Document
Performance of Advanced Heavy - Lift, High - Speed Rotorcraft Configurations Wayne Johnson*, Hyeonsoo Yeo**, and C.W. Acree, Jr.* *Aeromechanics Branch, NASA **Aeroflightdynamics Directorate (AMRDEC), U.S. Army Research, Development, and Engineering Comman d Ames Research Center, Moffett Field, California Abstract The aerodynamic performance of rotorcraft designed for heavy - lift and high - speed cruise is examined. Configurations considered include the tiltrotor, the compound helicopter, and the lift - offse t o o rotor. Design conditions are hover and 250 - 350 knot cruise, at 5k/ISA+20 C (civil) or 4k/95 F (military); with cruise conditions at 4000 or 30,000 ft. The performance was calculated using the comprehensive analysis CAMRAD II, emphasizing rotor optimizati on and performance, including wing - rotor interference. Aircraft performance was calculated using estimates of the aircraft drag and auxiliary propulsion efficiency. The performance metric is total power, in terms of equivalent aircraft lift - to - drag ratio L /D = WV/P for cruise, and figure of merit for hover.
Keywords: rotorcraft, tiltrotor, compound helicopter, lift - offset, performance Lift (JHL) Concept Design and Analysis (CDA) .
INTRODUCTION activities, work has continued to explore the NASA and the U.S. Army at Ames Research Center aerodynamic capabilities of compound helicopters, and have recently conducted a number of investigations of lift - offset configurations, as well as tiltrotor aircraft.
the ae rodynamic performance capability of rotorcraft designed for heavy - lift and high - speed cruise (Refs. 1 to NASA HEAVY LIFT RO TORCRAFT SYSTEMS 6). The motivations for these investigations include the INVESTIGATION tremendous potential impact of an efficient, large The Rotorcraft Sector within the Vehicle Systems rotorcraft on civil air transportation, and future military Program of the NASA Aeronautics Research Mission requirements for long - range, efficient heavy - lift VTOL Directorate established as its objective the improvement aircraft. This paper summarizes the results of these of public mobility and access to air transportation. The investigations, with emphasis on the performance of technology go als of the Sector originated from industry rotorcraft designed for heavy - lift and high - speed cruise.
studies and workshops during 2001 – 2004 that focused The NASA Heavy Lift R otorcraft Systems Investigation on a new class of vehicles known as Runway (Refs. 1 and 2) examined in depth several rotorcraft Independent Aircraft (RIA). References 7 – 8 showed configurations for large civil transport designed to meet that RIA can relieve runway and terminal area the technology goals of the NASA aeronautics program. congestion by replacin g small aircraft and short - haul The investigation identified the Large Civil Tiltrotor as flights that use primary runways. The primary runways t he configuration with the best potential to meet these would then be used exclusively for larger aircraft and goals. The design was economically competitive, with medium/long - haul flights. RIA would operate from stub the potential for substantial impact on the air runways and/or helicopter landing pads. This transportation system. operational conc ept would increase the capacity of the air transportation system. Reference 9 describes three With the increasing interest in large VTOL aircraft for RIA configurations analyzed by the rotorcraft industry military transport, i ncluding the on - going Joint Heavy (see Figure 1): the quad tiltrotor (Bell Helicopter), the reverse velocity rotor concept (Sikorsky), and th e .
Presented at the AHS International Forum on tiltrotor (Boeing). The studies identified the benefits of Rotorcraft Multidisciplinary Technology, Seoul, Korea, advanced technology and the resulting effects on October 15 – 17, 2007.
operating cost. In summary, Refs. 7 – 9 provide result of structure, drive train, and engine technology.
justification for the overwhelming positive impact that Cost models were developed, and used to estimate the RIA can have on the national air space. purchase price and direct operating cost of the heavy - lift rotorcraft designs. The sizing code was used to perform Usin g the RIA studies as motivation, work was focused sensitivity analyses, first to optimize the aircraft enabling technology for a notional civil VTOL transport (variations including disk loading, tip speed, and capable of carrying 120 passengers at a cruise speed of number of engines), and then to quantify the influence 350 knots at 30,000 ft altitude (Mach number 0.60) with of advanced technology.
a range of 1200 nm (without refueling) . This heavy - lift transport will be “neighborly” quiet when operating The code RC (Ref. 10) was the principal rotorcraft near communities, economically competitive with a sizing and performance analysis tool for thi s Boeing 737 aircraft, and will exploit available airspace investigation. RC was developed by the Advanced and ground space (excluding primary runways). Design Office of the U. S. Army Aeroflightdynamics Specific mission and technology goals were established Directorate, RDECOM. RC inputs include design to push the state - of - the - art in rotorcraft technology. strategy (engine sizing, rotor sizing, etc.), rotorcraft parameters (drag coefficients, tail volume ratio, et c.), The objective of the investigation was to select a heavy - and requirements and constraints (take - off, payload, lift rotorcraft system that has the best chance of meeting range, etc.). RC finds the aircraft that satisfies the the goals while being economically competitive. The designer inputs, then produces the rotorcraft description goals rela ting to aerodynamically efficient cruise and and conducts the performance analysis (Figure 3).
hover, structural efficiency, and low community noise were given highest priority. A NASA - led team of Technology in the sizing code is in troduced in terms of rotorcraft technologists analyzed three vehicle technology factors and performance models. Weights configurations suggested by the rotorcraft industry: a (at the group weight level of detail) are estimated from tiltro tor, a tandem - rotor compound, and an advancing parametric equations based on historical data. These blade concept configuration. These configurations were equations are calibrated to current technology level by deemed, as a first cut, to be technically promising. All comparing with exi sting aircraft. Technology factors are the candidate configurations were assessed for the same then applied to represent the impact of advanced mission and technology goals and detailed analysis in technology. In this approach, technology is a change multiple technology areas was conducted. Extensive from the statistical equation, attributed to a new engineering analysis was performed, including aircraft configuration or concept, new materials, new design design, performance optimization, blade and rotor methods, new operating procedures, etc. There are aerodynamics, airframe aerodynamics, loads and technology factors for blade and hub weight, vibration stability analysis, blade structural des ign, external noise, treatment, drive system weight, and fuselage, wing, and one - engine inoperative requirements, handling qualities, tail weight. Technology also influences performance, in and cost drivers. This approach was highly successful in particular rotor hover and cruise efficiency, hub d rag, attacking this complex design problem. and engine weight and performance.
An assessment of engine and drive train technology was Design Approach and Analysis Tools made in order to define and substantiate the sizing code The approach taken was to design large VTOL models. The engine model represented what could be tra nsports that are economically competitive with obtained from (or required of) modern technology today's regional jet airliners and meet the mission and engi nes.
goals. The principal cost drivers are weight, power, and CAMRAD II is an aeromechanical analysis of complexity. Advances in structural efficiency, helicopters and rotorcraft that incorporates a aerodynamic efficiency, control concepts, propuls ion combination of advanced technologies, including concepts, dynamics solutions, and prediction capability multibody dynamics, nonlinear finite elements, and should allow substantial reductions in empty weight, rotorcraft aerodynamics (Ref. 11). The trim task finds power, and fuel. Low power is ensured by low rotor the equ ilibrium solution (constant or periodic) for a disk loading and low aircraft drag.
steady state operating condition, and produces the The code RC performed the sizing of the rotorcraft, and solution for performance, loads, and vibration. The the comprehensive analysis CAMRAD II was used for flutter task linearizes the equations about the trim performance optimization, and loads and stability solution, and produces the stability results. The calculations. The process is outlined in Figure 2. The aer odynamic model includes a wake analysis to sizing code implemented significant weight savings calculate the rotor nonuniform induced - velocities, using (relative to current technology scaled to large si ze) as a rigid or free wake geometry. CAMRAD II has undergone extensive correlation with performance and The technology factors represent the reduction in VTOL loads measurements on helicopters, tiltrotors, and other cost needed to match CTOL cost trends, excluding the rot orcraft configurations. Complete aeroelastic models influence of complexity and the weight and power were developed for each of the configurations required for hover operation. Insurance, depreciation, considered in this investigation. and finance costs are driven by flyaway price. Baseline cost estimates for the heavy - lift rotorcraft designs were The rotor performance model in the RC sizing code was obtained using the above cost technology factors. A calibrated using the performance calculated by significant part of the differences between VTOL and CAMRAD II, and the sizi ng task repeated. An estimate CTOL costs must be the very different operations that of the drag of the airframe (based on trends for produced the cost data used to develop the models. The rotorcraft and turboprop aircraft) was used to define the remaining differences in cost must be attacked by aerodynamic model for the sizing code and the advanced technol ogy. Note in particular the importance comprehensive analysis.
