A feasibility study for advanced technology integration for general aviation

CHAPTER 1 SUMMARY A study directed toward the identification and evaluation of applicable advanced technologie~ for general aviation was per- form~d. An extensive data base was generated through visits to 31 general aviation manufacturers and 3 NASA research centers as well as through an exhaustive literature search. An evaluation tech- nique was developed which allowed candidate technologies to be ranked according to potential benefit. Finally, design studies were performed for a 6-passenger personal/business airplane and a 19-passenger commuter airplane. The General Aviation Synthesis Program (GASP) was utilized during the design studies t0r propul- sion system and vehicle sizing as well as mission periormance analysis.

In assembling the data base, extensive notes which were ac- qui red from the visits were eciited and are included in the report as an appendix. ~6 of the 137 technologies initially identified were evaluated and are discussed se~arately in the report.

The results of the technology evaluation indicated that propulsion, aerodyn3mic, and composite technologies are extremely

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attractive to general aviation. ~len these technologies were in- r; corporated into the design synthesis of the two airplanes, higher I · ' wing loadings and smaller ahplanes resulted. Fuel savings of 50% for the 6-passenger airplane and 40% for the commuter were

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realized.

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CHAPTER 1.

INTRODUCTION General aviation represents 96% of the civilian pilot force flying 99% of the civil aircraft and 84% of the total flight hours in the United States. During 1979 alone, factory billings for aircraft shipments amounted to $2.2 billion. However, propeller technology is largely of World War II vintage, reciprocating propulsion systems are cooled by excessively rich mixtures during climb, autopilot functions are fed back to the pilot as distract- ing control movements, and construction is typified by conven- tional aluminum sheet-stringer structure with protruding rivet heads in a large number of current aircraft.

This apparent conflict, where the fleet size is large and heavily utilized, the industry is enjoying record sales, while the equipment appears outmoded, points to some unique character- istics of the general aviation environment. Specifically, (1) the users (particularly single-engine and commuter users) are much more sensitive to purchase price than their heavy-jet, ..

commercial airline counterparts, (2) manufacturers are reluctant to incur increased production costs through product improvement for an apparently already satisfied market, and (3) the capital intensive nature of the industry may easily spell financial dis- aster for the manufacturer who misjudges the product improvement expectations of the user.

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However, the fuel price and availability problema which began in 1973, together with strong public reaction to noiae and emissions pollution, have resulted in the acceleration in develop- ment of several new and promising technologies.

2.1 BACKGROUND The recent development of many different technologies with potential application to the general aviation fleet of aircraft has reached a point where radically improved airplanes can now be foreseen. However, to realize these improvements applicable technologies must be identified and actively pursued in an orderly and timely manner. Also, synergistic effects resulting from the integration of appropriate technnlogies require identi- fication in order to better define research.

To illustrate this latter point, consider the development of an advanced natural lamina~ flow airfoil. Thorough evaluation of this technology would appear to dictate eventual full scale tunnel tests and flight teste. However. since surface roughness can impose severe penalties on the performance of this airfoil, a composite or bonded ~ing may be called for in order to eliminate the problem of rivet lines and butted skin joints. Here, a decision t~ pursue natural laminar airfoils through conventional manufacturing techniques may lead to an erroneous evaluation, while a decision to investigate a co~posite wing may be delayed if its potential benefit is seen only as an improvement in empty weight (when a conventional airfoil is utilized).

The key issues facina both research institutions (NASA

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Research Centers) and a capital intensive industry when confronted

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with the question of which techno1oaiea to pursue appear to lie

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I in how to identify thoae technologies which offer great potential for improving safety. performance, and cost as well as how to

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identify those with questionable benefits. An attendant result of such an evaluation appears to be the identification of those technologies with noteworthy (as opposed to highly ~ignificant) benefits which might be attained for a rather low level of development effort.

Silnificant pioneering work done in this area of techno1olY evaluation was recently completed by Berley (Ref. 18). The present research represents an effort to continue with this type of work on a much broader scale and in much Ireater depth.

2.2 PURPOSE In light of the above discussion, the purpose of this res~arch is threefold: (1) Identify candidate technologieF which appear to offer improvements in safety, fuel efficiency, performance, and utility of general aviation airplanes.

(2) Quantify the magnitude of these improvements.

(3) Investigate the synergistic effects of advanced technology integration on general aviation airplanes.

~ b c d 2.3 SCOPE Recognizing the different requirement, of different types of airplane. within general aviation, this reaearch effort w •• directed to an investigation of the impact of new technologies on a small airplane with a 6-passenger cabin (including pilots) and a larger airplane with at least a l2-passenger cabin (exclud- ing pilots). Performance guidelines for these two airplanes were very broad. General guidelines adopted for this study included high maximum 1ift-to-drag ratios on the order of 18, cruise speeds on the order of 250 knots, and landing speeds below 60 knotb.

Those tasks which were specified as a part of this study included the following: (1) R~presentative manufar.turers within the beneral aviation industry and certain NASA Research Centers were visited in or~er to integrate the views of government agencies and industry concerning new technologies.

(2) Promising new technologies were identified. and their impacts on two dif~erent airplanes were evaluated.

(3) Trade studies for the two specified airplanes configured conventionally witL aft tails and also as canards were performed (althou6h the canard studies met with difficulties).

(4) The Gen~ral Aviation Synthesis Program (GASP) was utilized in the evaluation of those technologies which affect vehicle weight and performance.

w 2.4 APPROACH The approach utilized to accomplish the required tasks lead- ing to successful attainment of the goal of this research may be broadly categorized as follows: (1) Identify and develop a data base.

(a) Conduct a literature search.

(b) Acquire, review, and tabvlate pertinent documents.

(c) Visit manufacturers and NASA research centers.

(2) Identify, develop, and test an evaluation technique.

(3) Identify and evaluate techuologies.

(4) GASP (a) Gain familiarity with the program.

(b) Modify as required to evaluate technologies.

(c) Benchmark against current technologies.

(d) Use to size configurations.

(5) Design and evaluate two advanced technology airplanes.

2.5 REPORT FORMAT Each of the four major tasks listed previously in Section 2.3 is liscuss~d in a separate chapter within this report, and an associated appendix is included where supporting documentation ..

for three chapters is prcvided. Here Appendix A is provided for Chapter 3 (visits). Appendix B is provided for Chapter 4 (techno- evaluation), and Appendix C is provided for Chapter 5 (designs).

GASP modifications ar~ discussed in Chapter 6.

Finally, Chapter: summarizes the results of the research ef- fort and Chapter 8 closes with conclusions and recommendations.

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CHAPTElll

VISITS !Q MANUFAC'l1JRERS AND RESEARCH FACILITIES 3.1 INTRODUCTION Visits to manufacturers and NASA research facilities formed one of three major thrusts of the present research. The purpose of this particular effort was to interview representative indus- try and research facilities with the goal of developing the information base and contacts required to support the project.

The 34 facilities which were visited provided a wealth of infor- mation and. as a spinoff. directed attention to many other sources of pertinent information.

This chapter presents an overview of the planning. initiation, proceedings, and results of the visitation phase. Appendix A is included as a supplement. where notes assembled by the research team during each visit have been compiled. ecl:ted. and are presented in abbreviated form.

3.2 MANUFACTURERS, RESEARCH FACILITIES, AND MEE'fINCS The work state~ent for this research pff~rt listed 28 manu- facturers and research facilities to be visited. As the project progressed, the list was modified to accomodate changes in order to provide the required data b~se for the ensuing technrylogy evaluation. In order to support the final visitation list without exceeding budget constraints, different modes of transportation

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were utilized. Figure 3.1 illu.trates the scope of the visitation phase by presentina the geographical location of the facilitiee involved in thi3 reeearch project.

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3.2.1 Manufacturing Facilities

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Table 3.1 provides a listing of those manufacturere vieited.

3.2.2 Changes ~ the Manufacturer Visitation Scbedule These changes represent the addition of five facilities to those already planned for. The following is a summary of the changes and the rationale for including them.

(1) Bellanca Aircraft Engineerlng. This firm was responsible for the development of the Bellanca Skyrocket, a high per- formance single engine aircraft. The inclusion of this firm was believed to be important because of their achievements in low drag airframes and composite construction techniques.

(2) Hartzell, Hamilton Slandard, and McCauley. An examination of the initial list of facilities detected deficiencies in the propulsion area. While engine manufacturers were in- eluded, representatives of the propelier manufacturers were not. These three companies were added to the facility list to broaden the propulsion technology base.

(3) Curtiss-Wright. This facility was added because of their work with rotary combustion engines.

3.2.3 Professional ~~eting~ These meet!ngs proved valuable to the research efforts of the project and provided significant insight to technology _l. __ ~~_........-. ... _~ ___ ~ __ ~_. ______ ~ ... _ ... , ___ ... _____ • ___________ _ , W~. .

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'J o • Klo Flight Research Lab • VL.;it.ed and Su .yec.

olo Visit Scheduled but. Cancelled • \' isited but not s\,rveyed o Meetings Fig. 3.1. Locations of Facilities and Meetings General Aviation Manufacturing Facilities Visited Table 3.1- Type i ~ , Loe,uion Facility Canpany I- I , I Wichit&, Kansas Airframe Beech Aircraft

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; Corporation Alexandria, Minnesota Bellanca Aircraft

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Corporation Middletown, Delaware Bellanca Aircraft Engineering, Inc.

e'ssna Aircraft Corporation Kansas Pawnee Division Wichita, Wallac£ Division Wichita, Kansas Wichita, Kansas Gates Learjet Corporation Gu1fstream American Savannah, Georgia Corporation Mooney Aircraft Kerrville, Texas Corporation Piper Aircraft Lakeland, Florida Corporation t

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Rockwell International General Aviation Bethany, Oklahoma Division Mojave, California Rutan Aircraft Factory Brittain Industries. Tulsa, Oklahoma Avionics Inc.

and Autopilots Cessna Aircraft Corporation Aircraft Radio & Boonton, New Jersey Control Division (ARC) iiow .""' ""1'- PaC'" ex !

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General Aviation Manufacturing 'aclliti.s Visited Table 3.1.

(Continued)

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Edo Aire/Mltchell Mineral Wells. Texas Avionics and Autopilots King Radio i Olathe, KanEias Corporation (cont)

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Narco Avionics Fort Washington, Pennsylvania Rockwell International Cedar Rapids, Iowa Avionics and Missile Group (Collins Avionics) Sperry Flight Systems Avionics Division Phoenix, Arizona Systems Bell Helicopter Textron Fort Worth, Texas Irvine, California Bertea Corporation (Division, Parker Hannifin Corp) Los Angeles. California Garrett AiResearch Industrial Division Hawthorne, California Systems Technology Inc. (STI) Propulsion Avco Lycoming Stratford, Connecticut Stratford DiviRion Cessna Aircraft Corporation ..

McCauley Accessory Dayton. Ohio Division Wood-Ridge, New Jersey Curtiss-Wright Corporation Phoenix, Arizona Garrett AiResearch Manufacturing Co.

Hartzell Propeller Piqua, Ohio Inc.

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p+ Table 3.1. General Aviation Manufacturing Facilities Visited (Concluded) Propulsion Teledyne Continental (cent) Hotors Aircraft Products Mobile, Alabama Division Unit~d Technologies Corporation Hamilton Standard Windsor Locks, Division Connecticut Pratt and Whitney Montreal. Canada of Canada Williams Res~~rch Walled Lake. Michigan Corporation implementation, current state-of-the-art, and advanced technology development efforts. An attendant benefit was the acquisition of reference material. The presentations provided exposure to many subject areas of interest and th~ acquired papers provided documented support for the opinions expressed by the speakers.

The NBAA meeting differed significantly from the others in tliat it provided product information as well as an exposure to the marketing strategi~s of manufacturers. Table 3.2 provides a list of the meetin~s attended.

3.2.4 Research Facilities ------>- ---- .. - .. ---- Three NASA r('searl'h facilitlt>s were;> visitpd as shown in Table 3.3. These centers provided the necessary technical background fOT emcrgin~ technologies which would otherwise be l~ Table 3.2. Professional Meetings Month Meeting Organizat ion (1919) Location

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• 1st Central Oklahoma American Institute Feb Norman, AlM Mini SyU\posium of Aeronautics and Oklahoma

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Astronautics • Society of Apr Wichita, Business Aircraft Meeting Automotive Engineers Kansas Display Systems University of Jun Los Angeles, Engineering Short California at California Course Los Angeles AIM Systems and Aug New York, American Institute Technology Meeting of Aeronautics ilt d New York Astronautics NBM National National Business Aug Atlanta, Georgia Convention Aircraft Association General Aviation Lewis Research Nov Cleveland, Ohio Propulsion Conference Center (NASA)

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unavailable (unpublished). Also, they wi1ling:y discussed both strengths and weaknesses of candidate technologies within their fields of expertise.

Table 3.3. NASA Researc~ Centers Center Date Location ..

June, 1979 Moffett Field, Calif.

Ames Research Center Hampton, Virgir.ia Langley Research Center August. 1979 Lewis Research Center June. 1979 Clevel1\nd. Ohio 1" e -~ r&d r"t& « 1m tt m _' .........

m 3.2.5 Facility Contac!!

Successful implementation an~ completion of the visitation phase relied heavily upon the contacts established with the facilities. Table 3.4 lists the final contacts at each facility.

3.3 MEETINGS All meetings within the v1sitation phase proved quite fruitful.

In every esse. the reception of the oroject staff by the facility representatives was exceptional, and an open and candid atmosphere prevailed throughout all discussions. Areas of mutual interest regarding current work on advanced technologies as well as the impressions and attitudes of the general aviation industry were discussed. To cover each of these meeti..gs individually is impractical in this text. Therefore, an abbreviated compilation of notes generate~ by the project staff at each of the facilities visited is included as Appendix A.

The following discussion is intended to serve as an intro- duetion to that appendix. As such, it provides a brief description of the subject areas discussed during the meetings. Manllfacturer visits are first discussed, followed by NASA visits.

3.3.1 Discussions Nith Manllfactur:.!E.S. Facilities 3.3.1.1 Airframe. The airframe manufacturers are, in every sense, the generalists of the industrv. Their final produc~, an op~ra- tiona1 airplane, must include products representing many techno- logies and disciplines. Th~s, these manufacturers ml'.st have an awareness of all aspects of the industry.

Table 3.4. Facility Contacts Facility Contact Ames Research Cent.r Seth Anderson Research Assistant Intra-Agency Programs Avco Lycoming Walt Schraeder Director of Advanced Technology Engines Beech Aircraft Corporation Bill Wise Vice President Advanced Technology Bell Helicopter Textron Hugh Upton Group Engineer Research Electronics Bellanca Aircraft Corporation Andrew Vano Chief Engineer Bellanca Aircraft August Bellanca Engineering, Inc. President & Chief Engineer Bertea Corporation John C. Hall Group Vice President Brittain Industries. Inc. Charles Walters President Cessna Aircraft Corporation Virgil Davis ARC Division Chief Engineer McCauley Accessory W.B. Voisard ..

Division Chief Engineer Pawnee Division Harvey Nay Chief Engineer Wallace Division Emmett Kraus Supervisor of Advanced Design Curtiss-Wright Corporation Bill Silvestri RC Engin~ Program Manager r Table 3.4. Pacility Contacts (Continued) Facility Contact Edo Aire/Mitchell John Nixon Chief Engineer Garrett AiResearch Industrial Division Dick Barcus Project Engineer Aircraft Systems & Control Manufacturing Company M.C. Steele Director of Engineering Gates Learjet Corporation Richard Etherington Director of Technical Engineering Gulfstream American Bob Stewart Corporation Assistant to the Vice President of Engineering Hartzell Propeller Inc. Ben Harlamert Vice President. Engineering & Chief Engineer King Radio Corporation Dan Rodgers Group Leader Special Programs Langley Research Center Bruce Holmes Aerospace Technologist Flight Mechanics Division Lewis Research Center William Strack Supervisory Aerospace Engineer PropulSion Section Mooney Aircraft Corporation Fen Taylor Chief of Aerodynamics and Performance Narco Avionics Norman Messinger Manager Advanced Development t rrt rtrt ,--MI!I!!

Table 3.4. Facility Contacts (Concluded) Facility Contact Piper Aircraft Corporation Grahame Gates Director Advanced Engineering Rockwell International Collins Avionics G.L. Benning Vice President of Advanced Technology and Engineering General Aviation Division Larry McHughes Director, Engineering Rutan Aircraft Factory Bert Rutan

Sperry Flight Systems w.r. Robertson

Manager, Engineering Systems Technology Inc. Irving L. Ashkenas Vice President Teledyne Continental Motors L. Waters Vice President Aeronautical Engineering United Technologies Corp.

Hamilton Standard Division D.F. Phillips Head of Technical Planning Pratt & Whitney of Canada Sid Monaghan Chief, Research and Development Support Williams Research Corporation Edward J. Lays Senior Applications Engineer ..

Key areas included in the discus.iona were: (1) Propulsion and powerplantl. New generation enginel, engine controls, propellers, and their integration into an airframe.

(2) Structures. Composites, metal~etal bonding_ I

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(3) Systems. Flight control systems, micro-computer based systems, and avionics.

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(4) Aerodynamics. New generation NASA airfoil sections, and computational aerodynamics.

(5) Configuration. The advantages and disadvantages of canards and tandem wings.

3.3.1.2 Propulsion. Much developmental work is occuring within this industry, and many promising concepts are emerging. The "new generati.on" of general aviation powerplants and propellers was a prime topic throughout the industrial and research community.

Topic<; discllssed with the propulsion repres(>ntatives were: (1) General Aviation Turbine Engine - GATE.

(2) General Aviation Propeller Study - GAP.

(3) Advanced propellers - configuration aspects, advanced airfoil sections, composite materials.

(4) Propulsion integration - airframe-powerplant integration.

(5) High speed propellus - p.:-c'I,fan.

(6) Posit ive displacement engines - diesel, n,tary combust ion engine. advanced reciprocatin~ en~ine concepts.

(7) Powerplant control - integrated controls. microprocessor- based c.mtrols.

3.3.1.3 Avionics. Ad~anced technology prpvails within this manu- facturing group ~ue to the competitive nature of this particular ~- _rd $. « # market. Survivability demands continued re ... rch and development in new technolo,y areas, re.ulting in a level of sophistication that often aurpaas.s military and commercial markets.

Areas covered during the discussions included: (1) Digital technology - analog veraus digital avionica.

(2) Integration - multi-fw\ction avionics packages and atandardi- zation, multiplexing.

(3) Displays - electronic displays and instrumentation.

(4) 'light controls - digital Lad electronic flight cOutrol systems.

3.3.1.4 Systems. Systems manufacturers offer areas of technology transfer to general aviation. For the most part, their main contributions lie in areas other than general aviation. Broad foundations in other fields allow "spinoff" technologies to filter in onder circumstances and costs that are acceptable to the community of manufacturers and users.

The areas discu@sed included: (1) Fli~ht control actuators - advanced hydraulics applied to general aviation.

(2) Turbocharging - advancements for general aviation.

(3) ~isplays - Heads up and head mounted displays.

(4) Fiber optics - signal transmission, airframe structural monitoring.

(5) Flight control systems - microprocessor based systems.

fluidics.

q It ahould be noted that co.menta by induatry repreaentatlvea were not ltmited aolely to the topic. liated. Due to the nature of the interviews. the perlonal interesta of the participanta, and the degree of integration required for production of general aviation hardware, much overlap between subject area. existed.

Appendix A should be eon.ulted for further information.

3.3.2 NASA Research Centers Appendix A also includes the summari2:ed notes from the visits to the Ames, Langley, and Lewis Research C:enters.

3.3.2.1 Ames Research Center Topics.

(1) Avionics (a) Preliminary Candidate Advanced Avionics Systems (PCAAS).

(b) Demonstrator for Advanced Avionics Systems (DAAS).

(2) Stall/spin aerodynamic tailoring.

(3) Cooling drag.

(4) Small Transport Aircraft Technology (STAT).

(5) Aerodynamics.

3.3.2.2 Langley Research Center ropic~.

(1) Crash dynamics - seats. restraints, structures, fire preven- tion.

(2) Composites - types, testing, characteristics.

(3) Aerodynamics (a) 3-dimensional (b) 2-dimensional natural lamin~r flow sections . low speed sectio~s • high speed sections (4) Avionics and control. - pilot-ATC interface, fluidics. pilot workload.

(5) Stall/.pin research.

3.3.2.3 ~~ Research Center 1opic •• (1) GATE (2) Positive displacement en,ines (a) Advanced reciprocating engines.

(b) Alternative engine systems • Diesel engines Rotary combustion engines (3) Propeller Technology - turboprop, prop fan, variable pitch fan. turbofan, GAP program.

3.4 RESUL'f~ 9F THE FACIl.ITY VISITS The visitation phase of the project proved to be quite successful. A strong data base was developed through the inter- views as well as from those sourc~s of information identified during the intervic~s. Also, the insight and opinions of the I various representatives pr~vided indications of the practicality and/or feasibility for advanced technology development.

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3.4.1 Technology and the Industry To general aviation. like other industries. a technology 1s of no use unless: -0 Pl. _' (1) 1t .. t1af1e. a need, (2) tbe developer CaD afford 1t, (3) 1t 1. prof1table to the u.er, aDd (4) 1t 1. 1ntroduced at tbe r11bt u ....

The molt iaportant factor affect1na technology fmpleaentation 1. u.er acceptance. Manufacturer. cannot afford to pursue a technololY merely to improve a product unle •• tbe u.er require.

1t. A market for tbe techDololY mult ex1st, and tbe as.oc1~t.d develo~ntal costs mu.t be acceptable to tbe manufacturer.

3.4.2 TechnololY and £2!l

Three major cost constraints exist for tbe general aviation manufacturer when incorporating new technologies. tn general, 'hese may be grouped into the broad categories of developmental costs, certification costs, and product liability COltS.

3.4.2.1 Developmental~. Developmental COStl can be extreme, particularly in high technology areas. Many companies cannot absorb these and must rely instead on developments within NASA research centers or industries external to general aviation.

Often. the production base or technology requirements of the automotive industry or the spinoffs realized from ~ther technolo- gies serve to reduce general aviation developmental costs. For example, the cost of the 8,000 aircraft turbochargers producod annually benefits significantly from the 1 million unit. demanded by the automotive and tru,king industry.

e 3.4.2.2 Certification~. WhUe the project val conceruct pri.ari1y with an ........ nt of .dvanced tecbDoloaie •• the fe •• f- bility of proai.iaa tecbno108ie. heiDI intelr.ted into future airp1.nes .ppe.red to be jeopardized •• .uch by certific.tion COltl as by techno10lic.1 ri.kl. Some manuf.cturer •• for example, pointed out that theae coata can amount to 10 to 100 time. the co.t of deve10pine the te,hn010gy itae1f.

Since this difficulty ~resented serious imp1icstigns fo~ the research at hand. an effort wat made to aolicit the opiniona of manufacturers regarding certific.tion procedures. This w.s done in order to evaluate manufacturer p~rception. of the process and is not intended to be an evaluation of the process itself. In every case discussed, members of the present research team noted that sufficient information t.~ allow an unbiased evaluation of the problems identified was lacking. However, the following points appear to merit further discussion and consequently are incorporated into this report.

(1) Interpretations of the same regulation by different FAA regional offices may sometime'J result in mark':!dly diffen~nt certification r~quirement.s for the same technology in differ- ent regions.

(2) Delays resulting from "excessive red tape" nrt! e:o:pensive to manufacturers. One frustrated manufacturer produced docu- mentation which indicated that an application filed !fIure than 6 months earlier had yet to generate a response.

(3) Personnel qualifications were addressed by one manufacturer.

In this particular case. the manufactur.~ was irritated because aL inexperienced individual with a civil ena1neer- ina backaround was aiven the responsibility for establishina compliance with propulsion/airframe certification require- ments.

(4) New technoloaies are often evaluated by old or outdated techniques. In one case. an attempt to utilize a finite element computer code (NASTRAN) to validate structural integrity computations was perceived by the manufacturer to be hampered by a lack of familiarity with the code on the part of the certifying officials.

The significant and common denominator which appeared when- ever a manufacturer chose to identify "difficulties with certifi- cation" was that the issue was usually an emotional one. Better corom"aications between manufacturers and FAA regional offices such that both parties recognize the difHculties faced by the other appears to be a fundamental requirement if the issues are to be resolved. The inescapable fact is that manufacturers pay a heavy premium in order to certify a new t.echnology.

3.4.2.3 Product Liability Costs. These costs are also high.

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One airframe manufacturer pointed out that 15% of a single engine airplane's price represented product liability costs. Another general aviation manufacturer pointed to insurance premiums of $3 million per year to illustrate these costs.

Nuisance suits seem to prevail within this industry, and several examples were pointed out by some of the manufacturers.

t • Whether justified or not, such suits represent a very real coat to the manufacturer.

The disturbing aspect of high product liability costa ts that they point to a history of large court settlements. This would appear to deter the incorporation of new technologies and, instead, promote an atmosphere !'f conservatism wi.thin the indus- try.

Two facets of product liability deserve mention. In one case, manufacturers may avoid new technologies for fear that the incorporation of improvements to systems may be interpreted to mean that deficiencie3 exist in previous models of the same system. This otherwise unne~e9sary exposure to suits can be a very real deterrent to the incorporation of new technologies.

The second point to be made is thpt the incorporation of a new technology always has attendant risks. Here, the experience gained with a mature system will no longer assist the manufacturer in avoiding the many unforeseen problems associated with new technologies. Product liability takes on added Significance under these conditions.

3.5 RECOMMENDATIONS FOR NASA The following suggestions and recommendations are the result of observations made during (he visits to manufacturers and research facilities.

3.5.1 Communications Better communications need to be maintained between NASA and ·f!&? ..

_.~ 44 _. A!P# .¥& . >or'" the industry with regard to fortbcomina research topics .nd the results of completed research projects. For example, several companies indicated that they were not aware of the final reeults of the RedhAwk. and ATLIT programs. While this points to a dis- turbing llck of awareness of technical publications on the part of industry, NASA will promote its 1m3ge tmmensely if industry oHidals are advised of the availability of these reports.

tnt':restingly, NASA publications designed to fulfill this task (SCAN's, STAR's, Tech Briefs, etc) apparently are not reaching the industry or are not being used by them.

3.5.2 Basic Research and Product Development Industry encourages NASA involvelllent with basic ref€tlrch and discourages any efforts aimed toward product development. Sin~e industry involvement with basic research is Sffiall (less than 5% of the total engineering budget for ont' of the industry's leaders), this research team emphasizes that NASA should continue its efforts in basic research. High risk technologies such as those associated with propulsion, and sophisticated aerodynamic analytical tools such as those associated with natural laminar flow airfoils and numerical optimtzation techniques deserve special attettion.

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Avionics manufacturers. on the other hand, appear wery of any NASA efforts which may be perceived to lead to standardization similar to those embodied in ARINC specifications.

3.5.3 Research Contracts NASA should not fund programs which diminish a competitive ---~'- ~~----. __ """'b"'! ...... _ ... - ___ * .. ___________ _ advantage which .. y already be enjoyed by a manufacturer th~ouah its own research efforts. In those ca.es where the problem is .are perceived than real, NASA should clearly and publicly define the limits of the research contract being awarded.

