Rotorcraft Analytical Improvement Needed to Reduce Developmental Risk-The 1989 Alexander A. Nikolsky Lecture

Charles C. Crawfonl , Jr .

Head, Aviation Tech17ology Branch Georgia Tech Research Institute Georgia Institute of Technology

Atlanta, Go .

Charles C. Crssford earned a Bachelors Dcgree in Aeronautical Engineering in 1954 and a Masters Degree in Aeronautical Eneineerine in 1970, both from the Georgia Institute of Technology. He served as a flight engineer - .. an active duty as an officer in the Air Force and later as a civilian, between 1955 and 1962, at thc Air Force Flight Test Center, supervising the test of all classes af aircrafl.

Mr. Crawfard transferred to the U.S. Army Material Command Headquarters in December 1962 to serve as Chief of the Technical Management Division and later as Deputy Project Manager for the Army's Light Observation Helicopter Program during the competitive development of the OH-4.41516.

In May 1966, he joined the Army Aviation Systems Command, establishing the Flight Standards and Qual- ification organization to develop the Army's capability to qualify its own aircraft. This organization has been responsible far the development of Aeronautical Design Standards and for the planning and managing of airworthiness qualification programs far all Army aircraft. Principal qualification efforts include: AH-I Series, CH-47 Series, Bell Model 214, and Black Hawk, Apache, and AHIP designs, including the cnpines which they utilize. The organization was redcsignated the Directorate for Engineering with both development and quali- fication responsibility far new engines and aircraft as well as engineering support of fielded aircraft.

Mr. Crawfard was a charter member of the Senior Executive Service, established July 13, 1979, and served as the Army Avistion Systems Command's first Technical Director. He was also the Director of the AVSCOM Research. Develonment and Eneinecrine Center until 1988. The Aviation Research and Technology Activity (with individual dirccturales en-luratcd ,,itl! UAS,\ Anne\, I ang1r.y imtl I . cv i~ Kr\r.arch (:cntcr* n#~<l the A+iotialn AnnlirdTccl~nol~oer llircrlnrntc ul Furl Et#sti\, Virxinia); tttc Amty Atinnirs R t \u~ reh itnd l)cvelepmcnt .\rli\it!; . . . .. IIBP Arm) ,%\ittion 1:nyinrcring Flight Acti,il); the 1)irrct~or.tlrs for Enginrpriug, ,\dvanrcd Sjatonli, and !.ire (:vrlr\ SnNuarr I<nginrcring; ontl thc 'lc.4 dnd Evalualiun managmnrnt nllicr, rcpanrtcd 10 hlr. Crnriord as the RD&E Center Director-

He is currently the Branch Head of the newly established Aviation Technology Branch at the Georgia Tech Research Institute. The branch's activities include engineering support of the U.S. Air Farce Special Operations Farces helicopters and rotary wing technology efforts in the area of structural integrity, preliminary design, and experimental aerodynamics.

Mr. Crawford has been a member of the American Helicopter Society since 1960, serving as Director at Large four terms, SecretarylTreasurer two terms, President, Chairman of the Board, and twice Chairman of the Vertical Flight Foundation. He was awarded an American Helicopter Society Honorary Fellowship in May, 1980. Hc received the Secretary of Defense Medal of Honor for Meritorious Civilian Service in 1988.

Preface butions to ou r beloved technology. For someone of m y hack-

hen 1 received notice of m y selection b y the Society to g o u n d to be so honored is very sincerely appreciated. w present the 9th Alexander Nikolsky Lecture, I was in- Although I have been a n active member of the American

dced startled. This honor is normally awarded to those stalwarts Helicopter Society for 27 years, I never had the privilege of

of the Society that have made substantial theoretical conhi- knowing Professor Alexander Nikolsky. However, his repu- tation i s familiar to most of ou r membership. A few have had the privilege of being his students. I a m quite indebted to one

Nikolsky Lecture presented at the 45th Annual Natiollal F~~~ of the of these. He is Dr. Robin Gray of the Georgia lnstitute of - American Helicopter Society, Boston, Mass., May 22-24, 1989. Technology. Years ago he served as m y advisor in graduate

3

4 C.C. CRAWFORD, JR.

school and has since frequently couns2led me on rotor wing technology. Of equal importance have been his frequent rec- ommendations of outstandine voune academic ~erformers who - < - thereafter have played key roles within our government in the business of advancine rotarv wine technoloev. Manv of the - - -, previous Nikolsky lecturers have reviewed Professor Nikol- sky's exceptional career and provide humorous accounts of specific activities. I will forego this, and simply point out that I sincerely believe that Professor Nikolsky, the renowned lec- turer himself, would grin at the idea of a practitioner rather than a theoretician presenting a lecture in his name. Toward this end, I am truly grateful.

Introduction The goal of this 9th Alexander A. Nikolsky Honorary Lec-

turership is to bring focus to what I believe is the most sig- nificant problem faced by vertical lift technology today: the need to design vertical lift aircraft that will "fly off the drawing board" without significant changes before achieving thc status of a viable operational vehicle. The problem is the shortfall in the realism and maturity of the analytical tools available to the designer for such an air item.

The potential of vertical lift vehicles is almost without bounds. They were required in order to land on and return from the lunar surface. Of interest to this Society, however, is the more common use of low disk loading rotorcraft. The public accep- tance of rotorcraft in roles such as an emereencv evacuation - , has grown dra~~~alically. For many victi111\ ol autc~mohilc ac- cidents, shipwrecks. ennhuuakes. and fires in hieh-r~sc hu~ld- - ings; rotorcraft have providid the life line to survival. In military missions, they are required to position troops to occupy and hold terrain, which is the fundamental purpose of armies. This application, which started years ago through simple medical ekua t ion and observation hission<, has now expanded to the role of being a complete weapons svstem, with a wide range of firepowermade thrbugh integrated target acquisition systems operating over the full range of their respective sensor technologies. -

- The economics of developing, producing, and operating ro- . - .

torcraft have been and will continue to he under constant scru- tiny. The obligation to reduce the acquisition and direct operating costs of the& air items rests primarily with the design team; and using communities that set requirements and bear the bur- den of bcingproponents. Thc matter of improving the cfficiency of the development process clearly spans a larger group of organizations. In addition to develooers and users. a concerted ~~~~~~

effort to improve this process must'be supported by academic institutions and both the legislative and executive branches of the federal government, including the National Aeronautics and Space Administration (NASA). It is truly a shared responsi- bility of these parties to nurture the fundamental research efforts that will serve the public interest in defense and comnicrcial applications of the resulting technology.

As an example-within the past couple of years, NASA conducted a study of the nation's wind tunnel facilities. The study considered the status and collective capacity of these facilities to support future aeronautics re~earch'~ro&ams. The study had NASA, industry, and Department of Dcfense (DoD) input. The results were presented to Congress with an agenda for renovation of these vital national assets. Congress approved and appropriated funds for the Wind Tunnel Revitalization Pro- gram. Specifically, the Full-Scale Aerodynamic Research Fa- cility (40 x 80 wind tunnel) at the Ames Research Center was designated to receive modifications to enable aeroacoustic ro- torcraft research. Therein lies a lesson on how the technical community and government can work together to address the problem of improving our national aeronautics research capa- bility.

Constant review and analysis of the weapons development process has been perfumed by the executive branch of the

JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

federal government. Studies have been performed and elaborate reports written at virtually every level, from special Presidential commissions through the General Accounting Office, to special groups within the DUD and its military services. Many impor- tant suggestions of how to improve the developmcnt process have been compiled. A large maioritv of the new ideas have - " . been managerial in nature, such as "design freedom," "spec- ification tailoring," "concurrency," and "program stability1 financial risk." Of equal importance have been the studics undertaken by the legislative branch of the federal government with the same objective, also resulting in constructive "man- agement type" suggestions.

Our Iiu;npean ;&I So~ ic t countcrpdrts h;~vc dcnionstratcd s vihrdnt and su.;ta~ncd interest in rotorcr;~ft rcscarch and ~lcvcl- opment. There was a time when we could say that thc foreign rotorcraft interests were mainly ticd to the momentum of the U.S. activity. That is no longer true. Our nation is facing a greater technical challenge in worldwide aviation markets than ever faced before.

