Document
4 0 T V " CDC- T •; • '-; '•,. G K .v;, "..' .-. v . C H - > R ~ ,. % - Z C F O P »i' NASACR-145117 By S. Jon Davis Prepared under Contract No. NAS1-13624 By Boeing Vertol Company Philadelphia, Pennsylvania for .JUL1977
NASA
National Aeronautics and i RECEIVED Space Administration December 1976 NASACR-145117 RESEARCH REQUIREMENTS FOR DEVELOPMENT OF - -IMPROV-ED-HEL-ieOPTER ROTOR EFFICIENCY By S. Jon Davis Distribution of this report is provided in the interest of jnfo£^tion exchange. Responsibility for the contents resides in the author or organization that prepared it.
Prepared under Contract No. NAS1-13624 by " Boeing Vertol Company — —Philadelphia, Pennsylvania for NATIONAL AERONAUTICS AND SPACE ADMINISTRATION ABSTRACT This report documents-the research requirements for developing an improved-efficiency rotor for a! civil helicopter; The various design parameters affecting the hover and cruise effi- ciency of a rotor are :surveyed and the parameters capable of producing the greatest potential ! improveme7ftlM^idelitified7~Re~s75,^ programs to achieve these improvements : :! are definediahd estimated ;costs aiid schedules are presented. Interaction of the improved- :i efficiency rjQtorJvithlplher;tej;h^ helicopter is noted, including '•; its impact pri engihe noise;, hover and cruise performance, one-engine-inoperative hover capability, -" and maintenance and reliability.
This report was-prepared by-the Boeing Vertol Company for the National Aeronautics and Space Administration, Langley Research Center, under NASA Contract NAS1-13624.
William Snyder was technical monitor for this work. The Boeing Vertol Project Manager was 'Wayne Wiesher. "• ~" ~ " IV In order to improve the performance capability of future civil helicopters, the goals of a 9.3-percent improvement in h.over.efficiency and a 20-percent improvement in cruise efficiency by 1985 have been established.'-!;-" . " : . - ' . -" • - ' ' . " " ' ' ••': ThelmprpWmentin'"hover? efficiency is 'to be obtained by concentrating on a reduction 5 !
' ;in the rndubed--power-^&-uise e^ the cruise profile-power : Component.
TecHnological'gaps'have Be¥nTdentifiedln the areas of presently available rotor test data which reflects the effect of variations in rotor design parameters on rotor efficiency and analyti- cal performancerprediction..capability,in both flight regimes.
Therefore, research and development programs involving considerable model-rotor testing in both hover and cruise~flighrare recommended, leading to the eventual design and test of a ,. full-scale, improved-efficiency rotor.
The improved-efficiency rotor: '• 1 • Resulte~rn-a~2j9-pereent increase in cruising speed, a 15-percent increase in specific range, and a; 12.9-percent decrease in fuel consumption.
• Produces a reduction in empty weight of 3.83 percent.
• Results in smaller-engines-being sized ,by_the _HOEI requirement.
• Produces a 4.42-percent reduction in flyaway cost and a 6.75-percent reduction in direct operating"cost; v V
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Page ABSTRACT iii FOREWORD . . . - . . ' i v SUMMARY . . , ...._.._. .. .. . . . _ . . . y .
LIST OF ILLUSTRATIONS . . . ix LIST OF TABLES xi LIST OF SYMBOLS xii 1.0 INTRODUCTION 1 2 0 FACTORS AFFECTING ROTOR EFFICIENCY 4 ; 2.1 Improvement in Figure of Merit 4 2.2 Improvement in L/Dg 8 j , 3.0 DISCUSSION OF THE EFFECTS OF DESIGN PARAMETERS ON ROTOR EFFICACY; ... .. . 15 3.1 Factors That influence: Hover Efficiency 15 3.2 Factors That Influence Cruise Efficiency . .- . . . . . . . . . 20 v .;'•:'>) t-;"^ • '•*: ' ; ' • . • • ' 4.0 CURRENT TECHNOLOGY ^ ; ' . - . r O ' . , . ,.. . . 29 4.1 Previous Research and Development ;.' 29 4.2 Technological Gaps and Problem Areas 29 5.0 RESEARCH AND DEVELOPMENT REQUIREMENTS 31 5.1 Improvement in Hover Efficiency (Figure of Merit) 31 5.2 Improvement in Cruise Efficiency (L/Dg) 31 5.3 Integration of Rotor Design Parameters 32 5.4 Schedules and Estimated R&D Costs for Improvement of Rotor Efficiency in Hover and Cruise 32 6.0 TECHNOLOGY AND DESIGN INTERACTIONS 37 6.1 Helicopter Noise Generation 37 6.2 Performance . ..... .......... 39..
6.3 Empty Weight 39 6.4 Drive-System Efficiency and Weight 39 6.5 Hover With One Engine Inoperative (HOEI) 39 6.6 Reliability and Maintainability 39 6.7 Production 40 Vll
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— ———— Page 6.8 Rotor-Blade and Control-System Loads 40 6.9 Vibration . 40 7.0 IMPACT ON INITIAL COSTS AND OPERATING COSTS. . ... 4L 8.0 CONCLUSIONS 43 8.1 Improvement in Hover Efficiency . _....-—_-_..-.- ~ _.____._.... _ - ^ . ^ 8.2 Improvement in Cruise Efficiency -. 43 9.0 REFERENCES 45 10.0 BIBLIOGRAPHY . 46 vm [LIST OF fLLUJSTR^tlONSF.
fcim.i'ii ..<-..,. . . ,„ .. . . _ - . . _ , . > ___™-J Figure Page 1 Comparison of reduction in energy intensity for the compromise design mission 2 2 Comparison of technological development in terms of cost/unit El saved for the compromise design mission - - . - . - - - . . . . . . . . . . 3 3 /2 3 i Required values of Cg /C^ and KJJ^D for given values of figure of merit . 6 4 Minimum obtainable induced-power factor, Kjjjj} 7 5 Maximum obtainable hover efficiency 9 6 Typical improvement vectors for figure of merit 10 7 Power-required split in cruising flight 11 8 Profile-power-required split in cruising flight . . . . . . . . . . . . 13 9 Rotor lift/effective drag 14 10 Improvement vectors for figure of merit with elliptical-planform blade tip . . . . . ... , . . 16 11 Effect of tip shape on rotor induced-power factor 17 12 Effect of twist on hover efficiency and on rotor induced-power factor . . . . 18 13 Effect of camber and airfoil section on rotor hovering efficiency . . . . 19 14 Effect of roughness on rotor hovering efficiency 21 15 Effect of Reynolds number on coefficient of drag 22 16 Effect of change in blade twist on aerodynamic efficiency . .- -.—r ; . . 23 17 Estimated aerodynamic efficiency of the full-scale reverse-velocity rotor 26 18 Trend of lift and aerodynamic efficiency with percent lift offset . . . .27 : ix . : '' ' _/ ; Figure —_— • Page 19 Program schedule and estimated research and development costs for increased figure of merit 33 • • • 20 Improvement of figure of merit and expenditure as a function of time 34 21 Program schedule and estimated research and developmemVcosts j for increased rotor aerodynamic efficiency.(L/Dg) 35 22 Improvement of rotor aerodynamic efficiency and expenditure as a function of time 36 Table Page 1 Effect of Change in Blade-Tip Planform on Aerodynamic Efficiency 24 2 Effect of Rotor Parameters on Rotor Noise 38 3 Economics of Existing and Future Helicopter Fleets in the United States and Canada 42 4 Summary of Research Recommended for Improved Rotor Efficiency (Hover Figure of Merit) 44 5 Summary of Research Recommended for Improved Rotor Efficiency (Rotor L/D ) 44 E XI OF; Btu British thermal unit (1 Btu (mean) = 1055.87 J) Cjj -average~blade-profile=drag-coefficient, c^ = 8 P 2 3 j: : • - . ' : ' : : - - . . . ; I : i . O7rR pV T !
i • . - • • : : ! . . • ;- - • - ' jcj {"""' averagfe blade-lift^coefficient,eg = 6 Gp/a : ' ' ' ' :. '.
