotorcraft requirements (Y1) Component Development and Concept Validation Choose component development/advanced concepts that provide increased capabilities for power to weight, lower cost, and increased reliability (Y1-4) Choose concept(s) to achieve variable/multi-speed operation of the propulsion system (Y2) Analyze, design, and fabricate variable speed concepts (Y2-3) Conduct sufficient parametric tests to ensure scale effects are understood (Y3-4) T RANSMISSION : S CALE S UB -S YSTEM V ALIDATION , F ULL S CALE T ECHNOLOGY D EMONSTRATION Justification: Advanced configurations necessary for heavy lift rotorcraft propulsion will require scaled sub- system development and validation followed by full system testing of the advanced variable/multi-speed propulsion system.

Small Scale Demonstration/Transmission Arrangements Choose scale demonstration concept arrangement(s) (Y3-4) Develop scale demonstration system configuration test arrangement (Y4) Design and fabricate scaled variable speed concept(s) (Y4-5) Validate speed change propulsion system operation using scaled loading to correctly represent vehicle characteristics (Y5-6) Full Scale Propulsion Demonstration Choose full scale demonstration configuration (Y5) Using concept validated in scale test—develop design of full scale proof-of-concept propulsion system (Y5-6) Design and fabricate current or new system to test the heavy lift rotorcraft propulsion system (Y6-7) Conduct life cycle, over-load, and loss-of-lube testing for drive system qualification (Y7-9) Risk Reduction Tasks For High Performance, Structurally Efficient Rotor/Wing System (Structures, Figure B4) M ATERIALS C HARACTERIZATION A ND I NSERTION : I MPROVED S TIFFNESS , S TRENGTH , D URABILITY , D AMAGE T OLERANCE Justification: Material characterization data are prerequisite to advances in damage tolerance and durability.

Material Characterization Composite delamination Delamination in composite components, and debonding in bonded joints, for materials used in validation of analytical tools and TRL benchmark demonstrations (Y1) Delamination fracture toughness and fatigue resistance of the embedded sensor-to-composite bond for composite materials with embedded sensors (Y2) Delamination in composites with through-thickness reinforcement under mode I, mode II and mixed-mode I and II loading conditions, both static and fatigue (Y2-3) Material and failure mode for fiber-metal laminates (Y3), for selectively reinforced metals (Y4-5) Non-structural properties of multifunctional materials (such as damping, EM shielding, energy attenuation, RF radiance/attenuation) (Y6-7) Material Screening Identify advanced fiber metal laminates that have improved stiffness, strength, durability, damage tolerance and lower density (Y2) Identify multifunctional materials (structural plus electrical, damping, sensing, etc.) (Y2) Identify advanced in-situ selectively reinforced metals that have improved stiffness, strength, durability, damage tolerance and lower density (Y4-5) Manufacturing Process Development Very large, complex, integrated rotating components (blades & hubs) and non-rotating structure (such as bulkheads, frames, roof structure) (Y2) Develop methods for enhancing delamination resistance for all three fracture modes using through- thickness reinforcement concepts (Y2) Laminated metals (Y3) Highly anisotropic tailored components (Y3-5) Prepreg from fiber reinforced materials with self-healing matrices cured at autoclave temperatures (Y4-5) In-situ selectively reinforced metals (Y4-5) D URABILITY AND D AMAGE T OLERANCE (D&DT) Justification: Certification requirements for damage tolerance must be satisfied to achieve a viable light- weight rotor system and airframe. Improvements in durability and damage tolerance key to effective utilization of advanced materials.