This paper will suggest that the greatest contribution for reducing the risk of the developmcnt process may be a "hack to basics" technology initiative. In my view, this is the most critical missing ingredient. Its implementation rests with all, but the most painless birth of such a new initiative can occur by a recognition and appreciation of our ~roblems coming from the legislative side of'tke government. clearly, the benkts to the nation, both military and civil, are too important to sweeD the issue under the rug: Thc cver i,1ouari11g ialrre of rirne and the deplerion of available real esrare will boost the imponance of vertical lift aircraft to the top of the aeronautical technology heap for both military and civil applications.

1 will illustrate the nccd for a "hack to basics" rotorcraft tcchnology initiative by first reviewing the multi-discipl~ncd t c c h n ( ~ l ( ~ i ~ c~~rn~lcxi t ic i oirotary ~ i n ~ i i r c r n l i , not to 111cnt1o11 thc cvcr incrcaving technical co~nplcx~t~csol I I I I ~ S ~ N I ~ cquili~iicnt package integration essential fo r~a complete wcaponss~stem. However, a sound air vehicle is thc first order of busincss. Second, I will review by reference the higher risk elements of the design tcchnology and will compare the scopc of the pre- liminary design process for rotorcraft to their fixed-wing, but runway dependent, counterparts. Third, the shortfalls in the process will be illustrated by a review of the major design changes that occurred during the development of two major U.S. Army Aviation systems: the Utility Tactical Transport Aircraft System (UTTAS) and the Advanccd Attack Helicoptcr (AAH). Progress made in improving analytical tcchniques since their development will be pointed out. The need for further substantial improvement in this process is the capstone of this lecture. Sincc no "short tern-quick fix" is either logical or practical, 1 advocatc that the best course of action to insure viable military air items is the continued use of the proven competitivc "fly-off" acquisition strategy. 1 also plead that the work of developing and validating improved rotorcraft an- alytical methods be revitalized.

Interdisciplinary Complexities of Rotorcraft I believe that all Society mcmbers would agree that the tech-

nical complexities and interdisciplinary requirements for ro- torcraft far exceed those of fixed wing aircraft. In this regard, 1 mean no disrespect to the tremendous challenges faced by thc developers of the hypersonic "Orient Express" transport. However, the aerodynamic, aeroelastic, anddynamic interfaces of rotorcraft design represent the utmost challenge to our spe- cialist. The three principle areas that are the toughest to predict during the design phase (thus requiring significant optimization in flight tests) arc vibration, loads, and aerodynamic perfor- mance. Later in the oaoer. I will also build a case for significant . . . - risk in defining realistic empty weight for rotorcraft. Many papers have been written dealine with the interdiscinlina~ as- pects of rotorcraft design; however, 1 believe the &lurk has

JANUARY 1990 '1989 NIKOLSKY LECTURE 5

been best portrayed in a single ch& by Dr. David Peters of Georgia Tech. Figure I contains some minor rearrangements of his work. I will use it for a brief discussion to set the stage for the problems of rotorcraft design and optimization.

The basic dynamic components of a helicopter and their re- lationship to each other are shown graphically in this figure. It displays thc complex interdisciplinary intcractions that the rotorcraft technical community must accommodate. It considers only a single lifting rotor, but would be more complicated for multi-rotored configurations and anti-torque rotor systems. The upper three blocks represent the main lifting rotor. The rotor is divided into three basic theoretical models: blade aerodyn- amics, blade structural dynamics, and external flow field. These three theoretical models are not only highly coupled to each other, but are also coupled with the other areas of the dynamic analysis, airframe dynamics, airframe aerodynamics, power plantldrive system dynamics; and, of course, the flight control system. The complex interaction between these disciplines makes dynamic analysis one of the most difficult rotorcraft problems.

The forces that drive the vehicle result from the angle of attack on the rotor blade airfoil sections. This angle of attack is made uo of three comoonents. One comoonent is duc to

These equations describe the blade motion when subject to the applied airloads, (which can be a difficult problem in itself if composite blades are used). This blade motion in turn feeds hack to the blade airfoil theory by contributing to the angle of attack. This feedback loop represents the inherent aeroelastic stability problem that highly flexible airfoils produce. In ad- dition to the airloads, the blade airfoil theory also calculates bound circulation. This sheds a vortex wake, thus producing an induced external flow field. In addition to the lifting airfoil theory, and the blade structural dynamic theory, an induced flow theory (ranging from simple momentum theory to a com- plex free wake analysis) is needed to convert the time history of circulation into induced flow. These computations are ex- tremely complex. A complete tightly coupled solution may not be practical for engineering application. Thus, the coupling between those must be carefully adjusted to achieve an ac- ceptable balance between accuracy and computational effi- ciency.

In addition to high coupling among the three major rotor dynamic theories, the rotor dynamics are also coupled with fuselage dynamics. The blade structural dynamics is transmitted to the fuselaae in the form of hub loads-the shears and mo-

blade control enterei through the swasiplatc, another is ments whichdrive the fuselage motions by way of the trans- due to the hlade motion itself. and the third is due to the flow mission mounts. A similar condition exists through the engine induced by the wake. The angle of attack due to these com- mount. The dynamic response of the airframe to these loads ponents is used by the blade airfoil theory (which can range feeds back into the root of the hlade as boundary conditions. anywherefromquasi-steady aerodynamics, through panel mod- For this reason, it is not accurate to calculate loads for a fixed els, all the way up to solving the Navicr-Stokes equations) to huh and apply them on a fuselage. The shaking of the fuselage calculate thecirculation about, and airloads on the blade. These from the rotor loads in turn shakes the rotor, and that changes loads produce forces on the blades, feeding back to the blade those loads, resulting in a change to the root chord boundary structural dynamic equations. condition. (I guess the cliche, "What goes around, comes

DESORED FLOGMU GOMDWOOMS

WAKE IMPINGEMENT

CIRCULATION

STRUCTURAL

DRIVE TRAIN

TAIL ROTOR &OTHER CONTROLS

b AERODYNAMICS * WAKE IMPINGEMENT & T H R U ~ ~ E R ~ SURFACE MOTIONS

I QUASCSTEADY -

RIGID BODY MOTIONS FUSELAGE WAKE 1 -

Fig. 1. Rotarcran multi-discipline interactions (for main lifting rotor).

JANUARY 1990

tremely complex. Also, due to thz structural complexity of most airframes, a large and detailed finite element model is required. Programs l k e Dynamic Analysis Methods for Vi- brations (DAMVIBES), sponsored by NASA's Langley Re- search Center and i n v d ~ v i n ~ fou; principal helicopter manufacturers, are helping to broaden our understanding of fundamental dynamic finite element modeling principles, and improved correlation between predictions and test results for furelaee natural freouencies has been obtained. However. the - ~ ~ - ~ ~ ~

cost and time requi;ements of dynamic model development is hieh. and this will onlv he compounded as composite materials - . make up more of the airframe's structure.

The maior source of rotorcraft vibratorv excitation is, of course, the rotor system. The complication of rotational inertial forces. the flexihilitv of rotor blades, the complex aemdyn- amicsbroduced by h e rotor system and interfcience with ihe airframe, and coupling of the system with both airframe and powerplant dynamics-all these make rotor system dynamic analvsis tough. Charles Niebanck's candid paper (Ref. 3) on thc i;ttcmplei appli~,ation of hlade inays tuni"b.t#, re(lucc rolo~ vibratory loads demonvlraler the nccd for more accurale anal- ysis capsbilities. In that paper, a blade with analytically optim- ized mass distribution (optimized to give the minimum normalized magnitude of the sum'of all six huh loads components) was compared with a production blade. A comparison of S-76 flight tests results between the tuned blade and the production hlade is given in Fig. 4, and shows that the tuned hlade actually produced (higher) vibratory loads-obviously better analytical tools are required. The reason that rotor loads accuracy is such an important part of vibratory can be seen in Fig. 5 (Ref. 2). The airframe response to rotor shaft loads receives a contri- bution from manv load comoonents. and an error in one com- ponent will result in an erroneous vihration prediction; the cumulative error in all the load components is usually too large to allow accurate vibration predictions. Yet another reason for the difficulties in vihration analysis is the need for accurate higher harmonic loads. The technical complexities of predicting rotor loads is discussed further in the next section.

Although Loewy pointed out that great reductions in vihra- tion levels have hcen made over the past 40 years, see Fig. 6 ,

4P A i r f r a m e Accelerat ion 100% N,

INDICATED AIRSPEED. KNOTS

Fig. 4. Tuned blade night test results.