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CHAPTER !

TECHNOLOGY EVALUATION The goal of this phase of the study was to establish a rank ordering of technologie& according to their benefits to General aviation airplanes. This involved three major task,l: (1) select- ing the method of evaluation, (2) identifying the candidate tech- nologies, and (3) ·avaluating the candidate technologies.

4.1 METHOD OF TECHNOLOGY EVALUATION In selecting the evaluation method, the following criteria were applied to the available techniques in an effort to deter- mine the technique best suited to the requirements of the present =esearch effort.

(1) The method must allow the evaluation of specific technologies.

(2) The method must allow dissimilar technologies to be evaluated in a consistent manner.

(3) The method must be within the means (cost and time) of the project.

(4) The method must be as objective as possible subject to the three previous criteria.

Many of the techniques used in technological forecasting did not fulfill the requirements of this research effort. For example, regression and trend analyses assume one has suitable data to con- struct extrapolations. Such data does not exist for many of the technologies considered in this SL~=Y because they have had only limited use in _eneral aviation airplanes. A direct technology 31.

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evaluation uSing the Delphi Method (Ref. 125) was rejected aa beyond the means of the project due to the large number of techno- logies to be investigated. The evaluation method finally adopted is based on a technology figure of merit concept and has already met with some success (Ref. 18). This method, here termed "evaluation technique," is a simple linear compensatory lDOdel and is discussed below.

4.1.1 Evalua~ Technique In the evaluation technique (ET), candidate technologies (eT) are evaluated relative to current technologies by assessing their impacts on a group of categories. The categories model the major factors involved in the operation of an airplane and are assigned weightings (w) according to their relative importan('e. The impact of a CT in a given category is quantified by a relative benefit (b). where b > 0 indicates that the CT offers an improvement in the category (relative to current technology) and b < 0 indicates that the CT causes a degradation in the category (relative to current technology). Summing the products of the relative benefits and category weightings yields a figure of merit (FM) for the CT which is a measure of how much improvement the candidate technology offers in the overall operation of the airplane. The figure of merit is defined by Equation 4.1 while Fig. 4.1 illustrates the general concept of the evaluation technique.

FM '"' (4.1) i where n is th~ number of categories, b is the relative benefit of technology i in category j, ij Wj is the weighting of category j. and FMi is the figure of merH of technOlogy i.

Referring to Fig. 4.1, one notes that the adoption of t:his type of technology evaluation method is not without difficulties.

First, the categories must be selected, and second, their weight- ings must be determined. So, while the ET inherently satisfied the first three criteria, the fourt •• criteria, that of objectivity, demanded that special attention be given to the category selection and weighting.

4.l.~ Category Selection The first step in defining the category group was the estab- lishment of the following ~election criteria: (1) The group of categories must be broad enough to model the major factors involved in the operation of an airplane.

(2) There should be a mini~um of overlap between individual categories.

(3) The number of categories should be as small as possible while satisfying the first two criteria.

Various selection schemes were studied, and it soon became

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obvious that all of them involved generating a consensus from a set of opinions. The Delphi Method, althou~h too involved for th~ overall t~chnology evaluation, ~eemed \lell suited to th~ cate- gory selection problem and was therefore used. The survey group for the cateogry selection consisted of the project staff and the .... tr

.... '

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/

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9 10 8 5 FH

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Turbop r,3 +3 x 10 +5 x 10 +2 x 5 -+ +44 ro -4 x 9 -+

o x 9 +1 x 10 +2 x 8 o x 5 +26

Wankel

.. ~

Candidate

-1 x 9 o x 10 -1 x 8 o x 5 -17

"~glets ~ Technologie

-

L.E. Slats -2 x 9 +2 x 10 -1 x 8 o x 5

- 6

-

NAVSTAR/GPS +1 x 10 o x 8 o x 5 -+ + 1

-1 x 9

0 !

! ! !

~ ~

.....

(Relative Benefit) x (Categcry Weight) Figure 4.1. Illustration of the Evaluation Technique Ii J the faculty members of the Department of Aerolpace Engineering at the University of Kansas.

The category selection survey I involved the following pro- cedures: (1) Obtain a category list, confidence level, and comments from each of the participants. The confidence level is a measure of a participant's confidence in his response.

(2) Analyze the group response and generate an "average" category list.

(3) Feed back the average category list along with participant comments.

(4) Repeat (2) through (3) until a viable "average" category list is obtained.

The Delphi Method worked quite well although the analysis of the group response was complicated by the fact that the survey dealt with symbols (in the form of category names) instead of numbers. The survey converged to an acceptable category list (17 categories) in four survey rounds. The final category list is given in Table 4.1.

Referring to Table 4.1, one notes that some categories appear to overlap. Examples are the categories of Fuel Efficiency, Reliability, and Direct Operating Cost. The problem is minimized when one considers that Fuel Efficiency is concerned wit:- fuel availability as well as with fuel costs. Lik~wise, Reliability is concerned with operational readiness as well as maintenance costs.

rt #be

{

\

With these faetors in mind, the eatelory list was eonsidered aeeeptable. Further diseussion of the eatelory selection survey may be found in Appendix B.l.

Table 4.1. TeehnololY Evaluation Catelories Catelory Definition Altitude at which the maximum Ceiling rate of climb ia 0.51 mls (service ceilinl).

The characteristics of an air- Crashworthiness plane which determine the level of occupant protection in the event of a crash.

Maximum continuous cruise Cruise Speed speed.

All costs directly attributable Direct Operating Cost to flying and keeping an air- plane operational. All scheduled and unscheduled main- tenance costs and fuel costs are included.

Pollutants produced during Emissions the operation of an airplane; does not include noise.

Airplane weight without fuel, Empty Weight crew, and payload.

Noise perceived at ground Exterior Noise level due to the operation of an airplane.

Airplane cruise efficiency Fuel EfficiE-ncy measured in air-miles per pound-fuel.

Noise perceived by the occu- Interior Noise pants of an airplane.

Table 4.1. Technology Evaluation Categorias (concluded) I ~ ..

Definition Catelory Pilot Workload The amount of time. concentra- tion. and effort a pilot must devote to the safe operation of an airplane. This includes the affacts of airplane handl- ing qualities.

The price paid for a new air- Purchase Price plane by the user. including avionics and equi~ment co~ts.

Range The distance an airplane can fly without refueling. allow- ing for appropriate fuel reserves.

Reliability A measure of the pr('babllity of failure of an airplane com- ponent or system.

A measure of the effects of Ride Qualities ai~craft motion on the smooth- ness and comfort of the ride experienced by the occupants.

Safety A measure of an airplane's inherent characteristics which reduce the probability of an accident.

Static Comfort A measure of an airplan~'s inherent comfort. This in- cluUp.s roominess, seat com- fort. ventilation, decor,

•

ease of entry, etc.

Takeoff/Landing Performance This parameter includes take- off and landing speeds, field length requirements, and rates of climb and descent.

4.1.3 Catelory Weiahttgas Having compll~ad the category selection survey, the Delphi Method was again used in the deta1'llination of the cstegory _tahU.

This survey, termed the "category rating survey". was cODducted on a much larser scale than the category selection survey with 42 participants representing general aviation manufacturers and user groups, university faculty, NASA centers, and project staff. A list of the participants is aiven in Table 4.2.

Upon reviewing thf! 17 categories to be weighted, it was noted that in general, different types of airplanes would have different sets of category weights because of differing operational priorities. Because the statement of work (for this study) speci- fied that two types of airplanes were to be investigated, two sets of category weightings were generated. In the rest of the text, the two airplanes are referred to as Airplane A, which is a six passenger (including pilots) airplane for business and/or personal transportation, and Airplane B, which is a 19 passenger (excluding pilots) commuter airliner.

The category rating survey was conducted as outlined below: (1) PartiCipants gave each cat=gory (for each airplane) a rating

\

I (R) constrained by 0 < R < 10 with a minimum scale increment f , of .5. R· 10 was aSSigned to the cat~gory considered most important and the other categories were rated in a relative manner with R • 0 mean~ng the category is of no importance relative to the most important one. Duplications in ratings were allowed because two or more categories could be of equal r T"-'le 4.2. Cat.aory latina Survey Participant.

Aircraft Ownere and Pilote Aasociation AVCO Lycoming Beech Aircraft Corporation Bell Helicopter Textron Be11ance Aircraft Corporation Bellance Aircraft Engineering, Inc.

Brittain Industries, Inc.

Cessna Aircraft Company - Aircraft Radio and Co.~I;rol Division Cessna Aircraft Company - Pawnee Division Cessna Aircraft Company - Wallace Division Commuter Airlines Association of America Curtiss-Wright Corporation Edo-Aire Mitchell Division Garrett-AiResearch - Industrial Division Gates Learjet Corporation Gulfstream American Corporation Hartzell Propeller. Inc.

King Radio Corporation Mooney Aircraft Corporation

•

Narco Avionics Division NASA Ames RL'search Center NASA Langley Research Center NASA Lewis Research Center National Business Aircraft Association ~~ ....... ...- ______ sz"""'~· --.-.;;;....~ .. -~.~.~ .. ~-

i

Table 4.2. Catelory Rating Survey Participant. (concluded) Parkel H8~nlfin Corporation - Control 5Ylt .. Divi.1on (formerly Bertea Corporation) P1.pe) Aircraft Corporation Pratt & Whitney Aircraft of Canada, Lld.

Rockwell International - ColUnl Dividon Rockwell International - General Aviation Divilion Sperry Flight SY8tems - Avionics Division Systems Technology Incorporated leledyne Continental Williams Research Corporation

* University of Kansas, Aerospece Engineering Faculty

* Projecl Staff

*

These groups each contributed more than one particl~ant importance. Participants also rated their degree of confi- dence in their response with a confidence level (Cl) con- strained by 0 ~ ~T. ~ 1 for each airplane. Comments on the category ratings were encouraged.

(2) Feedback to the participant3 consisted of information regard- iug the mean category ratings, the category rating distribu- tions (in the form of histograms). and p~rticipant CJmments faT each of the airplanes.

(3) The survey continued until the participant response reached a predetermined level of stability (less than 151. change in partir.ipant voting).

---------- - ------......--- The category rating survey required three rounds to achieve satisfactory convergence. The category weightings used in the evaluation technique are simply the mean category ratings trom the final survey rounds.

Table 4.3 presents the final category weightings for Airplanes A and B and Table 4.4 shows the category rankings in order of importance. Details of the category rating survey may be found in Appendix B.2.

4.2 CANDIDATE TECHNOLOGY IDENTIFICATION This section deals with the selection of the technologies to be analyzed with the evaluation technique and discuss~s the data base used. Since a large list of advanced technologies is easily created, several constraints were placed on the selection process.

An early list, defined as "Preliminary Candidate Technologies," and a final list, defined as "final Candidate Technologies," are both discussed here.

4.2.1 Data Base In selecting candidate technologies, one must first collect information regarding technolo~:~s in general. The project relied on three major sources of information as detailed below.

•

4.2.1.1 Literature Search. A computerized literature search was conducted using the Lockheed DIALOG system. In preparing for the computer search, a manual search was made of the NASA STAR index to identify key words, technology areas, and ~amp1e titles. This 11 a- • tz .• M·· . 7 . __ . __ .. _._------=-.,.....--_.--- Table 4.3. Evaluation Technique Category Weighting.

Weighting.

Category *

** Airplane A Airplane B

Ceiling 5.429 5.447 Crashworthiness 6.816 7.333 7.922 Cruise Speed 7.480 Direct Operating Cost 7.906 9.451 Emissions 1.606 2.156 Empty Weight 4.340 5.657 Exterior Noise 4.287 5.413 Fuel Efficiency 8.606 8.157 7.452 Interior Noise 7.037 Pilot Workload 7.349 6.998 8.515 7.761 Purchase Price Range 7.290 7.181 Reliability 8.854 9.547 Ride Qualities 6.110 7.238 Safety 9.422 9.544 Static Comfc~ L 6.244 6.797 Takeoff/Landing Pe:."formence 6.874 7.407

* Airplane A - 6 Passenger Business/Personal Airplane

**

Airplane B - 19 Passenger Commuter Airliner 4') $ S r --------------~ Table 4 .4. Evaluation Technique Catelory Rankinls Catelories Rank Order

* ** Airplane A Alrplli.&"l~

B Safety Reliability 2 Reliability Safety 3 Purchase Price Direct Operatinl Cost 4 Fuel Efficiency Fuel Efficiency ,I Cruise Speed Purchase Price " 6 Direct Operating Cost Cruise Speed 7 Pilot Workload Interior Noise 8 Takeoff/Landing Range Performance 9 Interior Noise Crashworthiness 10 Takeoff/Landing Ride Qualities Performance 11 Crashworthiness Range Static Comfort Pilot Workload 13 Ride Qualities Static Comfort 14 Ceiling Empty Weight 15 Empty Weight Ceiling 16 Exterior Noise Exterior Noise 17 Emissions Emissions

* Airplane A - 6 Passenger Business/Personal Airplane

'II'll Airplane B - 19 Passenger Commuter Airlinet resulted in six primary search topics (a-f) to which three specific topics (g-i) were later added.

(a> Aircraft Design (b) Aircraft Propulsion (c) Aircraft Structures (d) Flight Controls (e) Navigational Aids (f) Avionics (g) Canard Configurations (h) NAVSTAR/GPS (i) GASP (General Aviation Synthesis Program) Table 4.5 presents the initial results of the computer search in terms of the number ci abstracts printed for each search topic; note that some topics were combined to eliminate duplications.

The 1655 abstracts were reviewed individually to identify pertinent articles. This resulted in 107 articles which were obtained from NTIS. The specific search procedures used for each topic may be found in Appendix B.3.

4.2.1.2 Visits and Meetings. Quite a few articles were found through recommendations of various people and from the references given in other articles. Also, aviation related magazines and journals were helpful in finding information.

The bibliography lists the most important bOGks, articles, and papers obtained while the discussion~ of Chapter 3 and Appendix A summarize the information gained from the visits and meetings.

Table ~.5. Computerized Literature Search Resulta Abatra.~.cr ~"inted Topic . ~ ..

(a) Aircraft Design .. 33 (b) Propulsion 206 Aircraft (c) Structures 275 Aircraft (d) Flight Controls 780 (e) Navigational Aids 780 (f) Avionics (g) Canard Configurations 22 (h) NAVSTAR/GPS 89 (i) GASP 0 4.2.2 Preliminary Candidate Technologies The preltminary candidate technologies (PCT) were the result of the first formal attempt to identify technologies felt to have some application to General Aviation airplanes. Selecting the PCT proved more difficult than expected for several reasons, not the least of which was how to define a "technology." To select technologies in a consistent manner, three criteria were formulated.

(1) Operational results of a technology are not themselves , technologies: Examples: "stall/spin prevention" and "increased TBO (time between overhaul)" are the results of the application of various technologies.

SPES '2m n r 7

-' (2) Design variables (parameters) are not technologies. Examples: "high wing loading," ''high aspect ratio."

Non-specific ''wish list" technologies are to be avoided.

(3) Examples: "low-cost sensors," "low-cost weather radar."

TIlese criteria were followed as closely as possible but mar-

~

giual technologies were given the benefit of the doubt to insure that all applicable technologies were considered. The PCT selec- tion process resulted in 137 technologies which are given in Table 4.6, grouped according to technology area.

Table 4.6. Preliminary Candidate Technolo~ies 1. AERODYNArHCS (11) • Winglets • Variable geometry winglets • Spoilers • Fowler flaps • Low drag surface coatings • Leading edge devices • Advanced low and medium speed airfoils (turbulent) • Supercritical airfoils • Advanced natural laminar flo· ... airfoils • Improved stall/spin through aerodyna~ic tailoring • Active laminar flow control II. AIRCRAFT SYSTFHS (g) • Microwave anti-icing • Sonic/pulsating anti-icing -- '0- _~ __ 1

-

Table 4.6. Preliminary Candidate Technologies (continued) II. AIRCRAFT SYSTEMS (concluded) • Lithium hydroxide/hydrogen peroxide batteries , • Air cycle environmental systema • Imp~oved lead-acid batteriee • AC electrical systems • Passive anti-icing through icephobic coatings • Variable cycle environmental systems • Accurate fuel mon.ttoring and management • High s~eed brushless alternator • Single unit starter-generator • Air bearings III. COHPUTA'l'IONAL METHODS (~J • Computational aerodynamics • Computational structural deSign/analysis • Aeroacoustic modeling/analysis • CADCAM IV. CRASm~ORTHINESS (7) • Load limiting seats ..

• Improved restraints • Foam filled fuel tanks Burst/tear resistant fuel tanks • Energy absorbing floor • Anti-misting fuel treatment • frangible fuel fittings Table 4.6. Preliminary Candidate Technologies (continued) V. FLIGHT CONTROL SYSTEMS (21) • Fly-by-wire • Fly-by-light • Active gust alleviation • Active ride smoothing • Active flutter supression • Active controls for re13xed inherent stability (CCV) • Integrated low-cost wing leveler • Integrated yaw damper • Separate surface technology • Winglets for lateral-directional control • Direct side-force control • Direct lift control • Single level thrust/drag control • Force-stick controllers • Digital automatic flight controls • Fluidic automatic flight controls • Stick shaker/pusher for stall prevention • Stabilizer/elevator spoilers for stall prevention • Pneumatic actuators • Hydraulic actuators • Electro-mechanical actu~tors VI. INFORNATION SYSTEMS (28) • Flush antennas • Digital data links Table. 4.6. Preltminary Candidate Technolo8ies (cont!nued) VI. INFORMATION SYSTEMS (continued) • Single function CRT displays • Time-shared CPT displays 'HUD Micro HOO • Warning annuciators • Total panel-mounted avionics • Active outside imaging • Fluidic shed-vortex airspeed sensor • Airplane health/diagnostic systems • Fluidic rate sensors • Multiplexing • ARINC-type broadcast hierarchy • Integrated avionics and displays • Piezoyresistive preasure transducers • Fiber optics for data transmission • Liquid crystal displays Flat CRT displays

•

• Touch sensitive CRT • '~eather radar • Alternate weather detection • Radar altimeter • Onboard computing capability • 3-axis magnetometer acceleration sensor f

•

I

•

r t m i-A Table 4.6. Pre1tminary Candidate Technologies (continued) VI. INFORMATION SYSTEMS (concluded) • Improved stall warning • Laser gyros VII. MATERIALS/PROCESSES (12) • Metal/metal bonding • Fiberglass composites • Kev1ar composites • Graphite composites • Honeycomb core composite skin panels • Single crystal metal • Powdered metal • Isothermal forging • Diffusion bonding • Friction welding • Corrosion resistant coatings • Matched-die fiber reinforced plastic (FRP) VIII. NAVIGATION CONCEPTS (11) • V()R/DME RNAV • Scanning VOR/Dt'.E RNAV

•

• Omega • Differential Omega • VLF NAVCOM • Loran C • NAVSTAR/GPS l Sf '-n Table 4.6. Preliminary Candidate Technologiel (continued) VIII. NAVIGATION CONCEPTS (concluded) • Inertial navigation • Doppler navigation • MLS • Inertial amoothina IX. NOISE (6) • Noise absorbing materials • Improved mufflers • Variable engine/prop gearing • Q~iet propeller technology • Low level pressurization • Ducted propulsors X. PROPULSION (25) • Advanced diesel engine • Advanced rotary combustion engine • Advanced reciprocating engine • GATE engine • Auto engine conversions • QCGAT engine • Liquid cooling • Stratified chatge • Improved turbocharging • Variable timing • Electronic ignition ~o .-=-. ... ""'_ ...... ""'--_- __ ~ __ _

~----_tr_-_- ________ __

- oe.

Table 4.6. Preliminary Candldate Techaololle. (concluded)

x. PROPULSION (concluded)

• Automatic mixture control • Lean burn combustion • Density compensatlnl fuel injection • Total microprocessor enaine control • Variable bypa.s turbofan • Variable pitch fan • Efficient propeller technology • Prop fan • Cooled turbine blades • Ceramic turbines • Composite propellers • Torsionally (aeroelastically) tailored propeller blades • Single lever throttle/mixture control 4.2.3 Yinal Candidate Technologies As a group, the 137 PCT of Table 4.7 were not equally suited to the evaluation technique. First, the PCT exhibit different technology levels. In the propulsion area for example, complete advanced technology engines are compared against component tech- nologies. One cannot really compare the two although both may be iJ~portant. Second. the PCT exhibit different technology orders, meaning that while most technologies are first order (i.e., ~ the airplane) some arc second order (i.e., ~ ~ the airplane).

_________ ~ ____________ J

, 4 '1<'~;::"~~'~ ; Examples are "advanced natural l_tnar flow atrfoUa" (firet order) and "cOliputational aerodynaaics" (second order). The firat requires the second. Hence. the two technololies ~re not r .. lly

\

' cOliparable. In addition to these fUftdament~l probl .. s. four other , constraints are noted: (1) the IT cateaory weiahtinas are valid , only for six pa.senaer and commuter airplanes. (2) the project

i

scope requires 1990 technology implementation. (3) technologies

I

for which data are lackina cannot be evaluated objectively. and ~ (4) some of the technologies are already gain1ng acceptance in

!

general aviation.

These factol's gave rise to a second set of n 1 teria which

I

,~ were used to select the final candidate technologies (FCT). • (1) Try to reduce the differences in technoiocy levels.

(2) Delete second order technologies.

(3) Delete technologies not applicable to the six passenger or commuter airplanes.

(4) Delete technologies which will not be ready for implementa- tion by 1990.

(5) Delete technologies for which sufficiEnt data are lacking.

(6) Delete technologies already gaining acceptance in general A;)pl1cation of the abo",e criteria resulted in 56 final candi- date te,~hnologiE's which are presented in Table 4.7 and are subse- quencly discussed in Section 4.4. Appendix B.3 contains a complete discussIon of the candidate technology selection process.

table 4.7. Pinal CaDd1clate Tecbaololiu

I. AERODYNAMICS (!>

• Winall!u • Spoilers • Fowler flaps • Low dral surface coatioll • Active lasainar flow control • Leadinl edse devices • Advanced low/medium speed airfoils • Improved stall/spin through aerodynamic tailoring II • AIRCRAFT SYSTEMS (~.> • Anti-icing surtace coatings • AC electrical systems III. CRASHWORTHINESS (4) • Load limiting seats • Energy absorbing floor • Im,rcved restraints • Burst/tear resistant fuel tanks IV. FLIGHT CONTROL SYSTEMS (14) • Fly-by-wire • Fly-by-l ight • Active gust all~viatlon • Active ride smoothing • Active flutter suppression Table 4.7. Final Candidate Technologies (continued)

\

~ IV. FLIGHT CONTROL SYSTEMS (concluded) • Active controls for relaxed inherent stability (CCV) • Integrated yaw damper Integrated low-cost wing leveler • Separate surface technology (SSSA) Direct side-force control • Direct lift control • Single lever thrust/drag control • Fluidic automatic flight control system • Act~'!e stall prev.!ntion V. INFOF~TION SYSTEMS (~) • Digital data links • CRT displays • :ll;D • Micro HUD • Systems status d~~play • Integrated avionics and displays • Fiber optics for data tr3~smission • Laser gyros

•

VI. NAVIGATION CONCEPTS (§) • NAVSTAR/GPS • Inertial navigat~on • Doppler navigation __ ~ri_~~_st t. wt.tt 1 ._ Table 4.7. Final Candidate Technologies (concluded) VI. NAVIGATION CONCEPTS (concluded) • Microwave Landing System CHLS) • Loran C • Omega V!I. ~(~ • Quiet. efficient propeller technology • Low-level pressurization • Ducted propulsors VIII. PROPULSION (7) • GATE engine • Stratified charge rotary combustion en&ine • HCRLB reciprocating engine • Stratified charge reciprocating engine • Advanced diesel engine • Liquid cooling • Improved turbocharging IX. STRUCTURAL ~1ATERIALS (1) • Fiberglass composites • Kevlar composites • Graphite composites .

• t m-hOt-n

--------------~-----

_4M&

~---- .,.,,-~ -----

.. 4f.

4.3 APPLICATION ~~ EVALUATION TECHNIQUE The application of the ET involved three main tasks: the mechanization of the ET, the determination of the relative bene- fits, and the analysis of the ET-generated fiaures of merit. The reader may find Figure 4.1 useful in the following discussion.

Mechanization of the Evaluation Technique Because of the sheer number of calculations required to evaluate 56 technologies in 17 categories for two airplane types, a FORTRAN computer program known as TCHLST was written to compute the figures of merit. Since the considerations of paramount in.terest to any user (input, output, and efficiency) are particu- lar1y significant for the case at hand, input and output were simplified by employing a problem oriented language (POL) known as SCAN, while efficiency was maintained by using arrays, array pointers, and packed arrays. While SCAN is available from the University of Illinois for several mainframes, its limited use together with other peculiarities built into TCHLST which are system-dependent (word size and system software) have resulted in a decision not to include the program listing with this report.

The program capabilities, however, are briefly discussed here in order to support the resulting data.

!nput datl, if not managed properly, can quickly invalidate results. Hence SCAN was em?loyec ~o allow data inputs in the form: -----~=-~==~;.,..

Me m AIRPLANE A AERODYNAMICS 'WINGLETS' .5 0 0 .5 0 -2 0 .5 0 0 -1 .5 0 0 0 0 0 , SPOILERS' - • 5 0 2 1 0 1 0 1 0 • 5 0 1 0 2 0 0 0

I

I and incorporate commands such as PRINT INPUT PRINT TABULAR CATEGORIES OMIT CRASHWORTHINESS PRINT BARCHART ALL RESET OMIT RESET PRINT, etc.

in order to (1) allow easy data entry, (2) allow accurate input verification, (3) allow quick, accurate data analysis, and (4) provide for s~nsitivity analyses under different category weight- ings. Efficiency was maintained by array management where, for example, different addresses were packed into the left and right half of a word size (36 bits). Hence, although not portable, the program was very efficient. with costs on the order of $3 to compute. sort, format, and output original data as well as per- turbed data (which could be demanded in real time).

4.3.2 Determination of Relative Benefits In assigning relative benp.fits, three factors were found to be of great importance. They are the relative benefit scale, the technology baseline. and the technology application and inte- gration.

4.3.2.1 Relative Benefit Scale. The scale of relative benefits (-b < b < b ) must be broad enough to allow adequate ij max - - max differentiation between technology impacts but narrow enough that guessing is avoided. This is a strong function of the amount of information available for a given technology so in general, an optimum scale for one technology will not be optimum for anothe~

I

technology. A workable compromise was found to be -3 ~ b ~ 3 ij

I

where increments of + 1 are normally used but, if justified, increments of + .5 are allowed.

4.3.2.2 Technology Baselines. A technology baseline must be established before any relative benefits are obtained because by definition. a relative benefit compares a candidate technology to a current technology baseline. Since different classes of air- planes utilize different technolo~~~~. separate technology base- lines were established for the six passenger and commuter airplanes.

These baselines are presented in Table 4.8.