Low Velocity Impact Evaluate impact damage resistance and identify unique impact damage modes of new structural/construction concepts and joints (Y1-2) Conduct composite post-impact static and fatigue testing to determine critical damage modes that affect strength and life (Y2) Evaluate effects of panel size and curvature on the impact response of new structural/construction concepts and joints (Y4-5) Continue evaluation of impact damage resistance and identify unique impact damage modes of new structural/construction concepts, joints, stiffeners, cutouts, and large attachments (Y4-5) Improve and verify the analytical models for impact damage resistance and damage tolerance (Y6-7) Fatigue Investigate spectrum loading effects on typical flexible structures for rotorcraft (flexbeams, blades, wing components, etc.) (Y2) Manufacture coupon-size flexible components with embedded sensors and conduct fatigue tests to determine durability of the component and sensor (Y2) Manufacture and test composite stringer pull-off specimens with self-healing matrices and z-pin reinforcement (Y2) Fabricate and test rotor and wing components with implanted flaws to determine the effect of typical manufacturing flaws on fatigue life (Y4-5) Develop probabilistic methods for structural certification (Y4-5) Develop fatigue life methodology to establish accept/reject criteria for manufacturing flaws in rotor and wing components (Y6-7) Demonstrate enhanced damage tolerance and durability provided by through-thickness reinforced, self-healing, polymeric matrix composites via test and analysis of heavy lift rotorcraft structural components (Y6-7) S TRUCTURAL E FFICIENCY : I NNOVATIVE C ONCEPTS F OR L IGHT -W EIGHT , D URABLE, R ELIABLE , L OW M AINTENANCE S TRUCTURES Justification: Heavy lift rotorcraft will require significant weight reductions in rotor system, airframe, and wing to be feasible. Structural design concepts are required that make use of the advances in damage tolerance and durability.

Design Concepts Develop structural concepts (wing/fuselage skins, joints, cutouts, concentrated load points, frame skin interface discontinuities) For components made with fiber-metal laminates (Y2) Utilizing multifunctional structures (Y2) Utilizing smart/adaptive structures (Y3) For joining (bonded & bolted) anisotropic & isotropic composites and metals for tailored structure (Y2) For components made with selectively reinforced metals (Y4-5) Analytical investigation of buckling (and post buckling) strength of anisotropic panels for tailored structure (Y3) Design, fabricate, and test Anisotropic panels for tailored structure (Y4-5) Tailored box beam with anisotropic top and bottom panels and balanced laminated in the other panels (Y4-5) Structural elements and small components using selectively reinforced laminated metals (Y6-7) Tailored rotor blade spar section (Y6-7) Tailored wing box beam and rotor blade spar (Y6-7) Risk Reduction Tasks For High Performance, Structurally Efficient Rotor/Wing System (Aeromechanics, Figure B5) A ERODYNAMIC E FFICIENCY : E FFICIENT R OTOR A ND A IRCRAFT Justification: Aerodynamic efficiency is key to economically-competitive heavy lift rotorcraft. Aerodynamic database and analysis tools are inadequate for design of heavy lift rotorcraft. Airframe drag levels required have not yet been achieved by rotorcraft industry. Edgewise rotor hub drag at level of current helicopters would not be acceptable.

Optimized Efficient Rotors Rotor system optimized for aerodynamic efficiency in hover and cruise (Y2-3) Airfoils (appropriate LCTR/LCTC/LABC hover and cruise environment; maneuver capability in edgewise flight; high Reynolds number; high advance ratio for LCTC/LABC) Geometry (including planform and twist, integrated with dynamics; for hover and cruise environment) Minimize aerodynamic interference (including rotor/wing for LCTR/LCTC, rotor/rotor, rotor/tail, rotor/engine, rotor/prop interference for LCTC/LABC) Minimize download Aerodynamic performance tests and validated design/analysis tools Medium to large scale wind tunnel tests, with representative blade and control system dynamics; define and confirm aerodynamic behavior and predictive capability (Y3-5) Medium to large scale wind tunnel tests, define and confirm aerodynamic behavior and predictive capability utilizing advanced technology (Y6-7) LABC: Active control (including 2/rev IBC) to improve cruise performance LCTC/LABC Hub Drag Reduction Identify candidate passive (fairings) and active flow control methods (Y1) Analysis and component tests to develop low drag concepts (Y2-3) Large scale tests to demonstrate concepts (Y3-4) Efficient Airframe Analysis and optimization, including interference (Y2) LCTR spinner drag reduction (Y2-3) Small scale and component tests (Y2-3) Complete airframe wind tunnel test (Y4) LCTC/LABC auxiliary propulsion Efficient, low noise propellers for cruise propulsion (Y2-5) Risk Reduction Tasks For High Performance, Structurally Efficient Rotor/Wing System, and For Super-Integrated Vehicle Management System (Aeromechanics, Figure B5) D YNAMICS A ND L OADS : L IGHT -W EIGHT , S TRONG , S TABLE , L OW M AINTENANCE R OTOR S YSTEM Justification: Low weight-fraction for rotor system crucial to economically-competitive design. Consequence of substantially reduced weight and low disk loading is large, lighter rotors with novel hub and control concepts. Such rotors will have radically altered dynamic characteristics compared to current rotors.

