Study on the feasibility of V/STOL concepts for short-haul transport aircraft

NASA CONTRACTOR REPORT u m z

STUDY ON THE FEASIBILITY

OF y/STOL CONCEPTS FOR

SHORT-HAUL TRANSPORT AIRCRAFT

.

by K. R. Marsh

Prepared by FJNG-TEMCQ-VOUGHT, INC.

Dallas, Texas OY

f

NATIONAL AERONAUTICSAND SPACEADMINISTRATION l WASHINGTON, D. C. l JANUARY 1967 c TECH LlBRAliY KAFB, NM OObOL4b NASA CR-670 STUDY ON THE FEASIBILITY OF V/STOL CONCEPTS FOR SHORT-HAUL TRANSPORT AIRCRAFT By K. R. Marsh Distribution of this report is provided in the interest of information exchange.

Responsibility for the contents resides in the author or organization that prepared it.

Prepared under Contract No. NAS 2-3036 by LING-TEMCO-VOUGHT, INC.

Dallas, Texas for NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $3.00 Paws No.

INTRODUCTION ..........................

Purpose ..........................

ConceptsStudied. .....................

.....................

StudyGraumiRules STUDYME3?HoDoImY ..........................

...................

Performance Estimates Weight Estimates.

sTuDYREsuIxs ..........................

Configuration Design ....................

.....................

EconwlcAnalyses 22 Operations Analyses ....................

...........

Additional Study of Selected Designs ............... ;: Specific Research Requirements ................

Alrwwthlness Requirements.

STUDYCONCWSIONS ........................

REFERENCES...........................

LET OFTABLES Pap;e No.

Table No.

Control Power Requirements . . . . . . . . . . . . . .

Typical OptlmumDeslgnDrsg Estimate. . . . . . . . .

Zero-I&ft Estimated Drag Coefficients, 60 Psseenger Designs . . . . . , . . . . . . . . . . .

GE 1 Gas Generator Technology . . . . . . . . . . . .

Avionics Equipment ILst . . . . . . . . . . . . . . .

Critical Design Conditions for 60 Passenger Airplanes . . . . . . . . . . . . . . . . . . . . . .

60 Passenger Optimum Airplane Physical Characteristics. . . . . . . . . . . . . . . . . .

8 Estimated Weights - 60 Passenger Airplanes . . . . . .

Direct Operating Cost Breskdown - 60 Passenger Designs - 500~Mile Stage Iength . . . . . . . . . .

10 Canpwlson of 60, 90 ard 120 Passenger Airplane Physical Characteristic8 . . . . . . . . . . , . .

11 Estimated Weights - 90 a& 120 Passenger Airplanes .

LIST OF FIGURES l%ge No.

Figure No.

AIrplanesDeveloped .................

LTV Study Flow Block Diagram .............

Rubberized Propeller . Engine Performance ......

Typical Optimum Design Selection Process .......

Typical Turboprop Airplane ..............

Turboprop Airplane Trwxxnission System ........

Fan-In-Wing Airplane Transmission System Schematic - All Engines Operating ................

Fan-In-Wing Airplane Transmission System Schematic - 11 Propulsive Wing Airplane Transmission System Typical Propulsive Wing Cross Section ........

I2 Turboprop Optimum STOL Airplanes ...........

16 F&n-In-Wing Optimum STOL Airplanes ..........

18 Influence of Propeller Geometric and Operating Characteristics on Cruise Speed - Turboprop STOL Airplane .......................

Influence of Engine-Propeller Operating Character- istics and the Wing Flap Setting on Take-Off Performance -Turboprop STOLAlrplane ........

Ability of Spoilers to Provide Roll Control - Turbo- iV Ian! OF FIGURES (cant ) Ngure No. Pqe No.

21 Ability of Vertical Tall to Provide Yaw Control - Turboprop 2000 Foot STOL Airplane . . . . . . . . . .

22 Weight Sensitivity to Fuel Reserves - Fan-In-Wing V/STOL Airplane . . . . . . . . . . . . . . . . . . . 73 Weight Sensitivity to Skin Friction Drag and Propulsion System Efflclency - Propulsive Wing 2000 Foot STOLAlrplane . . . . . . . . . . . . . . . . . .

Weight Breakdown Comparisons - 60 Passenger Airplanes . . . . . . . . . . . . . . . . . . . . . .

7s Direct Operating Costs - 60 Passenger Airplanes . . .

25 76 Initial Airplane Costs. . . . . . . . . . . . . . . .

26 77 Maintenance Requirements . . . . . . . . . . . . . . .

27 78 28 Direct Operating Cost Breakdown - 500 Mile Stage Length........................

Perceived Noise Level Dlrectlvlty Contours - Turboprop V/STOL Airplane . . . . . . . . . . . . . . 80 Perceived Noise Level Dlrectlvlty Contours - Turboprop 2000 Foot STOL . . . . . . . . . . . . . . . 81 Perceived Noise Level Dlrectlvlty Contours - Fan- In-Wing V/STOL Airplane . . . . . . . . . . . . . . . a2 Perceived Noise Level Dlrectlvlty Contours - Propulsive Wing 2000 Foot STOL Airplane . . . . . . .

a3 a4 Perceived Noise Level Versus Distance . . . . . . . .

Perceived Noise Level Contours for Take-Off and IandIng - Turboprop V/STOL Airplane . . . . . . . . .

as Perceived Noise Isvel Contours for Take-Off ard Landing - Turboprop 2000 Foot STOLAirplane . . . . .

a6 Perceived Noise Ievel Contours for Take-Off arx3 Laxxllng - Fan-In-Wing V/STOL Airplane . . . . . . . .

a7 Perceived Noise Ievel Contour for Take-Off snd fardIng - Propulsive Wing 2000 Foot STOL Airplane . .

aa IXST OF FIGURES (Co&) Ngure No. Page No.

Effect of Non-Productive Time on Direct Operating costs - Turboprop VTOL Airplane . . . . . . . . . . . 89 Effect of Non-Productive Time on Direct Operating costs - Turboprop 2000 Foot STOL Airplane . . . . . . 90 40 Effect of Non-Productive Time on Direct Operating costs - Fan-In-Wing V/STOL Airplane . . . . . . . . . 91 41 Effect of Non-Productive Time on Direct Operating - Propulsive Wing 2000 Foot STOL Airplane . . . 92 costs 42 Effect of Annual Utilization on Direct Operating costs - Turboprop VTOL Airplane . . . . . . . . . . . 93 Effect of Annual Utilization on Direct Operating - Turboprop 2000 Foot STOL Airplane . . . . . .

costs 44 Effect of Annual Utilization on Direct Operating costs - Fan-In-Wing Vj'STOL Airplane . . . . . . . . .

Effect of Annual Utilization on Direct Operating costs - Propulsive Wing 2000 Foot STOL Airplane . . .

46 Influence of a Joint Civil-Military Buy on DOC . . . .

Influence of Gas Generator Costs on DCC . . . . . . , 48 Hypothetical Route . . . . . . . . . . . . . . . . . .

Hypothetical Route Block Time . . . . . . . . . . . . 100 Hypothetical Route DOC . . . . . . . . . . . . . . . . 101 Gross Weight Variation with Passenger Load . . . . . . 102 Direct Operating-Costs Variation with Design - - _ PassengerLosd....................

VI The technical feasibllkty of many V/SToL concepts has been proven by wind tunnel tests and flying prototypes. With this proven technical capability ax~I In antlclpatlon of proJected population move- ments, It Is now considered appropriate to study the applicability of V/STOL airplanes to short-haul transport requirements.

A feasibility study has been performed In which eighteen air- planes have been developed around three V/S!lDL propulsion concepts, four V/STOL operational capabilities Sna three passenger-load capa- bilities. Eachofthe airplanes developedhasbeenoptimlzedto give a nearmlnlmumdlrectoperatlng costonthe design stage lengthof 500 miles within the constraints of its selected V/STOL propulsion system, Vlsn>L operational capability and passenger-load capablllty.

This study has fouIld the turboprop V/STOL airplanes to have only modest cruise speed capabilities, relatively low direct operating costs, and comparatively light weights; and there are considerable data to guide the designer of turboprop V/STOL aircraft. The fan-in-wing V/STOL airplanes have a relatlvelyhighcrulse speed, hlghdlrectoperatlng costs, and high propulsion system plus fuel weights and hence gross weights; there are considerable data available to the designer of fan- In-wing airplanes though not as voluminous or complete as for the turbo- The propulsive wing V/STOL airplanes have high subsonic cruise prop.

speed capabilities, low direct operating costs, and relatively light weights; but there are only limited data to guide tk deslgaer of such aircraft.

Mxe research data on the V/STOL concepts evaluated In this study will permit better design optimization and reduce the technical risks associated with the development of these aircraft. A set of EMera Aviation Airworthiness Standards, applicable to t.& novel flight capa- bilities of V/STOL aircraft, should be developed.

Purpose The bulk of the population gain expected by 1930 will be in urban areas. At least three super-metropolitan areas - the northeast corridor, the Great Lakes area, and along the California-Pacific Coast - will exist They will each extend approximately 400 miles, and they will by 1960.

contain approximately 50 percent of the country's population.

The airports which will serve these super-arztropolitan areas will be forced to increase in size and move further from the population centers to find adequate space for servicing the long distance travelers and to avoid problems associated with community acceptance. As a result of this and the increasing congestion on urban higkrways, the short-haul traveler will be faced with a dilemma - the lack of a rapid, short-haul transport system.

V/STOL short-haul transport aircraft systems with aircraft capable of operating out of very snrallairports are consideredtobe one method of solving this dilemma of the short-haul traveler. !Che technical feasibility of many V/STOL concepts has been proven by wind tunnel tests and flying prototypes; but data were not available to estiblish the eco- nomic feasibility of various V/S!I!OL concepts for short-haul air trans- port applications. Consequently, IEV Aerospace Corporation under contract (Reference l)to NASA, Ames Research Center, has conducted an extensive analysis of turboprop, fan-in-wing and propulsive wing V S!FOL i propulsion system concepts to power short-haul transport aircraft. Air- craft were desiepled around these three propulsion system concepts, and their operation and costs were evaluated.

!Che basic aircraft were capable of carrying 60 passengers, and several 90 and 120 passenger aircraft were also developedandevaluated.

In addition to evaluating the economic feasibility of various V/STOL short-haul aircraft designs, an examination has been made of the research work required to develop these short-haul aircraft into suc- cessful commercial air transports, and an examination has been made of the ability of the existing airworthiness requirements to cope with the novel flight capabilities of V/STOL aircraft.

Concepts Studied Three different V/STOL propulsion system concepts were studied -- turboprops, fan-in-wings, and propulsdve wings (Figure 1). The turbo- prop airplanes used the trashed wing principle with wing tilt applied as 1 Addltio~l concepts were studied by Boeing and ticWed Companies under contract to NASA. Prelimimzary results of these three studies are contained in NASA ~~-116; Conference on V/STOL and STOL Aircraft, April 4-5, 1966.

required to meet the specific runway length design criteria. The fan- in-wing airplanes used the pure fan-in-wing principle where all of the gas generator hot gas was diverted to drive the wing fans for takeoff a~3 landing, aMi ducted straight aft in a conventional turbojet numru?r The propulsive wing airplanes, considered for STOL opera- for cruise.

tions only in this study, used a jet flap principle in conjunction with the propulsive wing to develop high induced lift coefficients for slow speed flight, and these airplanes have a high-bypass ratio turbofan propulsion system.

For this study, the term VTOL, when applied to a specific aircraft, implies that the particular aircraft is designed to takeoff vertically, fly its design stage length without refueling, and land vertically with all fuel reserves on board. The term V/STOL, when applied to a specific aircraft, implies that the particular aircraft is designed to takeoff with a short takeoff run, fly its design stage length without refueling, and land either with a short landing or vertically with all fuel reserves on board; but this aircraft also has the capability to takeoff verti- cally, fly a fixed distance less thsn the design stage ,length without refueling, and land vertically with all fuel reserves on board. The term STOL, when applied to a specific aircraft, implies that the particular aircraft is desagned to takeoff with a short takeoff run, fly its design stage length without refueling, and land with a short landing run with all fuel reserves on board. The STOL airplane has no vertical takeoff and landing capability. For this study, the design stage length for all airplanes is 500 statute miles, and the design VTOL stage length for the V/STOL airplanes is 50 statute miles.

For the turboprop and fan-in-wing propulsion system concepts, VTOL, V/STOL and STOL airplanes were developed; for the propulsive wing con- cept, only STOL airplanes were developed. STOL airplanes were developed for operation from 1000 and 2000 foot runways. All airplanes were opti- mized to give a minimum direct operating cost at the 500 statute mile stage length.

Study Ground Rules The ground rules used for this study were mutually agreed upon by determined by associated studies and experience.

NASAand LTV, and were NASA study ground rules. -The more importantgroundrules estab- lished by Reference 1 are as follows: Passenger accommodations. - The passenger plus baggage weight shall be 200 pounds.

Five abreast seating or more shall be used. Thin back seats with 32 inch pitch seat spacing will be used. The seat width will be 20 inches.

Two lavatories, 38 inches by 35 inches, will be provided.

One stewardess will be provided for 60 passenger versiona and two stewardesses will be provided for 90 and I.20 passenger versions.

The noise level in the passenger compartment shall not exceed 75 decibels or 70 decibels in the speech Interference level in takeoff and cruise, respectively.

Airplane design criteria. - All alrplenes will be optimized to give near minimum direct operating costs on a 500 mile stage length.

Space will be provided for 20 pounds of revenue cargo per seat.

Self-contained passenger loading stairrays, starting systems, and air conditioning will be provided.

Structural design criteria. - Components such as cross shafting and hot gas ducts shall be designed for infinite life.

The airplanes shall meet the structural strength requirements of Reference 2.

The landing gear will be designed for a limit sink speed of 12 feet per second.

Special slow speed flight control criteria. - The special slow speed flight control power requirements are presented in mble I.. If the desired values of control power Impose a severe penalty on an air- plane, the acceptable values may be used. These slow speed controlre- quirements are in addition to the trim requirements.

Performance design criteria. - The take-off and landing perfornmnce of all aircraft will be based on the assumed atmospheric conditions of an 86oF day at sea level.

All cruise flight perfornmnce is calculated assuming standard day atmospheric conditions.

- With all engines operating aad the Special VTOL design criteria.

aircraft trinrmed, the thrust-to-weight ratio shall be equal to or greater than 1.15 with no control input or 1.05 with 50$ of tb? mxlmum control capablllty about any one axis and 20$ of the m~urimum control capability about the other two axes. The control system must be able to give lOO$ of the maximum control moment required about any one axis while it is providing 50$ of the lllaximum control required about the other two exe6 (no thrust-to-weight ratio specified).

With the critical engine inoperative and the aircraft trimmed, the thrust-to-weight ratio must be equal to or greater t&n 1.05 with no con- trol input or 1.0 with 50% of the control required about any one axis and 20% of the control required about the other two axes.

Special STOL design criteria. - With the critical engine failed, the airplane must be able to attain a flight path angle of zero In the final approach configuration without a speed change.

The landing field length required will be the calculated total landing distance from a 50 foot height during the landing approach to the end of the landing roll-out times a factor of 1.67.

During the landing approach, the rate of descent shall not exceed 800 feet per minute at a height of less than 50 feet.

The maximum deceleration rate during a landing roll will be 0.5 g's. (It is to be noted that these last three ground rules limit the maximumapproach speed to 54 knots for 1000 foot STOL airplanes and to 86 knots for the 2000 foot STOL airplane.)

The takeoff field length required will be the calculated distance from the start of the takeoff ground roll to the point where the airplane reaches a height of 35 feet, assuming that a critical engine is failed.

Special approach design criteria (VTOL and STOL). - At the design approach speed and with all engines operating, the airplane must be able to increase its normal load factor by 0.3 by changing angle of attack or power.

At the design approach speed and with the critical engine failed, the airplane shall be able to encounter a ten-knot, sharp-edged, verti- cal gust or a ten-knot, horizontal speed change without encountering excessive buffeting.

At the design approach speed and with the critical engine failed, the airplane must be able to increase its normal load factor by 0.1 by changing angle of attack or power without encountering excessive buf- feting.

LTV study ground rules. - During this study, ECV has adopted the following ground rules because of limitations considered to exist for the 19'70 time period specified for this study.

Propeller diameters are limited to twenty feet.

Tip-turbine-driven-fan pressure ratios are limited to 1.3.

Where engines are interconnected by a hot exhaust interconnecting duct, no more than two engines can be exhausted into a commonduct.

STUDYMETEIoDomY Two general methods of study have been used in this program (Figure 2).

First, a series of pazwnetrlc studies were used to select the characteristics of airplanes that would perform the design mission within the constraints of the study grouA rules azxI the FAR airworthiness requirements. These alr- planes were then analyzed to detersdne their direct operating costs (DOS) on the design stage length of 500 statute miles. That comblnstlon of character- istics whdch resulted in a near minimum DOC for each set of field length and passenger load criteria was then selected as optimum. Numerous slnrpll~ng steps were taken during the parametric study. These simpllficatlone were checked to insure that they did not lmpalr the integrity of the study; and hence the data generated by these studies are considered adequate for com- paring one sirplane against another, providing both airplanes use the same V/STOL propulsion system concept and design criteria.

The parametric studies were followed by a detailed design, operational and econanic analysis of each of the optimum sircraft. From these analyses, the fan-in-wing V/STOL,turboprop VTOL, turboprop 2000 foot STOL and propul- sive wing 2000 foot STOL airplanes were selected to warrant additional study.

Additional analyses of these airplanes Included the development of 90 an3 I.20 passenger versions, studies of sensitivity of selected airplanes to the varia- tion of selected variables, and extended economic analyses.

Performance Estimates The cruise performance estimates for this study have been made at two levels of accountability. For the parsmetrlc studies, generalized per- formance estimation procedures, based on gross geometric and flight character- istics, have been used. These parametric studies prwided the data from which the characteristics of the optimum airplanes were selected. Detail estimates were then made of the Installed propulsion system performsnce and drag characteristics for each optimum airplane; and these data were used in com- paring the optimum airplanes with each other.