4.3.2.3 Technology Application and Integration. In evaluating a given technology, a decision must be made as to how the ..

technology will be used and to what extent other airplane para- meters will change. For example, spoilers do not offer much (if any) advantage over ailerons unless the freed trailing edge is used for additional flap span. Even then, one must decide if ~·"""--~~-·"·"'--"-~---''''''''-----'''·-_J:_-_ ... ______ 'Za· , •• ____ -..., Table 4.8. Technology Baselines I

* ** Airplane A Airplane B

• Reciprocating engine • Turboprop • Aluminum structure • Aluminum structure

"I

!

• Limited flush riveting and/or • Extensive flush riveting bonding and/or bonding • Conventional avionics with • Conventional avionics with VOR/DME RNAV but no auto- autopilot but no RNAV pilot • Plain or single-slotted flaps • Single-slotted Fowler flaps • Conventional controls and • Conventional controls and control surfaces control surfaces ;

i

• No anti-icing • Anti-icing in the form of hot air or boots

* Airplane A - 6 Passenger Business/Personal Airplane

**

Airplane B - 19 Passenger Commute~ Airplar.e takeoff and landing performance will be improved with a fixed wing loading, or if an improv~ment in cruise efficiency will be sought by increasing wing loading. Similar technology application and integration effects are exhibited by many technologies, especially in the areas of composite structures and advanced propulsion.

After reviewing the performance characteristics of current six g' tm' c passenger and commuter airplanes. it was decided that technologies offering either cruise or takeoff and landing performance improve- ments would be used to benefit cruise since the current airplanes seemed to have adequate takeoff and landing performance.

~.3.3 Sensitivity ~ Relative Benefits Uncertainty in generating relative benefits is unavoidable and therefore must be accounted for. Using the guidelines of Section A.3.2, three sets of relative benefits were generated for each technology and are termed "pessimistic." "likely," and "optimis- tic" (PLO) relative benefits. Their use and definition are analogoub to the pessimistic, most likely, and optimistic time estimates frequently used in CPM and PERT analvses. Likewise, a Beta distribution is assumed such that an "expected" relative benefit is obtained as given by Equation 4.2 (Ref. 219).

b + 4b + b L b _ P 0 (4.2) E 6 To avoid unnecessary computations, only the PLO relative bene- fits ~re fou~d and the expected figures of melit (FM ) are E obtained directly from the PLO figures of merit. It is easy to

•

show that the expected figure of merit is given by Equation 4.3.

FU of 4~ + FM

fl~ = -L 6 0

(4.3) The PLO relative benefits were gener~ted by an iterative process. First, information on each of the technologies was ob- tained from the project data base; Second, the PLO relative benefits for each technology were estimated. Finally, the PLO relative benefits were reviewed and updated as required. The PLO relative benefits are preltminary in that they are used only to test the , evaluation procedure.

Figure 4.2 presents a sample of the pessimistic, likely, optimistic, and expected (PLOE) figures of merit where the expected figures of merit were computed with equation 4.3. (Complete results may be found in Appendix 8.4.) This figure illustrates the characteristics of consistency and uncertainty which are dis- cussed below.

4.3.3.1 Consistency of Relative Benefits. To obtain a meaningful ranking of the technologies, each ~ust be evaluated in a similar manner with the same level of objectivity. DOing so results in a consistent set of likely relative benefits. While it is impossible to illsure absolute consistency, the PLOE figures of me~it provide information which allows on~ to check for reasonable consistency.

This is done by comparing the likely and expected figures of merit for each technology. If a technology exhibits a large difference in these figures of merit, it implies that the likely relative bene- fit was itself either optimistic <F'Mz. > ~) or pessimistic

(F't\ < ~) and needs to be re':ined. The data of Figure 4.5 is

typical and shows that the preliminary likely relative benefits

140 AIRPLANE A

• Optlm Istlc

• likely

• Expected

• Pesslm Istlc

~100

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Figure 4.2. Sample of the Pessimistic, Likely, Optimistic, and Expected Figure of Merit are quite consistent in that the likely and expected figures of merit are nearly identical for most technologies. The small incon- sistencies exhibited by a few technologies ~~re kept in mind when the final relative benefits were determined.

4.3.3.2 Uncertainty of aelative Benefits. It was hoped that the PLOE approach. in addition to providing a check on consistency.

would give a measure of the uncertainty involved in the evalua- tion of ea~h technology.

For example. examining Fig. ~.2 one might conclude that the degree of uncertainty in the evaluation of a technology is directly related to the difference in its optimistic and pessimis- tic figures of merit. This has not proved to be a legitimate assumption because technologies which impact nearly all of the categories exhibit large changes in figure of merit for small changes in the relative benefits which results in an exagger!ted perception of uncertainty.

The initial t.::sting of the evaluation techdque (PLOE stun1es) proved that the procedures established for evaluating the tech- nologies worked well and resulted in ccnsistent 8nalyses. The follow- ing sections address tht! final evaluatir.m of the candidate techno- logies in detail.

4.4 OBSERVATIONS ON THE !1TILIZATION OF THE r;VALUATION .TECHNIQUE This section briefly discusses some of the more salient points involving the use and interpretation of the evaluation tec.htdque.

Specific items addressed include: em • - if (1) the effect ~f qualitative values of different raters in aaa1gn- ing relative benefits to tecbno1oliea, (2) the atabi1ity of figurea of merit under different conditione, (3) the orthogonality of categories, (4) the effect of varying the weighting for Empty Weight, and (5) the Significance of figure of merit scores.

4.4.1 Q~a1itative Values ~ Raters As the evaluation technique was applied, tt became evident thkt different raters had a different range of figures of merit resulting from different relative benefits applied to each of the technologies. However, an examination of the several results indicated remarkable consistency between the relative rankings of technologie'l. It wa& thert!fore decided to u.se one rater and emphasize the refinement of relative benefits generated by the rater rather than to standardize the results of several raters at the expense of exhaustive refinement.

~ 4.4.2 Stability of Figures of Merit I After several data sets were analyzed. it became evident t~at the figures of merit: (1) exhibited consistency in ranking technologies as was

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I predicted by the PLOE studjes dhcussed previously in Section 4.3.3.

(2) could change significantly if its ant:/clpated application and integration changed (Section 4.3.2.3) regardless of whether

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thia change reeulted fro. a techDololY breakthrouah or a more innovative (le •• cons .. rvative) application.

In l~ght of (2) above, every effort has been made to define the baaeline and the perceived application of each technology aa .hown in Section 4.5.

4.4.3 Orthogonality 2! Categories

Th~ orthogonality of categories and their weighting. as developed by Surveys 1 and 2 and discussed in Section. 4.1.2 and 4.1.3 was investigated by performing a factor analysis of the rela- tive benefit matrix (56 x 17) for both Airplanes A and B. The Biomedical Computer Programs (BHDP) statistical computer package as described in Reference 218 and available on-line at the Univer-

I

sity of Kansas was used to generate the factor data.

The ~ntent of this study was to determine the possibility of grouping the 17 categories into a smaller, orthogonal set of com- ponents. If this could be done (and it was strongly suspected that such would be the case), then the possibility of reducing the number of categories into a smaller and more managea~le set needed to be investigated (see Section 4.1.2).

BMDP results for both Airplane A and Airplane B indicated that six factors would explain a significant amount of the variance in both relative benefit data sets (74% for Airplane A and 76% for Airplane B). Furthermore. the first three factoIS explained ~O% and 51% of the data variance for Airplane A dnd B respectively.

These three factors included 11 of the 17 categories for both 6f airplanes where ten of the 11 cateaorie. were coalOn to both.

Th! .ix factors are shown in Table 4.9.

An examination of this table, however, quickly identifie.

• extreaA difficulti.. 10 iapleaentlna an evaluation technique con- sisting of only six ortho80~1 component.. Specifically, the loss of differentiation between vehicle performance .... ur .. ents would defeat the purpo&e of the technology evalu~tlon. For example, benefits in cruise speed could not be differentiated from those accorded to empty weight. Likewise, improvements in noise control could not be differentiated from takeoff and landing performance. Although purchase price is gIouped with DOC and reliability for Airplane B. it is statistically tied to static comfort and crashworthiness fOT Airplane A. Yet, purchase price received a score other. than zero in 51 of 56 technologies for Airplane A and 44 of 51 technologies for Airplane B.

Armed with this information, it was decided that the 17 ~ate gories developed by Survey 1 ~rovided a better measurement of a technology's i~pact than th~ six factors resulting from the factor enalysls.

4.4.4 Effect of ~ Weight A concern expressed early in the developoent of the evaluation technique dealt with the quantification of ~ategory weightings through the Delphi Method. Since a major benefit of the method is to allow r~spondents to preserve their views without coercion from othl~rs, a decision to halt further rounds of the surveys is often L@@ XW44-:_ - -- Table 4.9. Six Pactor. for Airplue. A aDd B Airplane B Airplane A Categories Factor Categorie.

Factor 1 bnge Fuel Efficiency Cnti.e Speed Cailinl Ceiling Cruiae Speed Fuel Efficiency Ranle Empty Weight Empty Weilht 2 Exterior Noise 2 Exterior Noile Interior Noise Interior Noise Takeoff/Landing Takeoff /Land.lnl Performance Perfonaance 3 Reliability 3 DOC DOC Relinbility Emission. Purchase Price 4 Static Comfort 4 Safety Crashworthiness Pilot Workload Purchase Price 5 Safety 5 Crashworthlnes8 Pilot Workload Static Comfort 14 Ride Quality 6 Emissions Ride Quality made based on the stability of responses. This implies that the use of an average rating may be misleading if a large dispersion of responses is noted. (Note that "stability" is implied by a large numt-~~r oi unchanged responses from one round to tht: next, and does n~( ~ply that ratings have converged about a mean.) Such was the case for the Empty Weight category for Airplane B. As shown in the histograms in Appendix B.2., these ratings ranged almost uniformly from 0 to 9.5. Hence, results reflecting Fapty Weights of 0, 5.657.

and 9.5 (low. meall, hi) for Airplane B are included in Appen~ix B.S.

4.4.5 Significance!!! riaure !!! ~ kor ..

Before proceeelial to the actual ruu1ta of the evaluatioa technique. oae shou14 be .ware of the followilll .epeeta of the figures of meLit: (1) Values near Bero illPly 110 aip1f1cant i1Iprov8ll8llt ill effect on the user. For example. an illproveM1lt ill ranae _y be offset by high coat or maintenance.

(2) Positive numbers reflect net benefits to the user widle negative numbers reflect net peoalites to the user.

(3) Macroscopic figures of merit (~) are fairly stable. where FMi - 50 signals greater benefit from technology i than FM - 30 does for technology j. However, PMi - 50 is diffi- j cult to differentiate from say, ~ - 45. A major exception to this rule is noted in (4) below.

(4) Those technologies impacting many categories have higher figures of merit than those which affect only a few. Hence, the four technologies listed under the crashworthiness group have scores which do not truly measure their benefit to the user. This is due to the fact. that although they all received maximum scores (+3) for Crashworthiness, they did not impact DOC, Safety, Reliability, etc. (Also, Crashworthines8 received relatively low scores as shown in Tables 4.3 and 4.4.)

(5) The figures of merit reflect relative benefits as opposed to benefit/risk. Hence, although the pro~ulsion group reflects particularly high benefits to the user, the extremely high risk of development to the manufacturer is not accounted for.

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4.5 EVALUATION TECHNIqUE UlULTI Tbis .ec:ti01l discu .... the fiul fip •• of .. r1t obta1aed wh.n Equation 4.1 .e .pplhd to the 56 liMl c .... i4.t. t.chnolo- 11ea (PCT) identified 1n T.bl. 4.1. The 17 cat.lory _tahtt u.ed h.re are tho.e ahoWD 1a T.ble 4.3 and .r. ave rase we1sbts. T.cb- oolOIY IrOupi1l,e were •• tabliebed for convenience and have no bearinl on the relative benefits or the ftaurea of merit.

Each technology Iroup ie first discussed in leneral terms followed by a tabular rankina of those technololiea whicb were considered to be in the particular Ir~up. ~ brief discussion of each technololY then follows.

4.5.1 OVerall Technology Ranking Tab1es4.10 and 4.11 show overall rankinls for the 56 and 51 technologies considered for Airplanes A and B respectively.

Both lists are identical except for iive propulsi~n technololies which were not considered appropriate for Airplane B.

4.5.2 Aerodynamics Nine technologies were evaluated and are shown in Table 4.12.

As shown in Table 4.10, the first three technologies under Airplane A were among the top seven oZ 56 technologies. Winglets, when considered for a new wing, offer little b~nefit to the user of eith2r airplane. Spoilers (implying full-span Fowler Flaps) appear rather attractive for Airplane B as shown in Table 4.11.

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Tabl. 4.12 4erod~~c.

Flaur.. of ~rlt for Techno1olY (9) Airplane A Airplane B 56/-53/85* 51/-57/58*

Natural Laminar Plow Airfoila sa 25

Spoilers 57 38 Fowler Flap. 48 24 Low/Medium Speed Airfoil. 29 31 Leading Edge Device. 24 29 Improved Stall/Spin -- Aerodynamic Tailoring 24 15 Activp. Laminar Flow Control 4 -46 Wing1et.

- 3 - 4 Low Drag Surface Coatings -14 -15 It Total number of technologies/lowest PM/highest PM 4.5.2.1 Winglets. The advantages of winglets are not yet clear.

Although they promise to reduce induced drag while increasing wing bending moments when applied a •• modification to existing aircraft, there is little data to suggest that they would offer improvements to aerodynamic efficiency that could not be gained through increased aspect ratio and wing twist. Difficulties involving flutter an~ large sideslip angles require further study. Reference 156 describes theoretical improvements in range of 6% to 8% over a similar wing without winglet.. Although \ I

t

\

thia article citea larae atabili_iDa .a.ent. in ,.. at aidaalip

I

all.lee of 6 to 8 , the addition of vinalat. dao raau1ted ill • w.akly diveraent dutch roll aDda.

I

j The baaalin. a.aillet which thia teebDoloay va. evaluated for both Airplane A aDd Airplane B va8 a hiah-a.pect-ratio. twi.ted

I

wina. Small benefite in cruiee performance were offaat by waiaht

penalties to compensate for aubatential increaaea in root bendiDa

i

i moments.

4.5.2.2 SR9ilera. Spoilers alone offer only small advanteaes ~ , to general aviation aircraft by el~inating adverBe yaw. However,

J

they offer the promise of major improvements in performance by freeing the ent~re wing trailing edge for high lift device. (full span flaps). As mentioned previously in Section 4.3.2, this addi- tional maximum lift capability can be used to either ~prove takeoff and landing performance or, at the other extreme, improve cruise performance by increasing wing loading (W/S). Since in- creased wing loading offers substantial benefits in fuel effi- ciency and ride qualities with only a small penalty (if any) in cruise ceiling, this technology is evaluated from the standpoint

of increased wIs. It should be noted that since typical single

engine aircraft operate at w/s = 730 to 1220 N/m while commuters

•

operate at w/s = 2390 to 2870 N/m , Airplane A will reflect larger

~pacts from this technology (W/S for heavy transports are typically on the order of 4790 to 6700 N/m ).

Although cruise performance promises to reflect substantial gaIns, the incorporation of spoihra doE'S have attendant riaks.

loaliD .. r coatrol r .. ,.... .IUUi ....... , cOlltrol Hftr.u (clu. to

I

flov r .. ttact.Mlt). 1acr .... of authority with full epa flap.

deployed, au4 10 •• of control authority at aeaati.- ADIl" of

\

, attack are 'but a f. whicb er. 1cl_tUied in the literature • • poiler. i. attributabla not ao 8Uch to the lack of .poiler haacl- book data, but rather to the vaet a.ount of aileron haDclbook data available (Ief. 121).

Still unavaUable 18 a verified thiD-ai rfoU _thod for pre- dictina 2D characteristica of a wina-spoiler-slotted-flap confiaura- tion. However, Parkiu.on baa reported on relearch ulina a numeri- cal thick airfoil _thod utilizina a 2-aource .odel to predict c t and c va. a for an airfoil with a spoiler and alotted flap mo (Ref. 159).

Furthermore, Reference 111 reports favorably on fliaht test results of spoiler characteristics on a modified Cessna 177 Car- dinal (Redhawk program) and Reference 110 does likewise for apoilers on a modified PA34-200 Piper Seneca (ATLIT program). Reference 183 provides an overview of the state of the art of general aviation spoilers in 1974.

4.5.2.3 Fowler Flaps. Fowler flap. offe~ significan~ improve- ments in maximum lift coefficient (e ) when compared to the plain, L max split, or single slotted flaps which are typical of the 8ing1e engine general aviation (GA) fleet. However, their real benefits are more fully exploited when uBed a. full .pan fiapi in conjunction with apoUe~.. AlthcNab tlaia 14u 18 DOt ... (Wenacu 10, 122, 224).

dUficulUe ••• eocute4 wiCb iatea~.t_ the .po1l.~. have Mla,..

the v1de.p~ead u .. of thb CODCept.

Paul101l (Ref. 161) ~apo~tI aD iDu.... in '" of 961 __

.ax Fowler flap. were fully deflected o. a wiDd tUIUMl lDOCIel wiaa. 1ft a comparison between plain, elotted, aDd Fowler fla .. _ where the

plaift and slotted flap. had flap apana aptlroxi_te1y 501 ,.r the

wing semi-span while the Fowler flaps had flap spans a~orox~t.ly 75% of the willi leai-epaD, Paulsoft (Ief. 162) report a ~ C for L ux the Fowler flap approximat.e1y 48% Ireater than the plain or elotted flap. Under the conditions of thb latter investilaUon, the Fowler·· flapped configuration could either land at 8 speed 18% lower than its counterpart or reduce its wing area by 33%. Wentz (Ref. 224) reports a C • 3 for a one-quarter scale ATLIT wing equipped with L max full span Fowler flap •• For the evaluation at hand, Airplane A is presumed to have plain or linlle slotted half apan flapi and Airplane B il preau.ed to be equipped with half span Fowler flapl. The anticipated effects of full span Fowler flaps il reflected in the evaluation.

Hence, this technology is considered to be interlrated with lpoilera.

~.5.2.4 Low DraB Surface Coatinas. This technology is described in Reference 22, where an inve.tigation w~s conducted by the ~e1ns Commeri~al Airplane Company of nine liquid coatings and 60 filml adhesive Iyatems. !hil inveltilation is part of the Energy Efficient Transport (EET) element of the Aircraft Energy Efficiency (ACE!)

proara ad •• perf01:'Mll UDder coatract to lIASl-LlDale, ..... ch Cent.r. Tbr .. l1qutd eo&t1llp .... four fi1ll/ed_lY .. ar. cur- rntly UDdarao1Da further tutiq aDd evaluatioD.

MOtivation for this r .... rch va. aupplied by the 1973 fu.l crisis, aDd preUainary r .... rch by NASA of • T-33 v1Da reflected • 12X reduct1.oD in draa when .kin joint., hilll. 11 .• a. etc. were covered with a tIIIOOth thiD fila. Boe1q'a r .... rch iDd1cat ...

however. tbat only a 1.6% reductton ill draa coefficient (~) CaD be expected due to the exiatance of unaealed lap. 101' hilh lift devices and eGntrol aurfacea. St1ll, thie 1.6% reduction traDa- lates to a fuel savinls of 128,690 1 (34,000 aals) per Boeina 727 per year.

Analy.1a of the data reveals that although almort all of the drag associated with rouahness 10 elta1nated from wina and tails, only 25% of the draa associated with saps was eltainated. Hence.

application of a low dral surface cnatins to a wins with protruding rivets makes little sense. Prom a ai •• i~n profile consideration, Airplane A could benefit more from this technolosy than Airplane B because its mi.sion is cruise dominated whereas Airplane B's i.

characterized by more time in climbs an/ approaches where drag- due-to-llft is usually more slsnificant than profile draa.

However, even if one considers an advanced airplane charac- terized by flush rivets, bonded surfaces, or composite construction, the propulsion system will ir,variab1y include a propeller. and system efficiencies will dictate lower cruise speeds closer to L/DIIU tban for a tur1M»jet 01" tur1M»fau CfnIDtel"pftt. • .. 1f._le

"viDa. due to _11 I'HuettOlUl 18 ,..oftla dna an ,nbabl, .on

thaD off .. t by pau1Uea til _taht, 18it1al. coat, .. DOC f. both Airplay A .. I.

4.5.2.5 Active 1MIiur ~ CoDtn1. Thi. tachDo1oay 1a l'aceiYiDI r.aewed iDterut ill 11aht of .oariDa fuel coate. Aa Shev.ll pointe out ln bference 191, ftOthiDa el .. has the potential of reducing draa to a. areat .. extent. AD ...,1e in hla uticle, based on an aircraft operatiaa at leyno1d. aa.ber. of 20 to 70 x 10 with 200 paaaeoaers and Grana. of SSOO uutlca1 aile., illustrate.

hi. point. If thls vehicl. operated with 751 of It. wiaa and tail surfac.s effectively laminarized, dra; would be reduced by 30% and LID would increa •• by 30%. For a at-flar type of vehicle ~.8, 200 pa •• eaaere. JT9D engin •• ), Maddalon point. out that l_inar flow control (LFC) energy requireaents d.cr.a.e as AI incr .....

above 7 (lef. 129). For this type of air carrier operation, DOC decreased d.spite a 17% penalty in aaintenence CO.tl snd a 3% penalty in purchase price.

Although LFC offera a far greater potential for parasitic draa reduc~lun than low drag surface coatings, it atill proai ••• lignificant penalties for Airplane A and Airplane B. As with all systeas which reduce parssitic drag, the benefit. realized by a propeller-powered airplane may be expected to be leas than for a jet or fan powered vehicle. Also, fuel costa repr~.ent a aignitil'antly ... 11er perC,:Htage of oc,c for Alr.,lane A, atld the

I

mi.sion profile of Airplane B do.a not allow it to take full advantage of the syata. The low altitude enviroaaent of Airplane B and the a .. ociated hi,her susceptibility to inaacu. duat. aDC!

other debris aUlleat hi,her matnteaance coat. for Ai~lane B than for a high-altitude cruiae vehicle.

It should be reiterated at this point that one factor whicb this evaluation proceas significantly does not measure is risk.

For the case at hand. an airplane optimized for LFC operation from the beginning of the design process ~lll have markedly different characteristics than a non-LFC counterpart. Degradation or the LFC system promises significant (although yet unquantified) penalites in performance and operating costs.

4.5.2.6 Leading Edge Device~. These devices. which are found on almost every cont~porary commercial jet liner, are only rarely founrl in GA aircraft. Fixed devices offer the same low pe~for- manee advantages as retractable Krueger flaps or slats. However, crui~e performance is degraded in exchange for the simplicity of a fixed system. Reference 111 presents some flight test results of a Krueger system installed on the Redhawk. Although they displayed excellent characteristics (.~CL ·.59 for the Krueger flaps with max Fowler flaps deployed 40°), it was suggested that such a dystem ..

was probably too heavy and too complex for a It6ht GA single engine aircraft.

In exchange for the in~~~ase in C ,w/s could be increased L max ~ven further to lmpt-ove cruise performance. However, these benefits would he offset by a probable degradation in ceiling.

An interesting application of a retract.bl. leading edge device would be in their use on the outboard section of the aft wing of a canard configured vehicle. Such an appllcation could offset the effects of increased u~ash from the canard during low speed high angle-of-attack operations.

4.5.2.7 Advanced Low and Medium Speed Airfoils. Tbese airfoils are characterized by turbulent flow co~iition8 and were initially designed as low speed airfoils to provide low cruise drag, high climb lift-to-drag ratios, high maximum lift, and well behaved stall characteristics. Results of these efforts are documented by McGhee, Beasley and Whitcomb in Reference 139 and are shown in Figures 4.3 and 4.4, where the L5(l) series ~ncompasses both the GA(W)-l and GA(W)-2. As shown, these low speed airfoils have better c~ than more conventional NACA sections and max tIc ~ .13 produced the highest c~ Reference 139 also describes max work done to reduce pitching moments of a 17% thick airfoil as well as increase the lift-to-drag ratio of a 21% thick airfoil. These low speed airfoils all have a design c~ of 0.4. Two mEdium speed airfoils are described where c1 a .3, Re • 14 X 10 , and M • .72 and .68. The design family of airfoils is shown in Figure 4.5.

..

These airfOils, howevet, are characterized by higher pitching moments and higher drag than, for example, 6-series airfoils.

Accordingly, an effort by Hicks and Schairer to increase the maxi- mum lift coeffici~nt of 63 -215 airft:l is described in Reference 84. Figure 4.6 shows the modification and Figure 4.7 shows one of ;;s SY!l&<J?l% k . \¥40 several fiaure. deplct1nl typical re.ult.. Thl ... y b. compared to Figure 4.8 from aeference 139 which ,howe low and medium .peed airfoil data.

Airfoils Smooth M • 0.15 6 Re • 6 x 10 2.4

o NASA LSu)-SER IES IC d == 0. 40)

t

o NACA 23O-SER IES '

o NACA 44-SER IES

2.2 - Do NACA 24-SER IES ~ NACA 65-SERIES IC d= 0.40) t , 2.0 C l,max 1. 8 1.6 1. 4 - 1.- __ ,..10.
____ 1 ..... _~---L __ ...II __ --.L..I __ ...--1 1.2 . 04 . 08 . 12 . 16 . 20 . 24 . 28 tIc (frolT". Ref. 139) Fi~. 4.1. c~ va tic of NASA Low-Speed and NACA Airfoil!..

max M • 0.15 6

2.4

Re • 4 x 10

2.0

1.6

1.2

lid c = 1. 0

z

-

C = 0.40

-

z

I

1 I I

I

.24

. 12 . 16 .20

.08

tic

(from Ref. 139) Fig. 4.4. Effect of tIc on c and ~/d for NASA Low-Speed t Airfoils. max One difficulty which is posed for the thinner airfoils (tIc::: .13) is that these airfoils do not lend themselves to easy configuration integration in GA aircraft. Thicker airfoils <tIc:::: .17) offer more wing vohlme for fuel and control linkages. Also, they suffer a smaller weight penalty in construction.

- b __ , ............ ,,_ ,..-~.=-....~-..........'_~~ _________ ._ .. ___ ,, _______ • __ .. ___ _ lOW SPEED MEDIUM SPEED ~===== lS(lH1409

C GAIW)-2 -====-==--

LS( 1)-0413 ~-:---..

~ GAIW)~~_~ 0417 ~ lS( ~l-0417 REDUCED MOMENT

G~~-~~

lS(l)-0421 INCREASED LIFT! DRAG (from Ref. 139) I Fig. 4.5. Design Family of NASA Low-Speed Airfoils.

--- NACA 63 .215 ---- MOD B .:~.

•

y/c I I I I

: - 1 I

0 .1 .2 .3 .4 .6 .6 .7 .8 .9 1.0 x/c (from Ref. 84) Fig. 4.0. Hodification to NACA 63 -715 Airfoil.

________ d_t_. __ M __ . _____ tt ____________ ._. ________ ~ __ .~ _________ ~ ...

'"

u u I

•

o 1WlA-.",.

Ct a_, ..

,4, • .2 1.1 l I 'I --'----, I , ·,4 .. 0 • 11 'Ie 21. ,D1 ,02 .. ,tt ,w, ... "W .... c;. ea...

R - 2.5 x 10 M • 0.2 e (from Ref. 84) Fig. 4.7.

Experim~ntal Results of a Modification to an NACA 63 -215 Airfoil.