of maintenance costs, reflected in the large technology CAMRAD II calculations of rotorcraft performance factor.
have received extensive correlation with wind tunnel For the same mission, a VTOL aircraft will have higher and flight test measurements. Reference 11 presents gross weight and higher installed power than a CTOL correlation for several rotorcraft. Reference 12 aircraft. In addition, there are complexity fa ctors in the compares calculations with wind tunnel measurements VTOL model, including number of rotors and number of the performance of the Tilt Rotor Aeroacoustic of blades. Thus there is still a cost of VTOL capability Model (TRAM). References 13 and 5 present in the model, even when the maintenance and flyaway correlation of calculated rotor performance at high price technology factors are used.
advance ratio with NACA wind tunnel measurements (Ref. 14), XV - 1 flight test (Ref. 15), H - 34 full scale Mission and Design Conditions wind tunnel tests (Ref. 16), and UH - 1 full scale wind Based on the notion al vehicle capabilities and tunn el tests (Ref. 17). Reference 18 compares technology goals, the civil mission described in Table 1 calculations of coaxial rotor hover performance for an was defined. This investigation is not intended to NACA rotor (Ref. 19), an AFDD model rotor (Ref. 20), specify the market, but rather to identify enabling and XH - 59A flight test (Ref. 21). Correlation has also technology for civil applications of heavy - lift rotorcraft.
been performed for forward flight performance Note in particular the OEI requirement: at takeoff corre lations with XH - 59A flight test data (Ref. 22), in o conditions (5k ISA+20 C) the contingency power of the both helicopter and compound configuration.
remaining engines (133% OEI MCP) must be greater than 90% hover out - of - ground - effect power required Cost Models (the factor of 90% accounting non - zero speed and some Cost models were developed for VTOL and CTOL altit ude loss during the takeoff).
aircraft, based on statistical information for current operations. The cost metrics considered were flya way For maximum utilization, the aircraft must have a wide cost (purchase price, in 2005 US dollars) and direct range of capabilities. Although the aircraft were operating cost plus interest, DOC+I (in 2005 US designed to the mission defined in Table 1, hence with cents/ASM). The components of DOC+I were very little hover time, efficient hover and low speed maintenance (airframe, engine, rotor and drive), flight capability i s essential to the RIA operational concept.
crew, fuel and oil, depreciation, insurance, and finan ce This is reflected in the requirement for essentially OEI cost. A principal source for the cost models was Ref. 23 hover capability. The resulting designs optimize at and its unpublished extensions. The CTOL cost model balanced cruise and OEI hover power, so the cruise was based on the economics of U. S. airline operations. speed of 350 knots can be viewed as a fallout of the OEI requirement. Reasonable downwash and outwash from In order to compare VTOL and CTOL costs, the two the rotors hovering in ground effect is required for cost models were applied to a Boeing 73 7 - 700 at a stage effective utilization. For example, a download of 20 length of 500 miles. For the 737 in the VTOL cost lb/ft would produce an outwash with a peak velocity of model, the minimum complexity was used, and an over 90 knots. As a result of these con siderations, high installed power trend was used to get an equivalent disk loading aircraft (such as tiltwings) were not among turboshaft power (Ref. 1). The costs are substantially the configurations considered.
higher with the VTOL model. With th ese results it is possible to establish cost technology factors: Maintenance tech factor = 0.9/9.8 = 0.092 Flyaway price tech factor = 48.0/83.6 = 0.57 and drive system weight therefore increases by about a Technology Factors and Design Parameters factor of 2.2. In order to maintain aircraft empty weight Meeting the technology goals requires high speed, high fraction as size increases, the design approach must be altitude, and long range for productivity. The heavy - l ift changed, which conventionally has resulted in an rotorcraft must have low disk loading for good hover increase in disk loading with size.
efficiency, and low drag for efficient cruise. The actual Basic parameters of the rotorcraft were chosen for the disk loadings of the designs were determined based on heavy - lift configurations based on an assessment of minimum aircraft weight, power, and cost. For this current and future technology. The rotor blade loading heavy - lift rotorcraft investigation, the target airframe 2/3 (C / σ ) was chosen considering low speed W and wing drag was D/q = 1.6(W/1000) . This drag maneuverability requirements, with about an 8% level is higher than current turboprop aircraft, although improvement in maximum lift capability, relative to about 35% lower than is customary in the helicopter current technology. A relatively low hover tip speed industry (Figure 4). So good aerodynamic design was used, reflecting the importance of the noise goal.
practice should be sufficie nt to achieve the target for The cruise ti p speed was chosen to optimize the airframe drag. For concepts with edgewise rotors in performance. Hover download values consistent with cruise, hub drag must be added to the airframe and wing current technology were used. A low wing loading was drag of the aircraft. For this investigation, the target hub 2/3 chosen, for good low speed maneuverability and wide drag was D/q = 0.4(W/1000) , which is less than half conversion speed range. The same blade loading and of current hub drag levels (Figure 5). Achieving this hub wing loading d esign values were used for both tiltrotor drag level will require advanced technology, certainly and slowed - rotor compound configurations.
fairings but possibly also active flow control.
The weight technology factors used for the rotorcraft Configurations designs are summarized as follows: 79% for r otor blade Three aircraft configurations were the primary subject weight, 96% for rotor hub weight; 67% for drive system of the Heavy Lift Rotorcraft Systems Investigation: weight; 88% for fuselage and wing weight; 90% for 1) Large Civil Tiltrotor (LCTR) empennage weight.
2) Large Civil Tan dem Compound (LCTC) The definition of the technology level in the sizing code 3) Large Advancing Blade Concept (LABC) also involves performance and aerodynamics. For the These configurations were selected by industry as the rotor, the d esign blade loading C / σ was prescribed, W most promising candidates for the civil mission. The based on an assessment of what advanced technology conventional two - rotor tiltrotor configuration was could provide. Rotor induced and profile power in the considered, since a quad tiltrotor wou ld not present as sizing code were calibrated to the results of the much of a challenge in terms of rotor size. A low rotor comprehensive analysis calculations. Thus the sizing speed was used for the tiltrotor in cruise, to improve the code performance represented a rotor with optimum proprotor propulsive efficiency. The LCTC and LABC twist, taper, cruise tip speed, etc. However, current use edgewise rotors in cruise, hence the rotor rotation technology airfoils were used in the comprehensive must be slowed a s the flight speed increases, in order to analysis optimization. Some further improvement in keep the advancing tip Mach number reasonable. The aircraft performance can thus be expected from the use LCTC is a slowed - rotor compound: it has a wing and of advanced technology ai rfoils, especially if auxiliary propulsion for cruise, so the rotors are specifically designed for these aircraft. Airframe drag operated in an unloaded condition. The LABC uses stiff was specified as described above. Current technology coaxia l main rotors capable of carrying significant roll values were used for hover download. Some further moment, hence generating lift on the rotor advancing improvement in aircraft performance might be obtained side in forward flight. This configuration has been from download reducti on.
described as the advancing blade concept, or lift - offset The statistical weight equations used in the design code rotors. The LABC requires auxiliary propuls ion at high incorporate an influence of aircraft size, based on speeds, but has no wing.
historical trends. For rotorcraft designed to fixed disk The slowed - rotor compound considered had shaft - loading, tip speed, blade loading (solidity), and number driven tandem main rotors. Single main rotor and of blades, these equations im ply that rotor blade, rotor coaxial main rotors are alternate configurations. The hub, and drive system weight scale with gross weight as 1.26 1.39 1.12 number and arrangement of the main rotors affects W , W , and W , respectively. So for an performance through rotor/rotor and rotor/wing increase in gross weight by a factor of 2.0, the rotor interference; and affects the aircraft size because of blade, rotor hub, and drive system weight increase by antitorque and transmission layout issues. An alternative factors of 2.4 , 2.6, and 2.2; and the aircraft structural to shaft drive is a reaction drive configuration, typically the LCTC and then the LABC. At the design stage using jets at the blade tips. The reaction drive is use d in length, the LCTR cost is about 20% higher than that of a hover; in cruise the rotor is operated in autorotation. current 737. That is the cost of VTOL capability. The With reaction drive the transmission weight is greatly LCTR is more economical than the 737 for sta ge reduced, but the rotor cruise performance is lengths below about 200 miles.
compromised by the need for thick blades, and the Figure 11 shows the costs for the LCTR with and hover performance is poor because of high ene rgy without the cost technology factors, and Figure 12 losses entailed in delivering the air to the blade tips.
presents the corresponding DOC+I breakdown. These The heavy - lift rotorcraft designs are summarized in results emphasize and quantify the importance of Table 2. Three - views of the aircraft are shown in controlling the maintena nce costs for heavy - lift Figures 6 – 8. Recall that for these designs the blade rotorcraft.
loading, hover tip speed, and wing loadi ng were specified, based on assessments of the technology.