Parsmetric study methodology - cruise performance. - For the parametric cruise performance studies, generalized drag estimation procedures have been used. Skin frlctlon drag was assumed to vary with the three maJor aircraft components: (1) fuselage frontal area; (2) engine nacelle frontal area; eLlld (3) wing plus tall surface areas. Raplrlcal equations, found to give reason- able results for paremetric studies, were used to predict the skin friction drag for the various combinations of paremetric variables. To get approxl- mate ekes for tall conflgurations for the parametric designs, a survey was msde of the tall volume coefficients of flying VTOL and STOL alrcraft. l'hle survey showed that these coefflclents were reasonably constant; therefore, representative values of tall volume coefficient were canblned with the tall anus determined from preliminary layouts to get the tail areas for the para- metric analyses.

The drag rise due to ccmpressibillty was predicted by the method of Reference 3 for both the parametric ti detalled analyses. The predicted drag rise of the turboprop airplanes started at a Mach number of approximately The drag rise of the fan-in-wing airplanes was predicted to occur at a 0.7.

Drag rise prediction techniques are not Mach number of approximately 0.8.

applicable to prediction of drag rise characteristics of the propulsive wing confygurations; therefore, El!V has coducted high speed wixxl tunnel tests of a propulsive wing configuration. These test data have shown the drag rise Mach number of this conflguration to be 0.9; therefore, this drag rise %ch number was used for all propulsive wing configurations evaluated In this study.

The drag due to lift was estimated by the method of Reference 4 with a modiMcatlon applied as a result of ITV flight test experience. This method correlates drag due to lift with the aspect ratio of the wing ti It has been used for both the parametric and detailed analyses.

The basic gas generators have been rubberized as turboshaft, turbojet, and turbofan engines. Vendor data have been used to establish the character- The rubberized engine performance data istics of these rubberized engines.

were corrected for installation losses, hot duct losses (inciuding leakage), bleed air extraction, arA horsepower extraction to drive accessories.

Rubberized fan data have been used for the fan-in-wing arxl propulsive For the tip-driven-turbine-fans used in the fan-in-wing con- wing concepts.

These fans have been so cepts, a fan pressure ratio of Il.3 has been used.

limited by structural characteristics of this fan which must absorb the loads of the powering turbine attached to the fan tip. The fans on the propulsive wing concepts have a fan pressure ratio of 1.35, because related studies have shown this fan pressure ratio to be near optimum for such lift-cruise fans.

For turboprop concepts, the variation of thrust with fuel flow for com- binations of the largest and smallest turboshaft engines an3 propellers con- sidered reasonable for this study was detelmined. It was found that for the extreme combinations of these parameters (small engine and large propeller, or large engine and small propeller), the variation of fuel flow with thrust fell within a relatively narrow band for a given cruise speed (Figure 3 ); there- fore, an arbitrary line drawn between these extremes was assumed to def3ne the variation of fuel flow with thrust for the speeds evaluated. This method was then used to convert the turboshaft engine-propeller combinations Into pseudo-jet engines and thus simplify the analysis procedures for the para- metric studies.

Detailed analyses study methodology - cruise performance. - After characteristics of the opt3mum airplanes were determined Awn the parsmetric studies, more exacting performance estimating procedures were used to predict the performance capabilities of each of these optimum designs. Detailed drag estimates, such as shown in Table 2, were made for each optimum aizplane.

More exacting stability ar& control analyses were made, and the revised sur- were used in developing the de- faces, sized as a result of these analyses, tailed drag and weight estimates. Table 3 presents a stum~~ly of the detailed drag estimates for the ten 60-passenger airplanes developed for this study.

DetaIled predictions of the propulsion system Installation losses were msde for each of the three propulsion system concepts studied.

A set of propeller characteristics was selected, and the perfoae characteristics of this propeller were prograwned Into the digital computer so that these characteristics could be used in predicting the perfornusnce of turboprop de- signs. These propeller performance characteristics were obtalned fran Reference 5.

Special slow speed flight performance estimation methods.- The special slow speed flight performance estimation methods deecrlbcd herein have been used for both the parametric anl detadled analyses.

The analytical approach of Reference 6 was used'to predict the slow sped fUghtperfoI7sanc e characterlstlcs of the turboprop sirplanes evaluated in this study. Thismethod ass-s that the etresmtubelnfluencedbythe wing is filled with sir moving at the velocity of the stream tubes generated by the propellers. The methodusespower-offaerodynamic data,- the aero- dynamic forces and moments are calculated by combining these power-off aero- dywnic data with the dynamic pressure of the propeller stream tubes, but with a correction factor applied to caqensate for the ratio of the actual mass of air activated by the airplane to the assumed nwws previously described.

ECV Aerospace Corporation flight test evaluations of the XC-142A have shown the slow speed perfonmnce capabilities of this alrplane to be slightly better than would be predicted using the analytical method of Reference 6; therefore, the slow speed performance estimates developed for this study are (The XC-142A is a turboprop powered V/STOL wans- considered conservative.

port airplane which uses the tilt-wing concept to get Its VTOL capabilitles, aM Its geanetrlc characteristics are described In Reference 7.) The aero- dynamic characteristics of a 4% chord, full spsn double-slotted flap system have been used In predicting the slow speed performance characteristics of the turboprop designs evaluated In this study.

The analytical method of Reference 8 has been used to estimate lrduced This method predicts the ltiuced effects for fan-in-wing conflguratlons.

aerody-namic forces for fan-in-wing airplanes as functions of the fan geometry and performance characterlstlcs, the wing geometry, and the position of the Comparisons of the results obtained using the method of fan In the wing.

Reference 8 with witi tunnel data show the method to predict the induced aerodyne&c effects of fan-in-wing sirplanes sufficiently accurate for feasiblllty studies.

As a result of the lack of aerodyntunic data on propulsive wing con- flguratlons, LTV has used aerodynamic force axl moment coefficients fran tests of a non-optimum configuration (Reference 9). These coefflclente were corrected for aspect ratio differences between the model tested aad the de- signs avaluated, and these corrected coefficients are considered sufYlclently accurate for a feasibility study. Optimkatlon of the flap arrangement and configuration may allow better slow speed perfoxmmce than has been predicted for the propulsive wing airplanes evaluated In this study.

Weight Estimates The weight estimates used in this study were developed using statistical analyses, and they were used in both the parametric and detailed studies.

The weight estimation method consists of using statistical weight equations which have been shown to predict the weights of components of contemporary aircraft as functions of geometric and performance characteristics of the aircraft. These equations were derived using a digital computer routine which determines the best fit for a given set of statistical weight data.

This routine develops best fit weight equations as functions of geometric snd performance variables; and it also develops additional best fit equations by dropping one variable at a time and performing a least square analysis using the remaining vatiables. In this manner, the simplest equation fielding the highest degree of statistical accuracy is determined. Modifying factors, developed analytically, for special features or design compleldties of V/STOL aircraft, not accounted for in the existing statistical data, have been applied to the results of these statistical equations.

The same weight estimation equations for given components have been used, as appropriate, for all airplanes evaluated during this study, thus preventing inadvertent advantage being given to one concept as compared to another.

For situations where statistical weight estimating equations were not available or are otherwise inappropriate; vendor data, scaling curves or specified calculations, based on preliminary structural analyses, were used to arrive at the estimated weights of components.

Parametric structural loads and component sizing analyses have been performed as a part of this study to support the substantiation of the estimated weights.

costs This study has used direct operating costs (DOC) as the optimization criteria for the airplanes developed. The direct operating costs are those costs which accrue when an airplane is operating, and they consist of the depreciation of flight equipment, direct maintenance costs, and the cost of the flight operations. Direct operating costs for parametric studies have been estimated using LTV developed statistical DOCequations. These equa- tions estimate DCC as functions of selected airplane geometric parameters and performance characteristics, and they have been shown to be sufficiently accurate for assessing the effects of changes in design parameters on DOC for parametric studies.

Direct operating costs for the optimum airplanes have been estimated using the method of Reference 10 with some minor modifications. This method involves a relatively complex cost estimating procedure, arii it is expected to give cost data adequate for comparing one optimum airplane with another.

- The major factor in the depreciation Depreciation-flight equipment.

of flight equipment is the initial cost of the airplane. A traditional, detailed costing procedure has been used to estimate the initial airplane cost. This method predicts the cost of major airplane components using statistical cost data fran conventional airplane designs. For the V/STOL sirplanes of this study, correction factors have been applied to compensate for the additional compleldty of components that are different for V/STOL concepts. NASA speclfled that the depreciation periods an3 residual values specified in Reference 10 be used. The spares cost ratios were specified to be those of Reference 10 except that an avionics spares cost ratio of 50$ was to be used, and a spare parts price factor of l.3 was to be used.

Mrect maintenance costs. - In estimating the direct maintenance costs, a deviation was taken to the method of Reference 10. It was considered de- sirable to take advantage of recent experience available on turbine powered aircraft, arm3to account for the additional casplexity of V/STOL sircraft.

A compilation of maintenance experience of various transport airplane opera- tors on all types of turbine powered aircraft was made for each of the major systems of the airplane, with a breslsdown of the major systems made In accord with the Air Transport Association Specification 100. Plots were then pre- pared, for each of these airplane systems, of maintenance manhours per flight hour versus the system weight. Technical jtigment was used to adjust these resulting curves for the Increased maintenance canplexity of the same systems in V/STOL aircraft. These adjusted curves were then used to build the main- tenance dour per flight hour estimates for the designs developed in this study. The estimated weights of the designs developed were broken down in accord with the Air Transport Association Specification 100 format in order that these developed dntenance curves could be used.

Flight operation costs. - The flight operation costs include the fuel ati oil costs, crew costs, and insurance. NASA specified that the method of Reference 10 would be followed except crew costs were increased by 22$ to represent 1965 costs. Only a pilot snd copilot were assumed to be required for the alrplanes developed in this study.

Optimum Design Selection Process For airplanes hating a vertical takeoff capabillty, the combination of parametric variables having a near minimum direct operating cost were selected as the charscteristlcs of the optimum design. It was fouulthatthe vertical takeoff conditions were critical In sizing the propulsion system; therefore, each of the candidate combinations of parametric variables for the VT.OLairplanes had propulsion systems sized to meet these critical vertical takeoff cotiitlons an3 the required mission performance. Combinations of parametric variables were selected for cruise at altltties of 25,000 snd 35,000 feet. The primary parametric variables for these alrplanes were aspect ratio aal wing loading. Checks were made to assure thataddltionalpower above that required for VIOL would not lower the DOC of these airplanes.

For the STOL sirplanes, the thrust-to-weight ratio was an added primary parsmetr5c variable. It was not known whether the takeoff ad/or larullng or cruise thrust-to-weight ratio reqtirements would predominate; therefore, the thrust-to-weight ratio was varied for each of these sirplanes. For each STOL propulsion system concept, propulsion systems were sized to give three selected thrust-to-weight ratios, and the acccmpanying combinations of paramettic variables that would meet the design missions were identified. Again, cambi- nations of parametric variables were selected for cruise altitudes of 25,000 and 35,000 feet.

Mgure 4 presents an example of a typical optimum airplane selection process. Curves of this type were prepared for thrust-to-weight ratios of 0.6, 0.8 and 1.0, and the optimum combinations of wing loading and aspect ratio were selected for each of these three thrust-to-weight ratios. This plot then represents a plot of the optimum wing loadings and aspect ratios for one cruise altitude with the thirst-to-weight ratio as the primary variable; and it can be seen that the minim direct operating cost occurs at a thrust-to-weight ratio of approximately 0.9. Hence, a thrust-to-weight ratio of 0.9 would be optimum for this propulsion system concept, unless the takeoff a&/or landing thrust-to-weight ratio requirements dsmarkl a higher thrust-to-weight ratio.

The propulsive wing concept is unique since Its wing geometry is therefore, instead of varying dictated by the propulsion system arrangement; wing loading, aspect ratio anl thrust-to-weight ratio for this concept, the number of wI= fans and thrust-to-weight ratio are the primary parametric variables. The thrust-to-weight ratio and number of wing fans indirectly specify a wing area and aspect ratio for this concept.

STUDYRESULTS Configuration Design The aircraft designer is faced with a new challenge in his endeavor to successfully integrate all the requirements of V/STOL, and to a lesser degree STOL, aircraft into a useful vehicle. Unique propulsion system arrangements, sophisticated avionics equipments, and canplex control systems must be inte- grated into an airframe which also contains the canplexities of conventional aircraft; and this must be done at an acceptable cost and for a minimum weight.

The V/STOL aircraft is a closely integrated package in which no single system or canponent can be changed without affecting another. The efficiency of a V/STOL aircraft is a direct function of the degree of integration of its systems. A good V/STOL airplane cannot be obtained by simply ccnnbining an optimum propulsion system with an optimum cunplaent of avionics equipment and but rather its success depends on how well all an optimum control system, etc., the systems are integrated to function as a unit - not how well each subsystem operates independently.

General cmponent considerations. - Certain ccanponents of the V/STOL aircraft developed for this study were selected after special side studies were made. These canponents were the powerplants, the avionics equipments, and the control systems.

Powerplants. - In proJectlng the camercial enghe state-of-the-art Into the 1970 time period for airplane operation with comrclally certificated en&es by 1973, it has been necessary to use propulsion hardware srd perfor- mance which are now considered to be at military development levels. The can- ponents of the primary propulsion system used in this ettiy have been chosen at the secondary level of military developmtnt; Le., the equlpent would no longer be considered for an advanced military aircraft design. It was consl- dered that thle "derating" of military equipment establishes propulsion system component performance suitable for camercial operation In 19'7'3 with accep- table levels of rellabillty and mlntalnabillty.

It was assumed that a production version of a lightweight turbojet engine would be available for 1973 and that this engine would be acceptable as a pitch engine. For this role, the turbine inlet temperature was reduced from 2200°F to 1645°F to provide a performance margin for reliability and maintainability and to assure safe operations, since these engines also double as auxiliary powerplants.

A number of candidate primary engFnes were examined and the General Electric GE1 gas generator technology was considered representative of the engine technology that would be camnercially acceptable for 1973. Same general characteristics of this gas generator technology are presented in Table 4.

Avionics equipments.

- A survey of the electronic equipment manufactur- ing industry revealed that much of the airborne avionics hardware required for V/STOL aircraft is available today. Except for the all-weather takeoff and landing system and possibly the terminal area navigation system, the avionic equipment used for V/STOL aircraft will be essentially the same as for conven- tional aircraft. A camnunication and navigation equiment list was derived from XC-142A, canmercial helicopter, and conventional canmercial transport air- The FAA certification requirements were used as a craft equipment lists.

guideline for selecting the minimum equipment complement. The equipment used as a basis for determining weight and cost estimates is of recent design, and the specific equipment items are listed in Table 5.

Control systems. - The control systems for all airplanes developed in the study were designed to be capable of providing the "desired" levels of control power as specified by Reference 1 and presented in Table 1. It was found that for all airplanes, the use of "acceptable" levels of control power instead of "desirable" levels had onlv minor effects on the airplane designs because the furnishing of the "desired" levels of control power was not critical in sizing any propulsion system components.

Limited analyses were made to select the characteristics of stability augmentation systems. For this process, the airplanes have been broken into categories of airplanes having a hover capability, airplanes designed to oper- ate frcxn 1000 foot fields and airplanes designed to operate from 2000 foot The airplanes having a hover capability were determined to need dual fields.

rate plus displacement augmentation channels in pitch and roll, and a single rate channel in yaw. The airplanes operating from 1000 foot fields were deter- mined to require single channel yaw and roll dampers. The airplanes operating from 2000 foot fields were determined to require only a single channel yaw damper for their stabilization systems. Although extensive analyses and simulation studies would be required to confirm the results of these limited analyses, the results thus obtained are considered sufficiently accurate for a feasibility study.

Configuration descriptions. - As mentioned previously, the configurations developed for this study utilize three propulsion system concepts (1) the turboprop (2) the fan-in-wing and (3) the propulsive wing.

Turbopron nowered conceuts. - A typical 60-passenger turboprop airplane is shown in Figure 5. The turboprop airplanes have high wing arrangements and are powered by four turboshaft engines driving four propellers.

The wing is provided with leading-edge slats and full-span, 48% chord, double slotted trailing-edge flaps to give a high maximum lift capability and thus compensate for the high wing loading which is desirable for minimum direct operating costs.

A unit horizontal tail is mounted on the vertical tail, which consists of a conventional fin and rudder arrangement.

The fuselage has an oval cross-section and its length is established by the ccznbined requirements for cockpit space, passenger cabin and its facili- ties, and low drag.

Two doors are provided for access to the fuselage, and escape hatches are located on each side of the fuselage in the passenger cabin.

Two cubic feet of carry-on-baggage space per passenger are p&dad In Cargo and stow& baggage compartment access doors are the passenger cabin.

located at a convenient height on the lower side of the fuselage.

Space Is pr&d& for 1200 pounds of revenue cargo at a density of ten pounds per cubic foot,~st~bag~espaceisp~dedaes~ngeschpassengarcarrier,~ limits allowable without excess baggage charges.

The unique characteristics of the V/STOL turboprop airplanes are the tilting wing, the use of jet engines for longitudinal control awentation in all but the 2000 foot STOL, and a transmission system interconnecting all engines.

The general turboprop VTOL design hilosophy used by ITV is similar to LA that used in the develqent of the XC-l with one major exception, which is the use of jet engines for pitch control in place of the tail rotor. Two main reasons for this arrangement are the added safety margin provided by the redundant pitch systems; and, at the cruise condition, the drag is reduced by eliminating the tail rotor.

The engines are mounted directly aft of the propellers (Figure 6) keep- ing the cross-shafting unloaded except when an engine is out or unsymmetrical thrust is desired for control. The wing has no geanetric dihedral and the leading edge beam fran which the cross-shafting is supported is straight, thus eliminating the need for a gearbox between the left-and right-hand sets of engines.

For cruise flight, the turboprop airplane uses conventional ailerons, rudder, and a unit horizontal tail for control about the lateral, yaw and pitch axes, respectively.

For hover, pitch control is obtained fran the pitch engines, yaw control is obtained from differential deflection of the ailerons, and lateral control is provided by getting differential thrust fran the pro- pellers. During slow speed flight with the wing at incidence angles other than 0' or wO, a mechanical integrator is provided to canbine the outputs of these control producing devices in such a manner that the pilot always gets the roll, yaw and/or pitch manents that he has commanded.