CL „ ,.blade=element lift coefficient blade aerodynamic pitching-moment coefficient ty -9 rotor power coefficient, Cp = P/pirR V-p rotor thrust coefficient, Cip = T/p7rR V-pjp d ^D/dM rate of increase of airfoil-section profile drag with increasing Mach number DOC direct oper-ating cost, $/seat-km El energy intensity, J/passenger-km EW/GW structural empty-to-gross weight ratio ; : ' 3/2 FM ; helicopter rotor figure.of merit, 0.707 C /Cp T !
• HOEI hover, one :engine inoperative ^IND hovering-rotor induced-power factor rotor lift-to-effective drag ratio Maph -number -- advancing-blade-tip Mach number ^DD drag-divergence Mach number PPRQ profile-power component of rotor cruise power, kw Xll E , kw IND -- R rotor blade radius, m , R&D research and development RDT&E : research, development, test and engineering i Re ~ ; 'Reynolds-number ' t | . 1 ' . ! ' " ' ' : " sfc < ^pecific fuel consumption, kg/hr/kw T thrust, N Vfip ; rotor tip speed, m/s V ~ fbrward~speed, kph a-ppp rotor tip=pathrplane angle, deg 6TWIST rotor-blade twist, deg M advance ratio, V/Vrpjp p atmospheric density, kg/m a -rotor solidity,-bc/7rR i_ •__ :_„,.;.;.., xiii Previous.studies have.shown .that, on the basis of fuel efficiency, helicopters can be competitive with other forms of transportation for some missions. Current energy-consumption levels can be reduced, how'eyer, through the infusion of advanced technology into the design process.
; The stat^^fjrejference I]^aminbd five technological areas that promise to reduce ! helicopter energy consumption. These are sfc reduction, increased rotor figure of merit and cruise L/D^, paurasite-dragireduction, and reduced empty weight through the application of advanced-composite-materials.
Preliminary estimates were made of the development programs required to achieve speci- fied goals, the percentage of energy reduction for each technological area was also estimated and presented as development cost per unit energy intensity saved.
These results, shown in Figures 1 and 2, show that improving rotor efficiency (figure of merit arid cruise L/Dg); offers large payoffs in energy reduction at minimum cost.
This report documents the research requirements for developing such an improved- efficiency rotor fonajcivttihelicopter. The various jdejlgn parameters affecting the hover and cruise efficiency of a rotor are surveyed and the parameters capable of producing the greatest potential improvement are identified. Research and development programs to achieve these "improvements awndieflnedi and estimated costs and schedules are presented. Interaction of the improved efficiency rotor with other technological goals for an advanced civil helicopter is noted, including its impact qn engine noise, hover and cruise performance, one-engihe- inoperative hover capability, and maintenance and reliability.
s to 1- 4U 38.1 LU O cc.
LU o_ 30.35 1 • .
>
h ; LU 1 - — ~ 20 (D 16.6 cc LU 12.5 LU __ 9.2 6.5 O 5 8 w*U •0 1— 3.1 Q • I I LU S- n tt - ° SFC SFC FIGURE- L/D F. STRUCTURAL ALL E REDUCTION REDUCTION OF-MERIT TECHNOLOGICAL IMPROVEMENT "RED*UCT?ON EW/GW RATIO (CONV ENG) (REGEN ENG) IMPROVEMENT REDUCTION IMPROVEMENTS (REGEN ENG) ALL TECHNOLOGICAL IMPROVEMENTS (CONV ENG) Figure 1. Comparison of reduction in energy intensity for the compromise design mission ,3.0" -.2-°r 3.287 Ifl X O (1.874) X 3.2
- i 1.8 -
!> CO CO _ < 1.6 < 2.8- — 0.
^J ^ D _ m 1.4 ^ 1 2.4- - ^ Q I -• LU - Q 1 - 2 - LU < 2.0 CO ^> < 1 0 u LU CO '- - Qj 1.373 17375' tl t 1.6 z (0.784), (0.799) h- Qg (0.783) z > 1.214 co 1 2 ^ —' (0.692) 0 '"* to 0 6 l/^ w>w o o "3» — .
u3 0.8 ** z 0.4 0.407 6.384 LU Q.
O •T6."232) 0.270 (0.219) ^J Q. A . ^ - °- n 0 LU * " O 0.2 •• (0.154): _l LU LU > Q _ LU n
o
Q F STRUCTURAL SFC SFC "FFGU'RE" L/D ALL E e REDUCTION REDUCTION OF-MERIT -, IMPROVEMENT REDUCTION EW/GW RATIO TECHNOLOGICA (CONV ENG) (REGEN ENG) IMPROVEMENT REDUCTION IMPROVEMENTS .. (REGEN ENG) ALL TECHNOLOGICAL IMPROVEMENTS (CONV ENG) 'Figure 2. Comparison of technological development in terms of cost/unit El saved CXI for the compromise design mission
23TFACTORSi AFFECTING ROTOR EFFICIENCY
/'!
The following paragraphs provide some insight into the overall factors governing the improvement of rotor :effitiency in-hover and cruise.
2.1 improvement in Figure of Merit 1!
T^e-two-maJQpeomponents of-figure-of merit which have to be improved are the induced and profile powers._ The induced power is the theoretical power used to generate lift in the absence of j any airfoil profile drag. Momentum theory shows that the induced drag is minimized when a uniform distributibn of perpendicular induced or downwash velocity is achieved through the rotor. :Increasing the number of blades and/or having nonlinear values of twist result in more-uniform induced velocities with the associated increase in figure of merit.
Thfe other major component of actual hover power, the profile power, is dependent on : the best obtairiable"Uft~to dfag~fatio."This~is afunction of local Mach number. For the airfoils , in use today, blade sections would have to operate at C^ = 0.8 •+ 0.9 to achieve the highest lift- to-drag ratio.
In order to visualize the relative importance of the induced- and profile-power con> , ••-•• ponents to~figure -of meritj- a simplified analysis (ref. 2) is presented where: , .. ., v- 2.5984 :: ivTMTV"*~ ' ~' and ; " = induced-power component 2.5984 . = profile-power component.
•j~.r- 3111- -1*-t/-V**j^ ~" ', , The previojus equation is the-result-of rearranging the basic rotor hover-efficiency relationship, Cp ^ „ . . . - _ ™ .
into the form, n • _ FM =, •PRO -where
C
P
PRO
g The profiler-power component can be further rationalized into the format of reference 2 . by combining the basic rotery-wmg^efiHitibfis of : • Cj..___ ?. o-.Cp a : eg = 6(G-j-/a) ; . / . . . - i . . . . . . . . . . ^ 3/2 3/2 - , jnto.the parameter Sp /GJ (which is analogous to C /C in fixed-wing, performannft) so, r L D " . " ™»™™w™j(jw™<.-- — f\jL" —> - -—..»>,*.,».....,.™ i*.,.™, .. ~ ^*_ _i^__'™ ' "• ' ..,- that: ; ' ' ' ' ," •" ~" 3 2 /2 - . ; ' eg ! • _ - . 1.8371C ?
T \ This can be substituted into the profile-power component of the simplified hover-efficiency iequation to obtain,: ; „ : . . . , / : 2.5984., 3/2 3/2 ; 0.707C :Va(cg /c ) .
T d The factor iKjjvjj), or induced-power factor, is the ratio of actual rotor-induced power to -i the ideal rdtor-iniduced power (assuming a uniform downwash distribution). It can be minimized '•> by optirnizing-bladechord, twist distribution, and blade number, cj /^, the profile-power f factor, is influenced by im'proyed airfoil-section characteristics, etc.
f : ' • I : . " ' ' . - : • : • . : • _ . ' : . • • : • ; - .
3 /2 FigurF:Ti^a"pr6t"pfrth"e~refluireirvaue"s of KJJ^ and eg /c for given values of FM, D d 1 : ^ assuming a =;Q. lQ'*as;typical-;of^cphtempprary rotor designs. Superimposed on the plot are typical (cofaj&id&iij levels -for Asymmetrical (NACA 0012) and cambered (V23010-1.58) J r f •' i *—X —" "*~Q*"iHaX • *"—•—*--- .—------ v..- . . --. . . . *. . . . -^ _ / " _" -* - - - - - - - ; 3 /2 1 airfoil sections. Note that.the integrated .or total blade values of (eg /c ) for a rotor employ- d ! ing those sections will be less than the levels illustrated.