1989 NIKOLSKY LECTURE

COUPLED ELASTIC F-E INERTIA

PHASE, deg

Fig. 5. Calculated 3lrev lateral shear for baseline blade; V = 160 knots.

this has resulted from brute force vibration devices, rather than improved prediction methodology. Vibration analys~s is still more of an art than a science. Thus, virtually every new hel- icopter demonstrated higher than predicted vibration levels early in flight test programs. This led to extensive trial-and-error efforts during flight test to reduce vibrations to quasi-acceptable levels. It is most important to note that vibration differences in actual flight have been key factors in the evaluation of pro- totypes in every "competitive fly-off" by the military. Until a better grasp of the problem is reached, designers are forced to rely on basic factors such as proper placement of natural frequencies for the rotor blade, drive system, and fuselage, and adequate fuselagelrotor aerodynamic clearances. In the future, a combination of this plus extensive wind tunnel testing, and the use of new vihration control devices and concepts, followed by optimization of the whole system in flight test, will be needed if vibratoly levels are to be lowered to obtain a more reliable weapon system.

Rotor Loads Predictions After vibration prediction, rotor loads prediction is the high-

est technical risk area of rotorcraft development. Indeed, the inabilitv to accuratelv oredict rotor loads is a maior reason for < . the difficulty in vibration analysis. The state-of-the-art rotor loads orediction methods received much attention in the 1974 (Ref. 6) hypothetical helicopter rotor analysis comparison, where several maior airframe manufacturers compared predictions of natural frequencies of a given rotor in a vacuum: It was found that, even for this relatively routine task, there was significant disa;reemcnt bctueen thc compan~cs. William ~ou\man's re- view of rescarch in rotor loads (Kcf. 7) d<~c \ an excellent jab oicovcring the progress that has hecn madc qincc 1974, show- ing both vrcas of gaud agrucmcnt and arcds where correlation ispoor. It is concluded that even though advances have been made in rotor loads analysis, there is still a pressing need to develop our understanding of basic physics of the rotor system.

Predicting rotor loads is a more tractable problem than pre- dicting vihration levels in the fmelage for several reasons. First of all, the rotor loads are dominated by the steady component

8 C.C. CRAWFORD, IR.

0.7

JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

0.4

CABIN

VIBRATION 0.3

LEVEL (g)

0

1950 1960 1970 1980 1990 2000

YEAR

Fig. 6. Trend of helicopter vibration (relative to mil specs).

and the first oneor two harmonics, whereas the higher harmonic loads are more important for vibrations. Most of the dynamics which are not closely related to the rotor system, such as the fuselage dynamics, powerplant/drive train dynamics, and au- tomatic flight controls, are expressed in a reduced order model (if at all) for rotor loads analysis, and the state-of-the-art now is such that only rotor aerodynamics, rotordynamics, and some- times the wake, are integrated. Also, the finite elemont mod- eling of the airframe does not play as critical a role as it does in the vibration analysis.

Although analytical computer programs, such as Bell. Heli- copter Textron's C8l (Refs. 8 & 9), have made great improve- ments over the last decade, the accuracy of even the most sophisticated programs available may not be good enough for the technical challenges the future rotocraft will demand. As pointed out by Dr. Dan Schrage in his Overview of Technical Problems in Helicopter Rotor Loads Predictions Methods (Ref. lo), maximum sustained cruise speed (the flight condition most often used by industry to substantiate fatigue design) shows relatively poor correlation with flight test. See Fig. 7. The problems in rotor loads analysis can arise from errors in the rotor system modeling (rotor aerodynamics, dynamic, and wake), errors in the integration of the rotor system with the rest of the rotorcraft (hub modeling, aerodynamic interferences), or even the mathematical solution method.

The rotor aerodynamic and wake analysis face challenges such as dynamic stall, airframelrotor aerodynamic interaction, blade vortex interactions, and the pitching dynamics, to name a few. A lack of agreement on the basic aerodynamics of a rotating airfoil, such as three-dimensional stall, reflect the need for additional research in the fundamentals of rotor systems. ~ e w blade tips and airfoil sections, and highly cambered blades, will introduce problems with thecurrent analysis methodology. An example of such is the coupllng of the torsional modes with

the flap bending modes which can result in blade dynamic twist ("Mach Tuck") as discussed by Schrage (Ref. 10). Such phe- nomena indicate. that extensions of rotor loads analysis meth- odology will be necessary to support new rotor systcm configurations.

~ a i u r e to account properly for interaction of the rotor system with other facets of the rotorcraft often introduces crrors in the rotor loads. In order to simplify the airframe interaction, the hub is either assumed fixed or is modeled with some simplified analysis. If a complete finite element model were to be used,

0 50 100 150 200 250 300 350

BLADE STATION (IN)

Fig. 7. Comparison of AH-64 predicted and measured flatwise mo- ment diitributian.

JANUARY 1990 . the problem would become too computationally intense, but determination of correct blade root boundary condition model is not well understood. Yen and McLarty (Ref. I I) have shown the necessity of modeling the pylon impedance, as shown in the OH-58A loads in Fig. 8. Also, rotor configuration differ- ences such as teetering, articulated, or bearingless, makc dy- namic dissimilarity a problem. Thc dynamic behavior of thcse different configurations makes different aspects of the problem important for each. New configurations are hard to analyze due to composite materials (for which elastic coupling terms be- come important), branched load paths, or other aspects of novel design concepts.

Finally, the solution method is also a potential cause for erroneous rotor load predictions. Problems with assumptions in the "negligible" coupling of bladc modes, the problems introduced by load discontinuity at the blade root, and the lack of understanding of how many blade modes are needed for an accurate solution all give the mathematical solution additional comolcxitv. . .

The importance of accurate rotor loads prediction is seen in the strone imoact it has on rotor svstem weieht and fatigue life - - of dyna;ic c&nponcnts. If the loads are under-predicted, thc rotorcraft will havc either a restricted flight envelope or short- ened retirement lives of dynamic components; if they are over- predicted, the rotorcraft will be too heavy to perform its mission

---- HUB FIXED

- BLADES WITH NOMINAL IMPEDANCE AT HUB

CHORD

Fig. 8. Comparison of OH-58A blade bending moments with and without modeling of pylon impedance; V = 83 knots.

1989 NIKOLSKY LECTURE 9

efficiently. Errors in rotor load predictions result in the need for redesign of fhc rotor dynamic components, usually after first flight (see Design Changes Section.)

Aerodvnamic Performance The next most difficult aspect of rotorcraft prcliminary de-

sign is the aerodynamic prediction. I use that term rather than flFght to separate the impact of empty weight prc- dictions. Estimation of the powerplant installation losses is usually within reason. The aerodynamic prediction problems were addressed by Frank Harris in his 1987 historical per- s~cctive nresented at a National Suecialists Meeting of Acro- dynamic; and Aeroacoustics (Rei 12). This pap& superbly states the deficiencies in helicopter performanccestimation theory and provides penetrating insights into the following operating regimes: hover and vertical rate of climb prediction; flapping and feathering trim and rotor thrust in forward flight; and pcr- formancc with rceard to rotor ~rof i le drag. rotor induced drae. -. - and blade stall onset.

Hover performance data as shown by Harris demonstrate that the non-dimensional performance of four hclicopters designcd for the Army in the mid-1970's were essentially the same when tail rotor was excluded and adjustments were made for solidity differences. The initial prediction and the actual flight measurcment were significantly diffcrcnt. These differenccs arc repeated here in Fig. 9 for completeness. The flight data were taken during ~ r m ; competitive tests (Refs. 13; 14, 15, 16) and measured as accurately as possible at the timc. For a more recent case involving a modern technology rotor (OH58D), the predicted non-dimensional hovcr performance was 9 pcr- cent ootimistic (Ref. 171: and. for the most recent new com- .. . ~ ~ ~~~

posite;otor blade developed for the Army's use with the UH- 1H fleet. the resultine hover lift was 3.4 oercent low (Ref. 181.

Add to this short?all in predicting dower and the weight correlation for the hover case, a further limitation in predicting the vertical rate of climb. However, thc power to climb is sisnificantly less than momentum theorv would lead the de- signer to bllieve. Figure 10 from Ref. l i illustrates this point. In a time frame in which vertical rate of climb has such tactical importance to the military, and low empty weight is of such political importance to support arguments for lower acquisition costs; the absence of adequate prediction methodology is un- acceptable. This is aggravated by the fact that the flight test techniques for measuring vertical climb arc so difficult that its measurement is a treacherous situation. Notc the scatter in the test data of Fig. 10.