Fan-in-wing concepts. - For this study, NASArestricted LTV to the study of "pure" fan-in-a airplanes in which all the gas generator power is diver- ted to drive the lift fans for hover and slow speed flight conditions. Nuner- ous fan-in-wing arrangements were studied and certain fundamental characteris- tics of these airplanes were learned.

Discussions with powerplant manufacturers led to the conclusion that for the 1970-19'75 time period, it would not be possible to connect more than two gas generators into one exhaust manifold system. It was also found that the DOC for the fan-in-wing airplanes reduced as the wing loadings increased, and/ or the aspect ratios decreased; therefore, considerable effort was expended in selecting propulsion system arrangements which would aid in minimizing aspect ratio and maximizing wing loading. The ability to duct more than two gas generators to a camnon manifold or to deflect a portion of the pr5mary gas generator thrust vertically and hence reduce the required fan sizes would pro- bably have made the design integration problems less difficult. It is also possible that the use of turbofan engines for cruise thrust, instead of the turbojet engine that is considered a part of the "pure" fan-in-wing principle, could have lowered the fuel requirements for these designs to levels that would permit considerably snaller airplanes. Since these innovations were not studied, they can only be pointed out as possible areas to improve the capa- bilities of the fan-in-wing airplanes.

A typical fan-in-wing airplane is shown in Figure 7. It is a high wing airplane powered by six wing-mounted turbojet engines driving four tip-turbine fans installed in the wing and one tip-turbine fan installed in the fuselage nose. The wing has trailing edge flaps on the inboard section and canbination trailing edge flap/ailerons on the outboard section. The unit horizontal tail (UHT) is mounted on the top of the vertical fin in order to keep the UH!Taway fran the exhaust of the inboard engines.

The fuselage, except for the differences necessitated by the installa- tions of the nose fan, is similar to the fuselage of the turboprop-powered airplanes, as are the crew and passenger accanrmodations.

The unique characteristics of the fan-in-wing airplanes are the tip- driven fans mounted in the wing and in the nose of the fuselage. These fans Several constraints dictated the wing plan- govern the configuration geometry.

form which has an essentially constant chord. The first constraint was the minimum fan/turbine diameter. Another constraint was the routing of the fairly large diameter ducts for the hot gas in the wing. After analyzing many options, including routing in front of and in back of the wing beam, it was determined that the best possible routing was to keep the hot gas ducts between the front and rear box beams, thereby keeping the wing depth, the wing chord, and struc- tural box beam weight to a minimum. It is to be noted that the XV-5A uses this approach, but it has single fans per wing making the problem simpler. The chord of a minimum chord wing is then simply the fan diameter plus the front and rear box beam and the length of flap. To provide a taper would require adding area since the wing chord is already a minimum at each fan. Wing sweep back is used to keep the thrust axes of the wing fans in harmony with the wing aerodynamic center.

The engine nacelles are located so that the hot engine exhaust gases are directed to the fan between the wing leading edge and trailing edge box beams.

This is in accord with keeping the hot gas ducts from cutting through the beams. The outboard engines are mounted in the conventional "under slung" nacelle below the wing. The proximity to the fuselage requires the inboard nacelles to be located above the wing.

Several low wing designs of fan-in-wing airplanes were studied and these were found to have good features, such as more direct ducting paths for the hot gas which drives the nose fan. Fran further study of these low wing arrangements, it was concluded that the basic gas generators would have their performance severely penalized during hover due to the reingestion of the heated exhaust gases; therefore, the low wing arrangement was selected for the STOL airplanes only, with the high wing arrangement used for the airplanes required to operate VTOL.

For this study, the fan-in-wing airplanes having a VTOL capability were fitted with 15 percent chord flaps, and the designs having only a STOL capabil- ity were fitted with 25 percent chord flaps.

The front and rear box beams are designed to provide the same strength and stiffness as a conventional single box wing. This increases the weight of the box approximately 30 percent, which corresponds to a 15 percent increase in wing weight. Figure 8 is a schematic drawing of the hot gas ducting and engine locations. As can be seen, with all engines operating, the hot gas frcan each engine is divided so that each wing fan absorbs the hot gas output of l-l/3 engines and the nose fan absorbs the hot gas of approximately Z/3 of one engine. The engine sizes are established by the requirement to provide a thrust/weight ratio of 1.15 with the airplane trinnned.and no maneuvering con- trol input on an 86'F day at sea level.

Figure 9 illustrates an engine-out condition (the outboerd engine-out is used since it requires the most corrective action to meet the roll trim and For this condition, all the hot gas of the operating control requirements).

outboard engine adjacent to the failed engine is directed to the outboard fan.

The two inboard fans are powered to their maximum capability. The nose fan is limited to the amount necessary to provide the required 20 percent of the pitching-acceleration. The remainder of the hot gas is then ducted to the out- board fan on the side opposite the failed engine.

With one engine out, the remaining 5 engines are operated at an emergency rating of 110 percent of the gas generator gas horsepower takeoff rating.

An important feature of the fan-in-wing propulsion system is the variable inlet turbine which is designed to operate through a range of hot gas flow of approximately plus or minus &' percent frczn the naninal gas flow rate. This propulsion system feature 13 known as llgas power exchange" or llpower transfer."

The prjnciple involved is to have one gas generator zwpply hot gas to more than one power turbine.

Each power turbine is mounted on the tip of a fan.

Varying the turbine inlet area differentially fran one power turbine to another and holding total turbine inlet area constant m&es more hot gas flow through one turbine than the other, thereby changing the fan speed and hence providing differential fan thrust.

Each engine is provided with a diverter valve so that its hot gas can be diverted fran the fan/turbine to a straight-through nozzle providing conven- tional jet thrust for cruise flight.

During cruising flight, pitch control is provided by the unit horizontal During slow speed flight operations, pitch control is provided by dif- tail.

ferential thrust between the nose and wing fans. During cruise, directional control is provided by the rudder, and lateral control is provided by the ail- erons. During slow speed flight operations, directional control is provided by differential movement of the vanes which direct the exhaust of each of the wing fans. Iateral control is provided by the gas power exchange system pre- viously described, which gives differential thrust between the wing fans.

Propulsive wing concepts. - A typical propulsive wing airplane is shown in Figure 10. The low wing STOL airplane is powered by six turbojets driving eight wing-mounted turbines which are shaft connected to eight wing fans and two fuselage mounted turbines which are shaft connected to two nose fans.

The relatively low aspect ratio w3ng is fixed at a 5” incidence. A 20 percent chord fan air deflection flap is located at the wing trailing edge. Unit horizontal tails are located on the wing boans.

A conventional fin and rudder vertical tail is located on the fuselage.

The crew and passenger accamnodations are essentially the same as for the other concepts evaluated in this study.

I II 111111 I I Some of the unique features of the propulsive wing concept include: Twin forward-'facing nose fans, located in the forward section of .

the fuselage, which operate in the cruise as well as the STOLmode.

Outboard tails mounted on wing bocans .

Efficient jet flaps .

Efficient propulsion system for a wide range of conditions .

High cruise Mach number (0.9) .

The gas generators are mounted one in each boom and two on each side of the fuselage (Figure Ill). The wing turbines are sized and arranged so that each absorbs one half of the gas generated by one engine; therefore,, four engines drive the eight wing fans. The fuselage turbines are sized and arranged so that each absorbs aJ2 the gas power frcm one engine; therefore, each inboard engine drives a nose turbine/fan combination.

Each engine is connected to the corresponding engine on the opposite side by a hot gas duct with a shutoff valve. During engine starts, the shutoff valves are closed, permitting each engine to be started in turn. After the engines are started, the shutoff valves are opened to maintain complete thrust symmetry in the event of an engine failure. The propulsive wing concept also The total thrust loss due to a failed uses the "gas power exchange" system.

engine is quite small, in the order of 8 to 10 percent with one engine out and the other five engines operated at 110 percent of the takeoff rating.

Figure 12 shows a section through one of the wing fans. The propulsive wing consists of fans mounted vertically within the upper and lower surfaces of the wing, behind the leading edge inlet air duct. Each fan is driven directly by a turbine mounted in the aft section of the ~i-ng. The straight- through fan air flow duct exits through a variable area nozzle to ensure effi- cient fan operation under a wide range of power settings.

The installation of the fuselage nose fans is similar in concept to the wing fans., The turbines driving the nose fans are mounted in the mid-section of the fuselage and connected to these fans by long shafts, thus eliminating the routing of hot gas ducts the entire distance fran the gas generators for- ward to the nose of the airplane.

The wing structure is greatly influenced by the propulsive wing concept and represents a departure frcxn conventional wing design. The main wing torque box is comprised of front and rear truss beams plus upper and lower stiffened and stressed skin panels. The beams occupy the fKU depth of the physically thick propulsive wing, providing a stiff structure. The wing torque box is located well forward of all the hot gas ducting; and a gas leak, should one occur, would not impair the integrity of primary structure.

in cruising flight, longitudinal control is provided by the two fully- powered unit horizontal tail surfaces.

During slow speed flight, pitch con- trol is augmented by differential thrust between the nose and wing fans. In cruising flight, lateral and directional control are provided by the flap/ ailerons and the rudder, respectively. During slow speed flight, lateral and directional control are augmented by differential deflection of the wing thrust vector as well as by differentially varying the magnitude of the thrust vector using gas power exchange.

Sixty passenger optimum airplanes. - Drawings of the sixty passenger optimum airplanes developed for this study are presented in Figures lj through The critical design conditions for these ten 60-passenger aircraft are 17.

shown in Table 6. It can be seen that takeoff requirements predaninate in sizing propulsion systems of aircraft having a VTOL capability, mainly because of VTOL thrust-to-weight ratio requirements. Cruise speeds for minimum DCC and landing conditions primarily influence STOL propulsion system sizing. The optimum cruise Mach number of the turboprops is approximately 0.6. Fan-in-wing configurations are limited to 0.8 cruise Mach number, and propulsive wing con- figurations to 0.9. High-speed wind-tunnel tests have substantiated the ability of the propulsive wing concept to cruise at a 0.9 Mach number.

Turboprop airplanes. - Drawings of the sixty-passenger turboprop airplanes are presented in Figures l-3 and 14. The geanetric similarity of the airplanes is evident. The wings on the VTOL and V/STOL airplanes tilt through loo", and they are fitted with dual pitch engines in the rear. The loo0 foot STOL air- plane has a wing that tilts to an angle of 20' for landing, and the airplane is equipped with one pitch augmentation engine. The 2000-foot STOL airplane does not have a tilting wing or any pitch augmentation engines.

The propulsion systems of the VTOL and V/STOL airplanes were sized by the requirement for a thrust-to-weight ratio of 1.05 on an 86°F day at sea level with the critical engine failed and the airplane at the design VTOL weight. The canbinations of wing loading and aspect ratio were selected to give a near minimum direct operating cost on the design stage length. A higher wing load- ing would give slightly lower direct operating costs, but then the transition stall margins would becane critical.

The propulsion systems of the two turboprop STOL airplanes were sized by cruise conditions which required high thrust-to-weight ratios to give the relatively high cruise speeds for minimum direct operating costs. Twenty degrees of wing tilt were used on the 1000 foot STOL airplane to permit it to meet its landing performance requirements. Only one pitch engine is required on the lOC!C-foot STOL airplane since its horizontal tail has sufficient control power to meet the reduced pitch control requirements with a critical (in this case, the pitch engine) engine failed.

The propeller characteristics of the turboprop airplanes were optimized for takeoff performance, a ground rule used by ITV for this study. As a result, 140 activity factor, 0.5 integrated design lift coefficient blades, and a teke- off tip speed of loo0 feet-per-second were used for all 60-passenger turboprop airplanes. During the optimization process for the turboprop designs, these propeller characteristics were not varied; but the engine-propeller canbination was run at the optimum rpm in cruise. It was found that this optimum cruise rpm was about 75 percent of the takeoff rpm. After it was determined that for the turboprop STOL airplanes, cruise performance was critical for sizing the propulsion system rather than takeoff performance, it was considered appro- priate to see if better propeller characteristics could be .selected since the characteristics of the selected turboshaft engine show a rapid drop in avail- able power as the cruise engine rpm is reduced. The effects of propeller tip speed, coupled with the 100 percent free turbine tip speed, activity factor, and design integrated lift coefficient on cruise speed, were investigated.

Figure 18 shows the variation in cruise speed as a function of propeller geoanetric and operating characteristics. A change of approximately 45 knots can be realized by changing the maximum design tip speed from 1000 to 700 feet- per-second, the activity factor from 140 to 100, and the integrated design lift coefficient from 0.5 to 0.3.

The effect of Mrying power and the maximum propeller tip speed on take- off performance of the 2000-foot design is shown in Figure 19 for various flap settings. It can be seen that the takeoff performance can be considerably better than the design landing distance. Decreasing the maximum tip speed has little effect on takeoff distance for these airplanes.

It was also desired to determine the necessity of cross-shafting for the turboprop 2000-foot STOL airplane. The sensitivities of the roll and yaw requirements to true airspeed for this configuration were studied to deter- mine the approximate speed where cross-shafting would no longer be needed.

Figure 20 presents the net rolling manent available from a spoiler roll con- trol system versus true airspeed. The effectiveness of six percent chord spoilers is shown for trim angles of attack of 5, 10, and 15 degrees. For any speed above 65 knots, the maneuvering roll control requirement can be met at reasonable angles of attack and with a 6 percent chord spoiler system.

However, Figure 21 shows that the yaw control requirement is more critical than the roll control requirement. These yaw control requirements have been developed fran test data on a two-engine configuration as reported inReference 10. The yaw- ing mcanents available frm a plain flap rudder with boundary layer control and a double hinged rudder are shown as functions of true airspeed. With the re- quired spoiler control input to trim out the resulting rolling mment at a trim angle of attack of 10" and with the number one engine out and the propeller feathered, the resulting yawing moments at maximum power show a relatively large vertical tail area or a sophisticated rudder system was required to eli- minate cross-shafting below approximately 75 knots.

Since the yaw control available was considered marginal in the opera- tional speed regime of the 2000-foot STOL design, cross-shafting was used on all turboprop configurations.

Fan-in-wing airplanes. - Figures 15 and 16 show drawings of the sixty- passenger fan-in-wing airplanes. These airplanes are,equipped with six gas generators, four wing fans and a fuselage nose fan. The airplanes having a VTOL capability have high wing arrangements and a "tee-tail;" whereas the STOL airplanes have low wings and more conventional tail configurations.

The gas generators of the VTOL and V/STOL airplanes were sized at the design VTOL weights jointly by (1) the requirement for a thrust-to-weight ratio of 1.0 on an 86°F day at sea level with the most critical engine inoperative and the reduced simultaneous control inputs required being developed, and (2) the requirement for a thrust-to-weight ratio of 1.15 on an 86°F day at sea level with all engines operating and no control input other than that required for The fans were sized by the requirement for a thrust-to-weight ratio of trim.

The wing loading and the aspect ratio are 1.15 on an 86°F day at sea level.

the maximum and minimum, respectively, that can be obtained with sufficient structure and s-pace in the wing for the propulsion system canponents. Even higher wing loadings and lower aspect ratios, if they were possible, would give lower direct operating costs.

The propulsion systan of the fan-in-wing lOO&foot STOL airplane was sized to give the thrust-to-weight ratio required to permit flight at the design landing speed, including the specified margins for control and stall with the critical engine failed. The propulsion system of the fan-in-wing 2000-foot STOL airplane was sized to give a minimum direct operating cost with the takeoff performance requirements being almost as critical as the cruise requirements. The direct operating costs of the fan-in-wing STOL airplanes would be lower if their wing loadings could be increased or their aspect ratios decreased.

The fuel weight of the fan-in-wing airplanes was found to be one of the factors causing the weight of these airplanes to be relativeJy large. Since the fuel reserves required by V/STOL short-haul transport airplanes have not been firmly established, it was considered desirable to determine the influence this factor could have on the size of the fan-in-wing V/STOL airplane. The reserve fuel for this airplane is 33 percent of the total fuel weight, and approximately 7 percent of the gross weight of the airplane. Figure 22 shows the sensitivity of the fan-in-wing V/STOL airplane empty weight and gross weight to change in reserve fuel requirements. This figure shows a 20 percent change in the reserve fuel will change the takeoff and empty weight by approximately 3.5 percent.

Propulsive wing airplanes. - The propulsive wing airplanes, shown in Figure 17, are gecmetrically similar, differing only in size because the lOOO- foot STOL airplane has a higher design thrust-to-weight ratio. The propulsion system of the propulsive wing 100%foot STOL airplane was sized to give the thrust-to-weight ratio required to permit flight at the design landing speed, including the specified margins for control and stall with the critical engine failed. The prapulsion system of the propulsive wing 2OOGfoot STOL airplane was sized by cruise conditions which give a minimum direct operating cost on the design stage length. The number of wing fans used on the propulsive wing airplanes were optimized to give a near minimum direct operating cost.

Since only limited data were available to support the design of.the pro- pulsive wing airplanes, it was considered appropriate to evaluate the sensiti- vity of the propulsive wing 2OOC-foot STOL airplane to changes in skin friction Figure 23 shows the variation in the design and propulsion system efficiency.

takeoff weight with changes in the skin friction drag and the propulsion system efficiency for the propulsive wing 2000-foot STOL airplane, and it shows a 10 percent increase in the skin friction coefficient causes a 3.5 percent change in takeoff weight, and a 10 percent change in propulsion efficiency causes a change of approximately 5 percent in the takeoff weight.

Configuration design summary: - Table 7 summarizes some of the more im- portant physical characteristics of the ten 60-passenger optimum airplanes de- Reference to this table shows that the gross veloped during this study.

weights of these airplanes vary from approximately 53,000 pounds for the tur- boprop 2000-foot STOL airplane to wer 95,000 pounds for the fan-in-wing VTOL airplane. A breakdown of gross weights into the five major categories of structure, propulsion, fixed equiment, fuel, and useful load less fuel (Figure 24) shows that all designs have c-arable structural weight ratios which vary frm 27 percent to 29 percent of the gross weight. The actual weights of fixed equipment and useful load less fuel were almost the same for all designs; therefore, the percentage of gross weight assigned to these items varies inversely with the airplane's gross weight. The predominant factors in establishing the gross weight were then the sum of the propulsion system and These factors varied fran 23 percent for the turboprop 2000- the fuel weights.

foot STOL airplane to 40 percent for the fan-in-wing VTOL airplane. Table 8 presents a detailed estimated weight breakdown for each of the 60-passenger airplanes.