Reference 3 defines a maximum or upper limit to the practical amount of improvement, : in mducedipower that can be achieved. Figure 4 illustrates this minimum value of Kjj^p as a / S T A f T - x c .,; G o.
: ; C j v 120 r 110 • 100 • M = 0.5 Re = 8.7 X 10 NACA 0012 l\/[=_0.5 Re = 3.08 X ,i IO 1.2 K IND Figure 3. Required values of Cg and KrMr» for given values of figure of merit 1.24 1.20 1.16 o ? 1.12 1.08 1.04 1.00 0.002 0.008 0.010 0.004 0.006 Figure 4. Minimum obtainable induced-power factor, function of Ctr for a 4-bladed rotor. If we assume for purposes of illustration that an integrated . - . - „ £ , ^ , . . „ „ , _ „ _ , , . _ . . _ . . . * . . . . - W value of (cjj) /c^) = 100 is obtainable and use that value in combination with the values of Defined in Figure 4, we get the maximum FM curve illustrated in Figure 5.
. , -Figure* 6~iliustrates"the different" options available for improving rotor efficiency .in hoyer,.
i Point E is representative of a current, 1975-technology, square-tip rotor blade. For purposes of :: illustration' a desired improved FM level, in this case 0.80, is selected and vectors AE and BE — i. _.—~—.CL— .~ i~ - ~~ - •* r , -•, are drawn. Vector AE represents an improvement in FM obtained purely through a reduction • in inducedSpbwer (Kjjj0),l while vector BE achieves an FM = 0.80 through improved airfoil characteristics-with-no;change in-induced"powerr It should be noted, however, that vector AE assumes noi limit td.the amount of induced-power reduction achievable. If such a limit (as de- fined in Figure 4) is imposed, vector AE is replaced by vector CE.
: ... . _ 2.2 Improvement in LTD]?
Forward-flight power required can be divided into profile-, induced-, and parasite-power components. "" ~~ Pjrofile power isjiefined asi thepower required to overcome the profile or frictional losses incurred by the rotor as it \turns. As in hover, the basic profile drag is dependent on the lift/ i drag characteristics of the'airfoil sections employed, the surface roughness of the blade, and its operating Reynolds number range .--Also included are the profile-drag increments due to com- pressibility effects on the advancing blade and stall effects on the retreating blade.
Induced power, as in the case of hover, is defined as the power required to generate lift in the absence of airfoil profile drag.
• Parasite power is defined as the power required to provide propulsive thrust equivalent to the total parasite drag of the helicopter.
Figure 7 illustrates a typical rotor power-component split with forward speed. Note the ; : reduction m~irnportance_ofinduced .power»to the total power required as forward speed increases.
For purposes of evaluating the relative efficiency of a rotor in producing lift in forward 1 : flight, the profilejand induced^; pdwer can be combined and expressed as an equivalent rotor lift- [., to-drag ratio, w h e r e : . . - - . . - _ P )/V iND and L . = rotor lift force, N ' Ppj^Q = rotor profile-power component, kw = rotor u ?IND y?-d .ced-power component, kw ~V =~ forward speed, kph.
0.94 r 0.92 NOTES: 1. a = 0.100 2. K| = MINIMUM OBTAINABLE ND 3/2 0.90 3. C /C = 100 c d DC! - - LU 0.88!
O UJ §' 0.86 oi 0,84 0,82 0,80.
0.002 0.004 0.008 0.010 0.006 Figure 5. Maximum obtainable hover efficiency U) CO 1.0 1.1 1.2 1.3 1.4 K IND Figure 6. Typical improvement vectors for figure of merit I 1
A
PROFILE POWER o Q.
< O 60 - O HI INDUCED ; POWER ! LLI QC S 20 0.6 0.2 0.4 0.5 0.3 ADVANCE RATIO, Figure 7. Power-required split in cruising flight -• m •I r> y K '.i Q l \ - 'I . As^qjLnJbe seejLfrpm_Fig!Uire_Z, PEQ-Rle^ppwer .comprises a much greater share of the total M power required and therefore presents a greater potential area for L/Dg improvement than in- duced power.
Figure 8 shows a split.of the-profile-power component of Figure 7 into its subcomponents.
Figure 9 illustrateslthe effect on rotor L/Dg of the various individual profile-power com- •'iponents identified inTiigureyS.T I^^correspohds to the rotor L/Dg with the profile-power ;<! : component eonsis^rig pnljf-ofr^e-basie-profUe-drag subcomponent. The region between lines (I)and(2)iUiistrateslthe;redt.ietip;n in L/Dg resulting from the addition of the advancing-rptor-blade ."" compressibility subcomponent of profile power to line® The further addition of the retreating- suits in line(3) ;blade-stall subcomponent to line BASIC PROFILE DRAG ADVANCING-BLADE COMPRESSIBILITY EFFECTS RETREATING-BLADE 0.4 AD VAISICF RATIO,/!
Figure 8. Profile-power-required split in cruising flight CXI , QJ Q ADVANCING-BLADE COMPRESSIBILITY EFFECTS O •z.
UJ O LLJ CJ cc LU RETREATING- BLADE STALL I 0.2 0.3 0.4 0.5 0.6 ADVANCE'"RATIO, M Figure 9. Rotor lift/effective drag
DISCUSSION OF THE EFFECTS OF DESIGNPARAMETERS
3.1-Factors That Influence Hover Efficiency As pointed put in section 2.1, the two components of rotor hover power are induced and -profile power. For a given number of blades, induced power is influenced very strongly by ; 'blade planfprm;and-tip; shape (spanwise chord distribution) and twist. Profile power is affected ;!
by the choice-and distributipriJof- airfoil sections, the surface roughness of the blade, and the operating Reynolds number of the blade.
5.1. l~Tip~ sh ape.'—" Induced power is reduced by a more-uniform span wise distribution of induced velocity across the blade. This induced-velocity distribution is controlled by the spanwise-lift distribution,_which_can.be. changed by either twisting the blade or modifying its chord distribution. , Vector DE in Figure "10 is-an example of the improvement in induced power possible simply by modifications in the tip-chord distribution of a rotor blade. The reduction in Kj^j} shown was obtained by changing the outer 7.5-percent radius of a square-tip blade to a modified 3 l2 elliptical-chord distribution. Note also the improvement in 5g /c^ obtained as a result of overall operation of the blade sections at a more-optimum lift distribution.
Figure 11 illustrates the Kjjqj-j reduction obtained by progressively increasing the per- centage radius of the outer portions of the rotor blade changed to a modified elliptical-chord distribution. Note how KJ^D versus span/radius modified becomes asymptotic to the limiting value of Kr identified by the results of reference 3.
3.1.2 Twist. - Figure 12 illustrates the effect on figure of merit and Kj^p of twisting a square-tipped, constant-chord blade with a linearly varying twist. Use of a nonlinear twist distribution would result in still"greater reductions in 3.1 .3.,Airf oil section. — Figure 13 illustrates the effects of airfoil-section type (camber and thickness/chord variations) on hovering efficiency. The upper plot shows the difference in performance between a rotor blade with uncambered and slightly cambered airfoil sections.
Note that tiie cambered~blade exhibits improved FM due to increased section L/D values. Note also that as tipMach number increases, the cambered blade is affected by compressibility-drag increases more than the uncambered blade.
> ' . . . . . - . . . .
( | The lower plot shows -a; comparison in FM between the cambered rotor from above and \a rotor with-an~air-foilsection of ^reduced thickness/chord ratio and less camber. Note that the {resulting rotor exhibits higher overall hover efficiency than the uncambered, thicker-section ; rotor in the upper plot. This is due to the increase in section L/D caused by the camber com- bined with the improved compressibility characteristics resulting from the thinner airfoil section.