Assessment of Configurations Cruise tip speed was optimized based on cruise The NASA Heavy Lift Rotorcraft Systems Investigation efficiency. The disk loading was optimized, based on reached the following conclusions (Ref. 1). For the aircraft weight, power, and cost. Basically the optimum NASA civil mission, the Large Civil Tiltrotor has the disk loading produces a balance in power requirement best cruise efficiency, hence t he lowest weight and between cruise and OEI hover. Cruise efficiency defines lowest cost. The LCTR is the configuration with the the power available, then the disk loading is chosen that most promise to meet the NASA technology goals. The uses that power in hover (a larger rotor would increase LCTR design presented is economically competitive the rotor and blade weight, while a smaller rotor would with comparable fixed wing aircraft, with the potential require more power hence more engine and fuel for substantial impact on the air transportation system.
weight). The empty weight fraction is about 65% for all The keys to achieving a competitive aircraft are: low three designs. The fixed weight is comparable to current drag airframe and low disk loading rotors; structural commercial jet aircraft. The drag of the LCTR is weight reduction, for both airframe and rotors; drive comparable to good turboprop aerodynamic design. The system weight reduction; improved engine efficiency; LCTC adds the drag of the hub (less than current low maint enance design; and manufacturing cost technology levels), and the LABC does not have the comparable to CTOL aircraft.
drag of the wing. This LABC design was produced by The Large Civil Tandem Compound has good cruise the sizing code using a rotor cruise performance that efficiency, but less than the tiltrotor, and higher was better than that predicted by the comprehensive development risk than the tiltrotor. Single main rotor analysis.
and tandem rotor configurations we re comparable in Th e aircraft cruise L/D = WV/P (based on cruise power, efficiency and risk. Even if reaction drive produces the including losses, at design gross weight) was the smallest slowed - rotor compound rotorcraft, the high principal efficiency metric. For the mission considered, installed power compromises efficiency, and the the LCTR has the best cruise efficiency, hence the reaction drive system has higher noise and substantially smallest design gross weight and the smallest installed increased risk.
power (Table 2). Next in efficiency is the LCTC, and The Large Advancing Blade Concept has lower cruise after that the LABC.
efficiency than the tiltrotor for the NASA civil mission .
Figure 9 shows the flyaway cost and DOC+I for the Lift - offset rotors are much more effective for missions three heavy - lift rotorcraft configurations, and Figure 10 at lower altitudes, as will be illustrated below.
presents the DOC+I breakdown for the 1200 nm design mission. Th ese figures include the Boeing 737 costs for LARGE CIVIL TILTROTOR (LCTR) comparison. The block hours per year value was based The configuration of the Large Civil Tilt Rotor is shown on Southwest Airlines operations. The difference in in Figure 6. The aircraft has two tilting rotors at the dead time between the VTOL and 737 reflected the wing tips, a low wing, non - tilting engines, and a difference in operations. For the VTOL costs, the horizontal tail. A quad tiltrotor (two wings and four aircraf t parameters (empty weight, installed power, rotors) would have smaller rotors, but increased number of rotors and number of blades) and the mission complexity and increased aerodynamic interference.
parameters (fuel weight, block time and block speed for The conventional two - rotor tiltrotor configuration is a specified range) were obtained from the RC code.
considered here, which allows more exploration of the The VTOL cost model is driven by gross weight and implications of large size on the rotor system design. A power, so the LCTR has the lowest cost, followed by low wing is used for better structural load paths b etween wing, airframe, and landing gear. The horizontal tail is cruise tip speed on the aircraft design parameters. Two sized by trim requirements rather than stability, because transmission configurations are considered in Figure 18: the rotors can be used for flight dynamics stabilization a 1 - speed transmission (engine speed varying with rot or as well as control. A vertical tail is not shown, but could speed), which is the conventional tiltrotor approach; and be added if neede d for yaw trim. A hingeless rotor hub a 2 - speed transmission (engine at optimum speed is used. Excessive coning can significantly reduce hover regardless of rotor speed), with no transmission weight figure of merit, so a tip mass of 1.5 slug was placed on penalty. If the aircraft has a two - speed transmission, so each blade at 95%R. This increases the hover figure of the engine can operate at its opt imum speed in both merit by about 2%. Table 2 gives the aircraft hover and cruise, the improvement in propulsive cha racteristics. Performance, loads, and stability efficiency as the tip speed is reduced results in calculations were performed. Isolated rotor performance significant weight and cost reductions. With a single - in hover and cruise were calculated using a free wake speed transmission, the optimum cruise tip speed is only geometry model. Dynamics characteristics, including 15% less than the hover tip speed.
rotor and wing structural design and aeroelast ic stability calculations, are discussed in Reference 3.
TILTROTOR AERODYNAMICS AND INTERFERENCE The blade twist and taper were varied to optimize the To explore further the aerodynamics of tiltrotor aircraft, rotor for hover and cruise performance. The hover o including wing - rotor interference effects, a tiltrotor was condition was 5k ISA+20 C, 650 ft/sec tip speed, C / σ T designed to carry a 20 - ton payload for 750 miles at = 0.1557. The cruise condition was 350 knots, 30k ISA, o 4k/95 F, with a cruise speed of 300 kno ts (Ref. 4).
350 ft/sec tip speed, and rotor thrust as required to trim Figure 19 shows the baseline design, and Table 3 aircraft drag. The twist distribution had two linear summarizes the aircraft parameters. The gross weight is segments, inboard (0.0R to 0.5R) and outboard (0.5R to 146,600 lb. The aircraft has two four - bladed tilting 1.0R). The linear taper ratio w as varied while rotors at the wing tips, a high wing, and a horizontal tail.
maintaining constant thrust - weighted solidity (constant The baseline aircraft de sign parameters are disk loading 75%R chord). Figure 13 presents the results for twist of W/A = 15 lb/ft , blade loading of C / σ = 0.140, and W optimization, showing the typical hover - cruise wing loading of W/S = 100 lb/ft . The airframe and compromise. The performance was calculated for a wing parasite drag is D/q = 55 ft . This drag value is matrix of inboard and outboard twi st values, and Figure considered aggressive in terms of rotorcraft trends but 13 shows the envelope of all the points. The result was achievable from good fixed wing aerodynamic design an optimum twist of – 32 deg inboard and – 30 deg practice. A hingeless rotor hub is used. The rotors rotate outboard; and an optimum taper of 0.8 (tip/root chord).
with the top blades moving outward in airplane mode. A This choice emphasizes the cruise condition, because parametric study was conducted to understand the cruise dominates the civil mission considered.
effects of design parameters on the performance of the An exploratory investigation was conducted to design aircraft.
airfoils specifically for the hover and cruise operating In addition, a quad tiltrotor was developed (not design ed conditions of the LCTR rotor blades. The resulting using the RC code) from the baseline conventional airfoil contours are shown in Figure 14, and Figure 15 tiltrotor, having the same gross weight, disk loading, plots the opti mum boundary for twist variations with and airframe size. Figure 20 shows the quad tiltrotor both state - of - the - art airfoils and these LCTR airfoils.
configuration, and Table 3 summarizes the aircraft The rotor performance from the sizing code and the parameters. The rotor size was determined t o maintain comprehensive analysis are compared in Figures 16 and the same disk loading as the baseline conventional 17. These results are for current technology rotor tiltrotor. The front wing span follows from maintaining airfo ils. Figure 16 shows the hover figure of merit of the the same clearance between the rotor and fuselage, and rotor. Figure 17 shows the cruise rotor propulsive the front wing chord by maintaining the same aspect efficiency, and the aircraft equivalent L/D = WV/P. The ratio as the baseline conventio nal tiltrotor wing. The aircraft L/D = WV/P is 12.45 at the design cruise speed rear wing span is 40% larger than the front wing span.
(L/D = WV/P = 11.1 including losses, se e Table 2), The rear wing chord has the same chord as the front significantly above current tiltrotor aircraft efficiency. A wing from the tips to the middle of the semi - span and major contributor to this excellent performance is the then linearly increases to the center line. The quarter use of a low tip speed in cruise (54% of hover tip chord line of t he rear wing is kept straight. This design speed). One consequence of this low tip speed is a very 2 approach results in wing loading of W/S = 67.16 lb/ft different result for the twist optimization, compared to for the quad tiltrotor. The rear rotors and wing are current designs. Figure 18 shows the influence of the located above the front rotors and wing. The rotors rotate with the top blades moving outward in airplane wing. The aircraft will trim at larger pitch angle, so the mode. rotors will carry more lift.