Difficulties with implementation of the LS(l) series into a con- figuration are suggested by the highly aft-loaded airfoil pressure distribution and suggest that a moderate amount of tailoring may be required at the wing-body junction to preclude premature flow separation of both the wing and fuselage in this region.

For the technology evaluation, it was assumed that neither Airplane A or B employed low Or medium speed airfoils. Airplane B

$ .-

should suffer a smaller penalty in paraaitic drag since its mission profile 1s more dominated by climbs and approaches where drag-due- to-lift is more dominant than in cruise • • R • 4 x 10 e (x/d • 0.075 T o lS(})-(M17

o MSm-m17

2.0 1.6 1.2 .8 C z .4 -.4 I I I -.8 .03 .04 .05 em Cd (from Ref. 139) Fig. 4.8. Experimental Results of Low and Medium Speed Airfoilri.

4.5.2.8 Advanced Natural Laminar Flow Airfoils. Present efforts to develop a natural laminar airfoil are aimed at achieving a cR.

max at least as great as that associated with the GA(W) -2 while main- taining pitching moment coefficients lower than those of the same airfoil. Preliminary work suggests sectIonal reductions in drag , b t by a. much as 25% at desian Ct' Improvement. in performance are offset by a requirement to maintain a amooth surfac •• and relativ.ly sev.re penalti •• in drag appear with the introduction of roughD •••• Thia suggest a that fluah riveting, bonding. or compo.ite materiala may be required to eliminate protruding rivet h.ads. Also, the user may expect to be confronted with a requirement to maintain a relatively insect-free surfac •• This technology proved fairly difficult to evaluate. Airplane A c~uld benefit rather significantly from this technology, but purchase cost penalties could outweigh benefits if a baseline of present manufacturing processes is upgraded to total flush riveting. For purposes of this evaluation, it will be assumed that this technology will be integrated with a composite wing since the latter offers reduced manufacturing costs. In the absence of an insect-free sur- face technology, Airplane B will encounter DOC expenses and reliability penalties associated with its more severe operating environment.

4.5.2.9 Improved Stall/Spin Characteristics Through Aerodynamic Tailoring. Solutions to stall/spin accidents range from better pilot stall recognition training to better designed vehicles. This evaluation addresses only the latter solution (in a rather restricted sense.

Anderson notes that solutions to the design problem appear to lie in (1) providing good handling qualities up to and beyond maximum lift, (2) making the aircraft spin resistant, and (3) pro- viding stable static and dynamic stability characteristics with « good con~rol about the pitch, roll, aDd yaw axe. (Ref. 4). Three approaches which appear feasible are: (1) Control stall proareesion on the wina to provide Ireater post- " etall roll dupina and natural buffet stall wamina.

(2) Ltmit lonaitudinal control power to prevent complete wing stall.

(3) Minimize adverse cross couplinl by automatic (SAS) means.

Item 2 is addressed by attempts to reduce the relative hori- zontal tail power throulh (1) tailoring, (2) active controls (Ref. 29) or (3) employing a canard configuration (Ref. 185). Item 3 has already been successfully demonstrated by the military. Certainly, separate surface technology as discussed elsewhere in this paper should be investigated. Artificial stall warning devices appear to provide significant contributions, and Ellis notes that a large portion of stall, mush, and spin accidents occur in airplanes which are not equipped with these systems (Ref. 49). Stick shakers appear to be th~ most effective warning device. Aural warnings also provide a certain margin of safety, but visual-only angle-of- attack indicators appear ineffective.

Having discussed certain problems and solutions which appear in the literature, emphasis is now focused on methods of controlling ..

stall progression through aerodynamic tailoring. Feistal, Anderson, and Kroeger have reported on plomising results based on a leading edge discontinuity as illustrated in Figure 4.9 (Ref. 8S).