/ 120 i- 110 - Figure 10. Improvement vectors for figure of merit with elliptical-planform blade tip 1.20r O ? 1.15 1.10 ; 0 2 4 6 8 1 0 12 TIP SPAN/BLADE RADIUS - PERCENT Figure 11. Effect of tip shape on rotor induced-power factor 0.78 HOVER EFFICIENCY 0.76 0.74 E °' \t °- :uu 0.68 ! OC lr> & 0.66 I U- t 0.64 0.62 0.60 -5 -10 -15 ROTOR INDUCED-POWER FACTOR o DC LU Q- I o o Q UJ OC 0 -5 -10 -15 ^TWIST ~ DEGREES Figure 12. Effect of twist on hover efficiency and on rotor induced-power factor L CAMBER SMOOTH ROTOR BLADES — 0.8 — - 0.7 CN "" ^--^ "^V,,^ ?5 a.
<-> 0.6 ~ o" C = 0.9 L i- — ; 0.5 I l l l l l l l l l i- - LI c > 0 NACA 63 A015 (230 MEAN LINE) • i~ NACA63 -015 - : 0.8 a u j v S - : 0.7 ^* -- Li c - 0.6 c = 0.7 L LL j _ a • 0.5 I l l l l l l l l l LI « c ? 0 u a _ • 0.8 ) C • — -*^ ^ SB> /~ 0 - t \~ •7 a - 0.6 c = 0.5 L _ 0.5 I l l l l l l l l l LI 0.4 0.5 0.6 0.7 0.8 0.3 ROTOR-BLADE TIP MACH NUMBER AIRFOIL SECTION ™ 0.8 -rrz — ^.
N ••«.
«1 - c5- 0.6 c = 0.9 J L (^ _ 0.5 O I l l l l l l l l l ^^ n \j d hf r- —- 0.8 *~ ^"^^ LLJ 0.7 NACA 63 A015 (230 MEAN LINE) ^i — ^ _ NACA 63^012 (130 MEAN LINE) 0.6 Q7 O LLJ *- £E I l l l l l l l l l ' | -1— e> 0 £ - 0.8 ^^^^^™^^"^^^^^^'n ^•^«T" — -«.
}r 0.7 — o cc ~~ 0.6 c = 0.5 L 0.5 i i i V i i i i i i 0.3 0.4 0.5 0.6 0.7 0.8 ROTOR-BLADE TIP MACH NUMBER Figure 13. Effect of camber and airfoil section on rotor hovering efficiency 3..1 A-Surface roughness.-—-Figure 14 illustrates the.importance of maintaining relatively smooth surface conditions on the leading edge of a rotor. Note that the standard-roughness condition results in a much more severe degradation in hover performance compared to the slightly rbughened leading-edge surface" also shown. This is because the standard-roughness condition corresponds;to actual grit affixed to the blade surface, rather than the surface wavi- ; - ness of the'latter. Obviously the former results in much more turbulent skin-friction drag than ; the latter, j ; ^ >; \r~""*~~:"~i", . ' . . ".": '' ' "" ~' 3rlc5~"Reynolds-, numberr^ Figure~1:5 illustrates the effect of Reynolds number on airfoil-section profile^drag coefficient. ^Reynolds numbers which occur operationally in hover *~ ' ' ' " ' C i " 7 jean range_frorn_4 x; ICr foj^modelrscale blades to 2 x 10' for full-scale, wide-chord blades operat- ing at moderately high tipspeeds. Note that at the Reynolds numbers associated with full-scale rotors, the'minimum Cjj obtainable corresponds to the flat-plate turbulent skin-friction-drag trend line. Note also-that-the variation in Reynolds number referred to above results in a 47- percent variation in profile C-Q, Obviously, it is of great importance to have airfoil-section data available over a wide range of Reynolds numbers for use in correcting model-scale results to full-scale conditions and for analytically accounting for the rapid variations in blade chord associated with planform modifications such as those noted in section 3.1.1. ' ; 3.2 Factors That Influence Cruise Efficiency As note"d~earlierin "section; 272,Totof crliise efficiency, or L/Dg, is a function of both induced and profile effects. The induced component depends primarily on twist and blade plan- form. The'prpfile[componentJs..4®R6?J^??t.pn^j)][c»de_planfqnn airfpil sectipn characteristics, I : blade-surface roughness, and Reynolds number. System or configuration variables which can be applied to modify rotor L/Dg include higher-harmonic control and lift offset. As seen from Figure 7 in section-272, induced effects account for a substantially smaller portion of cruise ^ power thari.profile effects.- Thus, the greatest potential for L/Dg improvement lies in reducing the profile-Spower component.
3.2.1 Rotor variables. — 3.2.1.1 Twist: In;fprward flight, since the effect of induced power on overall perform < mance is small, the effect of twist on overall performance is correspondingly small. In fact, the , large amounts of stalffc design twist which are desirable from the point of view of increased ; : •2 hover efficiency prove: to be a-disadvaritage in forward flight, producing higher blade stresses :-, and more.vibrations than.a lower_twist.-Live, or aerodynamically adaptive twist, offers the ; possibility,of designing a rotor with a large amount of static twist for improved hover efficiency ;!
'• that is automatically reduced in cruise flight, lowering both blade stresses and vibration.
: . . _. IlLL^llJiLJijJilLLiJjalilLiliiJ..- -• :- . ..:. :....: ..
( * n Figure 16 illustrates the almost negligible effect such a twist change has on the cruise .
,,i JYDg of a rotor. The data shown was obtained by inducing a nose-up pitching moment on a , model rotor, thus tending to unwind it.
i CONDITION OF LEADING EDGE SMOOTH SHELLAC WITH BRUSH MARKS (STANDARD ROUGHNESS 10.8 0.7 CO 0.
i i i i i I i i !>.
d is"
iP.6 C, =0.7 LU i _ L.
cc 0.5 i I I I I I I I I I I I i 0 0.8 ;0.7 0^6 C = 0.5 L 10.5: I 1 I l U V I I °-0.3 ; 0.4 0.5 0.6 ; 0.7 !"0.8 ROTOR-BLADE TIP MACH NUMBER . ... .
-, Figure 14. Effect of roughness- on rotor hovering efficiency 1.0 Si AIRFOIL SECTION DATA t/c MAX SYMBOL AIRFOIL (%) DATA SOURCE O SCHMITZ N-60 12.4 HOERNER o 8 o P . AVA 367 16.3 HOERNER o HOERNER O ARC-CLARK-YH 11.7 HOERNER *3 ARC-RAF-34 12.7 0.1 A NACA64A-212 12 HOERNER TJ ..
O V NACA 63-21 2 12 HOERNER Cg = 0.7 o NACA 64 -215 15 NACA TN-1283 LLJ O NACA 0012 12 D2-23280 O D2-24066-1 O BV 23010-158 10 ill NACA REPT 586 0 NACA 23012 12 QS O O XC V, ^9R>-^ DC
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DRAG
i i i i i M
i i i i i i i i 1 1 1 1 I 1 1 1 i i i i i i i i 0.001 6 ,8 107 10' pVc REYNOLDS NUMBER, Figure 15. Effect of Reynolds number on coefficient of drag -n 0 a o aiV i 7 r ^ / O o ]BASIC C»
u_ o
u_ LU — iAC = 0.033 D- O M IVI ! 0 O O DC LU 0.4 0.3- ADVANCE RATIO, Figure 16. Effect of change in blade twist on aerodynamic efficiency . 3.2,1.2_Bleide.planformand.,tip. shape: The primary effect of .blade planform, or tip- shape modification, on a rotor in forward flight is the redistribution of the lift in a more efficient manner, allowing operation at more-optimum section C^'s and thus reducing the section profile drag (and therefore rotor profile power). Table 1 shows the improvement in cruise L/Dg obtained by revising the outboard 10 percent of a square-tip rotor blade to a modified eUiptical^ord to TABLEl EFFEeT:OF-CHANGErEN BLADE-TIP PLANFORM T ON AERODYNAMIC EFFICIENCY Percentage Improvement in Rotor L/Dg Due Advance Ratio, M To Elliptical-Planf orm Tip 0.180 10.86 0.260 10.64 ; 0.333 7.64 0.380 5.62 1 3.2.1.3 Airfoil-section characteristics, surface roughness, and Reynolds number: Airfoil- 'j section drag characteristics in cruise flight influence profile power in the same manner as in hover. For example, increased camber results in higher section L/Dg's with a resultant profile- ; drag reduction and an overall rotor L/Dg increase. However, this increase in L/Dg is likewise , accompanied by increased vibration and blade stresses due to the increased section Cjyj. For this reason, large values of camber, with their resultant performance benefits, are generally avoided in rotor-blade design. Also, as in hover, the surface roughness and operational Reynolds number are important in determining section profile-drag level, and therefore rotor profile ' power.