Objective is design that is inherently stable and takes advantage of active load control to minimize loads and hence weight.

Design: Blade and hub concept (Y1) Solution to stability issues (LCTR whirl flutter, LCTC high advance ratio) Solution for strength and weight LABC: Structural coupling concept to maximize transient blade tip clearance Design: Load control (Y1) Requirement for active control of loads; approach (flight condition limit, active load control, other) Aeroelastic stability demonstration Small scale wind tunnel test of representative system (Y2-3) Load active control and automatic limiting Develop control concepts (Y1-2) Including hub moment control and rotor state feedback Compatible with handling qualities and maneuver requirements Small scale tests Medium or large scale wind tunnel tests to demonstrate capability and stability (Y3-7) Flight tests to demonstrate load control, including maneuvers (Y4) A CTIVE C ONTROL : N OISE A ND V IBRATION C ONTROL Justification: Noise and vibration requirements can not be met by design parameter selection or passive techniques. Active control will be required.

Design: Vibration and Noise Control (Y1) Requirement (noise, vibration, gust, performance control) Control method (IBC or on-blade or airframe) Develop and Demonstrate Active Control Control development for selected approach (Y1-2) Including integration with flight control system Wind tunnel tests (Y2-7) Including integration of noise and vibration control with load alleviation, flapping control, gust response control, performance improvement Flight tests to demonstrate effectiveness and reliability UH60-IBC and/or SMART (Y1) Advanced control method (Y4) Risk Reduction Tasks For Low Noise Aircraft (Acoustics, Figure B6) N OISE C HARACTERIZATION OF L ARGE R OTORCRAFT Justification: Low external and internal noise are essential for a truly market responsive vehicle. Heavy lift rotorcraft will have unique acoustic characteristics, and the goals are very aggressive.

Flight test measurements of noise of existing heavy lift vehicles (Y1-3) Exterior noise and interior noise Low frequency, long range propagation effects Identify dominant noise sources Psychoacoustic effects for low frequency noise (Y1-3) Low frequency noise simulator tests to develop human response database Assess/develop metrics to model response to low frequency noise exposure Demonstrate initial noise prediction capability (Y2-4) Exterior noise and interior noise Models of human response Conducted with aeromechanics and handling qualities tasks Advanced noise prediction capability (Y2-8) Wind tunnel tests of aeroacoustics unique to heavy lift Low noise procedure prediction/optimization methods Low frequency propagation modeling for community noise impact Interior noise prediction methods, including effects of passive/active models N OISE R EDUCTION C ONCEPTS FOR L ARGE R OTORCRAFT Justification: Low external and internal noise are essential for a truly market responsive vehicle.

Combination of noise reduction technology and safe flight operations is required to achieve community acceptance.