For assessing the forward flight condition with trim condi- tions. the aerodvnamics communitv is unablc to oredict ac- curately blade flapping and rotor thGust. In assessini how good simole momentum theorv is. Harris made com~arisons of the- orehcal predictions with-test data gathered by 'Mc~loud, Big- eers in Ref. 19. In this referenced reoon, the first-harmonic - flapping was brought to zero by adjusiments of cyclic control positions. This enabled correlation with over 1,500 data points, over many advance ratios and Mach numbers for seven different rotors. When these tests are submitted to rigorous computer analysis for correlation of theoretical and test results, then; and only then, can the aerodynamics community demonstrate the adequacy of today's theoretical methods.

By reference, Hanis also shows that frec wake analysis shows a better "theory-test correlation" for rotor thrust prediction in trim flight. Given today's computer code capabilities, Harris also advocates usine actual non-uniform induced velocitv to - compute rotor thrust trim for forward flight trim conditions.

To accuratelv comoute rotor oerformance for hieh advance - ratios, focus is required on three interrelated factors, all of which limit rotor performance:

(1) Rotor Profile Drag (2) Rotor Induced Drag (3) Retreating Blade Stall Onset

10 C.C. CRAWFORD, JR. *

YAU-M

JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

Fig. 9. UTTASIAAH hover OGE performance (predicted vs. actual).

Another theoretical problem area identified for the aerodyn- amics community is accounting for induced drag at advance ratios greater than 0.25.

The areas where helicopter aerodynamic theory and actual behavior do not correlate have been eloquently described by Harris in much greater detail than I am comfortable in dis- cussing. However, I must emphasize that the problem areas identified and discussed by Harris are consistent with my thcme of buildirra stronaer rheorelical methods to ensure that rotor- crafrperf&n aspredicted, without significant changes during the develo~ment orocess. onlv simde outimization. We can . . then be realistic in seeking greater advances in helicopter per- formance.

Major Air Vehicle Design Changes in Flight Test Before I review specific design changes, I think it appropriate

that I first review the scope of each development phase in terms of number of flight assets and actual flight hours utilized in the developmentlqualification process. Both the Army's UlTAS and AAH programs consisted of a "competitive fly-off" phase emphasizing the capability of the basic air vehicle followed by a maturity phase which fully optimized and qualified the com- plete weapons system. In the case of U'ITAS, the overall de- velopment spanned seven yean and nine months (from award of the development contract to the approval to enter full rate production). The scheme of this development is shown in Fig. I I . The basic engineering development utilizes three prototype helicopters accumulating approximately 700 contractors' flight test hours and approximately 750 government engineering and operational test hours. The government test effort included that required for the "fly-off," as well as the all important con- ventional airworthiness assessments.

This does not include a fourth company-owned prototype which supplemented the government funded development pro-

gram, hut was fabricated relatively late during the first phase. In the initial funding request to the Congress, the Army had asked for five prototype aircraft in addition to a Ground Test Vehicle (GTV). The Congress restricted the funded program to the three prototype articles mentioned plus the GTV, how- ever, the companies found it necessary to construct a fourth flight article with their own funds to ensure adequatc dcvel- opmental hardware. Although they came on the scene much later than would have been desirable, these company-owned prototypes played a key pmhlem-solving role. The GTV was utilized for over I200 hours of tie-down operation.

During the maturity phase for the winning UH-60A, an ad- ditional 600 hours of contractor flight testing was accumulated on the prototypes, plus 150 flight hours on the first two pro- duction aircraft. Government testing during this phase included another 150 flight hours utilizing the prototypes and 1100 fl~glits utilizing early production aircraft. GTV operations continued for another 700 test hours. The totals: 1450 hours-contractor, 2000 hours-government, and 1900-hours GTV.

In the case of the AAH, a "competitive fly-off" was also used, however, the scope of the developmental effort prior to the "fly-off" was significantly less. It relates to what today would he considered a Demonstration/Validation effort. The scheme of the AAH development effort is shown in Fig. 12. Phase 1, the developmental work prior to the Government Com- prehensive Test (GCT), was primarily air-vehicle oriented. Weapons operation was limited to a rough cut assessment for structural implementation. Phase 1 utilized two prototypes ac- cumulating approximately 300 flight test hours and approxi- mately 100 govemment engineering flight test hours. The government operational test effort was only 25 flight hours. A GTV was utilized for approximately 365 hours of powered tie- down operation. During Phase 2, which constitutcd Full Scale Development (FSD) for the winning AH-64, three additional

JANUARY 1990 1989 NIKOLSKY LECTURE II

MOMENTUM

'HOVER

FAIRING FOR OH-58D 0

Fig. 10. Momentum theory overprerlicts power required for vertical elinah.

prototype aircraft were manufactured. These were primarily for was particularly true in the areas of vibration, flight perfor- develovmental and intenration of mission eauivment which in- mance. and flving aualities which were all immediatclv ao- corporstcd a new ~arg ; Acquisition and designation System (TADS)/Pilot Night Vision System (PNVS), a laser semi-active guided point targct anti-armor missile (HELLFIRE) and a new 30mm arcd wcapon. The contractor's Phase 2 flight test effort encompassed an additional 980 hours of testing utili7,ing the two original prototypes and 715 hours of tcsting using the three new prototypes. Thc total government flight testing, enginccr- ing and operational, on all prototypes during Phasc 2 was 880 hours. This included 200 hours for Operational Test 11 (OT 11) which was the key prercquisite for a production decision. Con- tinued GTV operation of over 1050 hours contributed signifi- cantly to the quality and reliability of the rotor-drivc system. The loss of one test aircraft due to a mid-air collision, and early Phase 2 budget instability, clearly impacted the overall schedule. The final airworthiness qualification, using early pro- duction aircraft, included 170 contractor flight test hours and 140 governmcnt engineering flight test hours. The period be- tween the initial contract award and the decision to enter pro- duction encompassed eight years and six months. The flight totals are: 2175 hours-contractor, 1145 hours-government, with 1415 hours-GTV.

As mosi will recall, the design changes required after first flight were significant. In some cases, changes needed to meet Army requiretncnts could not he incorporated in the relatively short time available to prepare for the fly-off. Therefore, some of the required changes were simply "proposed" for the pro- duction configuration with limitcd analyticallwind tunnel suh- stantiation orovidcd as Dart of the nroduction orooosal. It is . . interecting to notc that those de\igns 11131 ~~~corpor:itetI improve- mcnts hcforc the tlv-oft !rcrc 1011kcd uoon ;I.; hav~ne s~ell~flc;tnt lower risk of ultimately meeting minkurn ~ r m ~ nee&. This

. - . parent to Army evaluation pilots. The YT700-GE-700 e&ink, common to all designs, performed exceptionally well during the entire "fly-off" process due to its relative maturity. This had resulted from its exhaustive developmental tcsting prior to installation in an aircraft.

Thc design changes to he discussed herein have been limitcd to those areas whcre the change may have been avoided if a more encompassing preliminary design phase, bettcr analytical methodology and substantially more wind tunnel testing and simulator activity, had been available. Adequate analytical tools were simply not available; nor did time permit the othcr activ- ities. The UTTAS schedules were naively belicvcd to be re- alistic but the more stringent AAH development schedule fclt the urgent pressure of the need for a more capable attack air- crafi; following the cancellation of the AH-56, Cheyenne. While affordability was the most visible reason to thc public for thc AH-56 cancellation, technical problems exasperated by the lack of adequate analytical methodology were also a major contrib- utor.

For thc record, the most significant design changes together with the problems uncovered in flight tests that dictated the change are outlined in Figs. 13, 14, 15, and 16. It is interesting to note the types of changes that are common. These arc:

(a) Increases in thc main rotor height above thc fuselage- to minimize vibration and aefodynamic interference between the rotor and fuselage; and otherwise enhance safety through increasing the clearance over the canopy, the tail boom, and ground personnel. (Fig. 17 illustrates the dramatic differ- ences for the UTTAS vehicles.)