Of critical importance, however, are the section-compressibility characteristics of the airfoil such as drag-divergence Mach number and the rate of increase of CQ with increasing Mach number (dG^/dM^r (Improvements in this technological area are of benefit to both hover ; and cruise iflight. Howevef,;as; evidenced by the increasing dominance of the advancing-blade " compressibility effects on total profile power as displayed in Figure 8 of section 2.2, such im- , provements are of paramount importance in cruise flight.
i;.; ,..'.'...t.l.' !;.!!.'.:..! '„'.:. _ . ; i ; : . . ; ; . . ' • '. .. ...
; 8 Traditibnallyj-hightdrag-divergence Mach numbers have been obtained by using relatively •thin airfoiLsections; but with the recent advances in supercritical and transonic airfoil technology, it should be possible to design moderately thick airfoil sections with favorable compressibility ; '!'characteristics which, at the same time, do not compromise the low-speed high-lift characteristics 'excessively. : 3.2.2 Rotor-system and.configurati.on variables. — 3.2.2.1 Higher-harmonic rotor-system control: As noted in the previous section, advancing-jblade compressibility effects represent a sizable proportion of the rotor profile power at higher flight speeds. Instead of tailoring airfoil sections to obtain more-favorable compres- sibility characteristics,._^^ertmtiv^appj:oach is to Jpwej the rotor.tipspeed in forward flight, thus reducing the advancing-blade Mach number.
• ' This tipspeed-reduetiony however, increases the rotor advance ratio at a given forward speed, resulting .iri; increased .amounts of retreating-blade stall and an enlarged reverse-flow ; region. Thus, by simply reducing tipspeed, we have replaced advancing-blade-compressibility problems with retreating-blade-stali problems. It is possible, of course, to relieve retreating- blade stall;simply by unloading the rotor with a wing; i.e., compounding the helicopter.
The other recourse is to use higher-harmonic control in a reverse-velocity-rotor system, obtaining substantial lift in the reverse-flow region and redistributing the rest of the lift between the advancingHand "the^ foYe^and^ift'sections of the rotor disk in order to obtain a substantial rotor L/Dg increase. Figure 17 illustrates the L/Dg's to be expected from such a configuration.
The upper;L/Dg trend refl.ects.rQtor,operation with no attempt at trimming out rotor-blade flapping. Application of control power to eliminate blade flapping results in the lower L/Dg line illustrated. The L/Dg levels achieved are particularly noteworthy, considering that current- technology rotors-are-operating at L/I>g's of about 6 at advance ratios on the order of 0.4. The main disadvantages of such a system are the complexity of the hub and control systems and the requirement to operate at:pr near autorotation, with auxiliary propulsion being needed.
3.2.2.2 Lift-offset-rotor configurations: A second approach to the reduction of advancing-blade-epmpressibility effects is in the use of the highly lift-offset rotor. In this concept, - rotor tipsp!eed is reduced with the_ attendant increase in advance ratio (and inherent potential for retreating-blade-stall problems), and retreating-blade stall is dealt with by dumping lift on the retreating side of the rotor and increasing the lift by a like amount on the advancing side. The proportion of lift reduction on the retreating side to lift addition on the advancing side results in a given,fraction oflateral lift offset relative to the rotor center of rotation. Figure 18 illus- trates a typical trend of L/Dg and lift as a function of percentage of lateral lift offset. The major disadvantages of this type of system are in the large aerodynamic rolling moment generated by the lift offset and" the" necessity; asTinthe RVR concept, for auxiliary propulsion at higher speeds. The rolling moment can, however, be dealt with by various configuration approaches.
" Four such 'configurations are: 1 i • ' , ' , ' • • . r'>u. .
, 1. The coaxial superposition of two counterrotating, highly lift-offset rotors (as in the Sikorsky <' ABC concept).
' ; 2. Two oppositely rotating, highly lift-offset rotors in a tandem configuration.
, 3. Two oppositely rotating, highly lift^offset rotors in a lateral (side-by-side) configuration.
.25 16 (-» UNTRIMMED.
OPERATION o o _ TRIM uu 8 O i FLAPPING Q O oc 4 ID I J_ 0.4 0.8 1.2 1.6 2.0 ADVANCE RATIO, Figure 17. Estimated aerodynamic efficiency of the full-scale reverse-velocity rotor i V = 453.7 KPH (245 KN) | M = 0.7 M = 90 ! 1(90) °' .
= 0 DEG = 0.1110 I l l Q
II
;x o ; 12,000 50,000 m I 10,000 m Q 40,000 51% LIFT INCREASE X 8,000 X < 30,000 ^ 6,000 I- UL 20,000 4,000 CC cc O O 10,000 2,000 O <r I I 20 40 60 80 PERCENT LIFT OFFSET - PERCENT RADIUS.
Figure 18. Trend of lift and aerodynamic efficiency with percent lift offset ; 27 :.4. JJse_pf_a..single, higfalyjjfjioffset mtp^ii^algeome.trically. asymetrical single-main-rotor with tail-rotor configuration. : The potent® disadvantages of these various configurations can be summarized as follows: Approach 1 results in a relatively heavy rotor mast and complex control system due to : ~ :*^—™~ *—~~~...-_—.— ~~—>- , *.*——__^.:._z.'iv.1 "ii7jz'_"";"u__L_ri :™"J~.rj..~~i^.~ ".:;.'..? ." ' , \>'~'t71v " r the need to provide rotor-blade clearances and differential-pitch inputs to the counterrotating j coaxial roioreahditoabsorb:AerDllingmOTn^ "-—^- ',: Approach 2:results in a heavy fuselage structure since the portion of the fuselage between the rotor^rriasts miost absorb the opposing rolling moments, in essence converting the fuselage s into a large torque tube. I Approach 3-results in a configuration which takes up a lot of space due to its basic geometry and has a heavylwmg and strut structure since the rotor lift is concentrated at the tips of the) wing/strut and the major portion of the vehicle weight is at its center. Thus the wmg/strut;¥tnjicture mustiabsbW concentrated vehicle weight, and opposing aerodynamic rolling moments.
Approach 4 requires extremely careful design and careful control of the center of gravity and weight and balance.
C U R E N t
4.1 Previous Research and Development 4. Lj_^proyementhv jRotor Figure of Merit. - During the first 30 years after the first successful helicopter flights in the 1930's, figure of merit had only increased in percentage from the high 60's to the low 70's. But inthe last few years, motivated by the U.S. Army to develop the lifting-capability of "cargo-carrying helicopters, the slope of figufe-of-merit improvement versus time has-been increasing.
Limited investigations in the past have been carried out to determine the effect of vari- ations in airfoil-section camber and thickness on hover efficiency (ref. 4). Current interest at Boeing Vertol revolves around investigations into the effect of tailoring the tip-chord geometry to increase hover efficiency by reducing the induced-power component ,Cref. 5).
! 4.1.2 Improvement in Rotor L/Dg. — Work is in progress in industry, NASA, and the -., Army to increases the L/Dg (presently near 6) to values approaching 7 or 8. The most interest- j'i ing of this work is that variable twist changes the span and azimuthal loading of the rotor and decreases the blade cyclic 'loads-. With variable twist, both the aerodynamic and structural I speed limits of rotors as well as increased L/Dg at a given airspeed can be obtained. The variable L twist can"be"put in'fnechanically (Kaman) or through blade aeroelastic features of such nature ' • as to favorably redistribute the loadings over the rotor disk (Boeing, ref. 6).
Experiments and analyses are also being conducted (Boeing/Army) to extend efficient L/Dg values to higher advance ratios (/u -» 0.6). Results of preliminary tests show that rotor propulsive:forceswith~adequate"lifraridL/Dg's of 7.5 can be developed by conventional rotors at forward speeds up to 463 kph (250 knots) by the use of high values of cyclic pitch.