Design: noise reduction (Y1) Approach (design features, active control); certification and community impact Design: vibration and noise control (Y1) Requirement (noise, vibration, gust, performance control); control method (IBC or on-blade or airframe) Low noise operation procedures (Y2-6) Identify viable low noise flight operations (includes procedures) / maneuvers (coordinated with handling qualities) Flight test of initial low noise operations using existing vehicles (Y4) Interior noise reduction strategies (Y2-6) Develop viable active/passive interior noise reduction strategies Small to medium scale component testing of active/passive interior noise reduction (coordinated with structures and propulsion) Develop and demonstrate active control Control development for selected approach (Y1-2) Wind tunnel tests (Y2-7) Including integration of noise and vibration control with load alleviation, flapping control, gust response control, performance improvement Flight tests to demonstrate effectiveness and reliability UH60-IBC and/or SMART (Y1) Advanced control method (Y4) Low noise rotor (Y2-8) Active/passive low noise model rotor concepts and wind tunnel testing Refined and optimized active/passive low noise rotor design and wind tunnel testing Down select for full scale rotor design development and fabrication (Y4) Flight test demonstration of low noise design approach and operating procedures Large scale demonstration of interior noise reduction Risk Reduction Tasks For Super-Integrated Vehicle Management System (Handling Qualities, Figure B7) H ANDLING Q UALITIES AND F LIGHT CONTROLS FOR VTOL OPERATIONS IN CAT IIIC CONDITIONS Justification: Airframe structural frequencies of large rotorcraft will impact flight control and maneuver frequencies, necessitating an integrated design. Near all-weather operations (Category IIIC) capability needed for commercial schedule reliability, safety, and competition with fixed wing airliners.

Adaptive/robust fully augmented control system, including handling qualities with low rotor and airframe structural frequencies and load limiting or control Develop and refine certification requirements (Y1) Flight control system analysis and design (Y2) Moving-base simulation to explore concepts (Y2-4) In-flight simulator tests to prove concepts (Y2-6) Cockpit automation for pilot supervision of complex RIA operations flight tasks in congested airspace System analysis, design, and simulation (Y2-3) Flight tests to prove concepts (Y3-6) OEI/AEI O PERATIONS Justification: Transport category safety required for certifiable operations.

Develop and demonstrate OEI/AEI requirements and procedures Analysis, design, and simulation (Y1-2) Development and demonstration of certification requirements (Y2-4) Strategic Direction The strategic direction provides guidance for selecting highest priority activities, aimed at the four highest risk areas of heavy lift rotorcraft development. Note that there are some important and difficult tasks that are not yet high risk, including rotor aerodynamic design and optimization, airfoil design, airframe aerodynamics, and airframe structures.

H IGH TORQUE , L IGHT W EIGHT D RIVE S YSTEM Innovative design is required for low drive system weight. Large size implies high torque and high weight fraction, hence drive system weight reduction is essential for an efficient and economical aircraft. The focus must be on design concept, advanced-technology components, and materials.

Low maintenance is required for low operating cost. Low maintenance must be a primary design requirement, even ahead of weight and performance.

High flight speed requires, or at least benefits from, a variable speed propulsion system design. First it is necessary to establish the speed range available from advanced engine technology, and to define the engine required for the heavy lift rotorcraft concept.

H IGH P ERFORMANCE , S TRUCTURALLY E FFICIENT R OTOR /W ING S YSTEM Innovative rotor and wing design is required, probably with unconventional dynamics. Large size implies high weight fraction, high speed introduces stability issues, and good rotor system performance is essential for an efficient and economical aircraft. The focus must be on integrated rotor/wing performance and dynamic behavior.

Structural efficiency is required for low rotor and hub and wing weight. The focus must be on design concepts for durability and damage tolerance.

Low maintenance is required for low operating cost. Low maintenance must be a primary design requirement, even ahead of weight and performance.

L OW N OISE A IRCRAFT New approaches are required to meet the challenge of low noise. Large size implies low frequency noise and expanded acoustic footprint. An understanding of heavy lift vehicle acoustic phenomena (low frequency and relative distance to community) is required, including psychoacoustics for low frequency. New rotor design guidelines and annoyance metrics must be developed. The focus must be on a combination of rotor design, active control, and flight operations.