Performance calculations were conducted for the design C5) I ncrease of wing span by 10%. To maintain the o cruise of 300 knots at 4000 ft, 95 F condition. same wing loading, wing chord was reduced Rotor/rotor, rotor/wing, and wing/wing interferences accordingly. In this case, the rotors stay at the same were accounted for using the vortex wake model. The wing span as the baseline.
current analysis does not include fuselag e model, which C6) Increase of wing span by 10% (same as C5), but the is known to be important for oscillatory interference of rotors move to the wing tips wit h same radius.
the wing on the rotor, but is not usually necessary for C7) Increase of wing span by 10%, rotors move to wing wing mean induced drag. No nacelle model was tips, with blade radius increased by 12.2% (maintaining considered, thus any end plating effect was neglected.
rotor - fuselage clearance; decreasing disk loading to 11.9 Typical wake geometries and blade and wing lift lb/ft ). To maintain the same blade loading, blade chord distributions for the baseline conventional and quad was decreased ac cordingly.
tiltrotor are shown in Figures 21 and 22 respectively.
Only the tip vortices, which dominate the interference, Figure 23 shows the performance results in terms of are drawn in these figures, but there was a full vortex aircraft lift - to - drag ratio L/D = WV/P, calculated lattice behind e ach blade and wing. The wing wake without accessory or other losses, all for the design model consists of a vortex lattice in the near wake cruise condition of 300 knots. Rotor/rotor and behind the wing with 32 aerodynamic panels, rolling up rotor/wing interferences are accounted fo r using a to tip vortices (with shed wake panels between) in the vortex wake model for both the rotor and the wing, and far wake. Thus, comparable models are used for both the performance results with interference effects are wing and rotor wakes in this investigation of the compared with those without interference effects. The interference.
interference effects changes the aircraft lift - to - drag ratio by up to 2.1% for the parametric variations investigated.
For the conventional tiltrotor, the aircraft was trimmed The reduction of rotor tip speed (C3) increases the using elevator, rotor thrust, and pitch attitude to obtain aircraft lift - to - drag ratio the most and the increase of longitudinal and vertical force and pitching moment wing span (C5) also has a beneficial effect. The change equilibrium of the aircraft. For some cases, rotor of rotor rotational direction (C1) decreases the aircra ft flapping was also trimmed to zero (to reduce loads) lift - to - drag ratio significantly and this effect can only be using rotor cyclic pitch. For the quad tiltrotor, the observed with interference included in the calculation.
aircraft was trimmed using rotor thrust and front and The total effect of changing the rotor direction of rear wing pitch angles. The rotor thrust was used to rotation was – 3.0% of L/D. Rotor disk loading change achieve longitudinal for ce equilibrium and the front and (C2 compared to baseline, and C 7 compared to C6) has rear wing pitch angles were always adjusted so each a small influence on the aircraft cruise performance.
wing carries half of the gross weight. Rotor flapping These results are for flapping trimmed to zero; in was also trimmed to zero using rotor cyclic pitch.
general, flapping trim does not change the influence of A parametric study was conducted for the conventional the parameters, but does reduce the aircraft lift - to - drag tiltrotor, w ith the objective of understanding the effects ratio by up to 1.4% (Ref. 4). The aircraft L/D = WV/P is of design parameters on the aircraft performance. The 7.4 at the design cruise speed (with interference, but following cases were considered.
without losses), and L/D = WV/P = 7.75 with reduced C1) Change of rotor rotational direction (baseline is top cruise tip speed (Fig. 23).
blades outward).
Figures 24 and 25 show the show the rotor propulsive C2) Increase of disk loading to 16.6 lb/ft (re duction of efficiency and wing drag (induced and parasite), rotor blade radius by 5%). To maintain the same blade respectively. These are the same calculations as in loading, blade chord was increased accordingly. Figure 23, except that individual performance components are compared. For the baseline case, there C3) Reduction of cruise tip speed to 350 ft/sec, to is a significant reduction of the wing induced drag increase the propulsive efficiency of the rotor (baseline because of the favorable co mbination of the rotor wake cruise tip speed is 626 ft/sec).
and the wing; and a slight increase in rotor propulsive C4) Reduction of wing angle of attack relative to the efficiency because of the nonuniform flow field from fuselage (thus relative to the rotors) by 3 deg, to the wing interference. The change of rotor rotational investigate the effect of lift sharing between rotor and direction increases rotor propulsive efficiency somewhat. However , it also increases the wing induced drag significantly, thus an overall performance penalty reduction of rotor tip speed increases the aircraft lift - to - is observed. The reduction of rotor tip speed (C3) d rag ratio the most among the design parameters increases the rotor propulsive efficiency as well as investigated, and the increase of wing span also has a decreases the wing drag. Thus, the most performance beneficial effect on the aircraft performance. The improve ment is obtained. The tip speed value of 350 change of rotor rotational direction decreases the ft/sec was selected based on the optimum aircraft aircraft lift - to - drag ratio significantly and this e ffect can performance as shown in Figure 26. The rotor tip speed only be observed with interference included in the was varied from 250 to 450 ft/sec and the optimum calculation.
cruise performance was found at 350 ft/sec tip speed.
Fu rther reductions in rotor rotational speed did not LARGE CIVIL TANDEM COMPOUND (LCTC) improve the aircraft L/D.
The configuration of the Large Civil Tandem Compound is shown in Figure 7. The aircraft has two The reduction of wing angle of attack (C4) changed lift main rotors in tandem configuration, a high wing, sharing between the rotor and wing, reducing wing lift p usher propellers for cruise propulsion, and a horizontal by about 4000 lb and increasing rotor lift by about 4000 tail. The length of the fuselage follows from the lb. The r educed wing lift decreases wing induced drag specification of the payload, and the disk loading was and the increased rotor lift increases rotor induced drag.
optimized to balance the cruise and hover power. As a However, the reduced wing angle of attack also changed result there is no overlap of the rotors . The horizontal wing tip vortex trajectories in such a way as to increase tail is sized by trim requirements rather than stability.
beneficial interference effects, thus t he rotor propulsive Table 3 gives the aircraft characteristics. Performance, efficiency was not changed much. The net effect of the loads, and stability calculations were performed. The reduction of wing angle of attack is a performance comprehensive analysis modelled the auxiliary improvement. The increase of wing span (C5) decreases propulsion as forces app lied to the airframe. Rotor/rotor the wing drag, but slightly decreass the rotor propulsive and rotor/wing interference were accounted for using efficiency.
the vortex wake model. A hingeless rotor hub is used.
Figure 27 illustrates effects of the aerodynamic In hover and low speed flight, standard tandem interference effect on the quad tiltrotor cruise helicopter controls, plus aircraft pitch and roll attitude, performance. The power required changes due to could b e used to trim this aircraft. In cruise the aircraft interference are shown. The interference effects was trimmed using ailerons, elevator, and differential between the front rotors and the front wing and between propeller thrust; plus propeller thrust, and aircraft pitch the rear rot ors and the rear wing reduce required power.
and roll angles. Front and rear rotor collective pitch The front wing has a beneficial influence on the rear angles were set to values optimiz ed for cruise rotor power. The rear wing also has a beneficial performance (optimized rotor thrust). In addition, rotor influence on the front wing (positive interference flapping was trimmed to zero (for load control) using velocity reduces total induced velocity, and thus reduces rotor longitudinal and lateral cyclic.
wing induced power). The front rotors increase both rear rotor power and rear wing power, although the The blade twist and taper were varied to optimize the effect is not significant. A large effect is from the front rotor for hover and cruise performance . The hover o wing to the rear wing. Much of this wing - wing condition was 5k ISA+20 C, 650 ft/sec tip speed, C / σ T interference effect is, however, simply the be havior of a = 0.1491. The cruise condition was 350 knots, 30k ISA, tandem - wing configuration, distributing the induced 205 ft/sec tip speed, 138764 lb gross weight. The twist losses between the two wings. It is the substantial distribution had two linear segments, inboard (0.0R to deviation from elliptical loading for the combined wing 0.5R) and outboard (0.5R to 1.0R). The linear taper system (see Fig. 22) that produces a non - ideal induced ratio was varied while maintaining c onstant thrust - power loss.
weighted solidity (constant 75%R chord). Figure 28 presents the results for twist optimization, showing the Thus aerodynamic interference has a substantial hover - cruise compromise. For each value of outboard influence on tiltrotor cruise performance, and should be twist, the inboard twist values are 3, 0, – 3, and – 6 deg.