Figure 4.10 depicts an essentially flat-topped lift curve out to

<. • e

(L. H. SHOWN) SEGMENT NO'S • 5 4 3 2

•

~~~,.o~'~~I"~~~~~~'~~'~~~~

I I I I I i

WING _I_...J _ -1. _.1... _I_...L - J.... C TIP

to'" IN GLOVE 6

HERE FOR DATA "' SHOWN IN THIS !!

~ __________________ ~, __ ~ ______________ ~ m PAPER) 0

84" ------.J 21"~. a"

(213 em) --"'1(53 em,r- (110 em) CROSSECTION OF DROOPfO LEADING EDGE GLOVE USED APPROXIMATE GAW·1 UPPER SURFACE (from Ret . S. '}) Fig . 4.9. L ea ding Eug lodifi ation for tall/Spin All viation an angl - f -atta ch of 32° . Si g nifi ca ntly , rollin g mem r.t excursions s we ll as y win mom nt w re acc pLabl throu hout the ra nge of an 1 -of- tta k t s t d. 41cks nd H nn hav r port d on anum rical optimization t hniqu wh r low p d wing (GA(W)-2 sec tion) n ~-~ ----~- ------------- ,

\

could theoretically be modified to move the initial It all to the wina root without affecUns either Ipan lo.d1nS or iDduced draa

I

charecteriitici (Ref. 83).

I

I

, For this Itudy, only the techDololiel reprelentinl efforts to i control separation are evaluated. Both Airplane A and B achieve

J

!

substantial benefits in safety at little coat.

!

Full Scale Tunnel, Tail Off, V - 89 kt ...

, ..

• .a A I ! I • II • •

• •

....

• •

. ...

(a) Aircraft with modified (b) Basic aircraft.

leading edge.

f:l:'om Ref. 55 Fiq. 4.10. Flat-Topped Lift Curve from Leading Edgp. Modification for Stall/Spin Alleviation.

~.5.3 Aircraft Systems Only twu systems as shown in Table 4.12 were evaluated. Sur-

•

prisingly, Icephobic Surface Coatings did not fare as exceptionally i well as one might expect. Here, miDCir penalties in weight (which t!

affected ceiling and range to a lesser extent) and cost almost t~tally negated the +3 relative benefit awarded for sGfety. Airplane B appears to benefit substantially when pneumatic boots or hot air systems are replaced.

\- • -r ~----~~--~~~~~----~-- .......... ~ ......... - ..... ~ ........ ----------------- ....... -- Table 4.13. Aircraft S,.t ...

F1aure. of Merit for TechnololY (2)

Airplane t

Airplane I

56/-53/85 51/-57/58 Icephobic Surface eoatings 5 33 AC Electrical Syst(~. 0 17

*

Total number of technololies/lowest FM/highest FM 4.5.3.1 Icephobic Surface Coatings. Very little was found 1.n the literature on icephobic surface coatings. Although Reference 18 mentions olefin plastics on helicopter blades, no mention is made of the success or failure of this research. Certainly, the rewards obtainable f~om a surface coating of this type demand further research. When one consideres that vertical and horizontal tails, wheel pants. struts, etc., could all be protected from ice accumu- 1ation in addition to the more conventional and limited GA applic- cation to wing leading edges only, the potential benefits of such a system become staggering. It is suspected that in-house, ua- published rese3r~h has already been performed by interested agencies. However, no information relating to efforts or results

•

surfaced in the data gathering trips performed as part of this research.

For this technology evaluation. the baseline for Airplane A is not considered tl) be equipped with an ant:l-icing system, whHe the baseline for Airplune B is presumed to be equipped with a pneumatic system con.i&ting of either boots or ducted hot air.

rt Airplane A is penaliaed for -iaht aACl co.t but raceiv .. e. hiper benefit in aafety. On the other had. Airplane B racetv .. an l~prQveaent in reliability, coat. and weiaht with only ... 11 tmproveaenta in aafety.

4.5.3.2 !£. Electrical Syat .. s. Major reaearch activity in ahifUna from DC to AC electrical syat .. a va. not detected either 1n the literQture or during dMta gathering tripa. However, at leaat one manufacturer cited a potential for large we1,ht savings in their aircr.ft product l~ne.

For the r!'e5ent, neither Aircraft A',. nor Aircraft B's tech- nology basel! ~~ is expecled to benefit to any great extent. Airplane A is characterized by Um.it~d weight penalties aaeociated with wirin~. while Aircraft B is leno",-!! t' favor panel-mcunted va remote- mounted avionics due to cost conaiderations (Ref. 167). Although newer C.:llllllluters are uaually equipped with autopilou and flight director systems, it is suspected that tre majority of operational commuters are still hand flown on cYnv~ntional round dial instru- ments with cross ?oj~ters. Hence, the potential weight savings are not as great as for a cunent business jet.

4.5.4 Crashworthiness rour technologies as shown in Table ~.14 vere investigated.

TIaesl' technologies are considered to he extremely attractive. As explained earlier, their figures of meri.t are relatively low because they affect only a few categories and specifically do not influence DOC. Safety. or Reliability. Also. as shown in Table 4.3 and 4.4.

the catelory vellhtlal for Cra.bworthtDe •• va. 6.816 <'11 of 17) for Airplane A aDd 7.333 (19 of 17) for Airplane B.

T.ble 4.14. Cra.hworthine •• Flauree of Medt for Technology (4) Airplane A

Airplane I

56/-53/85* 511-57/58 Load Limitjnl Seat. 17 J.9 Energy Abeorbins Floor 14 15 Improved Reltrainta 14 lS Burst/Tear Resi.tant Tauk. 8 9

*

Total number of technologies/lowest FH/high-st FH 4,5.4.1 Load Limiting Seats. Efforts to improve occupant .urvi- vabiltty in the event of a crash are becoming more refined wtth the corre1at!ons of tp.st data obtained from the NASA Impact Dynamics Research Facility (IDRF) and the FAA Civil Aeromedical Institute (CAM!) with analytical data from a modified <omputer algorithm known as MSmtLA (for Modified Seat-Occupant Model for Light Aircraft). Fasanella and Alfaro-Bou note good agreement between CAMI sled data clnd fllll scale results from the lOR!:'.

f1S0MLA (which USCJ a spring-dampeT medel a8 opposed to the finite element model h, ~(MLA) alse showed good agreement (Ref. 53).

Having noted that there is much which can be done towards .::4proving survivability through improved seat design, three different types of improved seats .a .hown in Figures 4.11 (b). (c). and (d) are being investigated. The wire bending loa d limitp.r is shown in Figure 4.11 (a).

() Wlr b ndin g 10 d l imi r (b) C ilin g u sp nd d at . t . ,

-

;; , I t ..

t ., 1 Jt ~ Ol I •• , • " 1: 1 I{ ( ) (d) r m un d s t k r c i n s · t n . . 11 . L d Limi in Cone p u " During stroking, the wire loop ia tranalated aloDS the wire under a ~'oustant forca of 4.4 kN (1000 lbf). T.ats as of April 1979 indicate that ~he ceiling mounted aeat reduced longlttidunal pelvic accelerations by 40% during a 15 m/s (50 fps) sled pulse and ver- tical accelerations by 50% during a 13 m/s (42 fps) sled pulse (Ref. 53). This seat. which has a 9.1 kS mass, suffers from added installation weight and a poasible 10s8 of ~troking distance resulting from cabin deformation at impact. From an appljcations/integration viewpoint, it appears that the floor mounted seat (10 kg or 23 Ibm) may offer more promise. However, preliminary test results indicated a requirement for further development work for the floor mounted and rocker action seats.

Thomson aud Goetz (Ref. 210) note in a similar report that the human tolerance of 25 g's can be met with a 30 em (12 in) stroke dissipated over 0.1 sec. Such a system places an upper bound of 12.2 mls (40 fps) on the seats.

Since the typical general aviation seat of 11 kg (25 lb) dissipates energy through seat deformation and leg buckling, load limiting seats offer great potential for improving vehicle crashworthiness and occupant survivability.

4.5.4.2 Energy Absorbing Floor. Several schemes which offer pro- mise in application are described in Reference 210 and depicted in Figure 4.12. Assuming a 15 cm (6 in) available stroke (through floor deformation), Thomson and Goetz point to an upper limit of 8.2 m/s (27 fps) which can be dissipated by the floor structure through controlled collapse. Despite information obtained from ~.

c' .~ airframe manufacturers which indicates that this type of occupant protection is already being introduced in a few product lines, wide- spread incorporation is not yet noted . Teats and data correlation with analytical models indicate that three computer codes (KRASH.

ACTION and DYCAST) demon~trate good quantitative results in pr - dictin dynamic vehicle response to impact loadings (Ref. 79). The continued development of these tools promises to give further in- sight to floor and cabin deformation. Fasanella and Alfaro-Sou report, for xample, that IDRF data showed a GA aircraft cabin during a 27 mls (89 fps) impact d forming and become 21 em (8.3 in) wid r od 23 cm (9 in) low r. The appar nt uph av ul of the center fl or s tion was actually th r sult 0 th cabin side walls forc- in th f loor down t h wall Junction. Incorporation of the energy ab orbin f loor into futur ircraft tak 5 on add d si g nificanc in vi w f:; u h r e nt t st d ta.

BEFORE a

IMPA C T~

Am R L:;;;;).

IMPA CT ~ C ORRUGATED LON G lTUDI AL CO RRUuA1[D SUBFLOOR CYLINDERS + FO RMABU CO RRUGATrD WEB WEB . &EADED INTERIOR FOAM K£EL WEB (SANDWICH) BULKH[AD .

AND / OR WITH fOAM NOT C HE 0 CO RNE R lXTlRIOR FO AM WI :H FOAM Fi~ . . 12 . En · t -·\' bs t ' bln ~ Fl • r r: n c'pts .

Both Airplane A and B will gain significantly in craabworthi- ness from thia technology at an undetermined coat in weight.

(Note that all of the five concept. depicted in Figure 4.12 incor- porate an energy-dissipating foam.)

4.5.4.3 Improved Restraints. Wben an ana1ysi& of business air- craft accident investigations showed that the use of sho~lder belts reflected a marked increase in survival, the researchers concluded that this use is considered to be the single most impor- tant means of improving occupant crash impact protection (Ref. 194).

This report also noted that restraints attached to the seat pro- vided little protection when the seat failed and separated from lhe floor. Further research by others indicates that a single shoulder strap allows the occupant to rollout of the restraint in some accidents. Hence, a restraining system consisting of two shoulder straps, two lap belts, and crotch strap, all connected at a single point centered on the lap belts appears to offer major improvements in crashworthiness. (Fig. 4.l3(a) depicts such a system.) Such a system would incorporate locking inertia reels in the shoulder harness and thereby afford the passenger more freedom of movement during normal operations than a diagonal single shoulder strap would. Tests to date indicate that a tensile force of .54 kN (122 lb) may be encountered in the harness system during a 27 m/sec (89 fps) impact. Hence, metal-to-metal buckles are preferred to metal-to-fabric cam systems.

This technology promises significant benefits to both Airplane A and B. AntIcipated penalties lie in increased costs.

...

~._tg * (a) Impro;_d Restraint (b) Conventional Restraint Fig, 4,] 3. Fi v straint vs Conv ntional Restraint - , 4 , 4 Ef orts t d v lop a 4 , r ~ w n coura d by of shIt 'ar 1 ta nt fu 1 tank r r p- lop d und r th lie HO\v v r , th t ok d v rm y h ..

n iv (or g ne r 1 viation Arm\' pr r rlS r t 0 h vy 3nd xp , tank In Ref. that a 0 lit r Ii 11 ptf'r lkq~ 1 [n mpt t imp1i( th o -t on .11 $1 , OCJU n tt m" v >r1 n, nd t i n limin tl'd ri s y. t'm , tall' t1 \"a r it' In w m nufa til d .. nd t nk \J i t h rnn lbl r 1- , : 11 v ,

-

catapult. Reference 48 hotes that the lightest tank which per- formed sa(isfactorily with no leakage after a 29 mls (96 fps} impact was one manufacture~ by Uniroyal Corporation from a single ply with a fabric weight of 43.23 gram/m • None of the 2- and 3-ply tanks failed . bu:: two other (lighter) I-ply tanks and the ori- ginal tank did. Figure 4.14 shows the installed tank, and Figures 4.15 and 4.16 show the crash site and test where the test vehicle impacted an eart~en hill equipped with sunken steel tubes and rock piles. Figure 4.17 shows the external damage done to the wing where the single ply tank did not fail. (Note that a ~ystem to shear the fuel line from the tank was incorporated, and the frangible fittings did not leak.)

•

Fig . 4 . 14 . ~ dified Fu _l Tank I ns t alle d i n Wi ng .

. . " .; (~ : 10 0 I · \" . \ . I " • ~ • ~ " .

o Fi . 4.1. Airfram on C tapult with Ea rth n Hill, St 1 Tub , nd R kp i l es \is ibl >.

Si n ~ nnual ./\ a id . nt s v ' r th t n y ar p dod b twn ,100 with po t ra h fir s 1 ) a nd 1 76 numb r d fr m 4, 200 t th pot n- r ' p rt J in pproxjmat 1 3 a 0 th acid nt annu 11 ir r f t o ff r d b th s tanks tial f r i mI r ov d f t y f r G i ~ . i ~ ni i a nt.

..

. tapp ' :-: po di ad 'unta g fo r b th i r p lan ' A.n 1\. Th ' l-I l y t t t3nk \. l ' h <:tl .4 k . (lIb) in lutijn . 61b) . In ri g ina l tank \. 1 h ~ ' . 4 k t h.? it L l n J ' wh 11' t h I s ) p ' r t . nk, pr du t ion • 1 \'l lu m' p' nalL > lit r O .

1 1 tanks may be expected to have a volume difference of less than 3.8 liters (1 gal) from conventional bladder tanks. These weight and volume pe na1ites are expected to be greater when compared to a wet wing con~ept. However, a wet wing may be expected to suffer from post crash impact failures at least as often as conventional bladder tanks.

Fi . 4 . 1. Tvpl 1 rm pa t f irfram Fo llow i ng Ca t ap ult Ac l e r atio n.

10 ..

Fig . 4 . 17. Win Dama Wh r a I - Ply T~ nk Did 4 . 5. onLro .!. S ystems Fourtt' n tcchnolo gi as sh ".'0 i n T bl 4 . 15 w valua t ed for n s Ai rp1a A nd IL H r e , tho.; i n t gr t d , l ow - cost ..

appear a Lr btl do s n t c t iv for Ai p1 n Ta b1 4.10) app r 5 for i p la n B. rhe rna or dif e r nc or h r ti n gs [ lit! 5 in h' hat Airp1un t\ is n quipp d with a n ' uto- is . n pilot whil B w'h simU r sc n rio i off r d (or S par a e Su rf c ' s j t s b n' to Pilo t Work10 d and Rid nd J) C p Ou li y f th initi 1 c SL nalt i s . Airplan B, ur iG IN AL PAGE IS OF P( R QUALITYi on the other hand. suffers no penalty in costs (autopilot replacement) and smaller benefits in Pilot Workload and Ride Quality.

Table 4.15. Flight Control Syst ...

Figures of Harit for Technology (14) Airplane A Airplane B 56/-53/85* 51/-57/58* 39 8 Integrated Low Cost Wf.ng Leveler Separate Surface Technology 22 24 13 16 Active Stall Prevention Lift Control 11 12 Direct Integrated Yaw Damper 3 -0 5 Act Ctls for Relaxed Stability -3 Single Lever Thrust/Drag Control -3 Controls -12 -1 Fluidic Automatic Flight Active Ride Smoothing -14 Side Force Control -14 -11 Direct Active Flutter Suppression -3e -38 -42 -35 Active Gust Alleviation Fly-by Wire -47 -33 -47 -30 Fly-by-Light

* Total number of technologies/lowest FM/highest FM

-- __ WM~·_··~ _______________ • _____ =* __ ·w ... ·~e .. ~~~ ____ ~ .............. ____________________________ __ i 1- , I Active controls, fly-by-wire, and fly-by-light are extremely unattractive due to coat, where the flight environment con.idered for Airplane A and B (low cruise speed) does not take full advantage of the benefite available to aircraft with high cruise speeds.

4.5.5.1 F1Y-!l-Wire. ~is technology is considered to be fairly well developed but very costly. Papers presented at an AGARD Conference in 1974 (Ref. 2~ and 57) noted then that the methods of designing a fly-by-wire (FBW) system were well understood but that s&fety considerations alone (without regard for certification requirements) cause the systems to be very complex and expensive.

The economic justification for such a system is very much dependent on the aircraft mission. Hence it is difficult to visualize their incorporation into the GA fleet witnin the next ten years, especially when one considers tile availability and cost of Iliaintenance for such systems. Some points regarding such a system deserve mention, h~wever. For p.xample, tha integratioa of FBW is considered manda- tory if and when active fl!ght controls prove economically justi- fiQbl~. Such flig~t c.cntrols include those designed for rel~~ed longitudinal static st3bility. gust alleviati~ot and flutter su}pres- sion. All of these technologies promise weight savings and improved operational efficiencies which lower direct operating costs (DOC) for conunerc~.al air transport. An intri.guing r?tiearch program which is curr~ntly being pursued by the NASA/Langley Research Center together wit, Princeton University involves Langley's Di.gital AvioniLf' Rese~rr.h syster .. (DARE), and Princeton's Avion1:.& ] O~l Research Air:raft (AlA). The AlA is a fully inatru.ented. five- d.lree~f -freedom, FBW, lov-Vinl, &inal. eDline GA aircraft. Thia vehicle, which hal already b.en uaed to 1nv.atilate flyinl quali- ti.s, humaD. factora, and control, will be used with DARE to expand on the above investigations as well as thos. dealing with advanced di,ital control concepts (Ref. 45).

4.5.5.2 FIY-!l-Llght. This technology complements the capabilities and possible active controls applicr-tiona of a FBW system because of the following characteristics attributable to fiber optics: (1) Quicker data transmission than possible with wires.

(2) Large bandwidth capability offers the potential for replacing several wires with one fiber.

(3) Fibers are non-inductive and non-conductive.

(4) Provides better signal isolation than wires by decreasing "crosstalk."

(5) Does not present either electrical or fire hazards.

However, fly-by-light (FBL) systems represent a somewhat Bignificant cost penalty over FBW systems. As ~uch, their integra- tion into the general aviation fleet appears even more remote than FBW ~ystems. At the same time, it shoulj be noted that Bell ..

Helicopte~ has logged several hundred flight test hours on their Model 206 equipped with a fiber optic yaw SeAS wH~h includes an optical encoder, ~iber optic link, and an optical receiver aud decoder. Bell also employed five complete fiber optic systems in the control of the 8~ shplate of an iron bird configuration

--

ne:t«m: • .~ ...

demonstrator. In this confiauration. failura of thrae syst ...

still allowed control authority over the main rotor.

4.5.5.3 Active ~ Alleviation. Thi. technoloay could offer

-

.

potentially large performance benefits resulting fro. reduced structural weiaht for those high 8pe.~ aircraft which operate out of short fields on the order of 610 m (2000 ft). These aircraft will require relatively low wing loadings on the order of 1.9 kN/m (40 Ib/ft ). and structural weight penalties may be incurred when decreased field length capabilities are sought. Active gust alleviation holds promise for configuratio'ns subjected to gust load factors on the order of five or above. When integrated to reduce structural weight (and strength), however, the system becomes safety-of-flight critical and requires costly redundancies and certification testing. A unique passive system for Single engine GA aircraft is described in Reference 173 where auxiliary aerodynamic surfaces were attached to a Cessna 172 to sense angle-of-attack and drive the flap system through a direct linkage.

Although this syqtem successfully attenuated gusts up to 3 mls (11: tPF'J in the frequen~y range betwe.en that of the phugoid and short-per1o~ mode, it did so at the expense of reducing C • m a Consequently. further btt;'~:;"'O; where a linkage to ,.~

•

troIs can be investigated appear warranted. An b;;~.-,,~ (.h,L ~,!'iHt i.,.

be made here is that gust alleviation hecomes attractive when significant structural weight savings can be realized. Sinc~ these savings USt1311y represent the elin.1nation of penalties incurred through strenJo:th ceql!irt-III~I.ta ... :hich are dictated by ..

aclvene pa.t load I ill cruile ., IYlt_ or techDololY which t 1ncrea.e. ~ of a ViDI (aDd thereby allow. for hiaher ViS) vill 1UX probably be IIOre COlt .ffecUV8 for tbe low-.peed GA fl .. t.

Active SUit allaviatiOD i. not expected to offer qUADtifiable benefitl to eitber Airplane A or B due to tbe relatively low cruile lpeed. (low SUit factors) of tbe.e vebi,~le •• 4.5.5.4 ~tive Ride s.oothinl. rbi. technology i, differentiated from that of active SUit alleviation in the lenle that this 1,1- tem doe I not p~ovide for reduced Itructural weilht.. AI luch it is not safety-of-flight critical.

Primary benefits result from improved handlinl qualities and ride comfort. Reference 34 IUllests that additional cost penal- ties for a cOUlDereial transport would be approximately 2% to 5% of total testing and certification expenses of a new vehicle with an integrated system. Retrofit of the system would incur higher cost penalties. Feadbility studies using a deHavu'land DH6 have been performed and indicate that total system weight should not exceed 2% of the aircraft's gross weight while total power requirements would not exceed 0.3% of total engine power.

However, an. effective ride smoothing system will require relatively large direct lift and side force surfaces located near the aircraft center-of-gravity (Ref. 34). No major reliability or maintenance problems are forseen.

For the present evaluation, both Aircraft A and B w11l incur C08t penalties in exchange for improvements in ride and handling qualities.

"~ 4.5.5.5 Activ. Flutter Control. Thi. techDoloay'. attractive- nes. i. degendent laraely on the aDOunt of weight which could be .aved at the expense of tor.ional .tiffn.... Wben abort haul

."

airplan •• with their attendant low W/S and high Ai are con.idered, advocate. of ~~t1ve control point to wtng structural weight .avina.

on the order of 40% attributable directly to au.t alleviation system.. T!l1s weight .avi1,lg •• however, may reduce torsional stiffness and flutter spe~d, and mandate active flutter control systems.

GA aircraft, on the other hand, already operate at low W/S and relatively low cruise speeds. As AR is increased, ~ore atten- tion will have to be focused on problems associated with aeroelas- tic phenomena such as flutter. Active flutter control is attrac- tive when otherwise realizable structural weight savings result in an unacceptable degradation of torsional stiffness.

I

For the prescnt evaluation. both Airplane A anA Bare pre- sumed to possess suffici~nt torsional stiffness to negate a require- ment for flutter suppression. As wIth active gust alleviation, this sy~tem is safety-of-flight critical since its anticipated use would be to reduce structural weights by n; .clng aerodynami(.

loads. It must be realized that the evaluation ratings ~ill chan!-:c l'unsiderably if .111 unanticipated increace in torsional stiffness is requ lred as a result of increases in a ... peet ratio.

4.5.5.6 Active Controls for Relaxed Inherent Stability. This concept holds promise for reduced DOC through reduced fuel con- sunption. Like all active cun~rol systems, however, it poses Fl9 very M.gh penalites in terms of front end (purchase) coata. Un- like guat alleviation and flutter suppresaion, where aerod~ic efficiencies result from decreased structural weights. this technology offera a direct payoff by allowing for reduced drag resulting from relaxed (natural) stability constraints. This is a:hieved by sizing the tail to meet control constraints instead of stability constraints and usually results in smaller tails with dec'_~sed (parasitic) downloads when the main (forward) wing is moVt-:~ fryr-.iud. Results from studies of medium and heavy C011lD8r- cial transports were very promising. Reference 107 showed that in the cade where takeoff gross weight w~s kept constant, either payload was increased by 15% Or range was increased by 20%.

Reference 144 noted that in ~he case where the mission (payload and range) waR kept constant, relaxed static statility could yield a 10% reduction in gross weight and a 5% increase in cruise LID. Reference 118 notes that a study involvbg the NASA Jetstar airplane, where active gust alleviation and relaxed ntatic stability were investigated, showed tail Rurface areas reduced oy 40% and fuel consumption reduced by 21%. This latter example should be interpreted with caution, however, since th~ effect of the gust alleviation system was to allow for a substantial decrease in wing sweep with an increase in AR form 5.3 to 9, which in itself offered a significant improvement in LID.

4.5.5.7 Integrated Yaw Damper. Yaw dampers typically are employed on high altitude, high speed airplanes, although they also may be applied effectively for yaw damping at low speeds in general d t C' sa Aviation aircraft. Typically, they may be employed whenever a poorly damped dutch roll mode attributable to low values of the yaw damping derivative C are encountered. The integration of nr .0 this type of syst~ is difficult with present autopilots, how- ever, because yaw damper functions are typically fed back to the GA pilot through the controls as annoying distractions, usually dUL'ilg the approach-to-landing phase. This may lead pilots to turn the system off at a time when they need it the most. Hence, a system integrated through separate surface technology where the pilot receives no feedback and system failures do not result in critical situations appears attra~tive. Yaw dampers will allow aircraft vertical control surfaces (rudders) to be optimally sized for P~L~4r~lar flight conditions while retaining superior handl- ing qualities over th~ aircraft's entire flight envelope.

4.5.5.8 Integrated Low Cost Wing Leveler. Bergey notes that an automatic and low cost wing leveler that does not depend on auxi- liary power would be extremely valuable to the GA cOIlll1unity (Ref. 18).

While investigating inflight airframe failures for the period 1966-1975, Staple ford (Ref. 197) notes that a lack of spiral 8tability is probably a key f~ctor in determining the frequency of ..

loss of control. Two flight test programs (one by the FAA using a Beech Debonair A-33 and one by NASA using a Mooney H20) clearly demonstrated the serious problens encountereri by a non- instrument-rated pilot who ventures into IFR conditions. Staple- ford goes on to note that the absence of any airframe failures _~44g4 Q by the Mooney H2O during the ten years investigated and encompassing over six million flight hours points clearly to the benefits of a wing leveler. (The H2O was equiVped with a wiDI 1..,.1.1" as standard equipment.) The attractiveness of a wing leveler, then, is predicated on ~nhancing spiral stability and reducing the potential for airframe failures resulting from recovery procedures (or lack of them) from unusual attitudes. Under more favorable flight conditions, a wing leveler still offers significant bene- fits in reduced pilot workload.

At present, there is a low cost (~ $100) fluidic system avail- able to the GA community. Also, a separate surface system retro- fitted to a Cessna 172 has been satisfactorily demonstrated (Ref. 176). For the evaluation at hand, data on forecast auto- pilot use appears to indicate that half the aircraft in Airplane A's category will have autopilots, and almost all new commuters (Airplane B) will be so equipped. Hence, it will be arbitrarily assumed that Airplane A does not have a basic autopilot and that Airplane B does.

4.5.5.9 Separate Surface Jechnology. This technology has parti- cular significance to the GA community due to;the nature of controls (reversible) incorporated by most single engine and commuter airplanes. In such systems, autopilot functions are mechanized by tying a servomotor into the primary surface cable controls. Hence, all autopilot functions are fed back to the pilot and, with the system off, stability augmentation capabilities which ~-- ___ . __ " ________ d&~" __ ~ ______ ~ ______ ~ __ ~ ______________ _ ;a4¥A ~ . ',;;¥Af _ i2A&€44F" '-W!!f.4 _C .. _ Ii 0 ,&.

4 - ;~wm .. ia. £ '1

may be incorporated into the autopilot a~e lost. Separate surface technology, on the other hand, can provide continuous wing leveler and yaw damper functions with appropriate wash out circuits incorpora~ed 60 as not to interfere with pilot control, and hence remain totally transparent to the pil.;)t. An autopilot function could also be incorporated which, too, would be trans- parent. Such a system (SSSA) has been flight tested in a Cessna 172 in a wing leveler mode and a Beech U99 in a three axis autopilot attitude command mode with excellent results. Costs of such a system are expl~ted to be similar to present auto- pilots, with improved safety. diird-over failures result in the pilot flying the aircraft in an out-of-trim condition with no require- ment to override an autopilot servomotor since one is not tied to the primary cable system. References 175, 176, 181, and 182 provide more details on SSSA, and Reference 20 notes that the development of samaTian cobalt motors makes SSSA dppear even more attractive.

It should be noted that SSSA has recently been incorporated into the GA neet in the form of a yaw damper on the Hitsubishi Diamond I. As with the wing leveler evaluation. Airplane A is consid~red to not have an autopilot while Airplane B does.

4.5.5.10 Direct Side Force Control.

The primnry application of this type of system would be to augment active ride smoothing systems. Lapins and Jacobson (Ref. 119) noted that side force controllers are more effective than rudders alone in alleviating the effects of turbulence through active ride smoothing systems.

j • "!IP'·"-'~ ..

This same requirement was noted by Conner and Thompson in Refer- ence 85. When used for ride smoothing. such a.ystea will not be considered safety-of-fllght critical. Still, heavy cost penalties are expected for both Airplane A and B, although Airplane B should benefit considerably through greater passenger satisfaction in what promises to be a more turbulent flight environment.

4.5.5.11 Direct Lift Control. Severs1 applications of direct lift eontrol (DLe) are envisioned, and all Kre mechanized through spoilers. In th.~ lIu'1"e sophisticated systems, DLC will be required for ride smoothing in order lO offset vertical loads. In some large commercial aircraft, DLC may be required for ~dequate flight

path control as wls is increased. When such a system was employed

in the Redhawk (modified Cessna 172), more precise, easier, and apparently safer approaches resulted since flight path angles could be controlled without changing aircraft attitude. Like- wise, landing ground roll can be reduced substantially in those cases where the approach path is constrained by obstacles. This would allow descents beyond the obstacle without a requirement to lower the aircraft nose and increase airspeed. Also, when the spoilers are maintained in a partially deployed configuration, a more conventional approach may be flown at hip-her airspeeds without incurring a landing performance penalty. This latter application promises improved safety since a greater airspeed stal1 lnargin can be maintained.

4.5.5.12 Single Leve~ Thrust/Drag Co~!tro1. The use of spoilers to control flight path angle during approach to landing is known to improve the landing performance of moat pilots. When such.

system 1- mechanized through the throttle. pilot workload is reduced both during the approach and during any required go- arounds. Improvements in reduced pilot workload ,'ere noted even in the CAse where the throttle lever retains its individual func- tion but has the DLC spoiler control mounted on it as .!1 thumb wheel (as was done on the Redhawk), It should be mentioned that single lever thrust control alone offers promising advantages in reduced workload, Teledyne Con- tinental Motors has incorporated a governor into a single lever control which effectively combines the functions of throttle, mixture, and RPM. This system, although not yet in production, has undergone over three years of development and is presumed ready for production. In its present configuration, the throttle is used to set RPM, and the governor mechanization acts to control manifold pressure and fuel flow.

4.5.5.13 Fluidic Automatic Flight Control Systems. The fluidic three axis system designed for NASA by Honeywell and subsequently flight tested in an Aero Commander 680 FP showed excellent relia- bility and functioned very similarly to conventional autopilot systems (Ref. 222). Altitude hold, however, was degraded above 1,830 m and power re.covery of the fluidic servo amplifiers was only 40%. Yet, systems such as th~se demonstrate a clear capaLility for further developemnt. The continuing decrease in the cost of microprocessor logic, however, has done much to offset the initial _ ¥I 444!44!!;:;g:nJ? # advantages of lower coat and high reliability attributed to fluidic ayatema. On the other hand, developments in low-coat fluidic sensors auch aa the low apeed a1~apeed indicator, the vortex rate aensor, and fluidic atall aenaor, offer significant benefits at reduced coat. One of the most advantageoua a~st~4s to result from efforts in fluidica was a winf leveler which incorporated a laminar flow proportional fluid amplifier with a very high signa\-to-noise ratio that vas developed by NASA Langley personnel.

4.5.5.14 Active Stall Pr~vention. This technology offers great potential for reducing stall/spin accidents. Work by Chevalier where a spoiler is added to the lower· surface of the horizontal tail to prevent the attainment of the stall angle-of-attack (Ref. 29) appears encouraging. Since it appears that a signifi- cant number of stall-related accidents occur in airplanes which are not euqipped with artificial stall warning systems (Ref.

49), the neeJ for an effective stall warning/prevention device appears well justified. Several warning concepts (stick shaker, audible horn, visual angle of attack indicator) are technologically mature. Of these, the stick shaker appears to be the most effec- tive. The possibility of incorporating a stick pusher also appears attractive.

One detraction to active prevention systems lies in their accurate operation. Although the incorporation of active preven- tion in light GA single engine aircraft appears warranted, the question of degraded control authotity below the stall angle-of-attack detracts from the otherwise undisputed utility of such a systea.

Cost appear A to be another possible detraction. In order to achieve widespread acceptance, such a aystem must either be mandated or offered at very low cost.

3.5.6 Information Systems.

Eight technologies were evaluated under this grouping and are shown in Table 4.16. As illustrated, Digital Data Llnks, Integrated Avionics and Displays, and Systems Status Displays are all fairly attractive technologies. The Micro HUn, although presently reflecting no benefit to the user, deserves mention here, This system, discussed further in Section 4.5.6.5, promises to provide all HUD functions without either a CRT or the associated optical system. Presently, its main penalty lies in cost.