S UPER -I NTEGRATED V EHICLE M ANAGEMENT S YSTEM Broad spectrum active control is required for an effective heavy lift rotorcraft. Large size implies a significant influence of low frequency airframe elastic modes on flight dynamics. Active control is required to achieve the goals of low rotor-induced vibration and noise. Safe operation in one-engine inoperative conditions is essential for civil rotorcraft. Rotor load limiting and active control are needed for full utilization of the structural capability in the rotor and airframe. Hence an expanded integration of the vehicle management system is required: a flight control system for good handling qualities and gust response, active control of vibration and noise, and rotor load limiting and active control. The focus must be on load limiting and system integration.

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Risks addressed Advanced Technology Thrust Full-scale Propulsion X Early Technology Thrust P4 P4 System Ground Test Full-scale Structure X X S4 S4 Ground Test Large-scale Rotor System Wind Tunnel R5 R5 X X X Test Flight Simulation Test Q4 Q4 X X Noise and Control F5 F6 Flight Test X X X Integrated Large-scale d X X I5 X I5 X Wind Tunnel Test Requirements Design system Prototype Development D7 Fabrication Propulsion Requirements Design D7 Super-integrated vehicle management Fabrication Aero/structurally efficient rotor/wing system Low noise aircraft High torque, lightweight drive system BLACK represents Early Technology; RED represents Advanced Technology; Numbers = TRL Figure B1. LCTR Technology Readiness Level Benchmarks.

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 d d Configuration features d d Engine system d Transmission system d Structural concept d d Blade and hub concept d d Load control d d Vibration and noise control d d Noise reduction d d Handling qualities d d Cost reduction d d all other program elements g decision ground test wind tunnel test flight test d w f Figure B2. Concept development.

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Design d d Transmission system Engine system d d Component development and g g d d g d concept validation Small scale demonstration of transmission arrangements g g d g Full-scale transmission g g d d demonstration TRL Benchmarks Full-scale propulsion system P4 P4 ground test Integrated large-scale wind d I5 I5 tunnel test g decision ground test wind tunnel test flight test d w f Figure B3. Risk reduction tasks for high torque, lightweight drive system (propulsion).

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Design d d Structural concepts Blade and hub concept d d Materials g g g g Material characterization Material screening Manufacturing process g g g g development Durability and Damage Tolerance g Low velocity impact g g g Fatigue g g g g Structural Efficiency g g g g Design concepts TRL Benchmarks S4 S4 Full-scale struct ground test Large-scale rotor system R5 R5 wind tunnel test d Integrated large-scale wind I5 I5 tunnel test g decision ground test wind tunnel test flight test d w f Figure B4. Risk reduction tasks for high performance, structurally efficient rotor system (structures).

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Aerodynamic Efficiency Optimized efficient airfoils, w w w w d d d rotors LCTC/LABC hub drag w w d d reduction Efficient airframe w w w d d Dynamics and Loads Design: blade and hub d d concept d Design: load control d w Aeroelastic stability demo w Load active control and automatic limiting w w w w d d f Active control Design: vibration and d d noise control Develop and demonstrate w w w d d f f active control TRL Benchmarks R5 R5 Large-scale rotor system wind tunnel test F5 F6 Noise and control flight test Integrated large-scale wind d I5 I5 tunnel test g decision ground test wind tunnel test flight test d w f Figure B5. Risk reduction tasks for high performance, structurally efficient rotor system, and for super- integrated vehicle management system (aeromechanics).

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Noise characterization for large rotorcraft f f d Flight test existing vehicles Psychoacoustic effects g g Demonstrate noise d w prediction capability Advanced prediction d w f capability Design d d noise reduction vibration and noise control d d Noise reduction concepts g f f d d Low noise operations Interior noise reduction g g d Develop and demonstrate w w w active control d d f f Low noise rotor g w f w d TRL Benchmarks Full-scale struct ground test S4 S4 Large-scale rotor system R5 R5 wind tunnel test Noise and control flight test F5 F6 Integrated large-scale wind d I5 I5 tunnel test g decision ground test wind tunnel test flight test w d f Figure B6. Risk reduction tasks for low noise aircraft (acoustics).