The powerplant sizes required by these designs are considered reasonable for the 1370 time period.

Economic Analyses For the 60-passenger airplanes, the direct operating costs per passenger seat-statute mile were predicted as a function of stage length. A 2000-hour per year utilization and a non-productive time of lo.25 minutes were assumed, in rccordnnce with Reference 1. Figure 25 is a plot of these direct operating costs for each of the 60-passenger airplanes. These costs were predicted using the method described in Reference 10 with the modifications specified by NASA in Reference 1. For the turboprop airplanes, the direct operating costs (DOC) vary from approximately 2.2 cents per seat-mile for the 2000-foot STOL to approximately 2.7 cents perseat-mile for the VTOL at a 500-mile stage length.

As the operating stage length is reduced to 100 miles, the DOCvary fran 3.4 cents per seat-mile for the 2000-foot STOL to 4.1 cents per seat-mile for the VTOL. In general, the DCC of the VTOL airplane are approximately 23 percent greater, the V/STOL airplane DOC are approximately 15 percent greater, and the lOOO-foot STOL airplane DOCare approximately 9 percent greater than the DOC (It should be noted that these cost data were of 2000-foot STOL a-lane.

generated assuming that all designs were fitted with propellers optimized for the takeoff performance condition.)

For the fan-in-wing airplanes, the DCC vary frczm 2.8 cents per seat-mile for the 2000-foot STOL airplane to approximately 3.6 cents per seat-mile for the VTOL airplane at a 500.mile stage length. When the operating stage length is reduced to 100 miles, the DOC of the 2000-foot STOL airplane increase to over 5 cents per seat-mile and the DCX of the VTOL airplane increases to approximately 5.8 cents per seat-mile. The seat-mile costs of the lCOO=foot STOL airplane and the V/STOL airplane are approximately 7 percent greater than the seat-mile costs for the 2000-foot STOL airplane.

For the propulsive wing airplanes, the DOC vary frcm 1.9 cents per seat- mile for the 2000-foot STOL airplane to 2.3 cents per seat-mile for the lOOO- foot STOL airplane on a 500-mile stage length. At a 100-mile operating stage length, the DOS of the 2000-foot STOL airplane are approximately 3.4 cents per seat-mile and the DOS of the lWO-foot STOL airplane are 4.2 cents per seat- At all stage lengths, the DOC of the l-foot STOL airpiane are approx- mile.

imately 25 percent greater than for the 20000foot STOL airplane.

The initial airplane costs used in predicting the DOCare presented in Figure 26. These costs were predicted to vary from 2.17 million dollars for the turboprop 2COO-foot airplane to 4.63 million dollars for the fan-in-wing VTOL airplane. The fan-in-wing and propulsive wing airplanes were predicted to cost approximately 80 dollars per pound of empty weight. The turboprop VTOL and V/STOL airplanes were predicted to cost approximately 74 dollars per pound of empty weight, and the turboprop STOL airplanes were predicted to cost approximately 66 dollars per pound of empty weight. These initial airplane costs were based on a quantity of 300 airplanes being bought with no research and development work being required to extend the technical state-of-the-art.

In developing the direct operating costs, the maintenance manhours per flight hour used for each of the 60-passenger airplanes are presented in Figure 27. It can be seen fraa this bar chart that the airframe maintenance was predicted to be approximately the same for all designs, varying fran about 8 maintenance manhours per flight hour for the propulsive wing 2000-foot STOL airplane to almost 10 maintenance manhours per flight hour for the turboprop VTOL airplane. The major differences in total maintenance manhours per flight hour between one airplane and another were due to the propulsion system main- tenance requirements. The major factors causing the propulsion system main- tenance requirements to vary were the number and size of gas generators. The propulsion system maintenance manhours per flight hour varied fran approxi- mately 5 for the turboprop 2000-foot STOL a-lane to nearly 15 for the fan-in- wing VTOL airplane.

Figure 28 and Table 9 show the breakdown of direct operating costs into the ccxnponents of depreciation of flight equipment, direct maintenance, and flight operations. The depreciation costs per seat-mile were the least for the propulsive wing 2000-foot STOL airplane. This occurs because of its nani- nal initial cost and its very high cruise speed (Mach Number 0.9). The fan-in- wing VTOL airplane has the highest depreciation, primarily due to its high, initial cost.

The cruise speed of this airplane (Mach Humber 0.8) was unable to canpensate for its high initial cost.

The direct maintenance and flight operations costs were lowest for the propulsive wing 2000-foot STOL airplane and highest for the fan-in-wing VTOL airplane. The high speed of the propul- sive wing 2000-foot airplane cabined with its relatively low maintenance requirements and nominal fuel consumption kept these canponents of seat-mile costs to a minimum. The size of the fan-in-wing VTOL airplane was sufficient to keep these major canponents of seat-mile costs to a maximum. Turboprop designs have relatively low to modest initial costs, direct maintenance man- hour requirements and fuel costs; but the modest cruise speed of these designs counteracts these ingredients of seat-mile costs. Rematching of the propeller and engine operational rpm range, which has previously been shown to increase the cruise speed of the turboprop STOL airplanes by approximately 45 knots, can reduce the DOCby approximately 10 percent for a 5OGmile stage length.

Operations Analyses For this study, operations analyses have included the determination of the far field noise characteristics of each of the 66passenger airplanes, a determination of V/STOL air traffic control problems, and a determination of some of the requirements of an all-weather takeoff and landing system for V/STOL aircraft.

Far field noise environment. - The noise generated by V/STOL aircraft loans as one of the major stumbling blocks to community acceptance of V/STOL As a result, the far field noise characteristics short-haul transport systems.

of the ten g&passenger airplanes have been estimated.

Typical perceived noise level directivity contours are presented in These contours were predicted assuming the airplanes were Figures 29 to 32.

developing maximum power fram all engines while static at ground level.

In the development of these contours, it has also been assumed that there are no wind effects, the air is dry and there are no terrain features which would affect noise transmission characteristics. Figure 33 shows the variation in the maximum perceived noise level with distance from the airplane, as measured along the radial line at which the distance is the maximum for a given PNdb.

It may be noted that the PNdb for all aircraft at a distance of 1000 feet are approximately 112, the maximum level considered acceptable adjacent to air- ports. At distances greater than 2/3 of a mile, the turboprop airplanes have noticeably higher noise levels than the fan-in-wing and propulsive wing air- planes. The turboprop 2000-foot STOL airplane has no jet engines to augment longitudinal control during slow speed flight; therefore, it has much lower noise characteristics than the turboprop V/STOL airplane. Analyses have shown that these jet (pitch control) engines make large contributions to noise in the 300 to 600 cycles per second octave band and are primary contributors to noise in octave bands above 600 cycles per second. High frequency noise attenuates with distance at a higher rate than does low frequency noise; thus the noise differences between the turboprop 2000-foot STOL and the turboprop V/STOL airplanes at distances greater than one mile are evidence of the pro- pulsion system power output differences for these two airplanes; but the noise differences between these two airplanes at distances of less than one mile, are evidence of the high-frequency noise generated by the pitch engine on the turboprop V/STOL airplane.

Perceived noise level contours for takeoff and landing are presented in Figure 34 for the turboprop V/STOL, in Figure 35 for the turboprop 2000-foot STOL, in Figure 36 for the fan-in-wing V/STOL, and in Figure 37 for the pro- pulsive wing 2OOO-foot STOL. These contours represent the noise levels that would be detected on the ground along the airplane flight path; and it has been assumed that during takeoff, takeoff power is applied on all engines and Luring landing, it the airplane makes a climbout at a 20" flight path angle.

is assumed that the airplane approaches at a 10" descent angle with the required approach power.

Propeller noise estimates have been developed using the methods of references I2 and l-3 with a delta correction factor applied.

The delta cor- rection factor was the difference between the measured and estimated noise characteristics of the XC-142A using these same estimating methods. Gas generator intake noise was estimated using the method of Reference 14 and the spectrum distribution described by Reference 15; and the noise characteristics of the turboprop and pitch engine exhausts were estimat.ed using the method of References 13 and 16 with a correction factor applied to account for the dif- ferent exhaust velocities. The noise generated in fan intakes was estimated using the methods of References 14 and 17, and the noise generated by the fan exhaust was estimated by the method of Reference 13. The fan exhaust noise of the propulsive wing airplanes was increased three decibels to account for the additional noise generated due to the flap being deflected in the exhaust slip- stream.

- Traditionally, new aircraft have been required to Air traffic control.

fit the air traffic control system rather than modifying the air traffic con- During the enroute mode of flight, a trol system to fit the new aircraft.

V/STOL airplane will be similar to conventional aircraft; and conventional enroute navigation aids and air traffic control systems should suffice. How- ever, terminal air traffic control modifications should be considered for a V/STOL short-haul transport system.

Metropolitan canplexes, with many airports for conventional air traffic, will find their conventional air traffic control systems unable to properly cope with novel flight capabilities and requirements of a V/STOL short-haul air transport system. A V/STOL short-haul transport airplane will not be able to econanically tolerate lengthy air traffic control flight delays, and an attempt to mix V/STOL traffic with conventional air traffic could impose severe econunic problems on the conventional air transport system as well. The V/STOL airplane will go in and out of airports using very steep ascent angles, because, economically, the V/STOL airplane should climb to relatively high cruise altitudes even for short flights in order to minimize DOG. As an example, this study has found that V/STOL airplanes should cruise at their design cruise altitudes (25,000 to 35,000 feet) if the range is greater than 150 miles, and the optimum cruise altitude drops to approximately 11,000 feet if the range is reduced to 50 miles. It is projected that movements of V/STOL aircraft in congested terminal areas, which would include these large and rapid altitude changes, would severely tax the capabilities of conventional air traf- fic control systems; therefore, the capability to effectively handle V/STOL air traffic movements should be developed.

All-weather takeoff and landing system. - One of many requirements for gener&public acceptance of a V/mhort-haul air transport system will probably be that it be able to maintain regular and dependable schedules under all-weather conditions. V/STOL aircraft, as a result of their lower operational speed capabilities in the terminal area, will have a potential for operating safely to lower weather minimums than conventional aircraft; but this potential will be maximized only if a suitable V/STOL all-weather takeoff and landing system is available. DOGbenefits will also accrue to a V/STOL short-haul transport system that has an all-weather takeoff and landing systemthat can handle V/STOL traffic effectively because such a system would be expected to reduce the non-productive times associated with takeoff and landing functions.

I -

Tasks that a V/STOL all-weather takeoff and landing system will probably have to perform include providing obstacle clearance, keeping air traffic flowing smoothly while maintaining safe flight margins for all aircraft, and minimizing ground noise generated by the V/STOL aircraft during takeoff and landing. Many advances are being made in conventional all-weather takeoff and landing systems, and these advances will have increasing applicability to STOL aircraft at the longer STOL design field lengths. These conventional all- weather takeoff and landing systems will probably not suffice for airplanes having a VTOL or a short design field length capability because it is expected that the operator of such aircraft will want the system to permit multiple, simultaneous approaches from any direction and at varying glide slope angles - a capability not being designed into the conventional system.

Using helicopter flight operations, fixed based V/STOL aircraft simula- tor studies and flight experience on the XC-142A, LTV has predicted that the following information should be displayed to the V/STOL aircraft pilot and/or fed into the autopilot during an all-weather takeoff or landing.

a. Angular position with respect to the takeoff or landing point.

Absolute altitude above the takeoff or landing point.

b.

c. Distance from the takeoff or landing point.

d. Velocity with respect to the takeoff or landing point.

e. Angular rates f. Attitude, airspeed, and heading.

An all-weather takeoff and landing system having these capabilities includes not only the ground based IL!3 equipment, but also the as-yet-undefined airborne sensors and instrumentation and the aircraft control and stabilization systems.

Providing provisions for the airborne components of an all-weather takeoff and landing system and integrating these components into the aircraft may have a strong influence on the design of V/STOL aircraft, but the extent of this im- pact can not be predicted until detailed characteristics of the system are hlOWTl. It also can not be predicted whether any one V/STOL concept can per- form all-weather takeoffs and landing maneuvers better than any other concept, and this determination can only be established with extensive operational tests.

Operational analyses conclusions. - Noise loans as a potential coamnrnity acceptance problem for V/STOL short-haul transport airplane systems.

V/STOL short-haul transport aIrcraft will be able to use esdsting enroute air traffic control systems; but new terminal air traffic control systems will probably be required if V/STOL systems are to have practical economic characteristics.

Special all-weather takeoff and la&ing systems will probably be regzlred to take the mzzdmum advantage of the potentials offered by V/STOL aircraft. The operational analyses made in this study show that considerable additional research is required, but no one V/STOL concept appears to have an operational advantage over any other concept.

Additional Study of Selected Designs As a result of the econanic analyses, operations analyses, and technical judgement, four of the ten 60-passenger airplanes were considered to warrant additional study. The four considered to warrant additional study were the propulsive wing 2000-foot STOL airplane, the turboprop 2000-foot STOL airplane, the turboprop VTOL airplane and fan-in-wing V/STOL airplane.

The propulsive wing 2000-foot STOL airplane was selected for additional study because it had the lowest direct operating cost of any of the airplanes studied. Since the propulsive wu concepts are supported by very little test data and their technical feasibility has not been proven by flying aircraft, the turboprop 2000-foot STOL airplane was also selected since its direct oper- ating costs were the second best. The technical feasibility of this concept has been proven by flying aircraft.

The turboprop VTOL airplane was selected for further study because of, (1) the apparent military interest in VTOL aircraft, (2) only a slight weight and cost penalty when compared to V/STOL, and (3) the mass of data available to support the design of such an aircraft. The fan-in-wing V/STOL airplane was selected for further study for reasons similar to those used in selecting the turboprop VTOL with the exception that the V/STOL airplane was chosen rather than the fan-in-wing VTOL airplane because the VTOL was so large that it was not considered canpatible with the other designs being given additional study.

Extended econanic analyses were made of each of these airplanes, and 90- and XX-passenger versions were designed.

Extended econanic analyses. - The extended econanic analyses made of these four designs included evaluations of the influences of non-productive time, annual utilization, a combined military and civil buy, and gas genera- tor costs on direct operating costs.

Non-productive time. - For this study, the block time is measured from the time the airplane starts its taxi frcun the passenger loading ramp at the point of origin until it stops at the passenger unloading ramp at the air- plane's destination. The non-productive time or fixed time is the time lost in making air maneuvers plus the time spent for ground taxi, waiting for air traffic control clearances, etc. The major effect of non-productive time is to reduce the block speed and hence increase the direct operating costs. The following equation illustrates the influence of the non-productive time on the block speed: R Vb = to + tf Where: vb = block speed in miles per hour R = range in miles

I

the time to fly the route, with no non-productive time, to = in hours tf = the non-productive time in hours The time to fly the route, to, can be approximated by the equation R to = 7 cr Where Vcr = the cruise speed in miles per hour Thus the equation for Vb can be approximated: R Vcr v = R+V t cr f As the fixed time approaches zero, the block speed approaches the cruise speed, or when the range is large cqared to the product of cruise speed and fixed time, the block speed approaches the cruise speed. Ry contrast, as the and for a non-productive time not equal to zero, the block range decreases, speed is considerably less than the cruise speed. As an example, for a cruise speed of 500 mph, a range of 50 miles and a fixed time of 10-l/4 minutes, the block speed is only 37 percent of the cruise speed. If the range were 500 miles instead of 50 miles, then the block speed would increase to over 85 per- cent of the cruise speed.

Figures 38 through 41 show the variation of DOCwith stage length for non-productive times varying from 4 minutes to 15 minutes for each of the four airplanes considered to warrant additional study. These curves are for a utilization of 2000 hours per year. At the design stage length of 500 miles, the DOC for a &-minute non-productive time is approximately 93 percent of the DOCwith a 10-l/4 minute non-productive time for each of these four airplanes.

At a 50-mile stage length, the DCC for a h-minute non-productive time is approximately 75 percent of the DOCfor a 1%l/Lminute non-productive time for each of these airplanes. The variation of non-productive time has an increasingly important effect on DOC as the design stage length is reduced, but no one airplane is apprecFabJy more sensitive than any other to the varia- tion of non-productive time.

Annual utilization. - For all curves presented in this report, the annual Figures 42 utilization is 2000 hours per year, unless otherwise stated.

through 45 show the variation of DOCwith annual utilization for each of the four airplanes considered to warrant additional study. In general, it can be concluded that increasing the annual utilization from 2000 hours per year to reduces the DOCto approximately 80 percent of the DOC 4000 hours per year, for 2000 hours annual utilization at all stage lengths; and the variation of annual utilization does not show advantages for any one airplane.

Canbined civil-military buy. - In making an assessment of a combined civil-military buy on DOC, Reference 1 directed that the civil buy alone would be for 300 airplanes, but the canbined buy would include 600 airplanes. It was assumed, based on limited side studies, that the civil airplanes could utilize 75 percent of the non-recurring design, development, and testing per- formed on the military airplane when there was a combined buy. Figure 46 pre- sents the influence of the canbined civil-military buy as ccmpsred to the civil - only buy on DOCfor the four airplanes considered to warrant addi- tional study. It can be seen that the canbined buy reduces DOC approximately 15 percent for each of these airplanes with no one airplane having a decisive advantage over any other.

Insofar as the probability of a military buy is concerned, the following points should be considered. The military has shown a reluctance to buy STOL airplanes capable of operating frcan 2COO-foot airfields. The military services have conducted experiments on modifying existing airplanes to develop such per- formance capabilities. These operational capabilities have been demonstrated in spite of some undesirable flying qualities; but there has been no apparent move by any of the services to remedy these minor deficiencies and procure such vehicles.