Widespread use of such a system could dr1.ve costs down significantly and make this technology extremely attractive.

4.5.6.1 Digital Data Links. This concept entails the communca- tion of data (which is presently rendered verbally) between the airplane and various agencies in a digitized format much as present data links are used for encoding altimeters. Three types of systems are addressed in this section. The Digital Data

•

Broadcast System (DDBS) is intended to provide RNAV systems Wit~l infonlation required to satisfactorily navigate both preplanned direct routes and "whatf!ver charted routes which are retained as an integral part of the ultimate area navigation environment" (Ref 97). Data would be broadcast in repeating data streams for . <t1

--~--------~---- -'~- --------- ---..,.--.------------..- ......... - ..... - -

specified station or route coverage and would be accessed by tuniua in the appropriate VORTAC frequency. It should be ~~te6 that this system would marry RNAV to the exist ina VORTAC route structure iustead of providing the more flexible option of inputting geographic coordinates. The present status of this system i. unknown, and the last time it appeared in the literature used in the present research was in 1976 (Ref. 97).

Table 4.16. Information Systems Figures of Merit for Technology (8) Airplane B Airplane A 56/-53/85* 51/-57/58* Digital Data Links 38 Integrated Avionics and Displays 32 42 25 25 Systems Status Displays CRT Displays 11 21 2 2 Micro HUD Laser Gyros -16 -14 Heads up Displays -23 -24 Fiber Optics (Data Trns) -23 -26

* Total number of technologies/lowest FM/highest FM

The second system to be discussed here is the Discrete Address Beacon System (DABS), which formed a significant aspect of the upgraded third generation ATC system (UG3RD) as discussed by the DOT Air Traffic Control Advisory Connnittee in December 1969.

This system has continuously appeared in the literature and rdceive.

further attention in a report prepared a. part of the 1979 Summer Faculty Fellowship Program in Engineering System. Desian (Ref.

148). As discussed in that reference. DABS will a~ccmmod.te the following information between the aircraft and its ATC environment: (1) Clearances.

(2) Runway surface winds to include wind shear and wake vortex information.

(3) Weather information.

(4) Minimum safe altitude warning.

(5) Confirmation of assigned altitude.

(6) Automated Terminal information Service (ATIS).

(7) Runway Visual Range (RVR).

(8) Holding instructions.

(9) Approach and departure clearance.

(10) Conflict alert and resolution instructions.

(11) Instructions as tc proper heading, speed, altitude, and the time t~ execute the ATC instructions.

The third system La :'e discusbied is the Automatic Conununica- tion and Reporting System (ACARS), which is used in conjunction with existing \11F udio equipment and allows for both voice as well as digitized information cOIllnunication to enhance a~_r-ground operational control communications. Although t~is type of system has been available h.r ten years, it has not been impletllented as rapidly as hoped due primarily to its $5,000 price tag (Ref. 148).

r me It appeara aipif1c&11t at this time to po1t\t out that studte.

of future ATe enviroaments point to a conge.tad environment as a very real possibility, and this .ullesta that the pilot could well become .aturated with communications function. alone. At.

NASA-sponsored Avionics and Controls Research and Technology Workahis;. (Ref. 147). the team diacuss1ng "General AViation and Short Haul" recommended that a principal focus for research is suggested to: (1) Minimize or eliminate the requirement for communication (talk) to as great an extent as possible.

(2) Allow IFR flights of the future to be performed as easily as VFR flights are today.

In vjew of this observation/recommendation, DABS appears to offer the most significant benefits of the three systems discussed.

It should be noted that DABS is crucial to the Intermittent Posi- tive Control (IPC) conception for Collision Aviodance Systems (CAS).

A basic DABS unit may be expected to cost about the same amount as present transponders ($750) or $2000 with an IPC display (Ref. 97).

For the present evaluation, both airplanes are presumed to already be equipped with encoding altimeters but are penalized fllr ~L\e IPC display. Substantial benefits in improved pilot workload and safety are realized.

4.5.6.2 fathode Ray Tube (CRT) Displays. CRT displays may be evaluated against mecuanical, conventional instruments, or against other electronic displays. For purposes of this evaluation, they • #e will be evaluated agaiDlt the former becauae other electronic systems are not envision~J to be competitive froa a coat atand- point.

·' Still, some discuasion appears warranted for various electronic display methods. The following data of 1975 vintage ia condenaed from Reference 196 and shown here a8 Table 4.17.

Table 4.17. 1975 CharactEristics of Some Electronic Displays System $/Character Volts/Panel CRT < 1 10.000-25.000 Plasma Panel 100 < 3 LED < 4 LC • < 6 The display application also needs to be addressed and is summarized in Table 4.18 in terms of their 1979 status.

The versatile capabilities and low cost of the CRT tend to more than offset its disadvantage of relative size and high voltage.

althoughthe latter can pose a safety problem. A flat panel CRT.

developed by Northrop for the Army and whose rights were sub- sequently sold to Texas Instru~ents. greatly reduces the problem of size but was found to be prohibitively expensive due to high mdnufacturing costs (Ref. 219). Present CRT's rely on a $130 m111iun annual market 1n computer terminals (which is doubling every three years) to offer very low costs. Indeed. one source

- 3_

(Taunas, Ref. 219), aUlae.tm OIM coat. of $9.00 for black aDd white unit. and $36.00 for color in very larae quantitie.. When pre.ent CRT capabilitiea are uaed a. a baaeline, the relative performance of other electronic di.playa in aix major probl ..

area. may be briefly tabulated as in Table 4.19.

Table 4.18. Electronic Display Technology Per.pective Display Application Available In R&D Discretes, Keters, & Electromechanical Liquid Crystal Legends Galvanometers Incandescent Light Emitting Diode Alphanumeric Cathode Ray Tube Chemoluminescent Electromechanical Electrochromic Ir." 'lndescent Electroluminescent Light Emitting Diode Electrophoretic

•

Liquid Crystal Ferroelectric Plasma Cathode Ray 'rube Electroluminescent Vectorgraphic Plasma Panel Light Emit ting Diode Liquid Crystal Electroluminiscent Video Cathode Ray Tube Ferroelectric Laser Liquid Crystal Plasma Panel (data from Reference 219) Armed with this background information, CRT's are evaluated now against their me~t~nical conventional counterparts. These systems offer reduced pilot workload 1n that scanning tasks can be em r.: · - -~ - - ,. - ""l., .' .. '.1.1 : !

, Table 4.19. Six Major Display Problem Areas & Their Impact on New Technologle.

(X designates problem with respect to current CRT) Electro-

I Flat Electco- Liquid Electro-

Light Plasma chroai(.s luminescence Crystal phoretlcs Emitting Discharge CRT Diodes Luminous Efficiency X X ....

N '.

I.-> t-tatrix I X X X Addressing I

I

X X X X Duty Cycle

I 1

I.

Uniformity/ X X X Gray Scale X X !: X X Full Color X , Cost (includ- , X X X X X X X ing Electronics) I ------ (data fro. Refereoce 219) reduced from a 22 em (9 in) radius about the artificial horizon to 6.4 cm (2.54 in). This, of course, implies an integrated avionics package capable of providing continuous systems monitor- ing and, inherently, warning-by-exception. Since the CRT's eliminate mechanical flags, etc. <as many as five or more in an integrated attitude director) reliability will be increased.

(One avionics manufacturer indicated that possible sticking needles and flags on a product ~.on ADI were corrected through the i!leorporation of servomotors to drive the flags.) For those airplanes using remote mounted avionics, substantial weight savings can result by r~ducing the number of signal paths.

Both Airplane A and B should receive significant benefits in reduced pilot workload and safety. (Recall that a display device is ess~ntial to DABS/IPC implementation.)

4.5.6.3 Heads Q£ Displays .(RUD). HUD systems offer the potential for reducing pilot workload during approaches to landing under both VFR and IFR conditions. Currently, they display information such as course gJid~nce, airspeed, angle-of-attack, altitude, etc., via a CRT through a lens system/combining screen which collimates the display on the aircraft windscreen. Thus the pilot may devote hls attention to external cues from his landing environment and still receive flight data without having to re- direct (and re-focus) his attention to within the cockpit.

Associated with this capability, howe~er, is the problem of cost.

HUD systems are envisioned to cost on the order of $25 to $45 thousand f~r GA applications while they may ~~st savera! $100 thousand in mil~tary applications. Other data which tends to detract from HUn implementation are the observations that: (1) At the NASA-sponsored General Aviation Avionics Workshop, no general agreement could be reached on the question of whether HUD studies for GA should be pursued (Ref. 196).

(2) A USAF study concerning the use of HUD's as a primary instru- ment reference in the A-7D, F-15, and F-lllD in 1976 resulted in mixed responses from 123 pilots on the question of using FUn's during approach to landing. Although a significant majority of A-7D and F-15 pilots said the HUD enhanced IFR operations, only a small percentage preferred its use during approaches. Most F-lllD pilots preferr~d not to use the HUD for normal operations where weapons delivery was not involved (Ref. 15). The major complaint of the USAF pilots was that erroneous information could be displayed without warning.

For both Airplane A and B, improvements in pilot workload and safety were offset by penalties in maintenance cost, empty weight, reliability, anc purchase price.

4.5.6.4. t-1icro HUD. A micro HUD. developed by Bell Helicopter ...

Textron, offers HUD capability at an anticiapted 50% of the cost of regular HUD's. Although this cost penalty is stLll significant for GA, the system merits discussion due to its unique imp1ementa- tion. The system uses a micro-processor to drive a fiber optic symbol generator which presently has the capability to generate a large range of dynamic symbols and numbere. (When the signal generator was viewed during July 1979, the microprocessor was not fully imp1eaented and consequently only a limited set of symbols were generated. This generator is illustrated in Figure 4.18.)

The most unique aspect of the system however, lies in the absence of the large and heavy optical system required of conventional HUD's. This system uses a pair of eyeglasses with a minute mirror centered in the eyepiece as shown in Figure 4.19. Signals are transmitted to the glasses via a fiber optic bundle and collimated to display images as shawn in Figure 4.20. Two similar systems are currently being tested by the Army.

The benefits of this system over conventional HUD's lie 1n the elimination of (1) the high voltage and higher power require- ments of a CRT, and (2) the elimination of costly and healY associated optical equipment.

Although the same benefits for the HUD were credited here, severe cost penalties were also awarded on the basis of an estimated $20,000 price tag, which could not be supported by either Airplane A or 8.

4.5.6.5. Systems Status Displays. Comprehensive engine health ~?nitoring is significantly absent in present GA aircraft and, consequently, pilots are confronted with having to derive this information from existing displays (when sufficient raw data is available). The current DAAS system promises to rectify this situation by sensing: , . ~ , .

, .

f __ .

" MICROPROCESSOR SYMBOL GENERATOR rELECTRO MECHANICAL / EXCITER N """' ~\i HORIZON LINE --.J MUL'rrPLE FIBER OPTIC SCANNER Fig. 4.18. Preliminary Fiber Optic Symbol Generator (Operational) Fig. 4.19. Micro HUD with Fiber Optic Link and Mirror.

(1) Nanifold p ress ur e (2) Eng ine RPM (3) Fu el Flow (4) Fuel Quanti t y (5) Oil Temperatu r e (6) Oil Pressure (7) Cy linder h ad tempe r atu r for each cy lind er (8) Exhaus t gas t n. perature fo r ei'\c h cy lind r (9) E xhaust gas oxygen fur a~h cy lind r (10) Cow l f1ao position (11) Auxiliary fuel pumps ; ; ' 1, ;[\ L .\Gr ~ ~ OF P( H 01 "r ".

'- - - - .......

. '

, / S£NSO" I ... PUTI 7 Fig. 4.20. Micro HUD Image Display.

plus numerous other aircraft system parameters. This system will alert the pilot to any critical out-of-tolerance condition and advise corrective action in some cases. It also presents commands to the pilot for setting the mixture lever (Ref. 39). Such enhanced ca~abilities do much to relieve the pilot from scanning and provides warning signals on an exception basis. The incorpora- tion of an annunciator of some sort appears warranted for such a system, and it is noted that Princeton's Avionics Research Airplane (ARA) is so equipped. Reference 31 points to integrated multi- function displays as having the highest priority of ten eqLipment discip4ines for NASA funding/study. This report (done in 1974) identifies a desired or expected purchase price of such displays at $500-$1500 for a single enc.·ne airplane and $3000-$7000 for a turboprop system.

When sufficiently integrated, system status displays promise more advanced warning of possible engine failures, more informa- tion to assist in selecting optimal power settings, and reduced pilot workload. Engine TBO's may have the opportunity for being improved through better engine management on the part of the pilot.

These benefits appear to apply equally to Loth Airplane A and B.

4.5.6.6 Integrated Avionics and Displays. This concept has received considerable interest and publicity over the past decade, due largely to the antieipdted effects of increased dir traffic congestion and the attendant increased pilot workload associated with a more restrictive flight environment, particularly under single-pilot IFR operations. NASA efforts to identify and develop an orderly investigation involving industry, educational institutions. and other agene ies, ar~ reflected in the Langley- sponsored Avionics and Controls Research and Technology Workshop (Ref. 147) and the 1979 Summer Fac-ulty Fellowship Program in Engineering Systems DCti :gn which was cu-sponsored by ASEE (Ref. 148).

~ - --~---~---~---·""d_· ________ e __ ~~

Ames efforts have been centered on the Advanced Avionics Systems studies which recently resulted in a Honeywell/King Radio contract in August 1978 to design and build a Demonstration Advanced Avionics System (DAAS) which will eventually be flight tested in a twin engine aircraft. Highlights of this program include several studies and a workshop as follows: (1) An ATC environment forecast (Ref. 97).

(2) An electronics technology forecast.

(3) A GA advanced avionics workshop (Ref. 196).

(4) Preliminary candidate Advanced Avionics Systems (PCAAS) study contracts to Southern Illinois University (Ref. 133) and Systems Technology Incorporated (Ref. 208).

(5) In-house efforts at Ames investigated I among other things, different low-cost options to solve t~,lIv1gation algorithms and to define a low-cost option to gyro sensors.

Interesting results to number five above involve the incorporation of a phase-locked-loop to enhance VOR scanning capabilities and the incorporation of magnetometers to solve the attitude sensing problem. Preliminary results comparing INS data with derived data using magnetometers is presented in Figure 4.21.

..

Under the present contract, the Honeywell/King DAAS effort will incorporate the following functions (Ref. 39): (1) Automated guidance and navigation using VOR/DHE navigational facilities.

(2) Flight planning.

rte i

--

··-~D

i~~

... , .. , .. 2M ___ _ TIIIII._ AmTUDI -INS --- - - COMI'UTEO ATTITUDE _INI ----COMI'UTEO "

•

~5~--~--~----~--~--~----~--~--~ o ~ ~ ~ ~ ~ ~ ~ ~ TIME ....

Fig. 4.21. Comparison of Pitch. Heading. and Roll Computations.

(3) Weight and balance and performance calculation •• (4) Monitoring and warning.

(5) St~rage of normal and emergency checklists and operational limitations.

Pilot interface with this system i:s through a touch panel which in turn provides inputs to the Integrated Data Control Center (IDCe).

The system is organized around an IEEE 488 bus and includes five microproc~ssors. Conventional pilot displays are incorporated with the exception of two CRT's. One serves as an IDCC interface with the pilot and the other serves as an Electronic Horizontal SlL~ation Indicator (ERSI) with an electronic map capability.

Unfortunately, no cost forecasts for this system were readily available. The STI study (which was dependent on the availibility of a $6000 OMEGA system) penalized an unsophisticated avionics user $4900 for a $13,275 total avionics package and a sophisticated user $13,300 for a $62.645 system. The SIU study, on the other hand. forecast a cost of $23,850 for a full IFR system in a Cessna 402 and $9962 for ~ Single engine new installation. (The STI PCAAS system used co~kpit display formats similar to DAAS but with seven microprocessors, and the SIU system replaced all con- ventional instruments with plasma panel displays.)

It should be mentioned that the GA avionics community is not in total agreement with the single bus architecture, and at least two manufacturers which were v.' sited expressed concern over possible single point failures Hhich would render the entire system either inoperative or unacceptably degraded. On the other haud, the concept of an integrated avionics system with or without the bus structure offers certain advantages Which would otherwise be delayed. For example, the powerplant health monitoring capa- bility of DAAS is seen as a major advance over present syste.a.

Here, the advantage of DAAS is that it allows for more types of information to be displayed to the pilot through the IDCC, where present configurations do not allow for either the required panel display space or separate display costs to be efficiently absorbed.

A forseeable spin-off from a system which incorporates micro- processor logic systems lies in the further introduction of these systems into GA aircraft. Certainly, one of the most beneficial applications of such logic systems lies in the electronic control of propulsion fuel systems to allow for more efficient operation.

For example, ajvanced positive displacement engines will almost invariably require a high-pressure fuel injection system where the fuel is injected directly to each cylinder. Such a system will benefit substantially from an electronic control which senses induction air density and modulates as well as times the fuel flow. (This is to be contrasted to low-pressure continuous-flow ..

systems in operation today.) Since the possibility of failure of such an electronic fuel injection system may be unacceptably high, the system could be mechanically driven and only modulated by an electronic logic circuit in order to achieve the highest -.<.--..'--'*- .. _______ ___ ~_~ __ t .... rt _____ ... Ds ______ , ____ _.~ __ I.~-- efUciencies. The fact that DMS already inco"y-,ratu a propulalon

\

-~ t monitoring fun ~ion (no con:rol) could lead to a reduction of in- flight en~ine failures by providing advance warning or indication

t

of impending failure.

4.5.6.1 ~ Optics for Data Transmission. Fiber optics have

I

the capability to: (1) Transmit data faster than by wires.

I

(2) Replace several wires with one fiber since fibers have a larger bandwidth capability.

I

(1) Function without the inductive or conductive properties of wires.

I

, (4) Retain bette~ signal isolation between fibers than possible between wires.

(5) Function without introducing electrical and/or fire hazards.

(Ref. 158.)

However, no appreciable weight savings or benefits may be anti- cipated for retrofit systems, especially where data transmission rates are lower than one megabyte/ sec.

Hence, fiber optic data transmission systems are not envisioned to offer advantages significant enough to justify their cost 1n the GA fl~et.

•

4.5.6.8 Laser Gyros. These systems, developed by Honeywell, con- tain no moving parts and sense accelerations by measuring fringe shifts between two laser beams. Reliability is expected to be very high. but costs are expected to be equally high. Hence. thp.ir incorporation into the GA fleet is not envisioned at this time.

4.5.7 Havilation Concepta Six technololiea as ahown in Table 4.20 were evaluated in this technolOIY group. Microwave Landini Syst ... appear more attractive to Airplane B than ~ dn. to a larler number of In terminal operations. Of the three area navilation conCC?tI inves- tilated. none appeared to provide significant bunefits. This was due largely to the fact that both airplanes were presumed to already have full IFR capability. and Airplane A was also equipped with an existing RNAV system. The ~wo velocity informa- tion sources (Doppler and Inertial systems) were both extremely unattractive to Airplanes A and B.

Table 4.20. Navigation Concepts Figures of Merit for Technology (6) Airplane A Airplane B 56/-53/85* 51/-57/58* Hicrowave Landing Systems 9 NAVSTAR/GPS 8 8 Loran C -24 -30 Omega -26 -32 Doppler Navigation -37 -40 Inertial Navigation -53 -57

*

Total number of technologies/lowest FM/highest FM = 4.5.7.1 NAVSTAl/GPS. Th1s system 1s a global, aatellite-based navigation .ystea beinl pursued by the U.S. Air Force. Six of the originally-planned 24 aatellites are already in orbit (Ref. 148), but Reference 8 indicates that the total nuaber has been reduced to 18, with. planned operational date of 1987 to 1990. This system is ext~emely attractive to civil aviation because it provides a (forecast) reasonably priced navigation ~y8tem capable of (1) Extreme accuracies in x, y. z and t reference frames.

(2) Area navigation.

(3) Global coverage down to ground level.

(4) Collision avoidance.

When one considers that g general aviation GPS system is expecte~ to possibly cost as little as $2800 (Ref. land 102) to $3600 (Ref. 148) whil~ an ~~AV system comprised of 1 VOR. 1 DME and 1 RNAV computer will cost $3700 (Ref. 148) while still being dependent on an ~bility to receive an acceptable VOR/DME signal, the attractive- neS8 of GP~ becomes clearly evident. It should be noted that the VOR/DME.'RNAV system iescribed above still cannot provide altitude information, and velocity information is highly cependent upon accurate VOR/DME position fhes which must be differentiated.

Three levels of system accuracy are provided. where the first two are protected by the military. The third, envisioned to relnain unprotected for civil aviation use. was recently found to be "too accurate" and consequently will probably be degraded .

• to offer 200 meter (660 ft) accuracies with a SOl confidence level (Ref. 8). In order to realize how accurate GPS i8, one should note that tests pdrformed at 6 km (20,000 ft) by a C-14l using either one or both of two satellites which were then in orbic reflected errors on the order of 4 m (13 ft) in x, y, and z (Ref. 1).

It is significant that a GPS-equipped aircraft will never be threatened by a signal-saturated environment because airborne equipment is totally passive, i.p.. no signals are transmitted from the aircraft as is the case with DME. When a dig:l.tal data link is added (Time-Division-Multiple-Access Link) together with a ground- based GPS transmitter, accuracies sufficient for terminal area guidance results, togeth~r with a collision avoidance capability.

In view of the above, benefits are expected for both Airplane A and B in terms of safety. Since the baselines for both airplaneR incl\~de IFR avionics, only minor cost penalties are expected. It should be noted here that the forecast data on costs is considered fairly optimistic and did not appear to be supported by manufac- turers. Hence, a penalty of -1 was assessed. If, on the other hand. costs do attain forecast levels, then a +1 might be warranted. In such a case, the figures of merit for both airplanes would be on the order of +24.

4.5.7.2 Inertial ~avigation Systems (INS). Inertial navigation provides the attractiveness of a self-contained navigational sys- tem. Also, it provi1es a capability of computing reasonably accurate velocities which normally could not be obtained by other means w~.thout enhancement. e. 8. J ILS signals are too noisy to be differ- entiated with any confidence (Ref. 155). When INS is used with r radio ranging and a Kalman filter, an Aided IN results, which has applicability to short haul transports by providing 4-D or two- segment approach guidance (Ref. 193). However, INS purchase prices are on the order of $100,000 in larger commercial arieraft and are not expected to fall below $30,000 for the GA market (Ref. 97).

The Aided INS system is an exception since it is predicated on low cost gyros. Annual maintenance costs are also exorbitant. In a study of relative avionics costs to the user, one investigator penalized INS $9000/year while choosing not to assess such ~osts against other systems (Ref. 195).

Based on the above, both Airplane A and B will suffer severe penalties with only negligible gains in their operating environment. GPS, for example, can provide global coverage with velocity capabilities for a fraction of the cost.

4.5.7.3 Doppler Navieatio~. This navigation mode, like INS, is

I

self contained and therefore not dependent on ground based systeu.. When used in conjunction with current navigation aids such 2S VOR/DME, it offers the potential fer 4-D navigation in ...

the terminal area. Like INS, however. it is an expensive alter- native to an augmented GPS terminal environment. In the enroute phase of flight. the differentiation of DME signals when pro- ceeding tal from a DHE station or the integrat ion of R.~AV capabilities wi~h D~1E and, say, three VOR stations should provide a satisfactory C E « so ~ E velocity computation in either the airway or area navigation environment. Certainly, CPS should provide satisfactory velocities in its conventional (unaupented) mode duril\g enroute flight.

Doppler navigation systems are therefore expected to assess high cost penalites to users where the saae capabilities are expected to be available at much lower costs.

4.5.7.4 Microwave Landing Systems (MLS). MLS offers an attractive alternative to conventional Instrument Landing Systems (ILS) due to an ability to handle several ~ircraft flying at different air- speeds in the approach phase. This capability offers promise even today to the GA aircraft operating in high-density, mixed- traffic, terminal environments. Figure 4.22 shows a conventional ILS profile and Figure 4.23 show~ an MLS profile where the advantages of such a system are more obvious. If one visualizes a GA aircraft operating at 40% to 60% the approach speed of larger connnercial jets lind trying to execute an approach to the same ~~lway as his jet counterparts are, the advantages of the MLS system becomes evident. Certainly. signcficant savings in fuel expended in the terminal area may be expected, particularly by commuter operations whose flight profiles dictate that larger amounts of their total flight time between refuelings is spent in the approach to landing phase.

Although Madden and Desai (Ref. 130) noted in 1973 that it was possible to track a curved approach path using MLS and DME to accuracies within the resolution of ATC radar, Hoffman and Hollister (Ref. 97) noted in 1976 that pilots expressed a reluctance to: -.'--.- .... -.--~ ~~~~-.....--.....- .................... e ... • ...... _-.·, ... _______ .... _ ... - ndW ______ ,.. ______ _ "-"~-... .7 , ....... ,=''':~ ..

EXISTING IlS

UNIQUE Al'PROACH PATH NO CONTINUOUS OME spor RUNWAY RANGES FROM FAN MARKERS ""t:" ~- -: " -,

~

'" ~ '---

OVERSHOOT GUIDANCE .......... L ___ _

, • - 3° - - J..-s---- -

......

.t- AZ.MUTH ONLY '- _ - - EXTENDED ......

' - 'x __ RUNWAY - - - - \3.5 N MI APPRO~ , CENTERLINE - - \\: N MI RANGE _APPROX20 <---- ..

(Ref. 148) Fig. 4.22. Typical I1S Implementation.

~.

< I

MICROWAVE LANDING SYSTEM

CAO GUIOANCE AREA I So OEEP BOo WI DE 1--40 1 W ITH C ONTINUOUS DMf " CO VERAGE TO 20 . 000 Fl J:) RAN GE 20 { 30 N MI

~

~ N

-

"

"- NO UNIOUE I'PPROACH PA 1 H 2. 000 F T ROL LO UT AN O

--

\ - AIRCRAF T REOUIREMENTS M ISSED APPR OACH~

-

C BAND ANTENNA IDMNII C AND Ku BAND ANT . (FORWARD CO · AX ANO WAVEGUIDE INSTALL .

- ----- I -- __ Tls o - - :lao. _ -;-:.::... MICROWAVE - RCVR / XMTR , DME INTERROGI. TOR / fiCVR , ~ - n ~--J FlIG~T PATH COMPUTER ~ O . OOO;T- --- DATA LINK DISPLAY ...J7'i S OOO~

~

"V~N I.A \ :la'll o\f ' ----.:::::::: FA A PREFERRED GUI DANCE 20 DEE P . 1 20 ° WI DE I SOOI CO VE RAGE TO 70 . 000 ~T (Ref. 148) Fig . 4.23. Typic al MLS Implemen tation.

(1) fly steep approaches which do not level off to the more o conventional 3 prior to the final flare, and o (2) fly curved approaches beyond the point where the 3 glide slope is intercepted.

A reasonable plan appears to be ene where the curved and steeper flight paths are used to within 150 m (500 ft) AGL and a transi- tion is made to a conventional glide slope.

Current ILS systems are characterized by reliability, safety, and efficient service. Probably, its only drawbacks are ltmited available channels (40, although only 20 are used currently) and performance degradation resulting from terrain irregularities and heavy snow. (Terrain effectA stem from the fact that ILS glide slope signals require radiated energy to be reflected from the ground plane.) MLS incorporates a microprocessor as well as a digital data link and costs are forecast to be on the order of $2000 for GA and $34000 for the commercial jet carrier by the year 2000.

MLS on the ot~er hand. offer~ five times the channel avail· ability and is not affected by terrain features. Also. a spin-off from the versatility of the MLS is that noise abatement procedures can be more easily integrated than under present ILS systems.

Hence. MLS offers improvements in efficient terminal opera- tions and possible noise benefits for some penalty in cost to both Airplane A and Airplane B.

4.5.7 .• 5 LORAN!l. LORAN C is a low-frequency, lona-rante. a11- weather radio navigation system with absolute accuracies on the order of 0.5 km (1600 ft). This accuracy may be improved by a factor of 2 or 3 by implementing more sophisticated user equipment. Use of LORAN is predicated on having velocity infor- ~tion. and this implies either differentiating the LORAN C time difference measurements (TD) or implementing an INS. LORAN provides area navigation capabilities down to ground level, but at present does not have full coverage of the Continental United Stat~~ (:ONUS). GA user costs for a receiver (without INS) are expe~ted to be as low as $3000 (Ref. 195), while military systems may cost $20,000 (Ref. 148). The receiver unit weighs approximately 11 kg.

LORAN C is being considered as a possible GA navigation aid primarily because it is being considered as a replacement for the present VOR/OME network which is characterized by extremely high operations and maintenance costs. The LORAN C ground-based system costs are expected to be an order of magnitude smaller than the present VOR/OME network.

As presently conceived, Airplane A and B will both suffer cost peualties if LORAN C is implemented. LORAN C does not pro- vide precision terminal guidance, althou3h ILS systems do. In evaluating this tE'chnology, the system is considered without an INS. Hence, the cost penalty is not as severe but is traded for deficiencies in reliability and safe~y.

3- t As this report was being finalized, new information regarding a joint test effort between DOT. NASA and the State of Vermont ~urfaced. The interested reader 1q referred to Aviation Week and Space Technology. March 24, 1980, pgs. 51-58.

4.5.7.6 OMEGA. Like LORAN C, OMEGA is a low-frequency, long- range, all weather navigation system. Unlike LORAN, however, it does have complete CONUS coverage. Its gteatest deficiency is that of accuracy, and CEP's on the order of 4 km (13,000 ft) are reported in Reference 1. When an accurate INS is integrated with the system, errors on the order of 2 km (6500 ft) are quoted.

Consequently, no terminal guidance is provided. Cost for an inexpensive OMEGA system is forecast to be on the order of $30ro (Ref. 195) without the INS. However, it should be noted that 1977 receiver costs were on the order of $ZO,OOO to $59,000 (Ref. 208).

Like LORAN C, this system's attractiveness lies in operational and maintenance costs of ground facilities which may be two orders of magnitude below that of the current VOR/DME network. This is better underst.ood when one considers that only eight stations are used to provide complete global coverage.

For the present evaluation where neither Airplane A nor Air- plane B is used over water, OMEGA (without INS) offet 'c l1cnalties in cost and safety.

d $ 4.5.8 Noise The three technologies investigated here are shown in Table 4.21. Quiet. Efficient Propeller technolotlY appears fairly attractive to both Airplane A and B. due primarily to a promise for reduced interior noise. Ducted Propulsors appear attractive to Airplane B due to the maximum scores (+3) awarded for interior and exterior noise and this vehicle's unique flight profile.

Table 4.21. Noise Figures of Merit for Technology (3) Airplane B

Airplane t

56/-53/85 51/-57/58* Quiet Efficient Propellers 29 Ducted Propulsors - 2 18 Low Level Pressurization -17 -18

* Total number of technologies/lowest FM/highest FM

4.5.8.1 ~ Efficient Propeller Technology. The fact that current propellers already operate at efficiencies on the order of 87% may lead one to doubt the existence Ot significant payoffs in this particular field of endeavor. A review of the literature, however, quickly identifies shortcomings: most of the propellers in use today are based on WW II technology. with appropriate refinements resulting from "cut and try" processes epplied to a ~_rtr __ - t

c ~-~" -~'r, -----------------------------------------

basic design. More recently, however, the requirements for reduced fuel consumptionandnoise pollution have led to renewed interest in propeller technology.

'.

NASA efforts to aid in the development of improved pro- pellers are focused in the General Aviation Propeller (GAP) Technology program, with a goal toward reducing fuel use by 8% to 9%, lowering noise by 5 to 10 db, and improving safety. This ambitious program involves propeller and airframe manufacturers.