considered in the design optimization. Interference Figure 29 shows the hov er and cruise performance for effects improve the aircraft lift - to - drag ratio of the blades with linear twist, varying from – 15 to 0 deg. A baseline conventional tiltrotor. The interference large negative twist improves hover performance, but velocities reduce total induced velocity along the wing zero twist gives the best cruise performance. The design span, and thus reduce wing induced power. The twist of 0 deg inboard and – 12 deg outboard was interference effect has very small influence on wing selected base d on the hover - cruise compromise.
profile power and rotor propulsive efficiency. The Compared to – 9 deg linear twist, this choice gives 0.3% better hover performance and 1.6% better cruise blade loading and the figure of merit value decreases as performance. the blade loading increases. A parametric study was conducted to examine the di fferences in the figure of The results for blade taper variation are shown in Figure merit trend. The parameters investigated in this study 30. Although the taper of 1.0 produced the best aircraft are twist, taper, tip speed, and airfoils. The effects of L/D, the taper of 0.8 (tip/root chord) was selected to those parameters on the prediction of hover figure of reduce blade weight.
merit were examined by replacing the compound The rotor advancing tip Mach number was varied from tandem rotor qua ntities with the conventional rotor 0.7 to 0.9 to find the optimum rotor rotational speed for quantities. Figure 34 shows the parametric study results.
high speed cruise flight, as shown in figure 31. To The twist, taper, and tip speed increases the figure of maintain low rotor drag at high speed, it is necessary to merit at low blade loading, but decreases it at high blade slow the rotor. The optimum cruise performance is loading. The biggest influence comes from the a irfoil found at M = 0.80 (for the airfoils used). Further change. A significant reduction of the hover figure of at reductions in rotor rotational speed do not improve the merit is observed at high blade loading and the trend aircraft L/D.
becomes similar to the conventional rotor. It appears that the state - of - the - art airfoils used for the compound Performance results from the comp rehensive analysis tandem rotor design have a st rong influence on the are shown in Figures 32 and 33. These results are for figure of merit trend.
state - of - the - art rotor airfoils. The optimized design has a twist of 0 deg inboard (0.0R to 0.5R) and – 12 deg COMPOUND HELICOPTER outboard (0.5R to 1.0R), collective angle of - 2 deg, and To explore further the aerodynamics of compound a taper of 0.8 ( tip/root chord). The hover figure of merit o helicopters, a design and aeromechanics investigation of an isolated rotor is calculated for the 5k/ISA+20 C was conducted for a 100,000 lb compound helicopter condition, with 650 ft/sec tip speed. The results are with a single main rotor, which is to cruise at 250 knots shown in Figure 32. The calculation was conducted o at 4k/95 F (Ref. 5). This aircraft was sized based on the using nonuniform inflow with a free wake geometry.
basic design parameters (not using the RC code). Figure The figure of merit increases as the thrust increases up 35 shows the configuration, and Table 4 gives the basic to around C / σ = 0.18, and then decreases. The figure T aircraft parameters. The aircraft has a six - bladed rotor, a of merit is around 0.73 at the design thrust ( C / σ = T high w ing, a horizontal tail, and two auxiliary propellers 0.149).
located on the wing for cruise propulsion and anti - Figure 33 shows the aircraft lift - to - drag ratio at 30,000 torque in hover. State - of - the - art rotor airfoils are used ft. The calculation was conducted using nonuniform for the main rotor blades. A hingeless rotor hub is used.
inflow with a prescribed wake geometry. Rot or/rotor Blade inertial and structural proper ties are scaled from and rotor/wing interference were included in the the LCTC blade.
comprehensive analysis model. The speed varies from The baseline aircraft design parameters are disk loading 250 to 450 knots; with the rotor tip speed decreasing of W/A = 15 lb/ft , blade loading of C / σ = 0.14, and from hover to cruise speed (350 knots) in order to T wing loading of W/S = 100 lb/ft . These values are from maintain M = 0.8 and then 205 ft/sec ti p speed was at the optimum design for the LCTC, and is shown below maintained up to 450 knots. The rotor performance in to give good performance for the present aircraft. The cruise is presented in terms of aircraft L/D = WV/P, C / σ = 0.14 and W/S = 100 are appropriate for an calculated without accessory or other losses, and using a T aircraft that unloads the rotor at a relatively low speed.
propeller efficiency of 0.86 (from the sizing code). The The aircraft parasite drag is D/q = 40.5 ft . This drag aircraft lift - to - drag r atio decreases as speed goes up. At value, which was obtained from hist oric trends (Fig. 4), the design cruise speed, the aircraft lift - to - drag ratio is is higher than current turboprop aircraft, but lower than L/D = WV/P = 10.1 (9.3 including losses, see Table 2).
is customary in the helicopter industry. The baseline The maximum hover figure of merit of the compound design has a wing span equal to the rotor diameter (Fig.
tandem rotor occurs at around C / σ = 0.17 (Fig. 32), T 35). The hover tip speed is 750 ft/sec, and the cruise tip w hich is high compared with many conventional speed of 502 ft/sec which gives M = 0.8 at 250 knots.
at helicopter rotors. The figure of merit for a conventional The advance ratio is then V/V = 0.84 at 250 knots.
tip articulated rotor was calculated and compared with that A parametric study of key rotor design parameters was of the compound tandem rotor, as shown in Figure 34.
conducted with the comprehensive analysis. The The conventional rotor has 7 blades, existin g airfoils, baseline design has disk loading of 15, design blad e and typical solidity, twist, tip speed. The maximum loading of 0.14, wing loading of 100, collective angle of figure of merit of the conventional rotor occurs at low 0 deg, and shaft angle of attack of 3 deg. The blade twist was varied to obtain balanced hover and cruise conducted using nonuniform inflow with prescribed performance. The hover condition was 750 ft/sec tip wake geometry. The airspeed varies from 200 to 350 speed, C / σ = 0.1484 (assumed 6% downl oad). The knots, with the r otor tip speed linearly decreased from T cruise condition was 250 knots, 502 ft/sec tip speed. hover. The aircraft lift - to - drag ratio decreases as The twist distribution had two linear segments, inboard airspeed goes up. At the design cruise speed (250 (0.0R to 0.5R) and outboard (0.5R to 1.0R). Figure 36 knots), the aircraft lift - to - drag ratio is L/D = WV/P = presents the results for twist variation. For each value of 7.69 (without losses).
outboard twist, the in board twist values are – 3, 0, 3, and Design variations of wing loading (W /S = 100 vs. 120), 6 deg. A large negative twist improves hover blade loading (C / σ = 0.14 vs. 0.09), and disk loading T performance, but the zero twist gives the best cruise (W/A = 15 vs. 12) were examined. The larger disk area performance. The design twist of 0 deg inboard and – 12 will give lower hover power. The larger blade area or deg outboard was selected based on the hover - cruise smaller wing area correspond to loading the rotor rather compromise. T he aircraft lift - to - drag ratio the hover than the wing. Note that C / σ = 0.09 would be T figure of merit variations in Figure 36 are larger than appropri ate for an advanced technology helicopter, those for the LCTC (Fig. 28). Thus, the blade twist is a hence the rotor could carry the aircraft weight to more important parameter for the current design than for conventional helicopter speeds. Figures 41 through 45 the LCTC. However, the aircraft lift - to - drag is less show the performance results in terms of aircraft lift - to - sensitive to the inboard twist change for fixed outboard drag ratio L/D = WV/P, calculated without accessor y or twist. Thus, the benefit of bi - linear twist diminished for other losses, and using a propeller efficiency of 0.86, all the current design compared with the LCTC.
for the design cruise condition of 250 knots. For each The blade taper ratio was varied as shown in Figure 37. combination of disk loading, design blade loading, and The linear taper ratio was varied while maintaining wing loading, three collective angles ( – 3, 0, and 3 deg) constant thrust - weighted solidity (chord at 75%R). and six values for the d ifference between wing Although the taper of 1.0 produced the best aircraft lift - incidence and shaft tilt angle ( α – α = – 4, – 1, 1, 3, 5, w s to - drag ratio, the taper of 0.8 (tip/root chord) was and 7 deg) are shown. The collective and incidence selected to reduce the blade weight. angle variations change the lift share between the rotor and the wing. The rotor speed was that required for M at The rotor advancing tip Mach number wa s varied from = 0.8. Figur e 41 shows the effect of wing loading (W/S 0.5 to 0.9 to investigate the effects of the rotor rotational = 100 vs. 120) on aircraft lift - to - drag ratio for W/A = 15 speed on the cruise performance, as shown in Figure 38.