The XV-5A and the XC-142A airplanes have been bought by the military services in order to gain operational experience with vehicles having a VTOL capability. The military is trying to determine just how such vehicles might better improve the operational effectiveness of military units. There is little question about the military being able to gain effectiveness by using V/STOL vehicles, but the question that remains is, "Will the increased effec- tiveness justify the increased costs of such vehicles?" Costs (i.e., total system costs) will be so criticalto this decision that it is predicted that the military will be unwilling to ccmpromise a first generation V/STOL vehicle design for a potential joint civil-military buy. The operational costs for a civil version are also expected to be so criticalto the success of a commer- cial V/STOL transport that the civil operator cannot afford a canpromise in his vehicle in order to get a combined buy, and it is not considered likely that the design conditions for a military V/STOL airplane would result in an airplane that would have operational costs on ccamnercial routes that could be caupetitive even though lower initial costs would result from a joint buy.

Thus, it is concluded that a ccmbined civil-military buy will be doubtful for any first generation V/STOL aircraft, but it is expected that this pattern will change with subsequent generation aircraft.

Influence of gas generator costs. - Since the propulsion system is so critical to the successful design of a V/STOL aircraft, a study was made of increasing the costs of gas generators by 100 percent and reducing the gas generator costs by 50 percent for the four airplanes considered to warrant additional study. Figure 47 presents the variation of DOCwith these gas generator costs. Increasing the gas generator costs 100 percent causes the DOCfor all airplanes to increase approximately 13 percent, Reducing the gas generator costs by 50 percent reduces the DOCby approximately 6 percent. The influence of the propulsion system costs on DOC are not as severe for the pro- pulsive wing 2000-foot STOL airplane, but this slight advantage is not suffi- cient to be decisive in favor of this airplane as ccmpared to the other air- planes.

- - L- - IQ-pathetical route analysis. - In order to evaluate the airplanes con- sidered to warrant additional study in an operational environment, NASA speci- fied a hypothetical route (Figure 48) for which the performance and DOC characteristics of these airplanes were to be calculated. Two assignments of non-productive time were made for this study. One was a non-productive time assignment of 10-l/4 minutes for all route segments (as specified in Ref- The other, dependent upon the field length performance of the air- erence 1).

craft, was as follows: 1. For a short takeoff, three minutes were used - two minutes for taxi from the passenger loading area to the end of the runway and one minute for the takeoff and acceleration to the climb speed.

2. For a vertical takeoff, two minutes were used - one minute for taxi from the passenger loading area to the takeoff ‘area, and one minute for the takeoff and acceleration to the climb speed.

For a short landing, 7-l/4 minutes were used - four and one-fourth 3.

minutes for get-tip? into the traffic pattern and gettirg aligned with the runway, one minute for the landing itself, and two minutes for taxi from the runway to the passenger loading area.

1:. For a vertical landing, two minutes were used - one minute to des- cend and decelerate from the let-down speed at an altitude of 1000 feet to the landing touchdown (this is performed as a straight-in approach), and one min- ute for taxi from the touchdown point to the passenger loading area.

For the variable non-productive time analysis, the V/STOL airplane is operated with a short takeoff and a vertical landing on segments A-B and D-E, VTOL on segments B-C, E-F, and F-A, and STOL on segment C-D of the route shown in Figure 48. The VTOL and STOL airplanes are operated with vertical and short takeoff and landings, respectively, on all route segments. Figure 49 shows the power of non-productive time on the route block time when non-productive time has been computed as described. As nn example, this analysis shows that, with a four minute non-productive time, the turboprop tilt-wing VTOL airplane cruis- ing at 350 knots has a route block speed almost equal to that of the propulsive wing 2000-foot STOL airplane cruising at approximateb 520 knots with a 10-l/4- minute non-productive time. If the propulsive wing airplane could reduce its non-productive time to four minutes, its route block speed would increase from 313 miles per hour to 390 miles per hour. Only a 6 percent improvement in route block speed, from 350 to 370 miles per hour, can be realized by giving the fan-in-wing V/STOL airplane a VTOL non-productive time.

The DOC for flying the complete route are presented in Figure 50. For this chart, the variable non-productive time schedule assumes lo.25 minutes for all except the VTOL segments, and 4 minutes for the VTOL se@nents. The most notable point shown on this chart is that the DOC of the turboprop VTOL and propulsive wing 2000-foot STOL airplanes are approximately equal for the variable non-productive time assumptions used. The DOC for the turboprop powered 2000-foot STOL airplane are approximately 6 percent lower than those of propulsive wing 2000-foot STOL airplane and the turboprop VTOL airplane when the variable non-productive time schedule is used.

$JO- and XXI-passenger versions. - 90- and l20-passenger versions of these four airplanes were developed to determine the effects of size changes.

The physical characteristics of these designs are presented in Table 10. The major changes resulting frcm those increased passenger loads are the increase in the number of engines and propellers frcm four to six as the design pas- senger load is increased to 90 on the turboprop VIOL airplane, and the in- crease in the number of fans fran ten to twelve as the design passenger load is increased to 120 on the propulsive wing 2OOGfoot STOL airplane.

Figure 51 shows the ratio of the gross weight of airplanes designed for other passenger loads to the gross weight of the airplane designed for 60 passengers. This figure shows that as the passenger load is doubled for the fan-in-wing V/STOL airplane, the design gross weight increased by approxi- mate* 65 percent.

For the turboprop VTOL airplane, the design gross weight increases by 45 percent as its passenger load is increased by 50 percent; and as the passenger capacity is doubled, the design gross weight increase is 80 percent. This change in slope occurs because a transition is made fran four to six propellers in going fran 60 to 90 passenger design loads; but the six propeller arrangement is still adequate for the XX-passenger design load.

The propulsive wing 2000-foot STOL airplane gross weight increases approxi- mately 65 percent as the passenger load is doubled; and the turboprop 2000-foot STOL airplane gross weight increases approximately 60 percent as the design passenger load is doubled. Thus it is seen that the growth character- istics of these four airplanes are catrparable except for the turboprop tilt- wing VTOL which has a noticeably higher growth factor caused by increasing the number of propellers and engines. The detailed estimated weights for these 90- and XX-passenger airplanes are presented in Table Il.

Figure 52 presents the direct operating costs on the design stage length for the 60-, 90-, and X%-passenger versions of these four airplanes. This chart shows that increasing the design passenger load fran 60 to I.20 passengers decreases the DOCto approximately 66 percent for the fan-in-wing V/STOL air- plane. For the turboprop VTOL airplane, the DOC of the l20-passenger version is 76 percent of the DOC of the 60-passenger version. For the propulsive wing 2000-foot STOL airplane, the DOC of the XX-passenger version is 57 percent of the DOC of the 60-passenger version; and for the turboprop 2000-foot STOL point airplane, the DOC of the XXI-passenger version is 63 percent of the DOC of the 60-passenger version. Thus, increasing the design passenger load benefits the DOC of the propulsive wing 2000-foot STOL airplane the most, and the turboprop VIOL airplane the least.

Fran this study it is concluded that the propulsive wing 2000-foot STOL airplanes can best adapt to design passenger loads greater than 60.

Specific Research Requirements Fran these design studies, the following specific items of research required to assure the timely deve1qFanen-t of prQnising V/STOL short-haul trans- port airplane concepts are identified; and these research items are divided into two categories. One category includes those items of research that are applicable to specific V/STOL propulsion system concepts, and the other cate- gory includes those items of research that are applicable to all V/STOL con- cepts studied.

Research applicable to specific V/STOL concepts. - The following items of research are applicable to the specific concepts evaluated in this study.

Turbwrop V/STOL concepts. - Considerable data are available to guide the designer of turboprop V/STOL concepts, but additional research may permit more nearly optimum designs and thus slightly lower direct operating costs; but it is not antic.ipated that marked extensions of the technical state-of- the-art could be gleaned fran such research. The following are specific areas where research efforts are considered appropriate for turboprop V/STOL short-haul transport aircraft.

. Recent experience at LTV has uncovered the fact that an accurate methodology for predicting the static thrust performance of propellers does not exist. This study has shown the static propeller performance character- istics to be critical to the design of VTOL turboprop aircraft; therefore, data are required which will permit an accurate assessment of the effects of propeller characteristics on static propeller performance.

The turboprop V/STOL short-haul transport airplanes will have lower ;OC if they can cruise at higher speeds; therefore, data are required to accurately define canpressibility effects on airplane-propeller interference in cruise flight, thus permitting proper airplane and propeller tailoring for improved flight performance at moderate subsonic cruise speeds.

Data are required which will accurately define the limits on the propeller-wing relationships for acceptable transition performance of tilt- wing aircraft. Data should be able to answer the questions: How far from the fuselage side can a propeller tip be?

What is the influence of propeller overlap or gap?

How far beyond the propeller tip can the wing tip extend?

What are the limits on the longitudinal positioning of the pro- peller plane with respect to the wing?

Fan-in-wing V/STOL concepts. - Considerable data are available to sup- port the design of fan-in-wing V/STOL short-haul transport aircraft, but addi- tional research may provide means of reducing operating costs by refinements in designs and by extensions of the existing state-of-the-art. The following are specific areas where research efforts are considered appropriate for V/ STOL fan-in-wing short-haul transport aircraft.

Data are required which define changes that must be made to per- mit t&driven fans to operate efficiently at higher pressure ratios than the present limit of 1.3. This will permit the design of fan-in-wing aircraft with higher wing loadings and lower aspect ratios, both of which contribute to lower direct operating costs.

Data are required which will permit design of a producible fan louver'system that turns this exhaust air to high angles efficiently. Such a capability will provide better takeoff performance for the STOL fan-in-wing airplanes, and better transition performance for the VTOL and V/STOL fan-in- wing a-lanes.

Data are needed which will guide the design optimization of the gas pier exchange system and its control. For this study, it has been assumed that a gas power exchange system will exist, and it has been assumed that aXI. the gas power exchange control devices will have the sensitivity re- quired by a V/STOL lateral control system. The parameters,which these con- trol sensing devices monitor to detect such emergencies as an engine failure, must be determined; the characteristics of the devices they control must be selected; and the required responses and sensitivities of the total system The integration of the gas power exchange control system must be obtained.

into the flight control system must also be accanplished; therefore, data are required to guide these design and design integration functions.

Data are needed which will guide the design of a hot gas inter- conneci system connecting several gas generators through a co~lppon plenum. This will also require research data which will guide the design of a multiple engine control system.

Data are needed which will guide the design optimization of a hot gas du&ing system. In particular, specific research is needed which will guide the design of hot gas ducting joints, expansion provisions, insulation, shielding, support, and flow control devices.

Data are required which will permit a more accurate assessment of the chkge in aerodynamic characteristics due to changes in configuration variables. This should include such items as the variation of induced lift when a wing, with more than one fan per wing panel, has these fans operating at different fan pressure ratios for extended time periods. Also, the ability to determine the variation of induced lift with unconventional wing/fan arrange- ments is considered desirable.

F'rouulsive wing V/STOL concents. - Little data are available to support the design of propulsive wing V/STOL concepts. Additional research efforts are required to provide much of the basic data for this concept; and it can be expected that extensions of the present state-of-the-art will develop fram active research efforts on this concept. The following are specific areas where research efforts are considered appropriate for propulsive wing V/STOL short-haul transport aircraft.

Data are required which will guide the design of optimum methods for deilecting the fan thrust downward for slow speed flight.

These data should consider both the internal flow characteristics within the duct, and the external flow characteristics around the wing; and they should be con- cerned with the induced lift characteristics for slow speed flight as well as the cruise flight efficiency of the concept.

Data are required which will guide the optimization of inlet desigd for propulsive wing inlets. These data should permit an assessment of the inlet configuration on propulsion system performance at all flight speeds, induced lift at slow speeds, and the drag rise characteristics of the airplane at high subsonic cruise Mach numbers.

Data are required to guide the optimization and establishment of design'requirements for exhaust systems for propulsive wing concepts. The variations in the slow speed induced lift characteristics as well as the cruise flight characteristics as functions of the propulsive wing exhaust con- figuration must be known.

Data are needed to guide the design optimization of a gas power excha&e system, a cmon interconnecting plenum, and a hot gas ducting system as has been mentioned for the fan-in-wing concept.

Data are required to permit accurate assessment of the change in aerodykunic characteristics due to change in configuration variables; e.g., data must permit assessment of the induced lift as fan pressure ratio, inlet aspect ratio, exhaust aspect ratio and/or flap deflection vary. A similar assessment capability for the influence of these same parameters on cruise performance is desired.

Research applicable to all V/STOL concepts. - The following are areas of specific research which are required to support the development of any V/STOL short-haul transport concept.

Data are required which will accurately define the control power requirknents for all flight regimes and size aircrafts.

Data are required which will define the cockpit display require- This will include ments s!or a VTOL, all-weather (zero/zero) landing system.

both the readout of data required by the pilot and the data accuracy.

Data are required to guide the design of foreign object damage (FOD) irotection devices on propulsion system installations and the deter- mination of techniques to minimize the reingestion of hot exhaust gases by gas generators for all configurations. An understanding of the canplete recircu- lation fields around all V/STOL aircraft is therefore required.

Data are required which can be used to establish design criteria for J&L and STOL airport surfacing.

Data are required that will permit the noise generated by the pro- pulsio; system to be a variable in the analysis process of optimizing a V/STOL propulsion system.

Data are needed to better define the origin of noise for all V/ STOL s&&m concepts. These data should be of such quality that the engineer will know how noise might best be reduced at its source.

Data are needed to describe the noise attenuation characteristics of various structural fabrication techniques and materials.

Research data are needed which will guide propulsion system manu- facturks in the reduction of weight of propulsion system components and re- duction in specific fuel consumption, especially for operations at low power settings.

Data are required to better define the non-productive times appli- cable 40 each of the V/STOL concepts. A minimization of the non-productive time will require an understanding of the operational limitations that contri- bute to non-productive time for each concept. These include such items as the time to start engines or fans, change configurations, and make instrument and it is anticipated that this item of approaches to a V/STOL IFR system; research will require considerable flight operational experience with many V/STOL aircraft types.

Reference 1, in harmony with the existing Federal Aviation Regu- lation;:, required that the required landing field length be the calculated minimum landing field length divided by 0.6. This factor has been determined to be appropriate for conventional aircraft but research effort should be expended to assure that this is an appropriate field length correction factor for V/STOL aircraft. This is an important parameter in determining the con- figuration of STOL aircraft and, hence, its magnitude should be established.

It is possible that this factor may change with each concept and/or with each design field length.

Data are required which will permit the engineer to make accurate estimaies of the static and rotary stability derivatives for all V/STOL con- figurations in all flight regimes.

Airworthiness Requirements In reviewing the capabilities of airworthiness requirements to cope with the novel flight capabilities of V/STOL aircraft, Federal Aviation Regulations Transport Category Airplanes," ard part 29 "Airworthiness Standards: Part 25, Transport Category Rotorcraft," have been used. In “Airworthiness Sttslldards : Part 25, the zero thrufh etalling speed of the airplane Is considered an operational limit, and many other flight characteristics are based on this e.g., the minimum allowable takeoff and approach speeds are functions speed; Such requirements are not appropriate for of the zero thrust stall speed.

V/STOL aircraft because they are designed to operate safely below this speed.

V/STOL aircraft may be influenced by ground effects more than will con- ventional aircraft, because sane V/STOL aircraft are able to fly in air dis- turbances that they are creating. Hence, Federal Aviation Regulations must take cognizance of this unique flight capability and assure that the V/STOL airplane always operates in a safe flight regime, especially in the ground effect region.

Numerous V/STOL concepts use gas generators to drive thrust producing devices through interconnected transmission systems. The propulsive wing and the fan-in-wing concepts have fans driven through an interconnecting system of hot gas ducts, and the turboprop aircraft have propellers driven by turbo- shaft engines through an interconnecting system of gear cases and shafting.

The existing Federal Aviation Regulations are concerned with engine failures where propellers are connected directly to the engine; hence the Federal Aviation Regulations must be modified to take cognizance of these intercon- nected transmissions systems and establish regulations which assure safety after failures likely to occur anywhere in the propulsion system.

In addition, design standards for canponents of the interconnecting transmission system must be established. Special attention must be given to the installation requirements for hot gas ducting systems to protect ducting and surrounding structure fran damage due to heat.

Where conventional aircraft can put fuel in their wings and thus keep it away frcen the passenger canpartments, many V/STOL concepts will be prevented fran putting fuel tanks in the wings and, hence be forced to put it adjacent to the passenger canpartments. The fan-in-wing concept, as an example, has its wing filled with propulsion system canponents. Federal Aviation Regulations must take cognizance of this potential safety hazard and assure that fuel system design standards will maximize safety where the fuel is located adja- cent to passenger canpartments.

It is considered appropriate to recamnend that a new set of Federal Aviation Regulations be established for V/STOL aircraft.

STuDYcoNXJSIoNS The technical feasibility of VTOL and STOL turboprop airplanes Short-haul transport aircraft has been proven by several aircraft.

using turboprop propulsion systems have been shown to be relatively light and to have relatively low Initial costs; but the modest cruise speeds of these aircraft reduce their attraction as potential commercial aircraft.

The technical feasibility of fan-in-wing airplanes has been demonstrated by one aircraft, and this study has shoun that short-haul transport aircraft of the 1970 time period using the lrpure" fan-in-wing The fan-in-wing aircraft do principle would be heavy and expensive.

have a good cruise speed capability, and thls combined with the ex- ternal appearance of a modern turbojet airplane would have strong passenger appeal ln connnercial operations.

The technical feasibility of the propulsive wing airplane has not been established by any flying aircraft, and only limited wind tunnel data are available to substantiate the potential technical The propulsive wing concept has been feasibility of this concept.

evaluated only in STOL short-haul transport airplane configurations in this study; and these airplanes have been found to be light, to have low initial and direct operating costs, and to have a high sub- sonic cruise speed capability. The apparent economy of operation for aircraft using this concept will be attractive to air transport oper- ators; and the unusual, but modern, appearance of aircraft built around this V/STOL concept probably should appeal to passengers.

Additional research could improve the efficiency of aircraft designed around the three V/STOL propulsion concepts evaluated in this study by permlttlng the design of more nearly optimum configurations.

IX is anticipated that the fan-in-wing airplanes would be partlculary affected by additional research and design studies. It is probable that slgnlficant reductions in size, initial costs, and direct oper- ating costs can be obtained by using turbofan engines for cruise, by deflecting a portion of the thrust downward for hover with the remaining thrust diverted to augment fan lift, and by unconventional fan/wing arrangements. It is also considered probable that additional research on the propulsive wing concept would permit some reductions on Its already low weight, Initial cost and direct operating costs; and it Is considered that such data would show a VTOL version using this propul- sion system concept to be very competitive with other VTOL short-haul transport airplanes.