I Limited testing has also been conducted to determine the effects of blade-tip-planform '> modifications on rotor cruise efficiency (ref. 7).
f 4.2 Technological Gaps and Problem Areas ' - ! Existing gaps and problem areas in the technology of rotor hover and cruise performance ', have been identified as follows.
. - & • , - , • , . . - . - . .
4.2.1 Hover Performance. — There is no broad base of low-Reynolds-number airfoil- section data for use in.rotpr hover-performance analyses.
~ j """• """""~ •" • There is no one rotor hover-performance analysis which easily and accurately reflects the effect of bJade-tii)-planfQrm .shape on hover performance.
• ~ _ . - . - - . - . . . . . _ . .
• There is no extensive; base of rotor test data (either model or full-scale) which reflects the interplay of various combinations of twist, planform, blade-tip geometry, and airfoil ; sections on hover performance.
j 4.2.2 -Cruise Performance. -= -There-is-no one.rotor cruise-performance analysis which easily and accurately reflects the effect of blade-tip-planform shape on cruise performance.
t t ' . . . . ' ' , . . . ' ' , ' - . ' I • Therels no extenave^baseoffmodel or full-scale rotor test data which shows the effect of : i blade; parameters; such-as tip shape:and airfoil sections on rotor profile and induced power 1 (and therefore L/Pp). _ _
!' "i — - * — - ^ --. - -
: j • Ther6 is a lack of detailed knowledge on the interaction of aeroelastic and dynamic effects I on areas such as retreating-blade'stall and their subsequent effects on rotor profile power j (and therefore.rotor
^JJDRESEARCH AND DEVELOPMENT REQUIREMNT^]
•• - - -
Rotor technological gaps and problem areas are defined in section 4.2. The following represents a summary of the actions which must be taken in order to realize the goals set forth in reference X .
; f i • ' • ' i 5.1 Improvement in Hover Efficiency (Figure of Merit) As indicated by .Figure 6 in section 2.1, reducing induced power is a more powerful means for improvingjhe^hover efficiency of current rotors than reducing profile power. Accordingly, the primary emphasis should be placed on gaining a better understanding of the interaction of blade design parameters and induced-power effects, and a secondary emphasis placed on improv- ing and developing new airfoil sections.
; •_ . . . .
This can be accomplished by: • Considerable model-rotor hover testing of various combinations of planform, tip shape, twist,,and airfoil sections to obtain a broad range of both induced- and profile-power data.
• Development of improved analytical techniques for accurate hover-performance prediction of rotor "bladesTdesignedTo benefit" from the technological data base obtained in the first step. Such techniques must be capable of dealing with rotor blades having spanwise chord variations andnonlinear twist and operating at high thrust coefficients.
• Considerable airfoil-section testing at low Reynolds numbers to build up a data base for use n in the! hover-performance prediction of various planform and tip-shape blades with the improved analytical tools noted above.
• Additional airfoil-section development work concentrated in the area of improving compressibility-drag-rise characteristics such as drag-divergence Mach number (Mj-j^) and dC dM. - - .
D/ • Employment of the rotor-optimization trends obtained from the broad model-rotor test programTin "conjunction with the improved analytic performance and design techniques to produce a full-scale optimum hovering rotor.
; : .5.2 Improvement in Cruise Efficiency (L/Dg) •'< As discussed'in~se'cti'b"n~3:2"th~e~r«iMtion"of profile power iri cruise flight is of the utmost importance in increasing rotor L/Dg. Such a reduction can be achieved by concentrating on obtaining a better understanding of the mechanisms inherent in both the retreating-blade-stall ,and advancing-blade-compressibility components of profile power and using both rotor design parameters and vehicle configuration approaches to reduce their effects.
This can be accomplished by: Considerable model-rotor testing (wind-tunnel) of various combinations of planform, tip shape, twist, and airfoil sections to obtain a broad understanding of the effect of these blade ;design parameters on rotor Development of improved analytical cruise-flight performance-prediction techniques which can accurately"reflect the interaction of the blade design parameters specified above on rotor forward-flight-cruise performance.
Development of airfoil sections with improved compressibility characteristics such as mcreas^dlirag^div^rgenceMach number • Investigation of the.interactions of blade aeroelastic/dynamic effects on areas such as retreating-blade stall and determination of the value of concepts such as live twist and higher-harmonic control in both reducing blade stresses and vibrations and increasing rotor ;L/Dj| by reducing the effect of the retreating-blade-stall component of profile power.
• Further study pf adyanced rotpr-system concepts such as the reverse-velocity rotor and configuration approaches such as the highly lift-offset rotor as means of reducing the retreating-blade-stall component of profile power.
; 5.3 Integration of Rotor Design Parameters Some of the approaches to be investigated for improving rotor efficiencies are compatible with both flight regimes; fdr example, improvements in airfoil-section compressibility character- istics can-be-of-benefit to both hover and cruise efficiencies. Some are not; the RVR system uses a double-ended airfoil section for obtaining lift in the reverse-flow region in forward flight, thus increasing L/Dg. This type of airfoil section, however, penalizes the rotor hover performance compared to a more conventional section. Thus, in developing an optimum rotor, care must be exercised in the integration of the rotor design parameters to obtain this combination for maxi- mum efficiency in hQver,and_cruise.
5.4 Schedules and Estimated R&D Costs for Improvement of -',' Rotor "Efficiency in Hover and Cruise Figures 19 and.21 are the schedule and estimated R&D costs for programs to improve i rotor hover and cruise efficiency.
i [ . - . : • - . '• •'- - :--••'- . ' . . . . ' . ' . ' • ' . - i . ; • ' . ' . . ' . . - i . ' - ' . ' .
: \"-' Figufes~20 and 22 are plots of program expenditures and improvements in figure of "merit and cruise efficiency as a function of time.
PROG RAM YEAR COST ($) ITEM 1 2 3 4 5 6 7 200,000 HOVER TESTS (MODEL) TO ISOLATE GEOMETRY EFFECTS 50,000 AIRFOIL TESTS TO OBTAIN CHARACTERIS-
•
TICS AT SAME REYNOLDS NUMBER AS HOVER-MODEL TESTS 75,000 DEVELOP COMPUTER MODEL THAT PREDICTS .
EFFECT OF TIP SHAPE, TWIST, PLANFORM, AND AIRFOIL 100,000 ASSESS THE LOADS AND PERFORMANCE OF THE HIGH-FIGURE-OF-MERIT ROTOR IN HIGH- SPEED FORWARD FLIGHT BY COMPUTER
- •
ANALYSES 200,000 DESIGN AND CONSTRUCT ROTOR MODEL THAT CONTAINS THE BEST BLADE SHAPES H AND AIRFOILS FOR HOVER AND CRUISE 200,000 CONDUCT MODEL TEST AND WRITE REPORT 4,000,000 DESIGN AND CONSTRUCT OPTIMUM FULL- : SCALE ROTOR FOR RSRA 3,000,000 INSTALL, TEST, AND EVALUATE OPTIMUM
_
ROTOR ON RSRA 150,000 WRITE GUIDELINES
'-
Figure 19. Program schedule and estimated research and development costs for increased figure of merit CXI 0.851- oc 0.80 FIGURE OF MERIT oc I 0.75 0.70 PROGRAM YEAR Figure 20. Improvement of figure of merit and expenditure as a function of time i ^ -i., vU- .-- PROGRAM YEAR 1 4 7 COST ($) ITEM 2 3 5 6 200,000 PREPARE COMPUTER PERFORMANCE PROGRAM TO CONTAIN LIVE-TWIST ROTOR CHARACTERISTICS COUPLED WITH HIGH FORWARD Tl LT AND CYCLIC PITCH 150,000 DESIGN, BUILD, AND TEST WIND-TUNNEL MODEL FOR OPTIMUM PERFORMANCE • • AND LOADS !