(h) Revised empcnnage configurations-a widc variety of horizontal tail appendices; some fixed, some movable, all

12 C.C. CRAWFORD, JR. JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

OCT 79

V FIRST FLT OCT 74

BASIC ENGR OEVEL (BED) I

MATURITY PHASE

PROOUCTION SOURCE SELECTION AlRFHAME FIRST FLT NOV 7 4 ~

BASIC ENGR OEVEL (BED) T F I R S T FLT OCT 78

YUH-6 IA I- PEP PRODUCTION

FOTE m GOVT

PVT PTU

A I OSARC . . ARMY I DEFENSE SYSTEM AC(1UISITION REVIEW COUNCIL GCT . . . . . . . GOVERNMENT COMPETITIVE TEST PEP . . . . . . . PROOUClBlLlTY ENGINEERING & PLANNING FDTE . . . . . . FORCE DEVEL TEST & EXPERIMENTATION PVT . . . . . . . PRODUCTION VALIDATION TESTS

Fig. 11. UTTAS program structure.

relocated; and revised vertical fin areas to enhance tail rotor moment on the advancing blade in high speed flight which thrust and reduce blockage. produces high loads and increased vibration.) Only the YUH-

(c) Main rotor blade configuration changes in the form 61 did not rcquire such changes, in that another type of of either an airfoil section change, a blade planform change, analytical breakdown occumd-the high speed lift capa- or both-to improve high speed maneuverability and elim- bility of the rotor had been underestimated. inate "Mach tuck." (This is a very large nose-down pitching (d) Tail rotor configurations-either physically changed,

A 1 DSARC I & I! MAY 71 A A 1 DSARC Ill NOV 76 A ASARC llla

80 81 76 Fiscal Year

AUG 84 OCT 84 j FIRST ARTICLE SEP ~3~ TT YEB 85 PRE.PRODUCTION TEST

mB i PVT CONTR GOVJ !

74

Calendar Year

A I DSARC . . ARMY I DEFENSE SYSTEM ACUUlSlTlON REVIEW COUNCIL GCT . . . . . . . GOVERNMENT COMPETITIVE TEST OT . . . . . . . . OPERATIONAL TESTING PVT . . . . . . . PRODUCTION VALlOATlON TESTS TAOS . . . . . . TARGET ACUUlSlTlON & DESIGNATION SYSTEM

Fig. 12. A A H program slructure.

7T 75 72 73

73 74 75 76 77 78 79

77

DSARC I SEP 72 ' II DEC t6ASAdC Ill NbV 8 l t VDSARC Ill MAR 82' CONTRACT AWAROYUN

80

,

78

PHASE I YAH-63 ENGR DEVEL I

(COMPETITIVE)YAH.6$ * T C O N T ~ T AWARD DEC 76 JUN 81 OT II

TADS SELECTION APR 8 0 1 W PHASE II FULL SCALE AUG 81 ENGR DEVEL I

I

CONTRACT AWARD DEC 81VFIRST FLT DEC V

PRODUCTION LONG LEAD FEB 81V -[,-I

79

81 82 83 84 85 89

JANUARY 1990 * I989 NIKOLSKY LECTURE 13

their control authority changed, or their operating environ- ment changed for the purposes of improving sideward flight capability to ease pilot workload when hovering in gusty wind conditions.

(e) Numerous fairings were added and components re- aligned for the purpose of reducing external drag. A more cost effective developmental program would extcnsively ad- dress this subject during eady wind tunncl testing. 1 will say more about the exhaustive wind tunnel test performed on Boeing's commercial 757 and 767 later in the paper.

Some optimization in tailoring of flight control systems will always be required during the flight test phase. Future ap- proaches will minimize such flight testing through theextensive use of simulators durine thc ~reliminarv desien ohase. How- .. . ., . ever, lhcse s~rllulat~oas will he no inure accuratc than thc an- alvlical tools which nroducc thc st:~bilitv derivatives :~nd control analogs elementary to their operation. Vibration reduction will always be a major thrust of rotary wing flight tcst; howcver, better analytical methods would make possible a more rapid understanding of the surprise sources and vibratory levels en- countered in early flight tests. Such an understanding leads to quicker and more cost-effective corrections.

the lack of adequate wind tunnel testing during the preliminary design phase. Though more complex than their fixed wing counterparts, very little effort was spent on wind tunnel testing. The testing done both pre- and post-first flight is shown below. In some cases, more wind tunnel activities occurred after the aircraft was in flight tcst than were performed during the initial design phase.

PRE-FIRST FLIGHT Airloil Sections Airvehicle Full Scale TIR Subtala1

POST-FIRST PLlGHl Airroil Sectialls Airvehicle Full Scale TIR Sublolal

PRODUCTION GRANDTOTAL

- - 1470 1615 160 97

1630 1712

455 NIA 2715 4045

'AH-63 Y A H ~ 6 4

84 514 293 432 50 -

425 946

96 -

321 886 24 -

45 1 886

NIA 623 876 1832

Wind Tunnel Testing Not withstanding the lack of maturity of the analytical tech-

niaues used in the desien of comnetitive rotorcraft for the Ar- m i ' s AAHIUTTAS p&grams p;eviously discussed, another reason for the large number of configuration changes that were With the exception of full scale isolated tail rotor tests, the required of each design after it entered flight test was due to wind tunnel testing was accomplished utilizing small test mod-

TOPUG FRT TEST PROBLEM DESUGN CHANGE BASIC AIRFOIL MAX LOAD FACTOR LOW ADDED DROOP SNOOT

.MAIN ROTOR BLADE OHOVER CAPABILITY LOW TIP CAP EXTENDED 4 INSBLADE

ROTOR HEIGHT • HIGH VIBRATION ROTOR MAST INCREASED 15 INS HlGH ROTOR BLADE LOADS

eSTABILATOR CONFIG *AIRCRAFT ATTITUDES IN LOW FIXED STABILIZER RAISED LOW SPEED FLT REGIMES MADE MOVABLE OVER A 43 DEG

HIGH SHAFT BENDING RANGE THRU AUTO CONTROLS

VERTICAL FIN AREA LARGE TRIM CHANGE REDUCED FIN AREA WITH CUT-OUT UNSTABLE SIDEWARD FLT

.TAIL ROTOR MOUNT HIGH FLT LOADS ROTOR SHAFT, SUPPORT STRUCT,

& CONTROL & CONTROLS REDESIGNED FOR INCREASED STRENGTH

LANDING GEAR HIGH DRAG CHANGED INCIDENCE ANGLE TO

SPONSON ALIGN WITH FUS FLOW TAIL ROTOR GEARBOX UNDETECTED OVERTEMP REPLACED GREASE LUB WITH OIL

CONTROL SYSTEM POOR CONTROL HARMONY LAT SENSITIVITY INCREASED AUTHORITY YAW/ROLL MECHANISM INCORPOR

.MAIN ROTOR SERVOS HIGH FLT LOADS REDESIGNED FOR INC STRENGTH MAIN ROTOR DAMPERS HIT DAMPER STOPS IN HlGH MOVED INBOARD DAMPER

SPEED AUTOROTATION OUTBOARD BY 3.5 INS

.MAIN TIRES MECHANICAL GROUND INCREASED SIZE & PRESSURE TO INSTABILITY INCREASE STABILITY MARGINS

AIRSPEED SYSTEM EXCESSIVE ERROR FLEW OVER 100 FLTS TO OPTIMIZE

.AFT NACELLE CONFIG HIGH TAIL VIBRATION REDESIGNED TO IMPROVE AIRFLOW & BUFFET PATH

.STABILITY AUG SYS VARIABLE FLT CONTROL SYS CHANGED FROM FLUIDIC TO CHARS W/ EXTREME TEMPS ELECTRICAL/MECHANICAL

Fig. 13. YUH-60A required design changes.

14 C.C. CRAWFORD, JR. JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

TOPUG FLT TEST PROBLEM DESIGN CHANGE

ROTOR AEROELASTIC STAB EVAL WITH/WITHOUT DAMPERS LAG DAMPERS REMOVED (INITIALLY REQUIRED BY ARMY)

TAIL ROTOR SIZE REDUCED SIDEWARD FLT INCREASED DIA FROM PROPOSAL, MANEUVER CAPABILITY FURTHER INCREASE FOR PROD

ROTOR WAKE INTERFERENCE YAW "KICKS" IN SOME ADDED BEANIE ON TAIL ROTOR FLT COND LARGER FAIRING FOR PROD

LATERAL CONTROL TRAVEL RAN OUT OF CONTROL DURING DOUBLED LAT SWASHPLATE LOW SPEED MANEUVERING TRAVEL

FLT CONTROLS RATE LIMITED TOO SLOW FOR MANEUVERING DOUBLED CONTROL ACTUATOR MAXIMUM RATE

VERTICAL FIN SIZE EXCESSIVE T/R BLOCKAGE IN REDUCED FIN CHORD SIDEWARD FLT

VIBRATION ABSORBERS EXCESSIVE VIBRATION ADDITIONAL FUS ABSORBERS EXTENSIVE TESTING OF ABSORBER INSTALLATIONS

PlTOT STATIC SYSTEM EXCESSIVE ERROR MANY LOCATIONS TESTED

HORIZONTAL TAIL SIZE MOVABLE TAIL ADDED COMPLEXITY

ROTOR HEIGHT

SPAN COMPROMISED AIR TRANS

FLEW FIXED HALF SlZE TAlL PROPOSED SAME FOR PROD

EXCESSIVE VIBRATION ROTOR MAST INCREASED 24 INS FOR PROD

Fig. 14. YUHdl required design changes.