The noise generated by V/STOL airplanes Is expected to be a major factor In obtaining coxuunlty acceptance of V/STOL short-haul transport sy8temf3. The far field noise characterlstlcs for the airplanes de- veloped ln this study are approxlmrtely the same at a distance of 1,000 foot from the airplanes with all engines at take-off power. Tpheatten- uation characteristics of propeller noise were considerably less than for the fan-in-wing and propulsive wing airplanes; therefore, the noise characteristics of the propeller powered airplanes were noticeably higher at distances greater than 2/3 mile. Additional experimental data ard analytical. analyses are needed to better define the ofigln of noise, to determine how to best suppress the noise at its origin, and to permit noise to be a primary design variable.

The non-productive time characteristics of V/STOL aircraft are crltlcalto the economy of operation of these aircraft; therefore, It is important to be able to Identify all the factors which contribute Examples of some of these factors Include the to non-productive time.

time lost In making configuration chsnges, the time lost In startlng or stopping any engines or fans that are used for slow speed operations only, the time lost In accelerating or decelerating through the transi- tion speed regime, the time lost in getting intermeshed with other alr- port air traffic, the time lost due to flying under Instrument flight The nebulous nature of rules rather than visual flight rules, etc.

these elements of non-productive time Is evident; and It Is projected that a considerable number of operational flight tests will be neces- sary to establish the ranges of magnitude that can .be expected for each of these variables. It is also projected that these magnitudes will be different for each V/STOL concept.

The airworthiness standards described by the existing Federal Aviation Regulations are considered inadequate to cope with the novel flight capabilities of V/STOL aircraft; therefore, it is recormnended that a new set of airworthiness standards be established specifically for V/STOL aircraft. The airworthiness standards for V/STOL aircraft should establish airworthiness safety objectives and hence be applicable to all V/STOL concepts.

REFERENCES "Study on the Feasibility of V/STOL Concepts for 1. Contract ~~~2-3036, Short-&u1 Transport Aircraft."

Part 25, Airworthiness standards: 2. Federal Aviation Regulations, "IYansport Category Airplanes."

"An Empirical Method for the Prediction of Airplane Drag 3. Johnson, R. P., Rand Report RM-IB~, 1963.

Divergence Mxh Numbers," 4. Sheridan, Hugo G., "Aircraft Preliminary Design Mthods Used ln the Weapons Systems Analysis Division," Navy Department, BuWepsReport No. R-5-62-13, June 1962.

"Generalized Method of Propeller Performance Estimation," Hamilton 5.

Standard Division of United Aircraft Report No. PDB6101, June 1963.

III, Cromwell, C. H., II; "A Stablllty Analysls of Tllt- 6. Payne, H. E., Wing Aircraft (Analytical)," Princeton University Report No. 477, Myl%O.

7. Ransone, Robin K. and Jones, Gay E., "XC-l&A V/STOL Transport Trl- Services Ilmlted Category I Evaluation;" Air Force Flight Test Center Technical Report No. 65-27, January 1966.

8. Goldsmith, Robert H. and EIickey, David H., "Characteristics of Aircraft with Lifting-Fan Propulsion Systems for V/STOL," IAS Paper No. 63-27, January 1963.

9. Mertaugh, L. J., Stancll, R. T. and Davidson, J. K., "Analysis of a Iow Speed Wind Tunnel Test of a High Msss Rate Vectored Propulsion Flow Model," IJY Report 2-53310/5R-2206, 31 July 1965.

10. Air Transport Association, "Standard Method of Estllaatlng Compsratlve Direct Operating Costs of Transport Airplanes,' June 1960.

11. mrgason, R. J. and Hammond,A. D.;" I&era1 Control Characteristics of a Powered Model of a Twin-Propeller Deflected Slipstream STOL Airplane Configuration," lusll TN D-1585, March 1965.

l-2. Hubbard. H. H., "Propeller Noise Charts for Transport Airplanes," NACA TN 2968, June 1953.

Von Gierke, H. E., "Handbook of Noise Control," McGraw Hill, 1957.

13.

14. &ncock, R. N., "Acoustic Evaluation of the ID-66 LFC - Suction Com- pressors," Northrop Corp. Report NOR-62-83, April 1962.

REFERENCES (COmlNUED) White, D. D., "General Procedures for Estimating Nolse Levels for New 15.

Turbine Engine Designs," Pratt and Whitney Report PWA-2321, &rch 1964.

16. Beranek, L. L. "Noise Reduction," &Graw-Hill, 1960.

Sowers, H. D., "Investigation and Wthods for the Prediction and 17.

Alleviation of Lift Fan Noise," USA !TFWZO!4 Report TR 65-4, August 1965.

CONl?ROLFOWEIRREQUIREMENTS TABm l.- VTOL 1.20 .60 .60 030 -50 025 l 54 -27 a45 .22 1.08 954 -96 .48 .4a .24 .40 .20 I20 --- STOL .40 .20 .20 .lO 60 -45 .22 .18 .18 009 90 .40 .20 036 -36 .18 032 .16 .16 .08 I20 TABLE2. TYPICAL OPl!IMJMDESIGNDRAGESTIMATE!

g

-*

Flat CharacterA F*R*

%

I f=Ih

Plate Wetted istic Component Area t:: I Lgth 10-b Sq Ft Sq Ft Ft

!

~- --_ I--1-_ --I--- ..--.-.a lloo j Wing (S = 69 Sq Ft) 12.5 .oo29 1.32 8.59 l l5 1.329 -00385 -00509 5.60 .@@9 L .00048 1Wing Camber -... .--.- .- ..“..-“..-I-.

z

244 9.59 .085 14.0 Vertical Tail a0286 1.23 .oo178 ___-- I Horizontal Tail 300 5*75 .085 8.4 l 00309 1.64 a1236 -- -. -- - t ) Nacelles 28.0 / .00257 1.091 .00280 i&o h0449 1.62 Inboard .00234 28.0 1.091 .00280 1.60 .0044g 1.78 -00257 l 00257 .00228 1.58

- I -

8.00 .01156 _.__..I.._ - 1.430 a0405 1.15 j.00466 -27 l OcQ39 ,.... .--- -..-.. _-.

I - 2.00 . oo28g i M~SC + Unidentif.

-1 y- I I.

i /I :TOTAL i. - 24.05 .03474 i =.

NOTES: f (Gear &tended) = 12 square feet .004go 'fe = j * Based on V = 350 hots at 35,000 ft I.___ -.__. -_...__ --.---~- _. i ZERO-IZFl?ESTIMATEDDRAGCOEFFICIERPS, TABIE 3.

60 PASSENGER DESIGNS -- --.

~I__--- Propulsion Design Wetted Design Equivalent wing 'fe System T.O. and Meld Area. Area- Flat Plate e Concept Ldg. Cap, Ft2 lengths FG Area - ft2 ability Ft.

-- -~ _ _

~..-- 1

VTOL 4911 24.05 .0348 6% -boPrqp -0049 -725 Turboprop V/STOL 4882 22.63 .0046 676 l 0335 Turboprop STOL 1000 610 21.52 .oo47 4589 l 0353 l 74 STOL Turboprop 2000 610 4589 21.37 l 0350 074 00047 Fan-in-wing VTOL 30.20 .0224 .OO42 1350 7119 -83 V/STOL Fan-in-ting 1000 .o251 .0042 .a4 5912 25.Og STOL 1000 1100 6281 26.78 Fan-in-tin@ .o244 .0043 -83 STOL Fan-in-wing 2000 .0261 895 5492 23.35 l oo43 -83 STOL Propulsive 1000 .o294 .0041 .80 768 5562 22.57 wing STOL Propulsive 2000 461.1 18.14 .o315 .80 575 l m39 wing

L

TABLJZ 4~ GE1 GAS GEWERATORTECHNOLCGY Turbine Inlet Temperature 22OO'F Compressor Stage Pressure Ratio 1.21 Turbojet Engine Configuration Thrust-to-weight ratio 8 Thrust per unit pound of engine air flow 75-85 pounds/ pounds/set, Thrust per unit of engine volume 350 pounds/ftj Turboprop wine Configuration SHIPto engine weight ratio 7 SHP/pound SHP per unit pound of engine air flow 140-150 SsHPlpound/ SHP per unit of engine volume 215 sm/ft3 TABLE 5.0 AVIONICS EQJJIFMEEC LIST Quantity Characteristics

I _. .~.. 4--t-

VHF Cammunications Transceiver Airborne Radio, Inc. (ARINC) Characteristics No. 546

r

Navigation Receiver 2 AlUNC Characteristic No. 547 Marker Beacon Receiver 1 Bendix MKA-28 equivalent ADF system 1 ARINC Characteristic No. 550 ATC Transponder 1 ARINC Characteristic No. 532D 1 ARIXC Characteristic No. 521~ Audio System 3 ., Cockpit Voice Recorder 1 ARINC Characteristic No. 557 Flight Data Recorder 1 ARINC Characteristic No. 542 compass system 2 Collins MC-102 or equivalent Weather Radar 1 RCAAVQ 50 or equivalent Terminal Area Navigation System 1 As yet undefined. May be a self-contained or precision NAVAID.

All-Weather Takeoff and Iatiing 2 As yet undefined. Will in- System clude receivers, sensors, displays, altimeters, and couplers.

TABIE 6. CR~ICALDESIGNCOND~IONS FOR 60 PASSENGEZAIRPLANES Cruise Cruise Design Field Alt. B&h Concept Length Ft No. Critical Design Crite@ VrOL Takeoff VTOL 35,m l 59 V/SToL Takeoff As VTOL 25,fJoO -65 Turboprop 1000 Ft STOL .615 Min DCC (Cruise Speed) 25,m 2000 Ft STOL .615 Mln DCC (Cruise Speed) 25,000 VTOL .8 Takeoff VTOL 35,ooo V/STOL .8 Takeoff asVTOL 35,oo Fan-in-Wing 1OOOFtsTOL .8 Iandiag 35,ooo 2000Ft STOL .a Takeoff (Fan Size) 35,ooo MinDOC (Cruise Speed) 1000Ft3!0L Landing 35,m -9 Propulsive Wing 2OOOFtSTOL Mn DOC (Cruise Speed) 35,000 J TABIa7. 6oPW OPrnmAIRpIARE: FBYSICAL CHARACERISTES 1 TURBOPROP TILTWING V/STOL 1 62.115 139,880 1 7,557 181 FT 179 FT 128 FTIOINI 5540 1 4 117 FT2 IN 19.23191.8 Il.06 1 74FT2lN 27FT8IN 3410 4 15FTllIN9 88.2 .88 TURBOPROP 1000 FOOT STOL 53,783 33,845 5,282 81 FT 3410 4 15FTllIN9 TURBOPROP 2000 FOOT STOL 52,758 32.908 5,195 81 FT 74FT2IN 27FT3IN 86.6 .88 1.15 FAN-IN-WING VTOL 95,327 60.660 19,865 98 FT 6 IN 71 FT 33 FT 7720 6 105 IN 3.73 70.6 79,587 47,622 17,190 93 FT 7 IN 58 FT 8 IN 29 FT 6400 6 87 IN 3.44 79.6 I.04 FAN-IN-WING V/STOL 5710 6 80 IN 3.73 71.7 .89 FAN-IN-WING 1000 FOOT STOL 78,919 46,861 17,282 90 FT4 IN 64 FT 35 FT7 IN -77 STOL 72,110 41.513 15,836 90 FT 4 IN 56 FT 8 IN 35 FT 7 IN 4600 6 71.5 IN 3.60 80.5 FAN-IN-WING 2000 FOOT 67,451 41,599 11.138 86 FT 74 FT 7 IN 32 FT 1 IN 4700 6 36.1 IN 3.74 87.8 .91 ADAM 1000 FOOT STOL 2540 .6 26.6 IN 3.36 95.5 .64 ADAM 2000 FOOT STOL 56,963 32,228 8,051 83FT4IN62FTSIN 29FTl IN SPEC RATED TE = RATED ENGINE THRUST, POUNDS w- DESIGN GROSS WEIGHT. POUNDS SPEC RATED SHP t RATED SHAFT HORSEPOWER WEE EMPTY WEIGHT. POUNDS FUEL WEIGHT. POUNDS wF= NECR= NUMEER OF CRUISE ENGINES L - AIRPLANE LENGTH = PROP DlAM ETER b WING SPAN = DIAMETER OF WlNG FANS h - AIRPLANE HEIGHT ASPECT RATIO WING LOADING. POUNDS PER SQUARE FOOT STATIC THRUST TO WEIGHT RATIO, ALL ENGINES OPERATINt. 86 F AT SEA LEVEL . . - . . . - . -..--.. . ._ _---. -- . . ..- --.--. .

-.-.--. .-,.,--.--. ,.-, ,.

TABIZ 8 ESTIMATED WFJGKPS 60 Passenger A planes -~ ~-__

- - -_- -_

?ropulsive

I

lkn-in-Wing

r

,

w I

-- WJ

1000 2ooo 1ooo 20& Loo0

FT m VP01

coMmm m! Fr FT V/sn: STOL STOI STOL ST01 5!iBL Wing Group 8884 Es 4350 536:

2562 -Fg

Tail Group 148c 62i 813~ Body Group 6852 716: Alighting Gear 2001 2951 2645 2522 1987 3697 3035 Flight Controls 215t 1416 2478 189: 743 4487 as@ 3315 Group Nacelle Group 2020 2630 2027 2335 2156 2640 $2; ;% 2;;: Engines 6;~ 4s: y&c % Air Induction 144 144 144 144 System - ExhEiust System 100 168 144 Iribricating Syste ~~ 140 it: 140 14c 14c 140 Fuel System 1031 66E 447 3u 1037 95c Engine Controls 128 128 128 128 128 l2e 128 Starting System 200 Propellers or Fan 4g lgi 3;ti 2;: 1;ZZ 3: system Transmission 1824 189 1406 I.241 3457 2302 2302 System * Auxiliary Power 200 200 200 200 2Oc 2oc 200 Plant Group Instrument Group 383 383 383 383 383 383 FIN. and Pneumati E2 305 300 285 354 350 330 Group Electrical Group I210 1150 1440 1140 1235 1535 1375 1310 Electronics Group 2$ 691 691 691 6% 691 691 6% 691 Furnishings Group 5906 591-9 $2 t;g :g ?E: Air Cond. and L710 E: :g; E; L527 Anti Icing Auxiliary Gear 40 40 40 40 40 40 40 40 40 40 Group TUIALEMPTY izz E6 G 5861 41513 gzzi

w

J599 WEIGH!I!

Water, Food,

633 g33 633 =zT 633 633 633 633

Beverage, etc.

Crew plus baggage 520 520 520 520 520 520 520 520 520 520 Passengers plus 2ooo L2ooo 2ooo 3loo 2cmo 2ooo l2ooo 2000 Kooo W3m3e Cargo I200 I200 l.200 I200 Fuel * ,",iz 7455 15994 Oil 250 250 250 250 w i!ia iEE TAKE-OFFGROSS '9587 etc.

1, 1 Roplalulvu wing USigllU.

TABIS 9. DIREEC OPERATINGC(l6TBREAKD(IwN 60 Passenger Designs !33044ile Stage Length Design, Direct Operating Costs

T

Propulsion P. 0. and Design system Field Depreci- Mainten- Flight ras- c'apability Concept Length-R ation ance 3peratione Total !urboprop VTOL .0088 .0267 .M9 .WO lurboprop V/STOL~ .oo78 .0091 .o247 .0078 tiboprop STOL 1000 .oo77 00078 00083 .o238 lurboprop STOL 2000 .0081 .o222 .0068 .0073 ?~-i.n-p&g VTOL .olJ.l .0143 .oogf3 .o352 'an-in-w- V/STOL .0084 .ol34 .oo87 .o3o5 'an-in-wing STOL moo .oo86 .0x27 .0301 .0088 'an-in-wine STOL 2ooo .0081 =.011g .0280 .0080 'repulsive STOL '1000 .0070 .o230 .oo67 ,093 Jing Fropulsive STOL 2000 .054 .0188 -055 .m79 ding TABLE 10.

COMPARISONOF 60, %IAND ~~OPASSENGERA~PLA~E CHARACTERISTES FAN-IN-WING TURBOPROP PROPULSIVE WING TURBOPROP VTOL v 9l-o~ 2000 FT STDL 2000 FT STOL* ESIGN PASSENGER LOAD 6. 9. 120 60 90 120 60 90 120 OlARACTERlSTlCS 160 90 120 ’ DESIGN GROSS WEIGHT - POUNDS 79.587 104,100 133.200 62.300 89,900 11 I.000 54,963 70.000 86.800 52.7 56 70.100 66.100 F ASPECT RATIO 3.44 3.31 3.22 10 I2 12 1.36 3.49 4.08 9 8 0 WINGLOADING-POUNDS PER SQUARE FT. 79.6 79.1 al.7 90 80 6.0 95.5 112.8 112.6 86.6 90 90 NUMBER OF FANS OR PROPELLERS 5 5 5 4 6 6 IO 10 12 4 4 4 PROPELLER OR FAN DIAMETER 87 IN IO1 IN 114 IN 18 FT4 IN 17 FT 9 IN 20FT 26.61N26.1 IN31.1 INlSFTll IN 16FT9IN 18FTll IN NUMBER OF GAS GENERATORS 6‘ 6 6 4 6 6 6 6 6 4 4 4 6.760 2.540 2.840 3.480 CRUISE SPEED - KNOTS * The cruise speed was 330 knots for the 60-passenger airplane when the lOO$ rpmprupellertlp speed was 1000 fps. The speed noted in the table is for a lCC$ rpn propeller tip speed of 800 ms, but all DOCcomputations except those shown In Figure 52 have used the lower cruise speed.