50,000 REVISE COMPUTER PROGRAM 300,000 CONDUCT FURTHER WIND-TUNNEL TESTS TO EVALUATE EFFECTS OF AEROELASTIC ADAPTIVITY ON OPTIMUM SOLIDITY, • • • AIRFOIL CRITERIA, ETC 175,000 ' REVISE COMPUTER PROGRAM .
7,000,000 DESIGN, BUILD, AND TEST FULL-SCALE
-
ROTOR FOR RSRA TO VERIFY PERFORM- • M • • • • • ANCE AND HANDLING QUALITIES 150,000 PREPARE DESIGN GUIDELINES BASED ON FULL-SCALE TESTS AND ANALYSES Figure 21. Program schedule and estimated research and development costs for increased rotor aerodynamic efficiency (L/Dg) 8.0 L/D IMPROVEMENT DC UJ Q.
UJ 6.0 Q LU CC TO H •Q UJ 4.0 LU a.
UJ X LU O ! 5 LU
£
2.0 < D PROGRAM YEAR Figure 22. Improvement of rotor aerodynamic efficiency and expenditure as a function of time
;P~TECHNOLOG Y AND DESIGN"INTERACTIONS'
The impact of rotor hover and cruise efficiency improvement on interacting technological areas and systems is discussed .in the paragraphs which follow.
6.1 Helicopter Noise Generation • _ : 6.ILL-Main-rotoriridise; — Table 2 shows how various rotor design parameters are related ! to different types of noise. Some of the more important effects can be summarized as follows: • Airfoils'sucK as~the Boeing~Vertol"VR-7 reduce low-speed slap because their shock-stall characteristics are better than the NACA 23000 series.
• Planfprm taper and twist reduce rotational noise because they move the blade-span loading toward the center of rotation.
An increase in disk loading increases the rotational noise due to the higher strength of the vortices.
Airfoil thickness ratio has a great effect on high-speed slap.
• Planform shape, especially tip planform shape, has a great effect on low-speed slap (0 to 30 knots) since its shape will affect the structure of the blade tip vortex.
All; of these effects on rotor noise must be given careful consideration when varying the rotor design.parameters.to increase rotor-efficiency.
6.1.2 Main/tail-rotor interaction noise. —Very little is known about the noise increase due to interactionbetweenthe main and tail rotors. It is known that tail-rotor noise is always higher in the presence of the main rotor than when alone. Under certain conditions, as when the helicopter is[flying aw_ay^fnpni the observer a pounding noise at the main-rotor-blade r passage frequency is measured. It is postulated that such pounding is tail-rotor noise being modulated jby the passage of the rhain-rotor-blade tip vortices through the tail rotor. This is an area that sliould be further researched so that such effects can be quantified and applied to the design of advanced rotors.: 6.1.3 Engine noise. - Improved rotor efficiency will have very little effect on engine- : jnoise level. ;The;increaseiin'fi'g^Tjof merit-will decrease required installed power by 11.4 percent; l >jSuch a reduction-in the-size~of-the p6werplant is estimated to decrease the sound-pressure level jof engine noise by one decibel; thus engine noise will not be reduced to any measurable extent by more efficient rotors.
,37 TABLE 2. EFFECT OF ROTOR PARAMETERS ON ROTOR NOISE Type of Rotor Noise Affected High-Speed Low-Speed Rotor Parameter Rotational Slap Broadband Slap Total Blade Area X Planform Taper X X Twist X (X) Tipspeed (X) X ® Airfoil Type
®
X X Planform Tip Shape X Surface Finish X X Disk Loading Blade-Span Load X Distribution X (X) Airfoil Thickness Ratio X X Design CL X X X No. of Blades X Freq only (X) Thrust
®
Chord X Radius X X X Aircraft Velocity (X) Items having major effect on noise 1 38 1 >.2 -Performance Improved rotor efficiency affects overall vehicle performance in several ways. The in- creased Hover "efficiency results in a lower hover-power requirement, with a resultant saving in installed power.; Compared to the 1975-teehnology-baseline compromise helicopter of reference - 1, the impr^^d^fficleiic^^otpr,helicopter_exhibits a reduction in installed power of 1.1.4 per- cent. This reduction in engine size is important because it lowers the overall absolute value of the fuel-consumption rate , improves specific fuel consumption at partial-power throttle settings, !
' and increases vehicle-specificTange.The improved-rotor Lr/Dj? results in a lower power required '. for a given! speed (and therefore lower fuel consumption) and a higher cruising-speed capability.
Overall, when comparedjoj^J9754echmolpgy-baseUne helicopter (ref. 1), the improved rotor results in a 2.9-percent increase in cruising speed at normal rated power, a 15.1-percent increase in specific range, and a 12.9-percent decrease in fuel (and therefore, energy) consumed.
6.3 Empty Weight Because of the performance improvements noted in section 6.2 with the resultant savings in fuel, and the iterative nature of the sizing process, the helicopter with the improved rotor .. exhibits a 3.83-percent decrease in empty weight. The EW/GW fraction remains unchanged.
' • 6.4 Drive-System Efficiency and Weight The improved-efficiency rotor should have no impact on helicopter drive-system efficiency.'.Drive-system weight wUl be reduced because of the reduction in installed power.
6.5 Hover With One Engine Inoperative (HOEI) The increased rotor efficiency in hover will result in a lower hover-power requirement, thus insuring smaller engines. The ratio of hover-power required to installed power will remain unchanged.
.-.-6.6 Reliability and Maintainability The maintenance of advanced rotor blades and hubs should be the same as, if not better than, present rotors because'befter materials and methods of fabrication will be used, along with ,a reduction in number of parts. If higher-harmonic control is used to obtain higher speeds, the ; ,control-systeni.design.sho iild.be .carefully surveyed for the effect of many more load cycles ; : applied during ;the life of the aircraft.: "' \ Control-systemltiiainTelia~n~cTwillliirobably''inclrease' during the initial use of advanced [rotors. However, since rotor-system (including upper controls) maintenance costs are only 15 , ^percent of total maintenance costs, small increases in upper-controls maintenance should not «: materially affect the overall cost of maintenance.
shaft to decrease drag. Therefore, maintenance considerations should be given to such internal configurations and the effects on transmission design arrangement.
......... 6.7 Production : The advanced rotor blades that will evolve should be no more costly to produce than present blades, even though they may have more complex planform and twist shaping, because : they will-be :constructed 6f composite-type fiber materials. Filament-winding techniques may even reducie cost. 'However, protection of the leading edge will require a complex formed-metal cap which 'can add to production cost. ^Research should be conducted to find methods for manufacturing these caps.
The upper-control! system may be more expensive due to the higher loads and greater number onload cycles if higher-harmonic control is used.
Bearinigless hubs will be used. This type of hub should reduce production costs through ; the use of jess-expensive material and fewer parts.
6.8 Rotor-Blade and Control-System Loads * ' !
Control-system loads WiU"beMgh^F¥nd the necessary static and fatigue strength m be designed into the parts parrying such loads. If higher-harmonic pitch is used, coupled with high once-per-rey cyclic, the;.. control loads may double from present loads. Both the control system and the blade design must consider such loading and research will be required to mini- mize such loads. ; ' . . . .6.9 Vibration The vibration level in the fuselage of helicopters equipped with advanced rotors will be low, approaching that of fixed-wing aircraft. This low level of vibration will come from the use of higher-harmonic control tareduce rotor vibratory loads at the source (i.e., on the blade) and from the use of vibration absorbers between the fuselage and rotor. Gust-alleviation systems will : further improve the ride quality. All three of these areas need more research and applied opera- tional experience."
IMPACT ON INITIAL COSTS AND OPERATING
The costimpact .of improving both rotor hover efficiency (FM) and cruise efficiency (L/Dj?) has been determined for the following categories: i . . .
• I n i t i a l ~ c o s t s ~ I — Research,Development/test, and engineering (RDT&E) — Initialinvestment • Direct-operating-cost - Reference 1 lists the RDT&E costs required to bring the rotor technologies up to the point where they can be applied to an advanced helicopter as: Hover-Efficiency-Improvement $7,975,000 Cruise-Efficiency improvement $8,025,000 Total Improvement $16,000,000 Thjese costs are not included in the-initial-investment (flyaway) costs of the vehicle.