consultants, and several universities.

Essentially, the major problem associated with propellers lies in the fact that they have been traditionally designed for performance, with little regard for noise. As one attempts to reduce noise, penalties in performance and/or weight usually mater- ialize. However. since the propeller produces approximately 85% of powerplant noise, the attenuation of this noise will do much to improve community relations and ride comfort.

A major hurdle in reducing noise lies in its accurate pre- diction. Succi (Ref. 205) recent~y reported on the accurate pre- diction of the sound field using the Ffowcs-Williams Hawkins equation modified for computational simplicity. Having achieved this, he then reported on different means of reducing flyover noise: (1) Reducing propeller radius by 20% resulted in an 8 dbA reduc- tion with a 4 1/2% loss in efficiency.

(2) Altering the radial load concentration from 80% to 60% by _~ __ .... ___ c~~ ______ ... ___ rt ... • .. = ___ ........ ___ ~ ____________ - (a) re-twisting resulted in a 4.2 dbA reduction with a 3.9% 108s in efficiency.

(b) changing planform resulted in a 4.8 dbA reduction and a 1% 108s in efficiency.

Two additional methods are also discussed (increasing the number of blades and the blade sweep) but results were not quantified as above.

Other developments which promise to increase fuel efficiencies lie in airfoil development. The ARA-D, reported in Reference 38, appears particularly promising for turboprop application. Briefly, this particular airfoil retains its takeoff performance instead of showing the typical reductions in thrust which accompany reduc- tions in activity factor. When one considers that higher cruise efficiencies are obtainable at lower activity factors (and what has traditionally implied lower takeoff performance), the ARA-D looks very attractive.

Increasing the number of blades has been recognized as an effective means of reducing noise. However, this presents a weight penalty which is compounded possibly by increased strength re~uirements resulting from increased vibration. In this light,

•

the Kevlar composite propeller developed and certified by Hartzell is extremely attractive. This propeller has a 50% blade weight reduction with a 100% increase in strength. Unfortunately, its price was increased by a factor of 2.5. As composites are better understood and manufacturing processes refined. significant weight and price reductions a,pear inevitable.

art s ' : !- In light of the above di.cu.sion. both Airplane A and B should receive significant benefit. in noi.. and weight reductions as well a8 safety, with a modest penalty in cost.

4.5.8.2 Low Level ?res.urization. This technology was envisioned as employing a low 6p on the order of 14 kN/m (2 psi) for the sale purpose of attenuating noise. However, the addition of such a system ~ill still require that a pressurization system be added at a distinct cost and weight penalty. Also, it was quickly noted that although this 6p could reduce an ambient altitude of 4.87 km (16,000 ft) to a ~abin altitude of 3.05 km (10,000 ft), it would also exert an outward force of 9.61 KN on 3 0.762 x 0.914 m door (2160 Ibf on a 2.5 x 3 ft door). Hence, even "low level pressurization" will impose significant loads on the aircraft structure, implying f~rther (structural) weight penalties in addition to those of the system.

Both airplanes may be expected to be heavily penalized for a reduction in interior noise.

4.5.8.3 Ducted Propulsors. Ducted propulso(s are investigated here as a possible means for reducing noise. Associated with much improved interior and exterior noise characteristics are improved takeoff and landing performance characteristics. However, cruise performance is degraded and significant- weight penalties may be incurred. Hence, ducted propulsors offer Airplane B greater benefits intakeoff and landLng performance due to its mission profile.

An intereatina propoaal toward attenuatina noiae is the Q-Fan as reported in leference 230. However. thb .yet. is llot con- "

sidered compatible with the low speed characteristics of Airplane I

, , A and B.

4.5.9 Propulsion Five propulsion systems and two related technologies were considered for Airplane A, while only two propulsion systems were examined for Airplane B. As sh;;,wn in Table 4.22, the GATE Engine had the highe~t figure of merit for Airplane A, and the Stratified Charge Rotary Combustion Engine was second. As shown in Table 4.10. the third technology for Airplane A had a figure of merit of 58. The ratings for both engines considered for Airplane B are encouraging since the baseline for this vehicle already included a turboprop.

In view of the particularly high figures of merit attained by this technology group, it appears appropriate to reiterate the fact that the present evaluation technique provides a measure of potential benefit to the user as opposed to a benefit/risk evaluat iOli.

4.5.9.1 GATE Engine. The General Aviation Turbine Engine studies which were begun in l~77 involved Detroit Diesel Allison. Garrett AiResearch, Teledyne CAE. and Williams Research, while management of the program was provided by the NASA Lewis Research Facility.

This study was ,!:·voted to investigating opportunities for advanced e $ Table 4.22. Pt'opulaion Figure. of Merit for Technology (8) Airplane A Airplane B 56/-53/85* 51/-51/85* GATE Engine 85 32 Stratified Charge Rotary Combustion Engine 84 21 Advanced Diesel Engine 36 Improved Turbocharging Liqu id Cool ing 13 Stratified Charge Reciprocating Engine 12 HCRLB Reciprocating Engine 11

*

Total number of technologies/lowest FM/highest FM technologies in small turbine engines below the 750 kw size class. and one key i,;sue involved the question of how to make the turbine engine cost-competitive with reciprocating engines. The results of the study were extremely encouraging and have been reported in the literature (Ref. 12. 65, 123. 204). Briefly summar- ized. turbine engines in gene13l traditionally suffer from penalties in purchase price and fuel consumption and consequently have not penetrated the cost-conscious GA market. Their dominance in larger 5ircraft can be attribut~d in lArge part to their attrac- tive features which include lower weight. apparently better safety, improved ride comfort, and improvec aircraft performance lSl expressed in terms of higher ceilings and cruise speeds. In addition, tre JP4 which they utilize has an approximate 10% advantage in energy content and a 10% price advantage when compared with avga~ (Ref. 204). Interestingly, three of the four study participants forecast competitiv~ purchase prices which were the result of lower component manufacturing costs and increased sales volume. The fourth participant elected to increase sophis- tication and efficiencies to the extent that purchase price was raised but operating costs were much lowe~.

Two major advantages of turbines which may not be readily apparent lie in its three-to-one weight advantage and potential for attaining greater cruising speeds. In a cruise dominated m1ssi~~ this results in a markedly smaller airplane for the same payload and range as a red procating-powered cOlnterpart due to the cascading effect that lower ~ngine weight and cleaner installa- tions have on vehicle size. As an illustration, consider that the lighter engine allows for a smaller and lighter vehicle, which allows for less drag. Thb allows for a reduction in required fuel volume which results in an even smaller airplane, which requires even less fuel.

For the present evaluation, Airplane A was considered to be powered b~" a conventional reciprocating engine. A maximum pur- chase price penalty (-3) was assess,d primarily becau3e the market forec3st used in the GATE studies appeared somewhat optimistic. Airplane B, on the other hand, was given a purchase "-" - ~- ~ -~ ____ """ __ l!iIr_= ___ ""_" ____ "'**"""~"""" price advantage due to an anticipated significant reduction in manufacturing cost. Recall that this airplane's baseline was presumed to already be powered by a turboprop. Hence. perfor- mance improvements in speed and ceiling were not allowed although imorovements in BSFC and purchase price were credited.

f

I

4.5.9.2 Stratified Charge Rotary Combustion Engine. Research in I this particular field shows promising results. The rotary com- bustion engine, possesses several inherent features which make it extremely attractive for application as an aircraft powerplant.

For example. it is sim?le. lightweight, and compact. Since power generation is not based on reciprocating pistons it is smoother.

Liquid cooling allows for quieter operation. safer cabin ~eat and a probable reduction in cooling drag (since present air cooled installations are typically somewhat inefficient with unnecessary drag penalties). The absence of valves and cams also promises quieter operation and improved maintenance. Whereas homogeneous charge fuel efliciencies were less than competitive some years ago, (BSF :: ,328 kg/kW/hr or .54 lb/hp/hr). stratified charge BSFC's of .262 kg/kW/hr (.43 lb/hr/hr) have recently been achieved and compare very favorably with present reciprocating engines. Projections for the future are for BSFC's on the order of .231 kg/kW/hr (.38 lb/hp/hr).

A significant feature of this system lies in its multi-fuel capacity. One engine presently operational at Curtiss-Wright burns lP4, diesel fuel, alcohol. or avgas. As such, it is extrel11('ly attractive in that transitions from avgas to JP4 with - ~~~~,,-, -- ....... - ....... --~~ .... -'"' ...... "".,.......-¥-I ... "" ... ~"T"-, changes in market trends which may be forced through fuel shortages will have minimal effects on this powerplant. The key to this versatility lies in the air motion within the combustion chamber which apparently is conducive to stratified charge operation.

However, certain drawbacks to this system may be deduced from present and forecast levels of the state-of-the-art.

Specifically, high pressure, timed fuel injection and time~ igni- tion systems are expected to be required. These represent fairly high cost technologies. Also, although the induction manifold is smaller than for reciprocating counterparts (leading to lower turbocharging requirements), an improved turbocharging capability will probably ~e required. This alone can represent significant increases in cost, particularly if the production base is not supported by auto~otive engine requirements.

For this evaluation, Airplane A is presumed to be powered by a conventional reciprocating engine and Airplane B by a current technology turboprop.

4.5.9.3 High Comprssion Ratio Lean Burn Engine (HCRLB). In- creasing compression ratio and leaning the fuel mixture are the most effective ways of improving fuel economy. When this process is applied to improve the performance of a current technology, homo- g~neous charge reciprocating engine, the type of engine identified by the acronym HCRLB results. High compression ratio offers increases in thermal efficicl':y of both air standard cycles as well as fuel-air cycles. However, when applied to a homogeneous - ---- -~-- - ---- charge engine, an upper limit is established by fuel octane number. (Present engines operate at CR's on the order of 8.5:1 and have BSFC's on the order of .255 to .268 kg/kW/hr (.42-.44 lb/hp/hr». When improved fuel injection (timed, moderate pressure) and improved cylinder cooling techniques are applied, leaner fuel/air mixtures may be obtained. Reference 169 points to BSFC improvements on the order of 4% to 5% when leaner operation incorporating improved fuel injection and cooling methods are applied to a TCM IO-520 engine. However, costs to the user are expected to be p40hibitive1y high as recertification costs are recouped. The reader is referred to Ref~rences 169 and 170, both by Rezy, Stukes, Tucker, and Meyers, for excellent dis- cussions and analytical results of very current concepts regard- ing the emissions and fuel consumption of reciprocating engines.

Only Airplane A was considered in the present evaluation, where B was considered to be powered by a turboprop.

4.5.9.4 Stratified Charge Reciprocating Engine. Much of the discussior, material presented in this section resulted frOlLl infor- mation received from Teledyne Continental Motors and is unpub- lished. Their assistance is deeply appreciated and acknowledged.

Interpretations of the general information acquired, on the other hand, are sulely those of the present reserrch team.

Reciprocating engine stratified ~harg~ systems are not new.

Three concepts presently being pursued by industry are (1) the Honda Compound vortex Controlled Combustion (CVCC) system, (2) the Texaco Controlled Combustion System (TCCS), and (3) The .54£;:;$ Ford Programmed Combustion System (PROCO). These three systems operate under different combustion processes and are briefly described below. .'

Stratified Charge systems differ from homogeneous charge (current) systems in that the fuel/air mixture of the former is : not homogeneous within the combustion chamber. Instead, a rich zone is maintained at the point of ignition, and the flame front

j

then progresses into a lean region which would not normally be J expected to sustain combustion ~lone. The overall effect is a leaner combustion process which yields lower emissions and better fuel efficiency. TWo ~ethods of charge stratification are currently visualized. In one case, stratification is achieved physically by Injtcting the rich mixture into a prechamber where it is ignited. The second method relies on obtaining an air flow pattern within the cylinder itself to maintain stratifica- tion and hence is called an open chamber system. The CVCC is based on a prechamber system ~hile the PROCa and TCCS employ open chamber methods.

At present, ~rechamber technology appears to have progressed to the point where a GA powerplant could be put into production ..

with a certain degree of conf;dence. Such a system, when compared to open chamber systems, appears to have less efficient output which would result in a heavier engine for a given power level.

Also, the chamber itself will probably require additional cooling,

j

and this may present a problem in an air-cooled configuration. On f ..

I' .,,~

'

I

i 1:;6 the other hand, the lean fuel-air r.atio can probably be extended, arid NO will be greatly reduced while HC and CO may be expected x to increase.

", The PROCO and TCCS. both open chamber systems, differ primarily in fuel injection time, PROCO uses early injection near BDC, while TCCS uses late injection near TDC. PROCO, as a result, offers the possibility of better air utilization but has no multi-fuel capability. TCCS, while rp.taining a multi-fuel capability, can suffer from a smoke problem due to incomplete air utilization. Both systems operate at very high compression ratios with TCCS at 12:1 and PROCO at 11:1.

Of the above systems, TCCS appears to be more attractive for GA application ~ue to its multi-fuel capability. Also, due to its ltite injection feature, turbocharging and compression ratios are limited by st~uctural considerations rather than combustion considerations. When tested in a jeep, TCCS showed BSFC reductions on the order of 35% under part load operation.

At the anticipat~d higher power loadings, differences between homogeneous and stratified charge operations indicate a 25% improvement with a turbocharged Tecs displaying BSFC's on the order of .25 kg/kl.J/hr {.4l lh/hp/hr) at BMEP' s of 552 to 689 kN/m (80 to 100 psi). Turbocharging the TeCS is expected to increase output without redllcing fuel economy or multi-fuel capability.

If air throttling is employed. its application will probably be used to i~rrove emissions control at low power settings.

With respect to present GA engines, TCCS will require a high pressure, timed fuel injection system and a variable timing igni- tion system. Less effective air utilization points to heavier .'

engines while complexity points to increased costs. Improvements to the reciprocating engine, then, tend to detract from its presently undisputed cost aevantage, while its advantage in fuel efficiency appears to be eroding. On the other hand, one cannot ignore the fact that this powerplant has established itself in the GA market, and it is difficult to visualize other engine types replacing it in the next decade.

4.5.9.5 Advanced Diesel Engine. The diesel engine appears attrac- tive for GA application due to its low-cost fuel (20% advantage) and its low BSFC's. Its fuel characteristics alone (where it is not octane/detonation limited) allow for higher ccmpression ratios and higher efficiencies.

At present, Teledyne Continental Motors, General Products Division, is invt.·stigating advanced two-stroke diesel eL.~ine conct::pts which are being 3pplied to a six cylinder, radial, 30U kW (400 hp) engine and a four cylinder, radial, 150 kW (201) hp) engine. The 300 kW engine is considered to involve higher risk than the 200 kW engine. and includes those associated with (1) ceramic comp' nents, (2) operation in an unfinned cylinder environ- ment (no cooling air), (3) high speed turbo starter/alternator, and (lb) catalytic combu~tors. Impressive performance characteristics of this engine includc- takeoff BSFC's on the order of .225 kg/kW/hr

"....- ..................... ---- ...... -~ ........... - ...... ---~--.-. ~.-- ~- .. -~--

"'#'"'" -,- -¥SlAP. f-- (.37 lb/hp/hr) and cruise BSFC's of .195 kg/kW/hr (.32 Ib/hp/hr).

Since this engine already enjoys a fuel price advantage, fuel c~.t.

alone could experience a 40% reduction over present day GA pro- '.

pulsion systems. Other attributes of this system include what has occasionally been called "uncooled operation." However, although the cylinders are unfinned. the injector pumps and after- cooler require cooling, and it is evident that cooling air of some sort must be provided for. This requirement is nevertheless anticipated to alJow for lower cooling drag losses than presently experienced by the GA fleet of air-cooled engines. Weight advantages of the system could not be ascertained. Alt~vugh the dry engine weight of the 300 kW engine is listed as 207 kg (457 lb) vs 262 kg (578 Ib) for a comparable GTSIO-520-H, the diesel engine is known to require oil cooling, and this suggests that inEtalled weight advantages may fail to materialize if the oil system weight is high. The purChase price disadvantage has been estimated to b~ on the order of 20%.

Airplane A should receive advantages in DOC due to fuel efficiencies but is penalized for purchase price. No data for emisaions characteristics was found, but the quieter ope~3tion attributed to the ab3ence of valVES was considered to be only a ~mall advantage.

4.5.9.6 ~iqui<! CooliI!.li' This technology is considered mature but due to risks involved with single point failures such as those associated with leaks, thermostat failures, and pump failures, has disappeared from the GA reciprocating en,ine powered fleet.

Still, the advantages attributed to its incorporation are signi- ficant enough that its possible re-fmplementation is investigated here. Only a few of the more salient characteristics associated with its advantages and disadvantages are listed here.

Advantages: (1) More uniform cylinder temperatures can be maintained, thereby relieving thermal stresses and improving TaO.

(2) Engines can be manufactured with closer cl .. ·r~nces by reducing the effect of thermal stresses, particularly those associated with idle power rapid letdowns. This could aid in TBO imp~ovement.

(3) More freedom is gained in engine/airframe configuration integration.

(4) Engines can be more compact due to the absence of fin spacing reGuirements.

t

(5) Critical cooling requirements such as required for exhaust valves can be more easily met.

(6) Fuel consumption durir.g chmbs should improve on the order of 10% due to the elimination of the requirement for cooling through rich operation.

(7) Cooling drag associated with inefficient methods o=~c, encountered in air-cooled installations can be significantly reduced. (This drag can be significant as reported in References 36 and 142.)

· .

k (8) Quieter operation may be expected due to the water jacket provided about the engine.

(9) Cabin heat can be provided much more safely through a liquid heat exchanger than through present methods. The latter suffers from the distinct possibility of introducing exhaust fumes into the cockpit should the exhaust pipe heat exchanger fail.

Disadvantages: (1) Single point failures resulting from a loss of coolant, thermostat failures in the closed position, or pump failures, are serious.

(2) Warm up times will be extended. In cold weather, this could result in severe engine wear, with tpe distinct possibility of condensed products of combustion contributing to cylinder corrosion.

(3) Deterioration of system efficiency through scaling suggests problems with system maintenance.

(4) A vapor pocket in the coolant system can result in localized interruptions in cooling.

(5) Engine maintenance costs may be expected to inc~ease due to an increase in manhours required to remove or disassemble the engine. Air cooled systems offer the possibility of removing single cylinders from the engine.

(6) Although dry weights may be expected to be on the order of air cooled systems, installed (wet) weights will be greater.

---, For the present evaluation. both reliability and s.fety were penalized by -1. thereby reflecting some optimism in the ability "- of technological improvements to offaet Disadvantage No.1. If this assumption should prove to be invalid. the severe penal tv

~

which would result will make liquid cooling extremely unattractive.

4.5.9.7 Improved Turbocharging. Improved turbocharging promises to improve fuel efficiency and increase cliMb and cruise speed.

by increasing thf critical altitude and aircraft ceiling, signi- ficant improvements on the order of 10% to 15% can be expected in range.

Howpver, procurement costs can be expected to be si8Oifi- cantly highe-: if the !.mprovements in turbochttrging are dictated for the GA fleet alone. Today, for example, procurement r.osts are kept low due to a very high production base provided by the auto- motive and truck engine markets. A major manufacturer, for example, cites a production base on the order of one million units/year, with only 8,000 units being used by the GA fleet. The specific te~hnology which promises to have the greatest impact on improved

turbocharger performance appears to be that of air bearings. •

It seems safe to predict that improved turbochargers will not find their way into the GA fleet until automotive requirements dictate their production. Howeve~. efforts toward improving fuel efficiencies (including the possibility of production of the PRoca or TCeS engines) may dictate their improvement.. According- ly, only a modest penalty in cost is assessed for this technology.

4.5.10 Structural Materials Three technologies were considered and, a •• hown in Table 4.23, appear ext'famely attractive. Here, all three were evaluated on the basia of anticipated maturity in certification procedures and small purchase price penalties. Hence, in view of presently increasing manpower costs which would be significantly offset by composite ~nufacturing processes, the user was forecast to suffer no significant purchase price renalties. The advantages of composites which lie in cleaner aerodynamic shapes and weight reductions, allow for improvements in ceiling, crllise speed, fuel eff:1.ciency t and range.

Table 4.23 Structural Materials Figures of Mer:!.t for Technology 3 Airplane B Airplane A 56/-53/85* 51/-51/58* 55 58 Fiberglass Composites Kevlar Composites 51 t.8 Graphite Compo~ites 43

* Total number of t~chnologies/lowest FM/highest FM

...

I 4.5.10.1 Fiberglass Co'llposites_. These comp.:>sites are attractive because they promise significant weight reductions at lower cost and risk than those associated \o/ith Kevlar or ::.raphite. although the latter two show greater tensile strength. Advantages noted also include the promise of reduced production cost, improved drag characteriatics, low corrosion, and araater etaenath than aluainua.

In OBe corrosion teat, epoxy-fiberalaas va. note' to not ~orroda vhen in contact with anodized, primed, or painted alU1linull sur- faces. In fact, ita inert and durable characteristics allow it to be used ~s one of lev~ral isolatorl for graphite, which reacts very stronaly al a cathode to almost every metAl. In one test, epoxy fiberglaa. va. exposed to a 5% salt spr3y for 5000 hour. and showed only superficial corrosion (Ref. 12~). In another, a helicopter blade fabricated by Bell Helicopters was cracked an~ then run for an additional 900 hours without further failure (Ref. 120).

Its use in primary aircraft structures has seen only limited application in Rutan's canard configured vehicles, the Windecker Eag1p, and the Bellanca Skyrocket, but this is seen to be largely the result of the over-designing required to certify a new produc- tion process, where weight advantages are quickly nullified.

4.5.10.2 Kev1ar Composites. These aramid fibers l~ve greater tensile strength than either fiberglass or graphite and are inherently inert. Current costs are 01 the order of $33 per kg as a woven cloth vs graphite, which costs on the order of $110 per kg in cloth form.

Hartzell's certified composite propE.'ller is made frotJI Kevlar.

Thi~ blade is 50% lighter and tyice as strong 8S its conventional counterpart but costs 2.5 times as much. As certification pro- cedures are refined to allow the full potential of composites to be exploited. both cost and weight ~hould be considerably reduced.

4.5.10.3 GTaphite Composites. Host of the liteTatuTe on coaposites published Teceatly appeaTs to deal with gTaphite. This composite •.

has the strength and stiffnes8 of steel, but only 60% of the density • Although it does not posless the tensile stTength of KevlaT, ita ) compTessive strength is greater. and offeT8 the potential fOT a 30% weight reduction in certain compouents. Although essentially ineTt when isolated, it acts as a strong cathode when in contact with Rluminum and must be isolated through the use of other mater- ia1s such as fiberglass. It is by far the most expenaive of the three compos4.tes invest igated. although costs have dropped from $550-$1200 per kilogram in 1968 to $55-$110 per kilogram currently.

Here. the lower figure is for sin.;le fiber filaments while the upper figure represents the cost for its woven cloth form. As in many of the other tectmo10gies, its cost is expected to drop substantially if the auto industry forms a production baRe for its use.

As its characteristics becom,_ better understood, its intru- sion into the GA market appears inevitable. At pres~nt. it is expected to be used extens1ve1y in the Lear Fan.

..

4.6 OTHER PROMISING TECHNOLOGIES As one might suspect, certain pr~lsing technologies were omitted from the present evaluation due primarily to application considerationa for airplane type as discussed previously in Section 4.2.3. In othE'r words. a technology ;tppHcable to the jet neet but not applicable to the single en;~inet li6ht twin, or commutE:f I.

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-

airplane fleet could not be evaluated utll1a1D.I the pre.ant technique .ince the catelory weights were developed for .pecific mis8ion profiles.

Two technologies raeeivins considerable attention in the current literature and which were not considered to be a part of the present research are the Prop Fan and the Quiet, Clean General Aviation Turbofan (QCGAT). Both of these technologies appear very promising for long-haul high speed cruise, and are discussed here.

4.6.1 Prop!!!!.

The no," familiar 8-bladed propeller which has appeared in several periodicals offers extreme promise for the present turbo- fan-powered commercial fleet. It was not evaluated as part of Airplane A-or Airplane B-related technologies due to the fact that it is presently being develof<!d for Mach numbers on the order of 0.8.

Presently, both an 8-bladed and a 10-bladed version have been investigaLed with extremely promising results. Fuel effi- ciency improvements on the order of 22% are forecast, with DOC improvements of 7% over comparable turbofans possible with improve- ments in design for maintainability (Ref. 46). The fuel figure is particularly impressive when one realizes that a 1% improve- ment provides an annual SQvings of 100 million gallons of air- craft fuel (Ref. 46) for the present civil fleet.

The major disadvantage of the prop fan is noise, and extensive studies are being performed to rectify this difficulty. When one (~ ~, ) considers that military long endurance aircraft can realize 35% fuel savings and reduce gross weight by 25%, the additional weight penalty for noise treatment appears to be a small price.

; .: • However, exterior noise still would not be resolved. (Notet'hat noise cancelling techniques already employed through blade sweep

I

-~ have decreased sound pressure levels on the order of 6 db.) The interested reader is directed to References 46, 100, 101, and 141 for further information regarding prop fans.

4.6.2 ~, Clean, General !viation Turbofan (QCGA~) AVCO Lycoming and Garrett AiResearch have both performed research in this area for NASA Lc.w1s with promislug tesults. As was the case for the prop fan, tn1s technology was not evaluated as part of the present study due to the lack of applicabJlity of turbofans to either Airplane A or B. In the research performed, Garrett chose to synthesize a stretched Lear 35 with a range of 33uO km (1780 nm) at an altitude of 12,200 m (40,000 ft) and M - 0.8 for its modified TPE 731-3. AVCO, on the other hand, had Beech synthesize an aircraft for its modified LTS 101 which resulted in a vehicle capable of 2780 km (1500 nm) at an a1ti-

tude of 10,000 m (33,000 ft) and M = 0.6. It is emphasized that

although the airplanes were synthesized, the modified engines were operational (although not flightworthy) and have been delivered to and tested by NASA-Lewis personnel. Results of the program ware extensively discussed at the General Aviation Propulsion Conference h~ld at the Lewis Research Fad ~ 0° -.:.y in Novemb-:r 1979, and indicated .' .. _I ~t"l t ace 0.£2 Lil i1SJtSL _

that _jor loala were Ht. Noticeably nrc aoala wera not aat

under moat conditione. However. all noi.e loala (which repre- .0 aented a 90% reduction in noiae footprint) aDd moat .. iaaiona goals were .. t.

Since a major thruat of the QCGAT prolraa waa to deaouatrate larle-engine technology transfers to amall-enaine •• it appear a that the program was quite successful. Certainly, the noise goals ac.hieved are extremely impressive.

4.7 SUHMAR.Y OF .!!!! TECHNOLOGY EVAWATION

S~veral significant po1r.ts made earlier in th~.s chapter are repeated here for emphasls. Flnally, a list of attraetive and un- attractive technologies will be disucssp~.

Features of the Evaluation Technique 4.7.1 (1) The rating yardstick used for Airplane A and Airplane B reflect a concensus of opinion both in the identification of relevant categories as well. as in their quantification.

(2) Responses to the three rounds of the survey to quantify the categories (Survey 2) ranged from 70% to 95%.

(3) Only comparable technologies should be evaluated under the present technique. For example, high pressure, timed, fuel injection should not be compared with Rtratified charge re- ciprocating engines.

(4) The relative figures of merit generated by different raters display significarot stability in ranking different techno- logies despite differences in magnitude noted between raters.

(5) Large changes in relative figures of merit between technolo- gies are not observed when relative ~enef1t. are perturbed aboHt an expected value. Howeve:, sut:"~ changes may be expected if the state-of-the-art improves and ~esults in an unforeseen application scenario.

(6) The four technologies in the crashwortl'.iness group appear sOlle- what underrated. This is attributable to the fact that improve- ments are restricted to only the crashworthiness category. and that this category had relatively low weights for both Airplane A and Airplane B.

(7) Small differences in figures of merit should not be inter- preted to rank one technology ovet another.

(8) The ranking of technologies by figures of merit does not re- flect risk in the present analysis.

4.7.2 Attractive Technologies (1) The GATE engine and the stratified charge rotary combustion engine dominate the rankings for Airplane A as shown in Table 4.9 and offer significant advantages to Airplane B as shown in Table 4.10.

(2) All propulsion engines offer benefits to both airplanes.

(3) Fiberglass, Kev1ar, and graphite composites are all extremely attractive provided certification procedures are refined and ~ modest decrease in price is noted.

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I I (4) Aerodynamic concepts such as natural laainar flow airfoils.

J I spoilers with full span flaps, and low/.ediua speed airfoils appear extremely atteactive.

(5) Integrated avionics systems which alleviate pilot workload such as UAAS and Digital Data Linke are very attractive.

A very important consideration which deserves mention at this point is that many technologies whle', were eliminated from con- sideration due to the criteria of Section 4.2.3 are directly inferred from the list of attractive technologies just defined.

For example, improved fuel injection, improved turbochargers, variable timed ignition systems, and ceramic technol~gy are all implied for the improved positive displacement engines. Also, improvements 1n ~;"mputatior.al aerodynamics are required in order to permit three dimensional wing-body-tail analyses as well as wing-spoiler-flap analyses.

4.7.3 Unattractive Technologies (1) Area navigation concepts such as LORAN Cand OMEGA are fairly unattractive to the user although they offer signi- ficant sav I.ngs in operations and maintenance expenses for the ground sites.

(2) Active controls, Fly-by-Hire, and Fly-by-Light are extremely unattractive to both Airplane A and Airplane B, as are HUD's, Doppler, and Inertial Navigation.

31_::_ :::; -- SASH" 4

X kE._

wa uta

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CHAPTER 1

DESIGN STUDIES i/ The design st;.I1ies illustrate the direct and synergistic impacts of advanced technologies bj ~o~paring advanced and current technology airplanes that have been d~signed to the same specifi- cations.