Y10 Y09 Y07 Y08 Y06 Y05 Y04 Y03 Y02 Y01 Design d d handling qualities Handling qualities and flight controls Fully augmented control g g g f f system Cockpit automation g d f d OEI/AEI operations Requirements and g g d d procedures TRL Benchmarks Q4 Q4 Flight simulation test Noise and control flight test f5 F6 Integrated large-scale wind tunnel test d I5 I5 g decision ground test wind tunnel test flight test w d f Figure B7. Risk reduction tasks for super-integrated vehicle management system (handling qualities).

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1. REPORT DATE (DD-MM-YYYY) 2. REPORT TYPE 3. DATES COVERED (From - To) 12/2005 Technical Publication 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER NASA Heavy Lift Rotorcraft Systems Investigation 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 5d. PROJECT NUMBER 6. AUTHOR(S) 1 1 Wayne Johnson , Gloria K. Yamauchi , and Michael E. Watts 5e. TASK NUMBER 5f. WORK UNIT NUMBER 21-065-40-10 8. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER Ames Research Center Moffett Field, CA 94035-1000 and A-0514419 Langley Research Center 1000 NASA Road, Hampton, VA 23681-2199 10. SPONSORING/MONITOR’S ACRONYM(S) 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) NASA National Aeronautics and Space Administration Washington, D.C. 20546-0001 11. SPONSORING/MONITORING REPORT NUMBER NASA/TP—2005-213467 12. DISTRIBUTION/AVAILABILITY STATEMENT Unclassified — Unlimited Subject Category 05 Availability: CASI (301) 604-5642 13. SUPPLEMENTARY NOTES Point of Contact: Wayne Johnson, Ames Research Center, MS243-12, Moffett Field, CA 94035-1000 (650) 604-2242 14. ABSTRACT 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 Civil 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 aircraft 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, structurally efficient rotor/wing system; low noise aircraft; and super-integrated vehicle management system.

15. SUBJECT TERMS Heavylift rotorcraft, Design, Tiltrotor, Compound helicopter, Advancing blade concept 18. NUMBER 17. LIMITATION OF 16. SECURITY CLASSIFICATION OF: 19a. NAME OF RESPONSIBLE PERSON OF ABSTRACT Wayne Johnson a. REPORT b. ABSTRACT c. THIS PAGE PAGES 19b. TELEPHONE (Include area code) Unclassified Unclassified Unclassified Unclassified (650) 604-2242 Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std. Z39-18

Overview

section operating conditions (lift coefficient, Mach collective was varied to trim the rotor to a specified View in PDF

f = ( 1.33(N -1) ) / ( 0.9N ) = 1.1083

f = ( δ L( θ ) ) / ( δ L( θ ) ) = 1.3414

W/A = 2 ρ V [ f f f ] = 2948 f

hover: P = W ( DL/2 ρ ) / M

APPENDIX A View in PDF

APPENDIX A

ROTORCRAFT SECTOR GOTCHA

Hover Efficiency Goal

V-22: W = 53000 lb, P = 12300 HP

CH-53X: W = 80000 lb, P = 18605 HP

1/2

1/2

Efficient Cruise Goal

2/3

Empty Weight Goal

e

weight (lb), and for this purpose

W = gross wt – fuel wt – payload wt – operating item wts – engine wt

e

or

e

system wt + transmission/drive system wt

gross weight = 53000 lb

engine weight = 2(1889) = 3778 lb

fuel + payload weight = 20000 lb

e e

e

e

Community Noise Goal

Flight Control Goal

Advanced Engine Performance Goal

Cabin Noise and Vibration Goals

Develop & evaluate -

APPENDIX B View in PDF

APPENDIX B

RISK REDUCTION FOR HEAVY LIFT ROTORCRAFT

Source & rights

Document details

section operating conditions (lift coefficient, Mach collective was varied to trim the rotor to a specified

APPENDIX A

APPENDIX B