TABLE 11 EsTlMATEDWEEElls ~andl20PassengezAlrplfines comes DeaWi Passe.wr P-4 WingGrouP Tail Group Body Group Alighting Gear 27?2 Flight Controls Group 1671 Nacelle Group 2l.81 Engines lge0 Air Induction System 90 I Exhaust System 240

G 140

Wbrlcating System 140 Fuel System 1454 Engine Controls 1.28 128 z: Starting System 200 Propellers or Fkn Systems 2% Transmission System * z2s g$ 1710 13% Aux. Parer Plant Group 200 200 200 200 Instrument Group 383 383 Hydraulic/Pneurnetic Group ;2; ;2 Electrical Group 1608 1461 1;; 1303 Electronics Group 691 691 691 Furnishings Group 7523 9fil 2;; Air Cond. and Anti-Icing 2167 2830 2044 Auxiliary Gear Group 60 80 60 80 TOTALEMETYWEZGHI! 6 80162 3IE 4 Vater, Food, Beverages,etc. -42% -38 -53 Crew Plus E&gage 660 --YE 660 660 Passengers Plus Baggage l8ooo 24000 Cargo 1800 2400 1800 Fkel* 24462 20665 igi ++ lbclude unusable fuel also * Hot gas ductlng, diverter valves, etc. for fan-in-wing and propulsive wing designs.

1000 2000 V/STOL FT FT STOL STOL

J

J J

J

J

PROPULSIVE WING

J \

TURBOPROP FAN-IN-WING RAMETRIC SPECIAL NASA

FAR AIRWORTHINESS

STUDIES c DESIGN CRITERIA

REQUIREMENTS

DIRECT OPERATING COST ANALYSIS OPTIMUM (MINIMUM 0. O.C.) POINT DESIGN AIRPLANES ECONOMIC 8 OPERATIONS ANALYSES DESIGNS WARRANTING ADDITIONAL STUDY I -UCVSXUDYFIxMBIOCKDIAGRAM FIWRE 2.

PROPELLER

0 20 FT 8140

A 14 FT 8140

cl 20 FT 2270

0 I4 FT 2270

0 PSEUDO ENGINE

-10 1 1 I I I 1 I I 1

I4 I6

0 2 4 6 8 IO 12

FUEL FLOW, Ibs/hr

- RUEBERIZXD PROPELLJ2R -ENGINE l%RFO~(X

FIGUKE 3.

I a

3,009s ,8

I.0

THRUST- WEIGHT RATIO

.022

.OZl

I

J

E g.020

I.0

.8

.6

gg

THRUST - WEIGHT RATIO

THRUST- WEIGHT RATIO

TYPICAL OFTIMUM DESIGNSELECTION PROCESS FIGURE 4. - I FIGURE 5. -TYPICALTURBOPFtOPAIEWLANE I I I I FIGUHE6.

-!lWWPROPAIRPiWE TIbtmassIm~~

Y

1\\

\

+

\

\ ‘\

\

‘1 +

‘.

ti

- TYPICALFAN-l%WIIVGAlRPhUE FIGUFUS 7.

__-._-. _.... . ._._ . . .- .

Cruise FAN-IX-WINGAIXPLANE~SMISSIONSYSTEMSCHEMATIC- FIGURE 8. - ALLENG3NES OElCEtATING Cruise Low Speed - FAN-IN-WING AlRPliUE TRANSMISSIONSYSTEMSCHXMATIC- FIGURE: 9.

OUIBOARDENGINE INOPERATrvE FIGURE 10. -!FYPICALPRORHSIVEW3XGAZDtPLME \ .

\ \ FIGURE 11. - PROF'ULSIVEWING AIRPLANE TRANSMISSION SYSTEMSCHEMATIC- ALL ENGSNESOPERATING

FAN TURBINE

.:.>:..m::;:.:+:.: . . .. ..

. . .. . .. .. . .. . .. . .. .. . .. . . :.:.: . . :.~‘.‘.s

............... . ...

::::::::::::~:::::::::.:.:.:.:::::::: ..................i5.*. ~.:.:.:.. . . ..

NOZZLE

FAN INLET

-!l?YPICALPFtOE'llISIVEUIlVGCRC6S SECTION* FIGUREl2.