and C / σ = 0.14. To obtain higher wing loading, wing T It should be noted that the rotor advancing tip Mach area was reduced by decreasing wing span with fixed number in cruise would be 1.02 with the hover tip chord. The aircraft lift - to - drag ratio increases as α – α w s speed. To maintain lo w rotor drag at high speed, it is increases (wing incidence increases or rotor shaft tilts necessary to slow the rotor. The aircraft lift - to - drag ratio forward) up to 3 deg for the collective ang le of 0 and – 3 increases as the advancing tip Mach number decreases, deg and up to 5 deg for the collective angle of 3 deg, reaching the maximum at M = 0.55, which at and then decreases. The best performance was obtained corresponds to V/V = 1.98. Most of the benefit of tip for the collective angle of 0 or – 3 deg and α – α = 3 w s slowing the rotor occurs at the initial 20 to 30% deg. Lower wing loading (higher wing area) increased reduction of the advancing blade tip Mach number. The the aircraft lift - to - drag ratio. The smaller wing area design point was selected at M = 0.80, which at corresponds to loading the rotor rather than the wing. A corresponds to V/V = 0.84. This values corresponds tip wing is a more efficient lifting device than a rotor for to about 20% reduction of the advancing blade tip Mach the current 250 knot compound helicopter, thus the number and 33% reduction of the rotor tip speed from larger wing area improves the aircraft performance.
hover condition.
Figure 42 shows the effect of blade loading (C / σ = T Performance results for the baseline aircraft are shown 0.14 vs. 0.09) on aircraft lift - to - drag ratio for W/A = 15 in Figures 39 and 40. The hover figure of merit of an and W/S = 100. To obtain lower blade loading, blade isolated rotor is calculated with 750 ft/sec tip speed and area was increased by increasing blade chord for a given the result is shown in F igure 39. The calculation was blade radius. Thus, solidity was increased bu t aspect conducted using nonuniform inflow with free wake ratio was decreased. The larger blade area corresponds geometry. The figure of merit decreases as the thrust to loading the rotor rather than the wing. Higher design increases. The figure of merit is around 0.78 at the blade loading (smaller blade chord) increased the design thrust (C / σ = 0.1484 with assumed 6% hover T aircraft lift - to - drag ratio because the smaller blade chord download). Figure 40 shows the aircraft lift - to - drag reduced rotor profile pow er.
ratio with different airspeeds. The calculation was Figure 43 shows the effect of disk loading (W/A = 15 both hover and 250 knots. The choice of these vs. 12) on aircraft lift - to - drag ratio for W/S = 100 and parameters was based on the LABC optimization (Ref.
C / σ = 0.14. To obtain lower disk loading, rotor 1), with the lower design altitude appropri ate for a lift - T diameter was increased, but to maintain the same blade offset rotor. The design lift offset is M/LR = 0.25 (rotor loading for the increased rotor diameter, blade chord hub roll moment M, divided by rotor lift times rotor was decreased. Thus, the blade areas are identical for radius LR). The rotor vertical separation is 6% of the the two cases. Disk loading has a small influe nce on the diameter. Advanced technology rotor airfoils are aircraft performance, although it will have an impact on assumed for the main roto r blades, permitting M = 0.9 at hover performance and the rotor weight. as well as achieving a mean blade drag coefficient of c = 0.0090 in both hover and cruise. The hover tip do The optimum required rotor shaft power and optimum speed is 700 ft/sec, and the cruise tip speed is 600 ft/sec, lift sharing between the rotor and wing are shown in which gives M = 0.9 and an advance ratio of V/V = at tip Figures 44 and 45 for the baseline aircraft (W/A = 15, 0. 70 at 250 knots.
C / σ = 0.14, W/S = 100) at cruise speed of 250 knots.
T Figure 44 shows the rotor shaft power for the baseline The wing is sized to carry 20% of the weight at the aircraft. The rotor power increases as α – α increases. cruise condition, with a wing loading of W/S = 120 w s At the optimum aircraft lift - to - drag ratio, the rotor shaft lb/ft . The resulting wing lift coefficient is 0.7. The power is a small positive value: b etween 500 and 1000 aircraft parasite drag is D/q = 50.0 ft . This drag value is hp. With the rotor in autorotation (zero rotor shaft aggressive for a ro torcraft, and would probably require power), the rotor thrust is large, hence the total rotor flow control for hub drag reduction. The power was drag larger and the aircraft L/D somewhat smaller. calculated using a propeller efficiency of 0.90 for the Figure 45 shows the rotor and wing lift for the baseline auxiliary propulsion. These values for the airfoil drag, aircraft. As α – α increases, the rotor lift decreases and airframe drag, and propulsive efficiency are aggressive , w s wing lift increases. A higher collective angle increases hence the calculated aircraft cruise performance is rotor lift and decreases wing lift. At the optimum lift somewhat optimistic.
sharing between the rotor and wing, the rotor carries 8 - The blade twist was varied to obtain balanced hover and 9% of the aircraft gross weight. The optimum l ift cruise performance. The hover condition was 700 ft/sec sharing between the rotor and wing varies with disk tip speed, C / σ = 0.0919. The cruise condition was 250 T loading, design blade loading, and wing loading.
knots, 600 ft/s ec tip speed. The twist distribution has four linear segments, with breaks at 0.25R, 0.50R, and LIFT - OFFSET COAXIAL HELICOPTER 0.75R. Figure 47 presents the results for twist variation.
The lift - offset or advancing blade concept helicopter has The optimum twist is – 3/ – 6/ – 15/ – 18 deg (inboard to the potential to achieve high - speed cruise with the rotor outboard, equivalent root - to - tip linear rate). The carr ying the aircraft lift. The lift - offset rotor uses stiff performance is much more sensitive to the outboard blades capable of carrying significant hub roll moments, twist (varied in Fig. 47) than to the inboard twist. In hence generating lift on the rotor advancing side in Figure 47, the twist of the last segment (0.75R to 1.00R) forward flight. This configuration still requires auxiliary has values from – 12 to – 24 deg; the lines are for the propulsion, but does not need to unload the aircraft twist of the third segment (0.50R to 0.75R) being equal weight onto a wing in cruise. A 150,000 lb helicopter to that of the last segment, – 3 deg more, or – 6 deg more.
utilizing coaxial lift - offset rotors was investigated, with The optimum is a compromise between cruise and o a design cruise speed of 250 knots at 5k/ISA+20 C. This hover.
aircraft was sized based on the basic design parameters The blade taper ratio was varied as shown in Figure 48.
(not u sing the RC code). Figure 46 shows the As for twist, four segments are used with linear taper in configuration, and Table 4 gives the aircraft parameters.
each. The optimum taper was 1.333/1.333/1.333/ 0.333 The aircraft has coaxial four - bladed rotors. A small (effective tip/root chord ratio). The performance is wing is required to support the auxiliary propulsion much more sensitive to the outboard taper than to the (likely propellers), and hence can carry some lift in inboard taper. In Figure 48, the taper of the last segment cruise. Lift - offset rotors require a hingeless rotor hub, in has values of 0.667, 0.5, and 0.333; for sever al values of order to carry the hub moment. Blade inertial and the taper in the third segment.
structural properties are scaled from the LABC blade (Ref. 1).
Performance results for the lift - offset aircraft are shown in Figures 49 and 50. Figure 49 shows the rotor The baseline aircraft design parameters are disk loading 2 effective lift - to - drag ratio ( L/D = TV/(P +P ), from e i o of W/A = 15 lb/ft , blad e loading of C / σ = 0.10, and T rotor induced and profile power) as a function of maximum advancing tip Mach number of M = 0.9.
at o airspeed, for several values of the rotor lift offset.
The design operating condition was 5k/ISA+20 C, for Figure 50 shows the corresponding aircraft L/D = [4] Yeo, H., and Johnson, W. “Aeromechanics Analysis WV/P, calculated without accessory or other losses, and of a Heavy Lift Slowed - Rotor Compound Helicopter.” using a propeller efficiency of 0.90. The calculation was Journal of Aircraft , Vol. 44 , No. 2, pp. 501 - 508, March - conducted using nonuniform inflow with prescribed April 2007.
wake geometry. The airspeed varies from 100 to 280 [5] Yeo, H., and Johnson, W. “Optimum Design of a knots, with the rotor tip speed varying to keep the Compound Helicopter.” HeliJapan 2006, AHS advancing tip Mach number below or at M = 0.90, and at International Meeting on Advanced Rotorcraft the wing lift coefficient fixed. At the design cruise Technology and Life Saving Activities, Nagoya, Japan, speed (250 knots), the rotor effe ctive L/D is 10.4 and e November 2006.
the aircraft lift - to - drag ratio is L/D = WV/P = 6.2. Lift - [6] Yeo, H., and Johnson, W. “Performance and Design offset is effective above 200 knots.