TOPIG FLT TEST PROBLEM PROPOSED DESIGN CHANGE

ROTOR DIAMETER WEIGHT INCREASE INCREASE DIAMETER BY 6 INS AIRFOIL SECTION WADV BLADE TUCK REDUCED THICKNESS OF OUTBOARD 25% ROTOR HEIGHT TAIL BOOM CLEARANCE INCREASE MAST BY 18 INS

FOR SLOPE LANDING ADDED 1.2 DEG FWD TILT MAST ANGLE .LOW CRUISE SPEED ADDED FAIRINGS FOR BLADE GRIP & HUB MlSC FAIRINGS *LOW CRUISE SPEED SPRING TORQUE TUBE BLADE FLAPPING .INCREASE MANEUVER .INCREASED ALLOWABLE FLAPPING

ENVELOPE ELASTOMERIC FLAPPING STOP FOR ADDED CONTROL POWER ABOVE 4 DEG FLAPPING

BLADE TIP SHAPE .LOW CRUISE SPEED PLANFORM MODIFIED

T/R BLADE SHAPE REDUCE TAIL POWER SWEPT TIP ADDED T/R THRUST AVAIL *SIDEWARD FLIGHT *INCREASED COLLECTIVE PITCH AVAILABLE

CAPABILITY AND ADDED COUNTER WEIGHTS T/R HUB LOW CRUISE SPEED ADDED HUB FAIRING WING CHORD REDUCE FUS DOWNLOAD 6 INCH CHORD EXT TO CREATE A FLAP

IN FWD FLT FOR ADDED LIFT HORIZONTAL TAIL POOR FLYING UPPER HORIZONTAL CHANGED TO PROVIDE

QUALITIES A 25% CHORD CONTROLLABLE ELEVATOR MlSC FAIRINGS LOW CRUISE SPEED ADDED WING FILLETS

MODIFIED MAIN LANDING GEAR FARllNGS REDESIGNED ENGINE COWL TO REDUCE

FRONTAL AREA & FLOW SEPARATION

Fig. 15. YAH-63 required design changes (all items only proposed for production).

JANUARY 1990 1989 NIKOLSKY LECTURE 15 . FRT TEST PROBLEM DESUGN CHANGE

.ROTOR HEIGHT .(I) CANOPY INTERFERENCE EXTENDED MAST 10 INS m(2J CONTROL POWER FOR ADDITIONAL 6 INS EXTENSION

DROOP STOPS

. . SLOPE LANDING

B L A D E CONTACT DURING . OPEN STOP ONE DEG W/ FULL RAISED ROTOR PUSH-OVER

LEAD-LAG DAMPERS H I G H CHORDWISE LOADS . REDUCED DAMPER STIFFNESS 20% .MAIN ROTOR BLADES UNACCEPTABLY HIGH SPEED SWEPT TIP TO REDUCE NOSE-DOWN

VIE CONTROL LOADS PITCHING MOMENT EMPENNAGE CONFIG INADEQUATE LOW SPEED . FINAL CONFIG ADAPTED AUTOMATIC MOVING

CONTROL STABILATOR MOUNTED LOW ON TAIL BOOM UNUSUAL ATTITUDE IN T-TAIL TRIED FOR A SIMPLE FIXED APPROACH

LOW SPEED TAIL ROTOR SIZE .INADEQUATE LOW SPEED . INCREASED DIAMETER FROM 100 TO 110 INS

I CnNTRnl . VERTICAL FIN SIZE INTERFERENCE IN RIGHT SIDEWARD FLT

REDUCED FIN AREA

.MOVABLE WING FLAPS EFFECT NOT APPARENT DIGITAL AUTOMATIC PITCH-UP IN HIGH

STABILIZATION EOPT SPEED FLT .XMSN SUPPORT H I G H FLT LOADS

STRUCTURE POWER CONTROLS & HIGH FLT LOADS

ACTUATORS TAIL BOOM TUNING EXCESSIVE VIBRATION IN

SOME CONDITIONS . TAIL ROTOR SENSITIVITY TO FLT VIES ACTUATORS

a AREA WEAPON FIRING RATE . EXCITED FUS VIBRATIONS ROTOR SPEED CONTROL . ROTOR DROOP W/

COLLECTIVE PULL

FIXED FOR SIMPLIFICATION/COST REDUCTION DOUBLED PITCH AXIS AUTHORITY

(10% TO 20%) REDESIGN FOR INCREASED STRENGTH & TO

CHANGENATURAL FREQUENCY REDESIGN FOR INCREASED STRENGTH

INCORPORATED SLOT TO REDUCE NATURAL FREQUENCY

RELOCATED FROM UNDER COCKPIT AREA TO TAIL ROTOR VICINITY

REDUCED FROM 875 TO 625 SHOTS/MIN ~ ~

ADDED COLLECTIVE RATE SIGNAL TO COLLECTIVE DISPLACEMENT AS ENGINE CONTROL OUTPUT

Fig. 16. YAH-64Arequired design changes.

els, quarter scale beingthe largest. Their efforts are summarized ing, particularly when rotating components are utilized. The below: industry therefore has not relied to a great degree on such testing

comlnon ~~~l objectives YUH-MI Y U R ~ I YAH-63 Y A H . ~ ~ to design acceptable configurations. Some historical compar- ~ a r i c ~ i r Vehicle Drag 1110 115 14 117 isons of the wind tunnel occupancy test hours for rotary wing ~ e r i p n Acro. Airloads. aircraft with recent fighter aircraft developments and the long

Enlpnnage 114 115 116 117 history of commercial jet airplane developments are shown in lralated Tail Rotor Fig. 19. The only exception to this trend has been the recent

~ h r u s t &Stability Full FUII FUII None development of the V-22 (not shown on the figure). Its design ( ~ o t e : AII cantractors perrormed drag reduction tests a r m Is1 nigh9 nrocess included a~oroximatelv 10,000 test hours of mdnv tvws

In addition, some unique tests were performed. For the record, they are listed here.

YUH-60: Rotor Tip Geometry, Maimail Rotor Interaction, and Impact of the IR Suppressor

YUH-61 Model Rotor for ForcesIMoments (10 ft. dia.), Tail Rotor AeroelasticStahility (multi-configs), 115 Scale Ballistically Damaged Rotor, Engine Exhaust Re- Ingestion, VibrationIShuffle w1Raised Rotor

i YAH-63 QuadricycleLanding Gear, Airfoil Tests (both pre-

& post-flight), Aft Nacelle Area, Gun Fairings, and Main Rotor Hub

YAH-64 Canopy Configurations, WingIFlap & Weapons Pylon, Empennage Environment, and Impact of thc IR Suppressor

bf different wind'iunnels. 1n.thc 'case of the Boeing 737ij67 designs, the wind tunnel testing accomplished prior to the de- cision to proceed with the development of the aircraft approx- imated 10,000 hours each, which is more than double that accomplished in any previous complete helicopter program. The bottom line appears to me to be: Tlieprobability of reduci~lg the risk of rhe full scale developmerlt of a new rororcrafr car! be sig~zificanrly reduced through more fhoro~rgh nind runrzel tesritzg in rlze preliminary design plzase so that the probability of design changes after first flight is minimized. This would follow the lead of the fixed wing industry.