5.1 DESIGN SPECIFICATIONS The statement of work specified that conventio~al and canard configurations be developed for two types of airplanes incorpnrati.ng appropriate advanced technologies. The preliminary specifications gi-,en are as follows: (1) Design conventional and canard configurations for: (a> A small airplane with at least a 6-passenger cabin (including pilots), and (b) A large airplane with at least a 12-passenger cabin (excluding pilots).

(2) Performance objectives are: (a) Maximum LID ~ 18

•

(b) Cruise speed ~ 250 kt (c) Landing speed ~ 60 kt For design purposes, the two airplane types were more narrowly defined. The "small" airplane is defined as a 6-passenger (incluJ- ing pilots) single engine business anMor personal airplane and ..

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.. .£ 2.

g; .. 0.£ .*,.

eorre.pond. to Airplane A of Chapter 4. The "larle" airplane i.

defined a •• 1t-pa •• euaer (exelud1na pilot.) eOlDU'ter airliner and

I

] , eorreapond. to AirFlane B of Chapter 4. The two type. of airplane.

are refered to a. the "6'pasaenaer" and "eOlaUter" 1n thb ehapter. !

•

5.1.1 6-Pa •• enger Design ~peeifieationa Charaeteriatic8 of exiatina 6-pasaeDger airplane. were used as guidelines in completing the 6-passenaer design specifications.

As shown in Table 5.1, the resulting specifications were split into "required" and "desired" sets. This was done because LID> 18 stronaly conflicts with V ~ 60 kt: Furthermore, since it was stall felt that landing apd takeoff distances are much more important than the speeds involved, the stall speed requirements were relaxE'd.

5.1.2 Commuter Design Specifications The characteristics of existing commuter aircraft were used as guidelines in establishing the commuter design specifications.

Again, the resulting specifications were divided into "required" il and "desired" sets. The "desired requirements for the commuter were relaxed extensively because of its operating envelope. Com- t muter Lircraft operate at low altitudes, making a desired LID of 18 extremely difficul~ to attain. Furthermore, a landing speed of less than 60 knots is somewhat unrealistic in the size of commuters typical of this study. Balanced field length, together with landing distance, is more important than landing speed. The resultfn· specifications follow in Table 4.5.2.

.. .J u ••••• & a a a Table 5.1. 6-Pas8enaer Design Specifications (1) Required: Seating for 6 persons (including 2 c~ew) , with baggase.

(2) V > 250 Itt at aros. weight.

cl'.1ise -

•

• (3) Ranse ~. 900 nm with 45 min. reserves carrying 6 passengers and baggage.

(4) Takeoff distance over a 10.7 m obstacle (ISA) and maximum ~ 610 m, sea level gross weight.

(5) Landing distance over a 15.2 m obstacle ~ 610 m, sea level (ISA) and maximum gross weight.

(6) Meet FAR 23 requirements except for Vstall as noted in (8).

(7) Desired: Maximum Lin ~ 18 cruise.

(8) Vstall ~ 60 kt in the landing configura- tion at maximum gross weight and sea level (ISA).

..

5.2 TECHNOLOGY INTEGRATION A major preliminary task performed during the course of the design study was the identification of those technologies to be incorporated into the designs. The technology evaluation of Chapter 4 was helpful in this respect and served as a guideline in selecting particular technologies.

,.".' .... £.3 .•.•.• 2 .............. 2 •. £ .......... IIIIII!I&lIjlll.II .... ~ .. ····ff:_!·;!~··~~~~~!!~.~ .. ~;~==.,~--~-::==.~ Tabl. S.2. eo..uter De.iID Specificationa Required: (1) SHUna for 21 peraona (including 2 crew) with baasaaG.

(2) Vi> 250 kt at maximum takeoff weiaht cn se - at 3050 m MSL (ISA).

(3) 2000 am ferry raDge where ferry ranle is determined for full fuel + 2 crew (only) at 250 kt with 45 min. reserves.

(4) Segmented range of nine 87 nm legs without refueling with a 60J load factor (12 pax) and with 45 min. reserves.

(5) SIngle leg range of at least 400 nm at 100% load factor with 45 min. reserves.

(6) Balanced field length of 1370 m or less at sea level elSA) and maximum takeoff weight.

(7) Meet FAR part 25 raquirements.

Desired: (8) (a) Maximum LID ~18 at cruise.

(b) Vstall ~ 60 kt in the landing config- uration at maximum gross weight It should be noted that the impact of many of the evaluated technologies cannot be determined by design studies. For example, while the impact of composite materials may be evaluated, the effects of load ltmiting .eat. and CRT displays are much harder to quantify. In general, only techDolo,ies which directly influence the miasion performance of an airplane are capable of being analyzed.

5.2.1 6-Passenaer T.cnnology .Integration As one would expect, the highest ranking technologies iden- tified in the analysis of Chapter 4 were in the areas of prop~18ion, structures, and aerodynamics. Specific technologies from ~~ch of these areas were selected and are termed "Primary Technologies" in that they directly influence airplane performance. Other TEchnologies whi.ch do not directly influence airplane performance were also int;')rr .'lrated and are termed "Secondary Technologies."

Table 5.3 lists the Primary and Secondary Technologies which were incorporated into the advanced technology 6-passenger airplanes.

The reasons for selecting the technologies in Table 5.3 are as follows: (1) Spoilers and Full Span Fowler Flaps Trade studies (see Section 5.3.1) showed that high wing load- ingB (1915 to 2394 N/m ) markedly improve cruise efficiency but degrade takeoff and landing performance. The obvious solution here is to utilize spoilers for primary roll con- trol and free the trailing edge of the wing for full span flaps. Fowler flaps are used because they provide high lift and low drag at small deflections (for takeoff) and higher lift and drag at large deflections (for landing).

Table 5.3. 6-Pa •• enger Advanced Tecbnololle.

* SecODdary TechDologie. ..

* Primary Technoloaie.

Spoilers for roll control Load lillitinl .eata EnerlY absorbing floor Full span Fowler flaps Improved restraint.

Advanced natural laminar flow airfoil Integrated avionics and Kevlar and/or graphite di.play.

composite structure Stratified charge. highly Integrated low-cost fluidic turbocharged rotary wing leveler combustion engine Quiet. efficient propeller technology it Primary Technologies are those which directly influence airplane performance while Secondary Technologies do not.

(2) Advanced Natural Laminar Flow Airfoil Preliminary data show that computer generated natural laminar flow (NLF) airfoils have low-drag characteristics approaching those of NACA 6-series airfoils and maximum lift coeffi- cients resembling those of tre turbulent-flow NASA LS-series airfoils, which are also computer generated. Furthermore.

i the advanced NLF airfoils are far less sensitive to rough- ness (in terms of maximum lift coefficient) than the NACA 6-series airfoils.

·1

~ (2) Kevlar and/or Graphite Composite Structure Advanced composites can significantly reduce structural weight while providing a stiff and aerodynamically smooth

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.urface. Several .tudie. indicate that .. jor co.t ".,ina.

are po.dbl. due to a reductiOll .i.tt the IWllber of parte aDd man-hour. raquired for production.

(4) Stratified Chari., HilMy Turbocbarl_ Iotary eo.buetion (aC) Engine Thi. 8Uaine i. cbaracteriaed by a biab power to weilbt ratio, good .pecific fuel consumption, and liquid coolina.

The ac 8Uaine was .el~cted instead of the GATE turboprop

primarily because the former is undergoing teaU while the latter 18 a paper study. Furthermore. the GAT! engine 11 evolvina into a higher horsepower category than required for the 6-pa.senger airplane.

(5) Quiet-Efficient Propeller Technology This includes computer generated hlade sections, optimized planforms. composite blades, etc. Cruise efficiency gains are relatively small (2-3%) but significant improvements in takeoff and climb efficiencies and noise characteristics are possible.

(6) Load Limit{ng Seats and Energy Absorbing Floor These technologies can reduce peak vertical g-loadings by 50% and therefore greatly reduce the risk of death or injury in the event of a crash. Little or no weight and cost penalties are expected.

(7) Improved Restraints Improved restraints are a , .. .i.nor technology in terms of air- plane weight and cost but greatly enhance sur/ivability by hk U £ .

1uuriaa that tbe occupant rnain. in bia ... t durina a cra.h. 1f loed It.itlna ... ts are uaed, effective re.traint.

beeoa. mandatory. Th. oaly prObl ... ideotified appear to b. tho •• a,soicated witb coafort aDd .... of u'e.

(8) Intearated Avionics and Display.

wnile not directly influencing perfor.ance. thi' tecbnoloay cara improve the utility and flexibility of an airplane aDd can be considered to incr ... e .afety .ince it ha. the potential to lub.tantally reduce pilot workloat4.. The degree of sophi.tication of the .ystem aa applied to the 6-pa •• enser airplane has not been defined but the capabilities of the technology in general are illustrated by the t~SA Ames PCAAS and DAAS studies.

(9) Integrated Low-Cost Fluidic Wing Leveler This wing leveler 1s visualized as a basic part of the air- plane. Because most airplanes are not spirally stable.

this technology can significantly improve safety. part i- cularly 1n IFR conditions by reducing pilot workload.

This system is commerCially available to the homebuilt market for less than $200.

5.Z.2 Commuter Technology Integration , The discussions of advanced technologip.s int.,·grated int., the 6-passenger designs apply to the commuter designs. Identical

\

technologies were integrated with two exceptions: (1) powerp1ant and (2) airfoil. Table 5.4 presents the advanced technologies in- corporated into the commuter design • • ..

Table 5.4. ColilDUter Advanced TE'cbnolosies

* *

Primary Technologies Secondary Technologies Spoilers for roll control Load limiting seats Full span Fowler flaps Energy absorbing floor Advanced low-speed airfo::!.l Improved rebtraints Kevlar and/or graphite Integrated avionics and composite structure displays GATE technology turboprop engine Efficient propeller technology it Primary Technologies are those which directly influence aircraft while Secondary Technologies do not.

The technologi.es not previously discussed but applicable to tne commuter are: (1) GATE technology t.llrboprop ~ngine.

This cngine was selected because of its high power-to-weight

•

ratio. low pur~hase price, and good specific fuel consumption.

Other advantages attributed to turboprops include high reliability and long engine lifetimes.

(2) Advanced low-speed airfoil.

The advanced NASA low-speed airfoils (LS-series) have characteristics similar to the original GA(W) series uf airfoila and exhibit aood max~ lift characteriatics toaether with tmproved pitch1na moment characteristica. Alao, the thickne.s of the low-speed section a •• iats in .. int.inina low wina weiahta while providina a hiah volume for fuel.

S.3 PARAMETRIC ~ STUDIES Parametric trade studies were conducted to establish the im?act of basic airplane parameters (wing loading. aspect ratio, power loading. etc.) ~n mission performance.

5.3.1 6-Passenger Trade Studies Three types of trade studies were ~onducted covering cruise.

taekoff, and landing ~erformance. Detailed derivations may ba found in hppeodix C.I.

The cruise trades were fonnulated to show the affective power loading ([n p/w1. hp/lb) required to cruise at 250 kt as a function p of wing loading (W/S), effective aspect ratio (Ae), zero lift drag coefficiept (CD ). and altitude. It quickly became apparent o that the high LID requirement necess·tated high cruising altitudes.

Figures 5.1 and 5.2 present the cruise trades for altitudes of blOO m and 9140 m. respectively.

The tlll<eoff ".rades were expressed in terns of wing loading (W/S) and power loading (P/W) required for a 610 ~ ~akeoff dis- tanee over a 15.2 m ohstacle. (This requirement was later changed to a blO m takeoff distance over a 10.7 m ohstacle.) The requirement proved to be a very strong {·.lnction of the maximum takeoff lift coefficient (C ) and a very weak function of L maA TO aspect ratio (A) and zero lift drag coefficient (C ). Th e se n o ch aracteris tics led to a fairly simple representati on of t ::e takeoff protlem as shovn in Figure 5.3.

Velocity • 250 kt

Altitude • 6100 m

Wing Loading w/s (N/m )

1000 1500 2000 2500

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Wing LOIcling - wlS tNlm )

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Wing Loadfng- wls (lblft2)

i , Fig. 5.2. 6-Passenger Cruise Trade Stu1iep (&t 9140 m) ""..-t·""'~""·,t .. ,"..,,,,,,,~,~ . . ...

Wing Loading",W/S (N/m ,

1500 2000 2500 3(XX) 3500

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NOTES: 610 m Takeoff Distance Over IS. 2 m Obstacle

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Fig. 5.3. 6-Passenger Takeoff Trade Studies " ... """"""".""',I"", ...... "e"" .... """" ..... .jI~ ......... ~;!Ioo ~_liIWoIH1! '. i ~Ur. •• ~'"ltl\l-·'.'·, ... 'UI"1 ... ~."' ... "' ..... _"o!i"',Iu,u .... _,.AA!",,··W'O\IIi"'".'~h''"''"''"':" ........ IniI ••• ~iII~'InI!.~~tII j "11.M'UI~!'Pt.. '1r.,.IIIIIi/lililfU .11'1'''11 _ .......... tIf, 1IJ''''1Ir.t1Mf' e.. oM et 1fhtrrtift 'tMNW!!!t!l!rd'W' * IIIIHLIII,IjpMI'" The landing trades show wing loadings and maximum landing lift coefficients (~ ) required for a 610 m landing ~ •

·

distance over a 15.2 m obstacle and for a stall speed of 60 kt.

Figure 5.4 gives the results of the landing trade study.

•

The trade studies of Figures 5.1-5.4 ar~ summarized in Table 5.5.

As a result of the trade studies, it was decided that relax- ing the Vstall ~ 60 kt requirement was justified since it would otherwise arbitrarily restrict the application of advanced tech- nologies. Airplane parameters and design goals for the advanced technology 6-Passenger airplanes were established as follows: Parameters: 2 2 w/s - 2155 N/m (45 Ib/ft ) A • 11 hcruise • 9144 m (30.000 ft) Goals: CD < .02 o e > .75 C > 2.5 and CD < .10 L max °TO TO C > 3.0 L max LDG 5.3.2 Commuter Trade Studies

---

Two types of trade studies were conducted for the commuter aircraft and consisted of cruise and balanced field length per- formance. Detailed derivations are included in Appendix C.2.

..

Wing loadlng--w/s (N/m )

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Fig. 5.4. 6-Passenger Landing Trade Studies ..

Table 5.5. 6-Passenger Trade Study Summary CRUISE TRADE STUDY

·

·

~-----------------------------------------------------~ Current high performance single engine airplanes typically have

Cn : .02. Ae ~ 5.5. and w/s : 1005 N/m which r~su1ts in Lin •

o

•

11 at 9140 m.

Lin • 18 requires much higher w/s (2l55 N/m ) and larger Ae (80) for ~ • .02. and C • .02 is very difficult to achieve with n o 0 wi S • 2155 N/m • TAKEOFF TRADE S~JDY

p/w required for constant w/s is a very strong function of

C • L maXoro

Because cruise requirements demand high wls, advanced flap

technology (full span Fowler flaps) becomes a necessity if reasonable Pl.W is to be maintained.

LANDING TRADE STUDY Vstall ~ ~O kt is a very limiting requirement compared to the 610 m landing distance requirement.

If the Vstall requirement is adopted. it becomes the critical

factor in selecting wIs. If the landin~ distance requirement

is used instead, the takeoff distance requirement becomes critical.

Similar to the 6-passenger studies. the commuter cruise studies • show the effective power loading (n P/W) r~quired to cruis~ at p 250 kt and 3050 m as a function of wing loading (W/S), effective aspect ratio (Ae) and zero lift drag coefficient (CD). Figure o 5.5 presents the cruise trade studies.

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c.. 0.01 ~ "'~ ... ~ ... _ 15 000750-

~ 04 ----= -:S;:ri 18 • =

:;::0. 1IIiiiiii ..... ~ • 20 :e

as ............. -::.:_____ O. OOSO.!

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= -------- ~

0.0025

o

30 40 50 60 70 80

Wing Loading -w/s UWft2)

Fig. 5.5. Commuter Cruise Trade Studies The balanced field length 8tudie. were developed in term. of ~ • effective a.pect ratio Ae, maxtmum 11ft coefficient in takeoff o configuration (~ ). and effective power loadina (n P/W). The :

1UX.ro p

studie. .how that balanced field length i8 dependent priaarily on ~ and (n P/W). Results of this study are pre.ented a.

max P TO Figures 5.6, 5.7, and 5.8, and were developed for values of ~ • .075 and Ae • 10.

o

Wing Loadlng,wls-Nlm

2(XX) 2SOO 3(XX) 3500 4(XX)

2(XX)

Coo - 0.075

Ae -10

r"Pf.N • 0.075 hP/lb

(0.0126 kW/N)

40 50 60 70 80

Wi"9 Loading, W/S -lblft

Fig. 5.6. Commuter Balanced Field Length Studies _ - .• _.-.¥ .44 __ £4.'; --,--.4S .. 021J . .. .22 ... . 2 Ji .. j 13 ,

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!

Wing LDading. w/s-N/m

2(8) 2500 3CXI) 3500 .., , F

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~ 5000

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Ae -10 600

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(0. 0168 kW IN) 400

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Wing Loading, W/S ---Ib/ft

Fig. 5.7. COMauter Balanced Field Length Studies 18~' .. LEU m £1 ..

21L ,SOd)#. £

Wing loading, W/S-N/m2

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3500 «lit

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~ 50 ~ 70 ~

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Fig. 5.B. Commuter Balanced Field length Studies.

__ u.

.12 ObHrvatlona reMiltina fro. an exaainat101l of theae fiaur ..

appear 10 Table 5.6.

Table 5.6. Commuter Trade Study S~ry CRUISE TUDE STUDIES : .025, Ae : 6.0, and Current commuters have values of ~ a w/s : 2155 N/m • at a cruise altitude of 3050 •• LID • 18 demands extremely low values of ~ (:.02), coupled a with high wing loadings (:3352 N/m ). As was the case with the 6-passenaer design a CD !:it .02 appears very difficult. to a achieve undeT conditions of high wing loadings.

BALANCED FIELD LENGTH Because of the high wing loadings desired in cruise flight, advanceo nap technology is required (full span Fowler. flaps).

in takeoff allows lower power loadings for takeoff High ~ max and thereby maintains a more reasonable difference between cruise and takeoff power requirements.

These re6ults dictated relaxation of two of the original specifications.

First. the requirement for V 11 < 60 kt is sta - not practicn~ for the commuter aircraft since these aircraft (~n

•

the 18-19 pas:;·;:l\$C!r ~.:"'ge) tYI'i.cally are not constrained by this requirement. A ~urvey for a commuter traffic model pre- sented in Reference 44 showed that 93% of th~ airports in use had effective runway lengths of 1220 m or greater. where e:tective l length is defined as the actual runway length corrected to sea level standard conditions. Hence. balanced field length as F L - ~ ~ .. - - . - - ~- - ." - - -- -

r

i oppoaed to .tall .peed i • .ore critical in the deter.lnat1oD of I I acce •• ible airport.. Second, the LID requir..-nt of 18 va. relaxed

j

.

} to a value of 12 due to a de.ired crul.e altitude of 3050 a. .

This altitude was .elected becau.. available data indicated that the bulk of the c~ter .. rut could be acc •• aed vlthout bevinl to cruiae at hilher Altitude.. Piaure 5.5 clearly 1adicat ••

that LID = 12 is a more reali.tic losl. Th. de.ilD loa1. e.tab-

li,hed for the commuter aircraft were: Paraaeter.: wls • 3112 N/m A • 12 Goala:

~ = .020 - .025

°

e ::: .75 5.4 CONCEPTUAL lJF.SIGNS This section describes the conceptual design of conventional and canard configurations for the 6-passenger and COMmUter airplanes which incorporate the advanced technologies discussed in Section 5.2.

..

Current technology baseline £irplan~1 were also eltabliahed to provide a basis for comparison of the advanced technology deligus.

The project staff was aided in this effort by the senior design class of the Aerospace Engineering Department at the Univer- sity of Kansas. The additional manpower proved valuable to the ,

I

I

g" w 4A4L_MM!

-.LtC;) &Os..e & $ sa design effort and assisted in identifyinl unforeseen difficulties s.socia~~~ wi!h technology integration and configuration analysis.

5.4.1 6-Passenger Designs A current technology baseU.r.a, an advanced technology «:on- ventionally configured vehicle, and a canard configuration were synthesized.

All three designs utilize the same basic cabin, which seats six passenget~ (all facing forward), and has a baggage area behind the rear seats. The cabin, which is illustrated in Figure 5.9, is somewhat idealized since it was common to all three designs.

A more involved design effort would probably result in a more rounded cross-section with contours more closely matched to the con- figuration in question.

Design of the 6-Passenger airplanes followed the methods outlined below. Detailed procedures may be found in Appendix C.

(1) Preliminary Sizing An initial gross weight was estimated using the Breguet range equation and the wing was sized ac~ordingly. Tail areas were sized with tail volume coefficients obtained hom similar types of existing airplanes.

(2) Weight a~d Balance Component weights were calculated giving an improved weight estimate and providing center of gravity information.

(3) Horizontal Tail Sizing The horizontal tail area required was determined from static margin and takeoff rotation requirements.

19j -------_._------- ._--------------, .

Nons: (I) Scele: I/lto

.

(2) All dimensions Ire Inside dlmenslo".

()) Cabin length • .66 m (It) MIX cabIn width • 1.17 m (5) Max cabin height • 1.27 m (6) •• 08 m Cabrn wall thickness (]) Seat pitch •• 66 m A-A B-B

c-c

•

Cab I n Wid t h.' • , 2 rn Cab I n Wid t h • 1.' 7rd Cab I n W i Ii t h • I., 2 III Cabin Height - l.nll1 Cabin Height. I.27m~ahln Height. I.27m Seat Width •• 46 I" Seat \Jldth • .!t(, III Seat \lidth • .51 f1\ A I s leW I d t h·. 25 ill -----_._ ... _------_ .. __ ._---._._----- .....

Fig. ~.9. b-Pass~n~pr Cahin l~yout 19/, ."J;.~a.o~iiI.e.yg_ .. aJZ2I .•.• !." •.• ~ .. tCj •. " .• 2. •...•. """Z~ ..• S"" •.• "4~2 •.•. Bn ...... $.tt.a .... _wJt •. ~ .. ~~~--~ .'

,- i (4) Aerodynamics ~irplane wetted area and zero 11ft drag coefficient were calculall!d.

(5) Final Sizing GASP (General Aviation Synthesis Program) was used to tlflytl the airplane. Engine aile and gross weight were adjusted to meet the design specifications. Prior to the aclual sizing.

GASP weight and drag routines were callb~ated agair-at the manual calculations.

The three 6-Passeng~r designs are discussed separately in the following sections.

5.4.1.1 Current Techn~ Baseline. The 6-Passenger current technology baseline (6PAXBL) provides a basip of compdrison for the advanced technology airplane.

In designing 6PAXBL, characteristics of current high r~rfo~- mance, single engine, general aviation airplanes were retained as much as possible. The major ex~entio~ was that a 7620 m cruise altitude was assumed so that the power required at cruise would not result 1.n an exceptionally large engine. Specifications for this airplane are those shown in Table 5.1 with the exception of the LID,::, 18 goal. 6PAXBL hilS the following characteristics: (1) Geared, fuel injected, turbocharged, air-cooled, 6-cy1indcr (opPOI:' ~(n reciprocating engin~.

(2) Conventional aluminum structure.

19'\ 4! Mi (3) Partial spaD, .inal.-.lott~ flap •• (4) Conventional control surfac ••• (5) Cruise altitude of 7620 m; pressurized to maintain ..

• a 2440 m cabin.

(6) NAC! 64 -A215 airfoil.

(7) W/S. 1053.4 N/. (22 Ib/ft ).

(8) A.. 7.S.

The design procedures discussed in Section 5.4.1 resulted in the 6PAXBL design shown in Figure 5.10 and Table 5.7 summarizes the characteristics and performance of this airplane.

Referring to Tsble 5.7 one notes that 6PAXBL is a high per- formance airplane which has speed. payload. and range characterls- tics more similar to those of the Bellanca Skyrocket than to those of more conventional genersl aviation airplanes. A compari- son of 6PAXBL to the advanced technology designs and to existing airplanes 13 discussed in Section 5.5.

5.4.1.2 Advanced T~chnology Conventional Configuration. The 6- Passenger conventional configuration. advanced technology airplane (6PAXAD) incorporates the technologies listed in Table 5.3. The trade studies of ~ection 5.3.1 were used to define the basic airplane parameters.

No real problems were encountered in the conceptual design of 6PAXAD but detailed design of the flap and spoiler systems would be complicated ~1 the small size of the wing. The general characteristics of this airplane are as follows:

\

19" s O O CO ^ p U+ p Y. U\ ~ L N V d _ M ^ N ^ J W ^ti n N n n p CID V 'O V P Q F. a 11^ Y O 1' _ M^ • Y r N N .^ n ^^ ^ u rW U N 4 .. .^ `L K i R ry v ^ ^ x •N s a a It z7 O ] C Y S O ^ o O ^ • ^ p ^ V^ ^ V ♦ { ] j ^ ♦ ^ \ M S O ^ w J H N t f ul d Y r^ D w N .rr u GJ F y ^ ; a n^ v\ r^ - Y @ ^.

C 11 m .i ^ t^ N ^ ^_ s.^^^ 1 ^ t? ___ w N F p _ NE ^ rv4 N ^ J ^ ,^ ^ M r L e ^^ ^ N ^ C J N C^ u - n 4 a T ^_.

U u .^ ^ ^ s ^ 'n ^ _ s a n Y A J C P U P A C V ^'v^ V V Y C `' I V V • ^ T 7 1 ^ p"4 .. S i ♦ ^ K .. d H a^ U ♦ ^ V ^ K . K L4 ^ W W I N x n d AC ^, u rl n V ^ Ff ~ y ^. S L.

~-- Table 5.7.

6-Pa.senaer Current Technology Baaeline Characteriatica WEIGHTS • .

Gross • l859.8 kg (4100 Ib) Payload • 544.3 kb (1200 Ib) Empty • 1051.4 kg (2318 Ib) Fuel (max payload) • 264 kg (582 lb) ENGINE Geared, fuel injected, turbocharged, air-cooled 6-cylinder (opposed) reciprocating Max Power • 317 kW (425 hp) Critical Altitude • 7620 m (25.000 ft) ~ 1b Cruise BSFC - .274 kW-hr (.450 hp-hr) Dry Specific Weight • .894 kg/kW (1.47 lb/hp) TBO • 1800 hr PERFORMANCE Vi· 250 kt (75% power, gross weight) cru se Range - 900 nm + 45 min. reserves (250 kt, 544.3 kg (1200 1b) payload) 2 LID • 11.7 (250 kt, gross weight) Vstail • 58.3 kt (flaps 40°) Takeoff dist. over 10.7 m (35 ft) obstac:e - 540 m (1771 ft) (flaps 10°)2 Landing dist. over 15.2 m (50 ft) obstacle - 351 m (1152 ft) (flaps 40°)2 1 7620 m (25,000 ft) altitude Sea level, gross weight (1) Geared, stratified charge. highly turbocharged, liquid cooled, 2-rotor. rotary combustion engine.

t

• • (2) Composite (Kevlar and/or graphite) structure • (3) Full span single-slotted Flower flaps.

(4) Spoilers for roll control.

(5) Cruise altitude of 9140 m; pressurized to maintain a 2440 m cabin.

(6) NASA advanced natural laminar flow airfoil.

2 2

(7) wIs. 2154.6 N/m (45 Ib/ft )

(8) A· 11 Here. the design procedures previously discussed in Section 5.4.1 resultecl in the 6PAXAD design shown in Figure 5.11 and Table 5.8 summarizes its characteristics and performance.

Table 5.8 shows that the 6PAXAD design meets the required specifications of Table 5.1 but does not have the desired LID ~ 18 or V 11 < 60 kt. A detailed discussion of 6PAXAD together with sta - a comparison with 6PAXBL is presented in Section 5.5.

5.4.1.3 Advanced Tech~ology Canard Configuration. The 6-Passenger canard configuration, advanced tehcnology airplane (6PAXC) in- corporates the technologies listed in Table 5.3 and the trade studIes of Section 5.3.1 were again used to define the basic air- plane parameters.

Several problems were encountered in the design of 6PAXC.

First, methods for sizing the canard and for calculating induced drag were lacking. Second, GASP will not handle canards. For £ ~USELAGE: Length. 7.53 m (24.7 ft) Wldt~ - 1.32 m (4.33 ftl Height • l.~~ m (5.38 ftl , Passenger •• 6 EN:;I~j£: l1ax Power - 205 kll (27S hpl Prop Di" - 1.83 m (6 ftl Crit Alt • ~144 m (30000 ttl I Blade •• WING: Area - 5.73 m (61.7 f(2) Sweep (.25e) - 1.7 deg Span - 7.92 m (26 ftl Dined .. ,1 - 6 deg Asp~ct Patio. 11 WHhout • 2 deg Tap~r R.tiQ •• 5 I Airfoil: NASA NlF I : FLAPS: Sir,gle slotted Fowler flaps I : cfl< •• 3~ bf/b - .n::.

I I I Tak~off 5ettinq • 20 deg landing Setting - 40 deg I I 2 2 HO~ TAIL: Area. 1.61 m (17.3 ft ) Taper Ratio· .S N Span • 3. 11 m (10.2 ft) o Sweep (.25c) - 3.17 deg o A:;pect Rat:o - (.

Airfoil: NACA 0009 ~.

2 2 VEIIT TAil: Arca· 1.71 m (13 ft ) Taper Ratio - .5 Span - 1.55 M (5.1 ft) Sweep (.25c) • 14 deg Aspect Ratio a 2 Airfoil: NACA 0009 (;

aj

~~ 3

c::::::::::-.

L~_ iJ ~ __ =@)~_ (0) Fig. 5.11.

Advanced Technology Conventional Configuration 6-Passenger Design I: ..

..

""""",,.,,·.-lIM'oI\I.!I>o1, 11!'IJ'·<IIJI.""liIlur.l!IIIf>4~~!llMlijjU'IIN~I~iIlItIliilllllililliilll"'j"f.n*"I.** ".''fllli'' ,1 U' Wth t 1)"" ,' .. e,.b. .- ..... ~--.' ,,.,,.

f Table 5.8. 6-P ••• eoaer Advanced TechDololY Conventional CoDfiauratioD Characteristic.

WEIGHTS Gross • 1258.7 kg (2775 lb) Payload • 544.3 kl (1200 lb) Empty • 582 kg (1283 lb) Fuel (max payload) • 132.5 kg (292 lb) ENGINE Geared. stratified charge. highly turbocharged. liquid-cooled 2-rotor, rotary combustion Max Power • 205 kW (275 hp) Critical Altitude - 9144 m (30.000 ft) ~ 1b Cruise BSFC - .231 kW-hr (.38 hp-hr ) Dry Specific Weight • .45 kg/kW (.74 1b/hp) TBO - 3000 hr PERFORMANCE Vi· 250 kt (75% power, gross weight) cru se Range • 900 nm + 45 min. reserves (250 kt, 544.3 kg (1200 1b) payload) 2 LID • 14.2 (250 kt, gross weight) V 11 - 64.5 kt (flaps 40°)2 sta Takeoff diet. over 10.7 m (35 it) obstacle • 610 m (1999 ft; (flaps 20°)2 Landing dist. ~ver lS.2 m (50 ft) obstacle • 406 m (1332 ft) (flaps 40°) 1 9144 m (30,000 ft) altitude Sea level, gross weight ~Cl £oE ,_

I

these reasons, 6PAXC cannot legitimately b. coapare4 to 6PAlBL or 6PAXAD, and as presented here, muat be couid.red only as a po.sible configuration and not a. a refined desian. The.e prob~ • • lema are discussed in greater detail in Section 5.4.3.

General characteristics of the canard desian are the ....

aa those of 6PAXAD as followa: (i) Geared, stratified charge, highly turbocharged. liquid cooled. 2-rotor. rotary combustion engine.

(2) Composite (Kevlar and/or graphite) structure.

(3) Full span single-slotted Fowler flaps.

(4) Spoilers for roll control.

(5) Cruise altitude of 9140 mj pressurized to maintain a 2440 m cabin.

(6) NASA advanced natural laminar flow airfoil.

(7) W/S - 2154.6 N/m2 (45 lb/ft2) (8) A· 11 The full span flaps called for may create a trim problem but the extent of the problem, if it exists, was not investigated d'Je to the previouoly mentioned lack of reliable analysis technique~.

Fi&ure 5.12 presents a 3-view of the advanced technology canard configuration.

5.4.2 Commuter Designs An advanced technology conventional configuration and an advanced technology canard configuration were synthesized. A third existin~ aircraft was selected as a baseline for co~parison.

r - --~~-~-

'H'H~"'"_I""'~'''"I'II:''+\'~'fflll/lII',,"~ • .

• • fUSELAGE: l~~9tn. 6.22 m (20.~ ftl Vldth • 1.32 N (~.13 ft) "eight. 1.42 m (4.67 ft) I P.ssengers • 6 ENGINE: "a. Power. 20S kW (275 hpj Prop Oi •• 1.83 • (6 ft) Crit Alt • '14~ m (30000 ft) I Ih,des • 3 WING: Area. 5.67 ~2 (61 ft1) Sweep (.2Sc) • 18.S deg Spotn. 1.89", (25.9 ft) Dihedral. 0 C:eg Asp"et llat io • I I WaShout • 2 deg Tap". Ru io •• 5 AirfOil: NASA Nlf flAPS: Single slotted fOWler flap, eflc •• 10 b,/b •• 7~:; Takeof' Setting. 20 deg landing Settlnq • ~O deg CANARD: Aru. I.J' m (15 ft2) t..per Ratio •• 5 Span. 2.9 m (9.5 ft) N Sweep (.25c) • 1.17 deg Aspec t llat io • 6 o IN Airfoil: NASA lS(H)-0417 V£RT TAil: Data Is for. single vertie.1 tail (winglet) Are ••• 6, _2 (7.~1 ft ) Taper Ratio •• 5' Sp~n • '.1£ ~ (J.79 ft) Sweep (.2Sc) - 17 deg Aspect Ratio. 1.)4 Airfoil: NACA 000,

f======= ~

c/'50~

G G 11

@ ®

Fig. 5.12.

Advanced Technology Canard Configuration 6-Passenger Design

I

L &

a

t ,

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The exi.tina aireraft selected ba. a 19 pas.eDger eapabl1ity aDd was •• leeted becaus. it i. one of very fev .ircraft specifically

I

• I d •• tlDed .s a commuter.

• Both advanced tecnnoloay conceptual de.iana utili.e the

t

! , .... fuselaa. interior. Accommodationa are included for 19 paalenaers (sit tina thre~ abreast) and two pilots. Additional space 1s included for cabin furnishina8 and a fliaht attendant.

A baggaae/cargo area is included behind the passenger area. with access through an upward opening cargo door. The pa.senger area of the aircraft is entered through a smaller door at the front of the cabin. Figure 5.13 presents this interior arrangement.

Design of the advanced technology aircraft followed the pro- cedures outlined below. Detailed data and calculations may be found in Appendix C.2.

(1) Preliminary Sizing The gross weight of each of the advanced designs was matched to the baseline aircraft. Tail areas were deter- mined initially by minimum static stability requirements.

(2) Weight and Balance A detailed we.ight and balance calculation was perfonned for each of the designs because interior furnishings and

•

accommodations for these aircraft constitute a large portion of the aircraft wetght.

(3) Horizontal Tail Sizing The horizontal tail was sized for static margin and rota- tion requirements.

~ - ..

• • •

[]

[]

--..-.-

--

Passenger Door Cargo Door a • 0.991 m (39 in) b - 0.451 m (17.75 in) C - 0.409 m (16.1 in) d - 0.511 m (20.1 in) ~

~ I

(typ i ca 1) e - 0.419 m (16.5 in) Scale: 1 in • ~8 in Cabin Length • 32.4 ft Hax Cabin Width - 75.6 in (inside)

-let-

--,,- Hax Cabin Height • 69 In ~inside) I d a Seat Pitch - 36 in " If , .. c -, II T

II

Sectlf)n A-A Fil. 5.13. eo..uter Cabin Layout (4) Aarodynaaic.

Airplane vatted area aDd aero 11ft draa coefficient vere calculated.

(5) Final Sialn, GASP was used to perfor. the mis.ion analysis of the air- craft. Here, gross weight was maintained whl1e engine slze was adjusted to meet design specifications. Two specific missions were analyzed. The fir6t involv.d a 8ingle leg mission at gross weight, while the second v •• characterized by nine 87 nm legs at a 60% load factor. Both missions were flown at 3048 s. Pdor to mission analyt;is, GASP was calibrated to obtain the drag and wci.ght characteristi.C8 which were predicted through manual calculations.

The two advanced desi~ns and the basebne a:l:rcraft "r~ discu8sed separat ely in the following s.~(~t i CHi 's • 5.4.2.1 Current .:r_e.chno1o~ .~as('lin!:.. The I.l:ir·cralt utl.U.u!d for the current technology baseline {COMBL) appears in F:lgu1:'i' :) .11 •• The characteristics of COtIBl. are: (1) Twin turboprop engines.

(2) Conventional aluminum structure.

(3) Partial span, double-slotted flu.!):!!.

(4) Conventional control suriacp:,;, {5} Cru1~e altitude. 3048 m.

(6) NASA 65 A2lS at the f(\ot. NACA 64 .\41i :Ii" t h,.~ tip, (7) W/S· 2lS7 N/m2 (8) A· 7.71

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