Investigation of Heavy Lift Tiltrotor with Aerodynamic Interference Effects.” American Helicopter Society 63rd CONCLUSION Annual Forum, Virginia Beach, VA, May 2007.
NASA and the U.S. Army at Ames Research Center are conducting investigations of the aerodynamic [7] Johnson, J., Stouffer, V., Long, D., and Gribko, J., performance capability of rotor craft designed for heavy - “Evaluation of the National Throughput Benefits of the lift and high - speed cruise. The NASA Heavy Lift Civil Tiltrotor,” NASA CR 2001 - 211055, September Rotorcraft Systems Investigation examined in depth 2001.
several rotorcraft configurations for large civil transport, [8] Stouffer, V., Johnson, J., and Gribko, J., “Civil designed to meet the technology goals of the NASA Tiltrotor Feasibility Study for the New York and aeronautics program. Further explorations have been Washington Terminal Areas,” NASA CR 2001 - 21 0659, conducted of the performance of tiltrotor, compound January 2001.
helicopter, and lift - offset rotor configurations, for both military and civil missions.
[9] Smith, D. E., Wilkerson, J., Montoro, G. J., Coy, J., and Zuk, J., “Technology Development for Runway These investigations have shown that heavy - lift, high - Independent Aircraft,” American Helicopter Society speed, long - range rotorcraft ca n be designed that are 59th Annual Forum, Phoenix, AZ, May 2003.
economically competitive and operationally effective, with the potential for substantial impact on both civil [10] Preston, J., and Peyran, R., “Li nking a Solid - and military air transportation systems. Table 5 Modeling Capability with a Conceptual Rotorcraft summarizes the calculated cruise performance of the Sizing Code,” American Helicopter Society Vertical configurations examine d.
Lift Aircraft Design Conference, San Francisco, CA, January 2000.
This work continues, including exploration of additional design conditions and multi - mission requirements, [11] Johnson, W., “Rotorcraft Aeromechanics further refinement and optimization of the designs, Applications of a Comprehensive Analysis,” HeliJapan application of higher - fidelity aeromechanics analyses to 98: AHS International Meeting on Advanced Rotorcraft the configurations, and closer coupli ng of the sizing Technology and Disaster Relief, Gifu, Japan, April code to the resulting performance calculations.
1998.
[12] Johnson, W. “Calculation of Tilt Rotor REFERENCES Aeroacoustic Model (TRAM DNW) Performance, [1] Johnson, W.; Yamauchi, G.K.; and Watts, M.E.
Airloads, and Structural Loads.” Ameri can Helicopter “NASA Heavy Lift Rotorcraft System Investigation.” Society Aeromechanics Specialists’s Meeting, Atlanta, NASA TP 2005 - 213467, December 2005.
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[2] Johnson, W.; Yamauchi, G.K.; and Watts, M.E.
[13] Floros, M.W., and Johnson, W. “Performance “Design and Technology Requirements for Civil Heavy Analysis of the Slowed - Rotor Compound Helicopter Lift Rotorcraft.” American Helicopter Society Vertical Configuration.” American Helicopter Society 4th Lift Aircraft Design Conference, San Francisco, CA, Decennial Specialists’ Conference on Aeromechanics, January 2006.
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[3] Acree, C.W., Jr., and Johnson, W. “Performance, [14] Jenkins, J.L., Jr. “Wind - Tunnel Investigation of a Loads and Stability of Hea vy Lift Tiltrotors.” American Lifting Rotor Operating at Tip - Speed Ratios from 0.65 Helicopter Society Vertical Lift Aircraft Design to 1.45.” NASA TN D - 2628, February 1965.
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[15] Hohenemser, K. H., “Full Scale Rotor T ests of the Air Force Convertiplane Model XV - 1 in the NACA 40 x 80 Foot Wind Tunnel at Moffett Field, California.” FM rotor hover figure of merit, ( T √ (T/2 ρ A) ) /P McDonnell Aircraft Report 3379, McDonnell Aircraft L/D aircraft effective lift - to - drag ratio, WV/P Corporation, February 1954.
(based on cruise power) [16] McCloud, J.L.; Biggers, J.C.; and Stroub, R.H. “An L/D rotor effective lift - to - drag ratio, T V/(P +P ) e i o Invest igation of Full - Scale Helicopter Rotors at High (based on rotor induced and profile power) Advance Ratios and Advancing Tip Mach Numbers.” M advancing tip Mach number NASA TN D - 4632, July 1968.
at P aircraft power [17] Charles, B.D., and Tanner, W.H., “Wind Tunnel Investigation of Semirigid Full - Scale Rotors Operating R rotor radius at High Advance Ratios,” US AAVLABS TR 69 - 2, T rotor thrust January 1969.
V flight speed [18] Lim, J.W.; McAlister, K.W.; and Johnson, W.
“Hover Performance Correlation for Full - Scale and V rotor tip speed tip Model - Scale Coaxial Rotors.” American Helicopter W gross weight Society 63rd Annual Forum, Virginia Beach, VA, May 2007.
W/A disk loading [19] Harrington, R.D ., “Full - Scale - Tunnel Investigation W/S wing loading of the Static - Thrust Performance of a Coaxial α rotor shaft angle (positive aft) s Helicopter Rotor,” NACA TN 2318, 1951.
α wi ng incidence angle w [20] McAlister, K.W.; Tung, C.; Rand, O.; Khromov, V.; and Wilson, J. S., “Experimental and Numerical η propeller efficiency, TV/P Study of a Model Coaxial Rotor.” American Helicopter ρ air density Society 62nd Annual Forum, Phoenix, AZ, May 2006.
σ rotor solidity (ratio blade area to disk area) [21] Arents, D.N., “An Assessment of the Hover Performance of the XH - 59A Advancing Blade Concept Demonstration Helicopter,” Report USAAMRDL - TN - ASM available seat miles 25, May 1977.
CTOL conventional takeoff and landing [22] Ruddell, A.J., et al. “A dvancing Blade Concept (ABC) Technology Demonstrator.” USAAVRADCIM DOC+I direct operating cost plus interest TR 81 - D - 5, April 1981.
ISA international standard atmosphere [23] Harris, F.D., and Scully, M.P., “Rotorcraft Cost LABC Large Advancing Blade Concept Too Muc h.” Journal of the American Helicopter LCTC Large Civ il Tandem Compound Society , Vol. 43, No. 1, January 1998. (Additionally, Harris, F.D. “An Economic Model of U.S. Airline LCTR Large Civil Tilt Rotor Operating Expenses,” unpublished.)
MCP maximum continuous power NOMENCLATURE MRP maximum rated power A rotor disk area OEI one - engine inoperative c mean drag coefficient for profile power do OGE out of ground effect C rotor thrust coefficient, T/( ρ AV ) T tip RC AFDD design code C rotor weight coefficient, W/( ρ AV ) W tip RIA runway independent aircraft D/q airframe drag divided by dynamic pressure VTOL vertical takeoff and landing Table 1. Civi l design mission.
1200 nm range, 120 passengers (26400 lb payload) 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 - best - range 30k ISA Descend at V - best - range (no range credit) [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% (Hover OGE P required) at 22k ISA, (OEI MCP) greater than (P required at V - best - range ) 4 engines Table 5. Summary of calculated aircraft cruise performance.
2/3 prop η Aircraft cruise condition (D/q)/(W/1000) L/D=WV/P Large Civil Tiltrotor (LCTR) 350 kts, 30k 1.50 12.45 o Tiltrotor 300 kts, 4k 95 F 1.98 7.4 o Tiltrotor (reduced rotor speed) 300 kts, 4k 95 F 1.98 7.75 Large Civil Tandem Compound (LCTC) 350 kts, 30k 1.88 0.86 10.1 o Compound helicopter 250 kts, 4k 95 F 1.88 0.86 7.7 o Lift - offset coaxial helicopter 250 kts, 5k ISA+20 C 1.77 0.90 6.2 Table 2. 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 V/V 1.69 2.88 2.32 tip 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 rati o 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) 1.50 1.88 1.29 body - interference - tail D/q 13.2 14.4 15.1 pylon D/q 10.0 9.4 9.5 wing D/q 14.1 15.8 – hub D/q – 10.7 13.5 2/3 hub (D/q)/( W/1000) – 0.40 0.45 Wing loading (lb/ft ) 80 80 – area (ft ) 1545 1735 – span (ft) 105 144 – Aspect ratio 7.1 12.0 – Weight empty fraction 65.3% 65.6% 64.7% Lock number 12.1 13.0 19.1 Total blade weight (lb), all rotors 5960 5168 1080 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 Cruise prop η / rotor D/q / rotor L/D 0.81 16.7 7.7 † e o Hover figure of merit (5k ISA+20 C) 0.78 0.73 0.68 † † sizing code used hover figure of merit = 0.79, cruise L/D = 12.0 e Table 3. Characteristics of baseline tiltrotor and quad tiltrotor designs.