The future is already a little brighter. Some wind tunnel testing is being conducted during LHX DEMIVAL to optimize and confirm aerodynamic design parameters. LHX wind tunnel testine will utilize the aoorouriate scale and full-size wind tun- nel midels to inve~ti~ate'individual performancelaerodynamicl

The use of the wind tunnel did play an imponant role in mod- aemelastic interactions, Amy RFP guidance indicated that these ification to the engine exhaust system and empennage config- models are expected to include: uratinn of the YAH-64. The differences can be seen bv ~

comparing the photographs of wind tunnel test models pre- (1) Small scale (I0 percent to 20 percent) airframe aero- sented in Fig. 18. dynamic models

The testing of rotorcraft models in a wind tunnel is signif- (2) Medium-scale (30 percent to 40 percent) rotorlair- icantly more complicated than fixed wing aircraft model test- frame interaction models

16 C.C. CRAWFORD, JR. JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

UTTAS RAISED ROTORS YUH-60 YUH-61

ORIGINAL ROTOR / EMPENNAGE COMPETITION TEST CONFIG

FINAL ROTOR / EMPENNAGE PROPOSED FOR PRODUCTION

Fig. 17. UTTAS raised rotors.

TYPICAL YA H-64 WIND TUNNEL TESTING

ORIGINAL ENGINE EXHAUST IR SUPPRESSOR

FUS WITH T TAIL FINAL EMPENNAGE CONFIG Fig. 18. Typical YAH-64 Hind tunncl testing.

JANUARY 1990 1989 NIKOLSKY LECTURE

ROTORWING FIXED WlNG MILITARY

0 FIXED WlNG COMMERCIAL J

Fig. 19. Comparison

(3) Isolated large-scale main rotor models (4) Large-scale anti-torque models

LHX wind tunnel testing will measure airframe drag, sta- bility parameters, and determine parasite power requirements. Wind tunnel testine will also include determination of main - and tail nltor elfic~enc~e\, power rcqutrcnlcnts. ~n;~xi~liuln thmst canahilltv. ~ r o ~ u l s i v c force. and hladc l(rads~bcndincl moments. ... . - Other areas to be evaluated are propulsion system losses, rotor system stability boundaries aerodynamic stability, and aero- dynamic interactions. LHX wind tunnel testing has been for- mulated to aid the design process, and substantially improve design aerodynamic loading problems that have plagued recent helicopter development programs. There may be a qucstion regarding adequate time to evaluate model modifications for optimization purposes, however, because of the sensitive com- petitive nature of these tests, no information on their exact scope has been made public.

I Weight Estimating Realism The history of the aircraft industry has been plagued with

overly optimistic weight predictions and their adverse impact . . . . on cost, schcdulc, and pcrti~rmancc. Rotorcraft are no excep- tion to fhls problcm. It is equally difiercnt tor hclth [he ~ndustrlal . . supplier a id the military customer to handle weight estimates

I of new designs that are based totally on parametric analysis.

I This tends to stifle cost-effective weight saving features. On I the other hand, when the initial estimate is based on a bottoms-

up analysis, its preparation involves an enormous number of

i man-hours and its accuracy is naturally subject to great opti- mism.

The optimism in the weight estimating process is driven by

I two almost unavoidable reasons. The first is the desire to win. Notwithstanding the desired preservation of the company, both the prestige of being a military supplier and the patriotism aspects of providing superior equipment to the DoD, fuel this desire. The second reason is an almost blinding faith in new technology. Most new technology efforts, explored eitherthrough the government's Tech Base programs or through the Inde-

of wind tunnel test times.

pendent Research and Developmental activities of industry, have either weight savings or weight avoidance as one major goal. These goals arc often achieved on an individual project basis however, collectively they never seem to sum up directly. I guess I'm saying that the impact of the growth factor concept when adding weight to a design is not as straightforward as when taking weight out. Also, if it is known that the military customer has some target weight objective, no matter how optimistic this objective may be, it is extremely difficult for either the supplier or the buyer to admit that this goal is not in hand.

The actual empty weight of any new aircraft is influenced by several other important factors. Changes in customers' re- quirements during the development process are quite common. Numerous design charges to insure a viable air item usually add significant amounts of weight. Very few "fixes" for re- liability, maintainability, aerodynamiclthermodynamics, struc- tural or other mechanical problems result in weight decreases. The cost implications of proposed weight saving features in the original design come to the surface as unfavorable. Tech- niques like the chemical milling of aircraft bulkheads (OH-6A) to save a few ounces are abandoned in the real world, because the expense of chemical milling does not make it cost-effective. During the preliminary design phase, many trade-offs are ac- cepted which in the real world of flying hardware are quickly found to be "flaky" and thus abandoned. This is particularly true when original trade-off, such as the lack of a subsystem redundancy or reduced design criteria are found to impact safety or to adversely affect maintainability. Finally, the biggest de- terrent to achieving a target empty weight is an unrealistic developmental schedule which does not allow time for redesign for the purpose of weight reduction. When low weight is of paramount importance, it must be recognized that such designs are an iterative and time consuming process, requiring schedule allowances.

I would like to substantiate these arguments by presenting comparisons of the weight history of the U P A S and AAH designs. The Army Helicopter Improvement Program (AHIP), which has a military designation OH-58D, is also included. Figure 20 presents the percentage change between the initial

18 C.C. CRAWFORD. JR. - JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

DEVELOP SPEC ACTUAL @ GOVT EVAL FIRST PRODUCTION ACTUAL

Fig. 20. Realities of weight

developmental proposal offered by each contractor for their aircraft relative to the:

a) spccification com~nit~ncnts ncg<~t~atcd for dcvclop~nc~it h) actual wclghts :tt the time 111 l l~c official guvcrnment

evaluation c) actual weights of the first production aircraft

Since the UH-61 and the AH-63 were not fly-off winners, producl~on wcight data ncvcr existcd. 'fhe reason for u s ~ l ~ g the ~nitial proposal cstimatc of thc prin~c manufdclurcr a> the COIII -

mon denominator for the percentages shown is quite important. Because these initial weight estimates were not generally known to the public, their values are included in Fig. 20. Notwith- standing the government's desire to acquire the lightest weight weapon system practical to meet mission requirements, the government's formal evaluation process contributes signifi- cantly to the improvement of realism in empty weight prcdic- lions. Government analyses of the group weight statement in great detail, looking for omitted items and unrealistic projec- tions, results in an increase in the full specification weight of about 4 to 5 percent, through the process of negotiating the development contract. Thcsc more realistic weight estimates, in conjunction with less optimism in the aerodynamic predic- tions of lift and drag and a more realistic assessment of pow- erplant installation losses, sometimes result in an increase in rotor diametcr before the detail design process begins. While this enlarged rotor is always heavier, it still results in a higher performing rotorcraft. A rotor diameter increase during the negotiation of the developmental contract with the Arn~y ac- tually occurred for both the YUH-60 and YAH-64. Ironically enough, they were the "fly-off" winners, having the best flight performance.

Figure 21 is the identical presentation of data in terms of the percentage change from the initial proposal for primary mission gross weights. These percentages also include the adverse im- pact of higher-than-estimated fuel consumptions, where appro- priate. It can be seen that the overall optimism in estimates of

prediction (empty weight).

the primary mission gross wcight ranges from 8 to 16 percent. Reduction in the mission weight of the production aircraft only occurred for the UH-60. This resulted from a tremendous em- phasis on weight reduction between the prototype and initial production aircraft. To enhance their probability of winning the production contract, Sikorsky proposed to eliminate ap- proximately 260 pounds of weight from the prototype. Thcy actually achieved an additional 536 pound weight savings, such that early production aircraft weighed approximately 800 pounds lcss than their development specification commitment. This gallant effort did not last. The nominal weight increase of each production aircraft was almost I pound per ship, resulting in the 980th Army aircraft being 866 pounds heavicr than the first production article. This increased weight is after the deletion of the aft cabin absorber, at 67 pounds, which was in the early production aircraft. Updates to the production specification for approved engineering change proposals result in reccnt aircraft being 35 pounds below the spccification empty weight (I 1,288 pounds). There are many viable reasons for this weight growth pattern; some customer-driven, some for production reasons, some to enhance safety, and many to improve readiness. The compelling point is that weiglzt growth is arl inevitable fact of life.

Power Margins The most practical way to avoid adversities in mission ca-

pability (as a result of overly optimistic empty weight estimates for paper proposals andlor aerodynamic efficiencies that are never quite achieved) is to design for high power margins from the outset. This approach is facilitated by the fact that engine power growth is also a fact of llfe. The succcssful growth of engines is well known history; and, when accomplished in a preplanncd manner, is most cost effective. At retirement from military inventories, rotorcraft engines have experienced a nominal growth of at least 30 percent from their original pre-

JANUARY I990 1989 NlKOLSKY LECTURE 19

INITIAL P R O B l L 15500 14600 14560 1.2'6900 3.947 EST!MTE

DEVELOP SPEC ACTUAL @ GOVT EVAL I7 FIRST PRODUCTION ACTUAL

Fig. 21. Realities of weight prediction (primary mission weight),

liminary flight ratings. A specific requirement of the Army's latest T800 engine devclopmcnt planned after LHX was a growth plan for 50 percent over a 10-year period. Both competing teams offer a 10 to 20 percent growth at the outset with vcry minimal change from their preliminary flight rating configu- rations. Why not use this capability?