The flyaway costs reflect only theT laborand material costs required to produce the vehicle after the desired level of technology has been achieved.
Compared;to the baseline helicopter (ref. 1), the initial-investment (flyaway) costs are 4.42 percent less. This is a result of the reduction in size and weight due to the more efficient • rotor systetnir ~ ~~~ I Direct operating cost (DOC) is reduced 6.75 percent from the baseline value. One of the major factors contributing to this reduction is a 12.85-percent decrease in fuel costs due to ..; the more efficient energy-consumption characteristics of this helicopter.
It is estimated that by 1990 22,000 helicopters of all sizes will be in operation in the United States and Canada^ These aircraft will have a total capital (initial) cost of $4.15 billion and will burn $32~5 rnMori worth of fuel per year (see Table 3).
If the cited savings in investment and fuel are conservatively applied to only one-tenth ; ; j of these 22,000 helicopters, the iiiitialrcost sayings can be (0.10) x (0.044) x (4,150,000,000) = ' >i $18,343,00"6| and jthe fuelfcost sayings can be (0.10) x (0.128) x (325,000,000) = $4,160,000 : J per yearrThus7the"savirigsTin"cajpitarcb^lIoW'cari'retiirh. the research costs, while the fuel- cost savings over 10 years "(in the I990's) will amount to three times the research costs. This is , a very good payoff considering only these two cost items. The value of increased productivity and smpj3ther ride will show up as wider acceptance and increased usage of helicopters.
- • -rj ,_TABLE-5._ECQNOM]^SjpF J.XISHNG^ND-EUTURE HELICOPTER ! : FLEETS iN THE-U^ITED STATES AND CANADA - Item "- Present Fleet 1990 Fleet* No. of Aircraft 5,500 22,000 Fuel/Yrjgal ; 107,000,000 430,000,000 : Fuel Cbst/Yr, $'/yc 80,000,000 325,000,000 Total Structural! Weight; Ib 4,600,000 18,000,000 Capital Cost, $ 1,036,000,000 4,150,000,000 Rental Price/Yr, $ 1,141,000,000 4,500,000,000 Place-Miles Available 2,143,000,000 8,600,000,000 Income/Aircraft, $/yr 200,000 200,000 _ Rent/Ifoc^Mfe -^ 50 Maintenance Cost, $ 103,000,000 410,000,000 *N6 inflation, ho improvement s 8.0 CONCLUSIONS^, £a-*.. :^j- ffifij-n - II- _ ay --,-,.-._,.»iiii_r.ii iiumkmiiip.-^'^ffiff^-^ .^1 *• . ..I The improvement of rotor hover "and cruise efficiencies and the R&D programs required to achieve jthem are summarized in the following paragraphs.
8.1 Improvement in Hover Efficiency The hover-'efficieney-improvement of current rotors depends primarily on the reduction of the induced-pdwer component; This reduction can be achieved by concentration of research in the areas defined in Table 4, carried out in conjunction with the development.of the following: '•"• "An improved rotor hover-performance analysis (computer program) • Design, construction; and test of a full-scale optimum hovering rotor • DocumentatioTi"orpidelines for the design of an optimum hovering rotor.
8.2 Improvement in Cruise Efficiency The cruise-efficiency improvement of current rotors depends primarily on the reduction of the proffle-powercomponentrThisTeduction can be obtained by concentration of research in the areas defined in Table 5, conducted in conjunction with the development of the following: An improved rotor cruise-performance analysis (computer program) Design eonstfuction, and test- of-a-full-scale rotor optimized for cruise r Documentation of guidelines for the design of a rotor optimized for cruise.
i t : TABLE 4. SUMMARY OF RESEARCH RECOMMENDED FOR IMPROVED ROTOR EFFICIENCY (HOVER FIGURE OF MERIT) Research Size Research Item Recommendation Priority Applicability Payoff High Blade Tip Shape Yes High All Medium Blade Planform/ Yes Medium All Taper Ratio High Twist Yes High All Medium Airfoil Sections Yes Medium All with Improved Compressibility Characteristics Medium All Medium Airfoil Sections Yes Operating at Low Reynolds Numbers TABLE 5. SUMMARY OF RESEARCH RECOMMENDED FOR IMPROVED ROTOR EFFICIENCY (ROTOR L/D ) E Size Research Recommendation Priority Applicability Payoff Research Item Yes High All High Blade Tip Shape Medium Blade Planform/ Yes Medium All Taper Ratio Medium Airfoil Sections Yes Medium All with Improved Compressibility Characteristics High Yes High All Aeroelastic Adaptivity (Live Twist) Effect on Retreating-Blade Stall, etc High All High Higher-Harmonic Yes Control All Low RVR Rotor System Yes Low All Low Highly Lift-Offset Yes Low Rotor System
Z39.0
1. Davis, S.J.; and Rosenstein, H. J.: Identifying and Analyzing Methods for Reducing the Energy Consumption of Helicopters. Boeing Vertol Company ; NASA CR-144953, National Aeronautics and Space Administration, Langley Research Center, Hampton, Virginia, November 1975. ' .
; I'"" ""'• ..... . ,"~~~~~",'~ ' ( . . . . . . - - . . - . - ; 2. StepniewskiyWr-Z;:- -Energy Aspects of VTOtrAircraft in Comparison With Other Ground and Air Vehicles..: Boeing Vertol Company; Second European Rotorcraft and Powered- Lift Airoraft Fo^rmn^ Bud^urg, Federal Republic of Germany; Deutsche Gesellschaft fur Luft und Rliumfahrt 'e.V'T^s^ch "510645,' D-5000, Koln, Germany, September 1976.
3. Harris, F.D.; and McVeigh M.A.: Uniform Downwash With-Rotors Having a Finite r I Number of [Blades. ; Symposium -oh-Helicopter Aerodynamic Efficiency, American ' Helicopter Society,! March 1975; 4. Dingeldein, R. G.: Consideration of Methods of Improving Helicopter Efficiency. NASA TN p!-734, National Aeronautics and Space Administration, Washington, DC, April 1961.
5. Rosenstein, H. J.: Wihirl Tower Investigation of the UH-61A Elliptical Tip Main Rotor.
Boeing Vertol report; D2rO- 11128-1, Boeing Vertol Company/Philadelphia, Pennsylvania, September 1976.
6. Doman, G.S.; Tarzanin, F. J.; and Shaw, J., Jr.: Investigation of Aeroelastically Adaptive Rotor Systems. Mideast Region Symposium on Rotor Technology, American Helicopter Society,August~l-976.~ • ' 7. Rosenstein, H. J.: Wind Tunnel Investigation of Modified Elliptical Main Rotor Tip.
Boeing Vertol report D210-1II17-1, Boeing Vertol Company, Philadelphia, Pennsylvania, August 1976.
8. McHiigh, Frank J.; and Taylor, Robert B.: The Reverse-Velocity Rotor - 300 Knots Plus Hover Capability. Mideast Region Symposium on Rotor Technology, American Helicopter Society, August"1976."
TlOJ BIBUOGRJI7HY]
• ^§i_;_^ .. . . „--„-... ... „ -.»-.-, -C-^*B** J J Harris, F: P.: Results of Tailored Geometry Tip Rotor Test (Hover). Boeing Vertol memorandum 8-7040-1-502, Boeing Vertol Company, Philadelphia, Pennsylvania, October 1975.
Landgrebey A. J.; Moffitt, R. C.; and Clark, D. R.: Aerodynamic Technology for Advanced : :) Rotorcraftr'^deast.R*0giion Sympdsium'c>n Rotor Technology, American Helicopter Society, 'August 1976. ~ - j — - ——--'—-..'- ; McHugh, Frank J.; and Harris, Franklin D.: Have We overlooked the Full Potential of the Conventib^rRoFo7?7"B^em^VelrtorC6rnpany; Preprint No. 903, 31st Annual National Forum, American Helicopter Society, Washington, DC, May 1975.
McHugh, Frank J.; and Shaw, John, Jr.: Benefits of Higher-Harmonic Blade Pitch: Vibration Reduction; Blade-Load Reduction, and Performance Improvement. Mideast Region Symposium on Rotor Technology, American Helicopter Society, August 1976.