Tiltrotor Quad Tiltrotor RC designed developed from tiltrotor Mission gross weight (lb) 146600 146600 Engines (hp) 2x18710 Rotor diameter (ft) 78.9 55.8 Disk loading W/A (lb/ft ) 15 15 o C / σ (geom, 4k/95 F) 0.140 0.140 W o C / σ (T - wt, 4k/95 F) 0.154 0.154 W Hover tip speed (ft/sec) 750 750 Cruise tip speed (ft/sec) 626 626 Solidity 0.0989 0.0989 Number blades per rotor 4 4 chord (75%R, ft) 2.79 1.97 taper ratio 0.7 0.7 Drag D/q (ft ) 55.0 60.3 2/3 (D/q)/(W/1000) 1.98 2.17 Wing loading (lb/ft ) 100 67.2 area (ft ) 1466 848 (front) & 1335 (rear) = 2183 span (ft) 96.4 73.3 (front) & 102.6 (rear) Mission, payload 20 ton range (nm) 750 o cruise altitude (ft) 400 0, 95 F cruise speed (kt) 300 Table 4. Characteristics of compound helicopter and lift - offset coaxial helicopter designs.
Compound Helicopter Lift - Offset Coaxial Helicopter developed from parameters developed from parameters Mission gross we ight (lb) 100000 150000 Rotor diameter (ft) 92.1 112.8 Disk loading W/A (lb/ft ) 15 15 o C / σ (geom, 4k/95 F) 0.140 W o C / σ (T - wt, 4k/95 F) 0.148 0.100 W Hover tip speed (ft/sec) 750 700 Cruise tip speed (ft/sec) 502 600 Solidity 0.0992 0.0871 (per rotor) Number blades per rotor 6 4 chord (75%R, ft) 2.39 3.86 taper ratio 0.8 1.333/0.333 Drag D/q (ft ) 40.5 50.0 2/3 (D/q)/(W/1000) 1.88 1.77 Wing loading (lb/ft ) 100 120 area (ft ) 1000 250 span (ft) 92.1 38.7 Mission, payload range (nm) o o cruise altitude (ft) 4000, 95 F 5k ISA+20 C cruise speed (kt) 250 250 Figure 3. Simplified conceptual design process.
Figure 1. Runway Independent Aircraft (RIA) industry concepts from 2002: Sikorsky Reverse Velocity Rotor (top), Boeing Tiltrotor (center), Bell Qua d Tiltrotor (bottom).
Figure 4. Aircraft drag trends (courtesy F.D. Harris).
Figure 2. Outline of iterative design process.
Figure 5. Helicopter hub drag trends (courtesy 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 11. Effect of cost technology factors on flyaway Figure 9 . Flyaway price (2005 USD) and DOC+I (2005 price (2005 USD) and DOC+I (2005 cents/ASM) for cents/ASM) comparisons for baseline designs.
LCTR.
Figure 12. Cost elements compared for LCTR with and Figure 10. Cost elements compared for heavy - lift without cost technology factors lift rotorcraft (1200 nm, rotorcraft and B737 (1,200 nm, 120 passengers, including 120 passengers); in legend [x,y], x is mai ntenance factor technology factors for rotorcraft costs).
and y is price factor.
LCTR Twist Optimization Maxima, XN Airfoils 0.82 Optimum: -32/-30 -30/-28 -34/-30 0.81 -34/-24 0.80 -30/-24 -38/-32 Twist rate, deg/R: inboard/outboard 0.79 Propulsive efficiency -40/-34 0.78 0.77 0.76 0.77 0.78 0.79 0.8 Figure of merit Figure 13. LCTR twist optimization.
Figure 16. LCTR rotor hover performance.
a) Root section: AFDD CTR1544, c = – 0.160, t/c = m 15.3% b) Mid - span section: AFDD CTR4475, c = 0.027, t/c = m 11.3% c) Tip section: AFDD CTR7500, c = 0.014, t/c = 9.0% m Fi gure 14. LCTR airfoil sections.
Figure 17. LCTR rotor and aircraft cruise performance.
Figure 15. Influence of airfoils on LCTR hover and cruise performance.
1 speed Figure 18. Influence of cr uise 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 and for 2 - speed transmission.
Figure 19. Baseline tiltrotor configuration (courtesy Figure 20. Baseline quad tiltrotor configuration (courtesy Gerardo Nunez of U.S. Army, AFDD ) Gerardo Nunez of U.S. Army, AFDD).
Figure 21. Wake geometry of conventional tiltrotor. Figure 22. Wake geometry of quad tiltrotor.
Figure 23. Aircraft lift - to - drag ratio of conventi onal tiltrotor (with flapping trim).
Figure 26. Influence of rotor cruise tip speed on aircraft performance.
Figure 24. Rotor propulsive efficiency of conventional tiltrotor (with flapping trim).
Figure 27. Interference effects on required power of quad tiltrotor (arrowhead indicates component receiving interference; numbers positive for power increase, unfavorable interference; units are HP).
Figure 25. Wing drag of conventional tiltrotor (with flapping trim).
Figure 28. Effect of blade twist on LCTC performance Figure 31. Effect of tip speed on LCTC performance.
(inboard twist = – 3, 0, 3, 6 deg).
Figure 32. LCTC hover figure of merit.
Figure 29. Performance of LCTC with linear twist.
Figure 33. LCTC aircraft cruise performance Figure 30. Effect of blade taper on LCTC performance.
Figure 34. LCTC hover performance buildup. Baseline is compound tandem rotor (using state - of - the - art airfoils).
Dashed line uses taper, twist, and tip speed of a conventional articulated rotor. Dotted line also uses airfoils of a conventional articulated rotor.
Figures 35. Compound helicopter (courtesy Gerardo Nunez, U.S. Army, AFDD).
Figure 36. Effect of blade twist on compound helicopter performance (inboard twist = – 3, 0, 3, 6 deg).
Figure 39. Compound helicopter hover figure of merit.
Figure 37. Effect of blade taper on compound helicopter performance.
Figure 40. Compound helicopter cruise lift - to - drag ratio.
Figure 38. Effect of cruise tip speed on compound helicopter performance.
(a) Wing loading W/S = 100 (b) Wing loading W/S = 120 Figure 41. Effect of wing loading on aircraft lift - to - drag ratio (W/A = 15, C / σ = 0.14).
W (a) Blade loading C / σ = 0.14 (b) Blade loading C / σ = 0.0 9 W W Figure 42. Effect of blade loading on aircraft lift - to - drag ratio (W/A = 15, W/S = 100).
(a) Disk loading W/A = 15 (b) Disk loading W/A = 12 Figure 43. Effect of disk loading on aircraft lift - to - drag ratio (C / σ = 0.14, W/S = 100).
W (a) Aircraft lift - to - drag ratio (b) Rotor shaft power Figure 44. Aircraft lift - to - drag ratio and rotor shaft power (W/A = 15, C / σ = 0.14, W/S = 100).
W (a) Rotor lift (b) Wing lift Figure 45. Optimum lift sharing (W/A = 15, C / σ = 0.14 , W/S = 100).
W Figure 46. Lift - offset coaxial helicopter (courtesy Gerardo Nunez, U.S. Army, AFDD) .
Figure 49. Lift - offset coaxial helicopter rotor effective Figure 47. Effect of twist on lift - offset coaxial lift - to - drag ratio, L/D = TV/(P +P ) (offset = 0.15, e i o helicopter performance (twist = linear rate, root to tip).
0.20, 0.25 [design], 0.30).
Inboard twist fixed at – 3/ – 6 deg; tw3 e xtends from 0.5R to 0.75R, tw4 from 0.75R to 1.0R. For each curve, tw4 varies from – 12 to – 24 deg.
Figure 50. Lift - offset coaxial helicopter aircraft lift - to - Figure 48. Effect of taper on lift - offset coaxial drag ratio, L/D = WV/P (offset = 0.15, 0.20, 0.25 helicopter performance (taper = tip/root chord ratio).
[design], 0.30).
Inboard taper fixed at 1.333; taper3 extends fro m 0.5R to 0.75R, taper4 from 0.75R to 1.0R. For each taper4 value, taper3 varies from 0.667 to 1.5.