An equal success story has been the relatively simple growth of rotorcraft transmission ratings. The term "simple" is used here in the context of minimal hardware changcs. The slope of a parametric transmission weight versus horsepowcr curve will predict a weight increase for a given horsepower increase which is at least two timcs the actual weight increase needed to qualify a successful fielded transmission at thc same higher rating. This should make the use of the planned growth of the T800 engine a simple matter if the initial transmission design criteria are not compromiscd.

T o quantify the need for a high powcr margin in the initial design, the optimizations in the aerodynamic and wcight pre- dictions will be reviewed for the five designs previously dis- cussed in terms of needed horsepower. Figure 22 illustrates an average power deficicncy of 5 pcrcent due to optimistic hover

i power required estimates and a 7-15 perccnt powcr deficiency

I due to unrealistic weight predictions. These factors arc com-

I bined in Fig. 23. Minimum and maximum boundaries for useful power margins (excluding vertical rate of climb considerations) are established in Fig. 24. A 15-20 percent power margin is a recommended "insurance policy" to assurc that critical mil- itary flight performance characteristics are achieved without compromising reliability, maintainability, or fleet readiness by scrimping on design fcaturcs (solely to minimize empty weight) which enhance the usefulness of the aircraft.

Improvements in Analytical Process Rotorcraft analysis codes have continuously improved since

the dcvelopment of the UTTAS and AAH. Currcnt Rotorcraft

Analysis Codes include C8l (Refs. 8 & 9), C60 (Rcf. 20), REXOR (Ref. 21). DART (Ref. 22), Normal Modes (Ref. 23). RDYNE (Ref. 24) and RACAP (Ref. 25). These codes have addressed a broad range of user needs in the various technical disciplincs.

However, limitations of current analysis codes include the load of internal interdisciplinary mathematical model consis- tcncy, inadequate modeling flexibility, theoretical and algo- rithmic documentation, and havc not bcen cxtensivcly validatcd. For example, these first generation programs have exhibited large differences in predicting helicopter loads. These differ- ences are largely a result of the lack of understanding of the uniquc aeromcchanical complcxitics of rotorcraft. Validation thrusts include data base initiatives for code correlation such as the DNW Wind Tunnel Tests of the Boeing 360, MDHC HARP, and UH-60A rotors, as well as the NASA-Army Black Hawk Flight Test Measurement Program.

lnitiativcs in providing comprehensive rotorcraft analysis codes have resulted in programs such as DYSCO (Ref. 26). CAM- RAD (Ref. 27 & 28). GRASP (Ref. 29), COPTER (Ref. 30), and 2GCHAS (Ref. 31). GRASP and CAMRAD have im- proved the state-of-the-art in nonlinear beam dynamic modcling and aerodynamic modcling for rotor air load prediction for use in an integrated comprehensive analysis.

However, DYSCO and 2GCHAS provide the most potential in providing a flexible software platform for the evaluation of new technology elements. An exccutive control system pro- vides standard module interactions, run time execution control, a user friendly interface, configuration modeling flexibility and a standard softwarc architecture. This provides a sound basis for future enhancements.

While the DYSCO and 2GCHAS codes provide a basis for future enhancements, they currently do not provide quantum leaps in module technology. Additional theoretical research is required in the fundamental disciplincs to achievc the ultimate goal. Only a better understanding of the structural and fluid

20 C.C. CRAWFORD, JR.

25% T

JOURNAL OF THE AMERICAN HELICOPTER SOCIETY

5%

0% I I:..

20% --

15% .-

10% -- I I:., Fig. 22. Power margin deficiency. a) Due to missing hover power (tap) b) due to weight growth (bottom)

Fig. 23. Power margin deficiency (due to missing hover power and weight growth).

JANUARY 1990 1989 NIKOLSKY LECTURE 21 . 25%

P

0

W 20%

E D

R E F 15%

I M

A 10% I

R T

G 5%

I

N 0%

DUE TO HOVER DUE TO WEIGHT COMBINED EFFECT

PREDICTION GROWTH

Fig. 24. Power margin deficiency (excludes VROC considerations).

mechanical phenomena, their associated engineering represen- tations, and'interactions in an integrated ~&~rehe i s i<e anal- ysis will lead to a flexible tool which may be used throughout ihe phases of rotorcraft design, analysis,-and test

Conclusions The lack of realistic and detailed rotorcraft analytical tools

has resulted in many configuration changes being required after the rotorcraft has entered flight test, in order to have a viable weapon system. These analyses short-falls are most prevalent in the areas of estimates for rotor vibration loads, and flight control system loads, overall vibratory levels, as well as.aero- dvnamic caoabilitv. The multi-discioline asoects of rotorcraft dynamics significantly complicates the construction of appro- oriate analvtical tools. The inabilitv of these tools to facilitate

design wilich would fly adequateiy "off the drawing board" was illustrated in a development of the Army's competing UTTASIAAH designs.

Extensive wind tunnel testina, as a predesian effort, is needed to support future rotorcraft de&lopmknts. while time consum- ing in the uo-front phase. this testing should minimize changes aitur ftrst fl~ght. Such changes hrr. extret~tely costly and oltcn rc~ult in lcss than o ~ t i n i ~ m f~xcs. Malor IixcJ-xinr wind ti~nncl efforts have averaged 12,000 to 20;000 hours oiwind tunnel occupancy time for fighters and 20,000 to 30,000 hours of testing for commercial jet transports. The UTTASIAAH design test efforts ranged from just under 1,000 to slightly over 4,000 hours of tunnel occupancy time.

History suggests that the optimism in empty weight and aerodynamic prediction for initial rotorcraft designs warrants a design power margin of 15 to 20 percent. Optimistic weight estimates are driven both by the dcsirc to win and by a blinding faith in new technology. Fixes for reliability, maintainability, aerodynamic, and structural or technical problems almost al- ways rcsult in weight increases. Weight growth is also often oromoted bv chanees in customer retluirements. The validitv , . - of some initial trade-offs, designed to savc weight, are found not to be in the best interest of safety and readiness. Optimistic

aerodynamic predictions are energized by the same trends. Ro- torcraft drive svstems are oarticularlv accommodatine in adaot- ing dcsign power rnargtns to dctuitl in-scrvicc power utiliz~tiun.

A rcvitalirdtic~n of the work < t i huildine and validatinc im- - - proved rotorcraft methods is much necdcd. Within the govern- ment. this effort has been spearheaded bv the develooment of the Amy's Second Genekition Comp;ehensive ~ c l i c o ~ t c r Analvsis Svstem (2GCHAS). Although this svstem provides an iniegratid framework forthc multi->isciplined natire of the rotorcraft analvsis, development of the individual technical modules to be hsed in ~ G C H A S requires extensive researchin their own right. The completion of this comprehensive ana- lytical system has been plagued by the scheduledelays normally associated with funding problems due to the lack of under- standing of its importance within the highest level of the mil- itary. In addition, the DoD, its military services, and the industry have all been hampered by their inability to convince congres- sional budget personnel that the technical complexities and risk of developing new rotorcraft are substantially greater than for their fixed-winged counterparts.

Since-there is no "quick fix" to the shortcomings of the rotorcraft analytical tools, the surest way to select the best available weapon svstem is through the "comoetitive flv-off" process. The .&$s Light observation ~ e l i c o ~ t e r ~ ro&m of the 1960's resulted in the world's most advanced small heli- copter. The UTTASIAAH competitive fly-offs ultimately re- sulted in successful weapon svstems. Althouah all specification . . - requirements are never met, with a fly-off the impact of trade- offs can be evaluated through the performance of actual mis- sions. The technical efforts of weight, flight performance, structural integrity, and the effectiveness of systems integration can be based on measurements not estimates. After all, com- petition fly-offs are the ultimate challenge to the industry.

Acknowledgments This lecture could not have been prepared without the as-

sistance of many members of the Society, who are employed by industry. It contains much historical information that is not