TILT WING V’STOL

TILT WING VTOL

L$ftIJ

rl

~~~~l~~

W = 62,300 LBS W= 62,115 LBS R=lO II?= 9.23 W/S=90LBS/FT2 W/ S=91.8 LBS/FT2 SHP/ENG=5960 SHP/ ENG=5540 FIGURE 13. - TURBOPROI' OPTYMUMVTOL AIDV/STCKsAIRFLANES

1000 FT STOL 2000 FT STOL

W= 53,783 LBS W= 52,758 LBS A?=9 A?=9 W S =88.2 LBS FT2 W/S* 86.6 LBS/FT2 SHP/ENG.=3410 SHP/ENG = 3410 FIGUHE 14. -!lXJRBOPROP Ol?CR4UMSTOLAIRpLANES

a

WSTOL

VTOL

rvIN.l

W= 95,327 LBS W=79,587 LBS cR=3.73 R-3.44 W/S=70.6 LBS/FT2 W/S=79.6LBS/FT2 T/ENG=7720LBS T/ENG=6400LBS FIGURE 15. - FAN-IN-WING OFTIMUMVTOL AND V/STtX AI-S -

2000 FT STOL

1000 FT STOL

Ih4 FTl

I 4c---- I W= 78 ,919 LBS W =72,110 LBS A?= 3.73 R- 3.60 W/S=71.7 LBS/FT' W/S -80.5 LBS,/FT2 TbNG=5710 LBS T/ENG=4600LBS -FAN-IN-WING Ol?JXMUMSTOLAI=S FIGm 16.

1000 FT STOL

2000 FT STOL

I-----74FT7 IN.-

l--86 FT-yT

83FT4 IN.+

t W=67,45 1LBS W=54,963 LBS fR= 3.74 A? 3.36 W/S=87.8LBS/FT2 W/S = 95.5LBS/ FT2 TlENG=4700LBS T/ENG=2540LBS FIGURE 17. -PROPULSIVE WING OPIXMUMSTOLA~IRPLUES TAKEOFFTIP SPEEXI - fbs APPROxILMATE OPJYXUMCRUISETIPSPEED - fps

(J) 370

s

A 360

-

; 350

w 340

g 330

= PROPEU.ERBLADEINT~~DESIGNLIFTCO~CIENT % AF-PROPEUERACTIVITYFACTOR - IN%UENCE OF FROPELIER03OMETRICAND OPEWTINCCHARACTER~DCS FIGURE l8.

ONCRUISE SPEED4'URBOPROPSTOLAlRPLANE,V~ONIiRP, AU! = 25,000 ET., Dp = 15.91 FT.

TOTAL OVER 35 FOOT OBSTACLE

GROUND ROLL

&-“X0”\\’

TIPSPEED

= 800 FT/SEC

@ 80% SHP

4 I-

TIPSPEED = 800 FT/SEC

@ 80% SHP

O .

FIGURE lg. - INFJJJEIVCE OFENGINE-FROPELJEROPERATINGCHARACTERISTICSANDTHE WINGFLAPSETllINGONTAKE-OFFPERFORMANCE-TURBOPROPSTOL tmmm, 86O~. DAY, sEA mnL 180 - 5O 160 -

IO0

I

.06c SPOILER ; 8. - TRIM !i

I;/

= 1 5o 60- ,MANEUVERING PLUS TRIM CONTROL REQ I I I I I I 4a/50 60 70 _ 80 90 100

TRUE AIRSPEED, knots

-20 - -ABILEf!YOFSPOILERSMPROVIDEROUCOI'?I'RO~ FIGURE 20.

TURBOPROP2OOOFOCYPSTOLAIRPIME, !cAxFhoFF POWER wl!rR OurBoARD ENGINEiFAIIZD, NOCROSS SHAFTMG

.--MANEUVERING PLUS TRIM

CONTROL REQUIRED

PLAIN FLAP BLC

GED RUDDER

I I I I I I I I I

I I I I

I I I

0’

100 110

70 80 90

50 60

TRUE AIRSPEED, knots

-ABILITY OF ~ICALTAILTO PROVIDE YAWCONpROldURBOpROP2O~ FIGURE 21.

FOOT STOLAIRFUNE, TAKE&OFFPOWER WlTHOUBOARDENGIBE FAIUD, NO CROSSSHAFTING, mpi W-= loo, 8F = 40’

PERCENT CHANGE IN FUEL RESERVES

FIGUEB22.

- WEIGEFSElVSI!lTVI!FY TO FUZL RESERVES - FAIO-lR=WING Vb!oL AI[RpficuIE, lcm$mJELm3ERvE=33$TcnALFuEL WEIGHT SFaNSITmY TO SKIN FRICTION DRAGAND F'ROPUISION FIGURE 23. - SYSTEMEFFICIFtNCY- pROPUL%VE WING 2000 FOOT STOL AIRPLANE

PROPULSIVE

FAN-IN -WING

TURBOPROP

WING

1000 2000

1000 2000 1000 2000 FT FT FT FT FT FT

VTOL WSTOL STOL STOL VTOL WSTOL STOL STOL STOL STOL

.- - FIGUHZ 24. - WEIGHT BIiEimmN coMpARIsoNs -60 PAL3sErmRAIRHAms . 10 PWWtKtU

-7

I I I I ncc ’ GNS 1 Jlwrnl . . . . cc .08 D.O.C.

S/SEAT MILE .06 I0 MI ~STAGE L~NGT~I

.O

FIGHT 25.

-DIElECTOPERA!L'INGCO!3!CS-6OPASSENtXR AlRPMNES;NON PRODUXTVE TIME = 10 l/4 laNmEs, VISION = 2000 HOuRs PER YEAR, PRODUXION RUN = 300 ASRPLANES.

PROPULSIVE WING TURBOPROP FAN-IN-WING i INITIAL AIRPLANE COST& MILLIONS OF DOLLARS AVERAGE INITIAL AIRPLANE COSTS ASSUMING 300 AIRPLANES PRODUCED WITH NO RESEARCH AND DEVELOPMENT COSTS INCLUDED FEURE 26. INITIAL AIRPLANE COSTS

I

PROPULSIVE I TURBOPROP

I FAN-IN-WING WING

STOL 1000 MAINTENANCE MANHOURSPE :R FLIGHTHOUR ,.,.,.,.,.,......... :.:-..:.:.:.

:.:.:.~:.~~:.:.:.:.:.:.:.:.

.............~.~.~...~.~...

FIGURE 27.

-MAINTENANCE REQUIREMENTS - bxl FLIGHT >>:s’:‘:.:.:*:.:.:.:+g.

B MAINTENANCE PROPULSIVE STOL STOL DIRECT OPERATING COST- CENTS/ 2 SEAT MILE FIGUIUZ 28. - DlREcT OFTJWIE COSTBHUIOXMN - 5ooIaIJ3~~ 270’

90’

-pWCRVEDl'JOISE INELD-CO-, FIGURE 30.

TuRBopRoP2OOOxKYJ!smL,1OOOFFsTIPsPEm 210’ 170- VW If30 190. 2olr -PERCEIVEDNOISE'IEVELDIREtX'IYlTYC~- FIGURE 33.

FAN-IN-WII'lGV/SToLAIEWUiE -PEtRCEIVEDNO~LEVELDIRECT~CONMIURS- FIGIRE 32.

F'ROPU=Ivg:WING2OOOFarrgMLAIRPLtUa

I .~ .~---I---

I I

AIRPLANES DEVELOPING MAXIMUM POWER STATIC AT GROUND LEVEL APPROXIMATE LIMITING VALUE t; FOR RESIDENTIAL AREAS AT

it

8 24

1000 FPS TIPSPEED

i

w

4 20

$

z

5 l6

IL

l APPROXIMATE LIMITING VALUE 2 12 FOR AREAS ADJACENT TO -

I-

CITY CENTERS

!!!

APPROXIMATE LIMITING VALUE FOR AREAS

L

80 90 100 ,110 120 60 70 DECIBELS (PNDB) PERCEIVED NOISE LEVEL *DISTANCE IS MEASURED AT THE ANGLE AT WHICH THE MAXIMUM PNDB’ OCCURS, MEASURED RADIALLY FROM THE AIRPLANE.

PERCEIVED NOISE IZVEL VS DIfil'IWE* PIG. 33.

1 , : ! I I : III1 Ill i [ 1 I r I 1 IIll ! i ; I i I I I Ii ! I I I 1 j r-r : 1 I j I i I I I I I I I I I I I I I I I I I ! III 1 III I I Ii ,I,.,.,,., ,, I PERCEXVEO NOISE LEVELCONTOURS FORTAKEOFFAND LANDING, FIGURE34. - TURBOPROP V/STOL AIRPLANE ‘I !

I I I I I I I I III ! I I I I I I I I I I I I I I I I I I I ! I j ! II/I I PERCEIVED NOISE LEVELCOIWXJRS F'ORTMEOF'FAND IANDING- nm35- TuRB0PR0P2000 Four STOLAIRPLME I / / I , I / I / I j I I I I I I I/ I I I!I I Illill II Ii/I//III ! j I j I 1 i I !H li1~‘1!IIlIl I I I I I I III I I ( 1 I 1 i IllIll II II!IIIIl/ III ’ I ;lI/lII 1 II ! !IlIIlII III II ,(I i ; I III 1 II I’ I I II (III I ! I I I I I l!!.;FEt!ll,!!!!l!.i..!l.!. .!/j~~!~~,,~~!,,,~~~,,,,j~,,~!.,!,~~~,~,.~~l~~!,l~~~Ill’ !I I: i PERCENED NOISE WLCONTOUFJS FORTAKEOFFAND LANDIt& FIGURERS- FAN-IN-WIING V/STOL AIRPLANE lmuRE37-PERCENED NOISE LEVEL COJ!lTOURS FOR TAKEOFFAND MUNG, PRoKlIsIvE WING 2Oool~. STOL AIRPLAm .lO Y I .06 I- 200 300 400 500 0 100 STAGE LENGTH - MILES - EFFECT OF NON-PRODlETIW TlME ON DIRFET Ol%&Y,l?mGCOSTS- FIGUREj8.

TuRBO~0PvToLAIRpLANE .lO .08 .06 .04 .02 n “0 100 200 300 400 STAGE LENGTH - MILES - EFFECT OF NON-PRODUZTIVFi Tm ONDIREXT OPEFUTINGCOSTS- FIGURE 39.

TUREiOPROP2OOOFOoT STOLAIRPUNE STAGE LENGTH- MILES -EFFECTOFRXJ-PROMICTIVETIME 0lVDIEiECTOPliZW!~Ctl6~~ FIGURE 40.

FAl'i-m-wnrCV/mLm 300 400 STAGE LENGTH - MILES FIGURE 41. - EFFECT OF NON-PRODWCTIVE TIME ON DIRECT OPERATINGCOSTS- PROPULSIVEWlNG2OOOFO(TrSTOLAlRPLANE .I2 .lO .08 .06 .04 .02 0 100 200 400 800 STAGE LENGM & MILES FIGUREk. - EFF!EZTOFANNUALUTILIZATIONONDIRECTO~lXGCOSTS- TUEWPROPVTOLAIRPLANE w zi .08 .06 .02 100 400 0 300 STAGE LENGTH -MILES - EFFECT OF ANNUAL UTILIZATION ON DIREXT OPEElATING COSTS- FIGURF: 43.

TURE!OPFtOP2O0OFOoT STOLAIRPLANE .12 .lO .08

!

.06 I- ii!

K i3 .02 - EF'FECTOFAl'WUAL~ILIZATIONOND~CT OPEMTRiGCOSTS- FIGUIGWC.

FAWN-WINGV/STOLAlRPLANE .08 UJ iz ii .06 0 100 200 300 400 500 STAGE LENGTH @ MILES FIGuRE 45. - EFFECT OF ANNUAL UI'ILIZATION ON DIREZT OPERATINGCOSTS- PROF'UUIVEW~G2COOFfXYTSTOLAIRPUNE

300A/C FORCIVIL BUY

6fKIAIC FORCIVIL-MILITARY BUY

CIVIL BUYONLY

I

JOINTCIVIL

Ml LITARY BUY

tzzl

.03-

;.;.;.;.>x.:‘>w:.: .:.:.:.:.:.:.:.:.:.:.:.:.:.

D.0. C. ..............

.............

...........................

................. . . . . .

:.:.:.:.>:+:.:::y::::: :::::::::::::::::::::::~ .:.:.:.:.:.:.:.:.:.:.:.:.~.

. ............ .

.............

$/SEAT MI I.E. o2 . ............ .

.:.:.:.:.:.:.:.:.~.~.:.~.~.

:.:.:.:.:.:.:.:.:.x.:.:.: .:.:.:.:.:.:.:.:.f:.~.:.~.

:.:.:.:.:.:.:.:.:. .........

::::::::::::::p:## :::::::::::::::::::::::::::.

.:.:.:.:.:.:.:.:.~.~.:.~.~.~ ..............

.:.:.:.:.:.:.:.:.~.~,: ’ @$. .

f

.Ol-

0' -l3FLUENCEOFA~CIVIL~~BUYCNDOC FIGURE 46.

5DDMlLE STAGELENGTH .05 .M) FAN-IN-WING TURBOPROP PROPULSIVE TURBOPROP WING 2000 FTSTOL TILT-WING D.0.C 2DDDFTSTOL $/SEATMILE .03 FIGURE 47. - INFLUENCE OF GAS GENEBlTORCOSTSON DW

I3DFR]

AtFUEL) [IFR] 300 MILES

50 MILES

I50 MILES

150 MILES

C (FUEL)

E (FUEL)

D ~FR]

NON PRODUCTIVETIME

@ IO l/4 MINUTES a SHORT TAKEOFF 3 MINUTES

VERTICAL TAKEOFF 2 MINUTES

SHORT LANDING 7 l/4 MINUTES

VERTICAL LANDING 2 MINUTES

NON PRODUCTIVE

TIME (ij

FAN-IN-WING V’STOL

NON PRODUCTIVE

I

El

TIME 0

-

TURBOPROP VTOL

TURBOPROP 2000 FT STOL

I

PROPULSIVE WING

2000 FT STOL

I

0 20 40 60 80 100 120 140 160 180

BLOCK TIME TO COMPLETE ROUTE, minutes

FIGUREi49. -HYFWEZICALROUTEBLOCKTIME 10.25MINUTES NONPRODUCTIVETIME

q

PROPULSIVE WlNG2000FT TURBOPROP

2.0-

TURBOPROP

TILT WING VTOL

r-

1.8-

WEIG MT

“GR oss

,-FAN-IN-WING

t

RAT IO

V/STOL

1.6-

TURBOPROP

2000 FT STOL

PROPULSIVE WING

2000 FT STOL

I I I

I

1.0

60 70 80 90 100 110 120

NUMBER OF PASSENGERS

*THE RATIO OF THE GROSS WEIGHT OF AN AIRPLANE DESIGNED

TO A PASSENGER LOAD TO THE GROSS WEIGHT OF AN AIRPLANE

DESIGNED TO CARRY 60 PASSENGERS

FIGUREl 51. -GROSS WEIGHT V~IONWITR PASSENGERLOAD .04 .03 TURBOPROP PROPULSIVE WING D.O.C. .D2 SISEATMILE . 01 ~GURE 52. -DIRECT OPRRXCIIJT: COSTS VARIA~ON~DESIGNPAS~RLOAD

N A S A C O N T R A C T O R N A S A CR-670(01)

R E P O R T

-

-

-

h ro I e U v?

A, z

by K . R. Marsh

Prepared by LTV AEROSPACE CORPORATION Dallas, 'Texas for Ames Research Center C . D E C E M B E R 1 9 6 7 N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D NASA CR-670(01) I

i

ADDITIONAL STUDIES ON THE FEASIBILITY OF V/STOL CONCEPTS I FOR SHORT-HAUL TRtlNSPORT AIRCRAFT By K. R. Marsh Distribution of this report is provided in the interest of information exchange. Responsibility for the contents resides in the author or organization that prepared it.

Prepared under Contract No. NAS 2-3036 by LTV AEROSPACE CORPORATION Dallas, Texas for Ames Research Center NATIONAL AERONAUT ICs AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical lnformotion Springfield, Virginia 22151 - CFSTI price $3.00 TABLE OF CONTENTS Page No.

INTRODUCTION. 1

Sensitivity of Airplanes t o Off-Design Operations . . . 1

Sensitivity of Airplane Designs to Alternate

Design Criteria . . . . . . . . . . . . . . . . . . 4

Propeller RPM-Engine RPM Wtch . . . . . . . . . . . . 6

LIST OF TABLES Table No. Page No.

1 V/STOL Short-Haul Transports -

Design Ground Rules . . . . . . . . . . . . . 14

2 Comparison of Airplanes Designed for 300-

and 5OO-Mile Statute Mile Stage Lengths . . . 15

Estimated Weight Breakdown - 60-Passenger

Propulsive Wing V/STOL Airplane . . . . . . .

ili LIST OF FIGURES Figure No.

Page No.

1 Effect of Altitude on N R P Cruise Speed .

. . . . .

LA Effect of NRP Cruise on the Requ red U l t mate Load Factor . . . . . . . . . 0 . 0 . . . . . .

1 7 2 D.O.C. Versus Cruise Altitude . . . . . . e . . .

D.O.C. Versus Cruise Altitude . . . . . . 0 . 0 .

3 19 4 20 D.O.C. Versus Cruise Altitude . . . . . . . . . .

Effect of Operational Range on Takeoff Distance . 2 1

6 Effect of Operational Range on Takeoff Distance . 22

Effect of Operational Range on Takeoff Distance .

7 23 0 Effect of Operational Range on Takeoff Distance 24

Propulsive W i n g V/STOL Airplane . . . . . . . 25

10 Effect of Takeoff Propeller Tip Speed, Engine

Overspeeding, and SHP on Payload . . . . . . . 26

11 Effect of Takeoff Propeller Tip Speed, Engine Overspeeding, and SHP on Takeoff Weight . . . . .

12 Effect of Takeoff Propeller Tip Speed, Engine

Overspeeding, and SHP on Average Cruise Speed . . 28

Turboprop VML Airplane Relative Direct Operating C o s t . . . . . . . . . . . . . . . . .

14 Relative D.O.C. f o r Tip Speed, Engine Overspeed and Horsepower Variations . . . . . . . . . . . .

Landing Drag P o l a r . . . . . . . . . . . . . . . .

16 Landing Drag Polar . . . . . . . . . . . . . .

17 Tail-Off, Nose Fan Inoperative, Power-on Polar

f o r the 60-Passenger Fan-in-Wing V/STOL . . . . 33

18 Tail-off, Nose Fan Inoperative, Power-on Polar f o r t h e 60-Passenger Propulsive Wing ADAM

2,000-Foot STOL . . . . . . . . . . . . . . . . . 34

i v LIST OF FIGURES (Cont) Figure No. Page No.

Effects of Size on Perceived Noise Level,

TurbopropVTOL. . . . . . . . . . . . . . . . . . 35

20 Effects of Size on Perceived Noise h v e l ,

Fan-in-Wing V/STOL, Takeoff . . . . . . . . . . . 36

Effects of Size on Perceived Noise Level,

Propulsive Wing 2OOO-Ft SML, Takeoff . . . . . . 37

22 Effects of Propeller Tip Speed on Perceived

NoiseLevel . . . . . . . . . . . . . . . . . . . 38

Effects of Parer and Propeller Tip Speed on

Noise, Takeoff . . . . . . . . . . . . . . . . . . 39

Comparison of Measured and Calculated

Perceived Noise . . . . . . . . . . . . . . . . . 40

25 Sideline Takeoff Ground Roll Noise Levels . . . . 41

V INTRODUCTION Under contract t o t h e National Aeronautics and Space Administration, Vought Aeronautics Division of LZV Aerospace Corporation developed a numberor V/STOL Short-Haul Transport a i r c r a f t designs around a set of common design These design c r i t e r i a are summarized c r i t e r i a . These designs in Table 1.

used turboprop, fan-in-wing, and propulsive wing propulsion system arrange- For t h e turboprop and ments f o r a t t a i n i n g t h e design V/STOL c a p a b i l i t i e s .

fan-in-wing propulsion system concepts, VTOL, V/STOL, and STOL a i r p l a n e s were developed; f o r t h e propulsive wing concept, only STOL a i r p l a n e s were developed. STOL airplanes were developed f o r operation from 1,000-foot and 2,000-foot runways, and a l l airplanes were optimized t o give a minimum d i r e c t operating c o s t on a 500-statute-mile stage length. The r e s u l t s of t h i s design e f f o r t a r e summarized i n Reference 1.

As a r e s u l t of t h e findings gleaned f r o m t h e work e f f o r t reported i n Reference 1, f u r t h e r s t u d i e s w e r e made of the performance of t h e s e V/STOL short-haul t r a n s p o r t a i r c r a f t when operated a t off-design conditions and of design changes r e s u l t i n g from using d i f f e r e n t design c r i t e r i a . Some of t h e basic aerodynamic input data that were u t i l i z e d i n developing t h e s e designs, and t h e noise c h a r a c t e r i s t i c s of some of t h e designs, were e v a l m t e d . These a d d i t i o n a l s t u d i e s a r e summarized herein.

STUDY I i E s U L T s S e n s i t i v i t y of Airplanes t o Off -Design Operations Reduced c r u i s e a l t i t u d e e f f e c t s . The airplanes designed f o r t h e study of Reference 2 were optimized t o give a minimum d i r e c t operating c o s t a t a 500-mile s t a g e length, and c r u i s e a l t i t u d e s were high (25,000 t o 35,000 f e e t ) ; therefore, t h e r e s u l t i n g design l i m i t equivalent airspeeds (EAS) were considerably less t h a n t h e c r u i s e speed capability of these airplanes f o r The study assumed that t h e r e would be no a i r operations a t low a l t i t u d e s .

t r a f f i c contrgl problems or operational problems t h a t would prevent these V/STOL short-haul t r a n s p o r t a i r c r a f t from operating a t optimum c r u i s e conditions. W i l e such an operation is desired, it may not be achieved during t h e time period being considered f o r these vehicles. Hence, t h e e f f e c t s of imposing lower c r u i s e a l t i t u d e l i m i t s were evaluated on some of these a i r p l a n e s . The e f f e c t s of lowering cruise a l t i t u d e on performance and d i r e c t operating cost were studied f o r the turboprop VTOL, turboprop 2,000-foot STOL, and propulsive wing 2,000-foot STOL airplanes.

The turboprop VTOL airplane w a s designed f o r a 285 knot l i m i t E A S and l i m i t load f a c t o r of 4.07. The turboprop 2,000-foot STOL with an ultimate a i r p l a n e was designed f o r a 282 knot l i m i t EAS and an ultimate load f a c t o r The propulsive wing 2,000-foot STOL airplane was designed f o r a of 4.07.

These design l i m i t 365 knot l i m i t EAS with an ultimate load f a c t o r of 4.05.

equivalent airspeeds and ultimate load f a c t o r s were s e l e c t e d a f t e r evaluating t h e effects of t h e 50-foot-per-second and 66-foot-per-second g u s t conditions on the operational l i m i t s and d i r e c t operating c o s t s of these a i r p l a n e s during t h e cruise, climb, and let-down portions f o r t h e design stage length.

Figure 1 . presents the e f f e c t of c r u i s e a l t i t u d e on t h e normal r a t e d power (NRP) c r u i s e speed f o r each of t h e s e t h r e e airplanes. Tne turboprop VTOL and propulsive wing 2,000-foot STOL a i r p l a n e s have a design c r u i s e a l t i t u d e of 35,000 f e e t . The turboprop 2,000-foot STOL a i r p l a n e has a design c r u i s e a l t i t u d e of 25,000 f e e t . From Figure 1 it can be seen t h a t t h e propulsive wing 2,000-foot SML a i r p l a n e can c r u i s e with N R P down t o a l t i t u d e s as low as approximtely 24,000 f e e t before encountering t h e l i m i t i n g FAS. The turboprop VTOL a i r p l a n e can c r u i s e with NRP down t o a n a l t i t u d e of approximately 22,000 f e e t before encountering t h e l i m i t i n g EAS. The turboprop 2,000-foot STOL a i r p l a n e can c r u i s e with NRP down t o an a l t i t u d e of approximately 19,000 feet before encountering t h e l i m i t i n g EAS. To use a n NRP c r u i s e c a p a b i l i t y a t a l t i t u d e s below these l i m i t i n g a l t i t u d e s w i l l require an increase i n t h e a i r p l a n e design u l t i m t e load f a c t o r and an increase i n t h e a i r p l a n e empty weight.

Figure la presents t h e required v a r i a t i o n s i n t h e design ultimate load f a c t o r if t h e s e t h r e e a i r p l a n e s are t o be permitted t o c r u i s e with NRF’ a t an a l t i t u d e lower t h a n those t h a t were found t o be c r i t i c a l . This f i g u r e shows that t h e ultimate load f a c t o r continues t o increase f o r t h e turboprop airplanes a l l t h e way t o a s e a l e v e l c r u i s e a l t i t u d e . By c o n t r a s t , t h e propulsive wing 2,000-foot STOL airplane reaches a m a x i m u m ultimate load f a c t o r a t an a l t i t u d e of approximately 5,000 f e e t . A t lower c r u i s e a l t i t u d e s , t h e ultimate load f a c t o r begins t o decrease. Although t h e c r u i s e speed c a p a b i l i t y of t h e propulsive wing 2,000-foot STOL airplane is considerably higher than t h e c r u i s e speed c a p a b i l i t i e s of t h e two turboprop powered a i r - planes, t h e lower aspect r a t i o of t h e wing of t h e propulsive w i n g a i r p l a n e is s u f f i c i e n t t o keep t h e load f a c t o r f o r t h i s airplane a t approximately t h e same l e v e l as t h a t which has been found t o be adequate f o r t h e turboprop airplanes.

Figure 2 presents a v a r i a t i o n of d i r e c t operating c o s t s (DOC) with t h e variation i n c r u i s e a l t i t u d e f o r t h e 60-passenger turboprop VTOL a i r p l a n e a t stage lengths of 150 and 250 statute miles. It shows t h e d i f f e r e n c e i n d i r e c t operating c o s t s when f l y i n g a t t h e l i m i t EAS compared t o f l y i n g a t t h e airspeed with NRP. The curves f o r c r u i s i n g with NRP are t h e dash l i n e s below the c r i t i c a l a l t i t u d e . (The design takeoff weights of t h e s e a i r c r a f t were unchanged; however, a s t r u c t u r a l weight penalty has been applied t o permit c r u i s i n g at t h e higher speeds that are compatible when using N R P a t t h e lower a l t i t u d e s . The a i r p l a n e f u e l a v a i l a b l e s have been reduced by t h e of t h e s t r u c t u r a l weight penalty.) The NRP curve f o r t h e 250-mile amount stage length condition is terminated a t an a l t i t u d e of approximately 12,000 feet because, a t a l t i t u d e s below t h i s , t h e airplane does not have s u f f i c i e n t This f i g u r e shows t h e f u e l t o permit f l y i n g t h e 250-mile s t a g e length.

benefits, i n terms of DOC, f o r being a b l e t o c r u i s e with NRP i f lower t h a n optimum c r u i s e a l t i t u d e l i m i t s a r e imposed.

Figure 3 is similar t o Figure 2 except t h a t it is f o r t h e turboprop 2,000-foot STOL airplane. These curves are similar i n shape t o those that were developed f o r the turboprop VML airplane, but t h e e f f e c t s of stage length are less pronounced and t h e v a r i a t i o n of DOC with c r u i s e a l t i t u d e does not have as s t e e p a slope f o r c r u i s i n g at a l t i t u d e s below the c r i t i c a l a l t i t u d e . Both Figures 2 and 3 show t h a t t h e DOC decrease s l i g h t l y as t h e c r u i s e a l t i t u d e is reduced from t h e design c r u i s e a l t i t u d e t o t h e c r i t i c a l c r u i s e a l t i t u d e . Below the c r i t i c a l a l t i t u d e , t h e DOC f o r N R P c r u i s e is approximately constant t o an a l t i t u d e of approximately 10,000 feet, and then it begins t o increase a t the lower a l t i t u d e s . Cruise a t t h e l i m i t EAS below t h e c r i t i c a l a l t i t u d e r e s u l t s i n increased Doc.

Figure 4 has been developed t o show the v a r i a t i o n of DOC with c r u i s e a l t i t u d e f o r t h e propulsive wing 2,000-foot STOL airplane. This curve shows t h a t t h e v a r i a t i o n of d i r e c t operating c o s t s with c r u i s e a l t i t u d e has only a n e g l i g i b l e v a r i a t i o n u n t i l t h e c r i t i c a l a l t i t u d e is reached.

The v a r i a t i o n of d i r e c t operating c o s t s with c r u i s e a l t i t u d e below t h e c r i t i c a l a l t i t u d e is not as pronounced f o r the propulsive wing 2,000-foot STOL a i r p l a n e as f o r two turboprop airplanes.

I n summary, then, these s t u d i e s have shown that if it is required that V/STOL short-haul t r a n s p o r t a i r c r a f t operate a t less than optimum c r u i s e a l t i t u d e s , it w i l l probably be p r o f i t a b l e t o compromise t h e s e a i r p l a n e s f o r c r u i s i n g at lower than optimum c r u i s e a l t i t u d e s by designing f o r a higher EAS.

E f f e c t s of varying t h e operating range. - Although t h e a i r p l a n e s

designed f o r t h e ground r u l e s s p e c i f i e d i n Reference 1 had a design s t a g e length of 500 s t a t u t e miles, it is realized such vehicles would seldom be operated a t t h i s s p e c i f i c stage length. Hence, t h e e f f e c t s of operating a t o t h e r s t a g e lengths on t h e takeoff performance were determined f o r some assuming that t h e lower s t r u c t u r a l load f a c t o r s would of these a i r c r a f t , be acceptable. Figures 5 through 8 present t h e r e s u l t s of t h e s e s t u d i e s f o r t h e turboprop VTOL, the turboprop 1,000-foot STOL, t h e fan-in-wing V/STOL, and t h e propulsive wing 1,000-f oot STOL airplanes.

Figures 5 through 8 present p l o t s of takeoff distance and gross weight The takeoff perform- versus t h e operational range f o r t h e s e four a i r c r a f t .

i s t h e t o t a l distance r e q u i r e d t o c l e a r a 50-foot obstacle on a ance shown sea l e v e l , 86°F day with one engine f a i l e d . Figure 5 shows t h a t t h e turboprop VTOL airplane, with one engine f a i l e d , has a VTOL c a p a b i l i t y s u f f i c i e n t t o permit f l y i n g up t o a 500-mile s t a g e length ( t h e design point f o r t h i s a i r c r a f t ) . If, instead of using a v e r t i c a l takeoff f o r t h e 500- m i l e s t a g e length, t h i s airplane, operated i n t h e STOL m o d e f o r takeoff, would have a takeoff distance of less than 250 feet t o c l e a r a 5O-foot obstacle. This a i r p l a n e could a l s o have an operational range of 1,000 miles and s t i l l r e q u i r e less than 300 f e e t t o c l e a r a 50-foot obstacle.

a s h o r t takeoff run when If it should be so desired, instead of using f l y i n g a s t a g e length of 1,000 miles, t h i s a i r p l a n e could have i t s passenger load reduced from t h e design number of 60 t o 44 and s t i l l use v e r t i c a l takeoff f o r the 1,000-mile s t a g e length.

The economy of t h e turboprop propulsion system is shown on t h i s f i g u r e i n t h a t only approxi- mately 7,500 pounds of f u e l are required t o extend t h e operational range from 50 miles t o 1,000 miles. It has been assumed f o r these analyses t h a t ade'quate space is a v a i l a b l e f o r such f u e l .

Figure 6 presents a comparable curve t o Figure 5, except it is f o r t h e turboprop 1,000-foot STOL airplane. It is seen t h a t a l a r g e change i n range has l i t t l e e f f e c t on takeoff distance. The takeoff performance presented i n t h i s f i g u r e assumes t h a t t h e a i r p l a n e does not use any wing tilt. A wing tilting c a p a b i l i t y of 20° is a v a i l a b l e ( t h i s 20° c a p a b i l i t y w a s p u t i n t o permit t h e a i r p l a n e t o meet its design landing requirements), and t h e use of t h i s 20' wing tilt could permit t h i s takeoff d i s t a n c e t o be consid- erably shorter. This f i g u r e again shows t h e e f f i c i e n c y of t h e turboprop propulsion system i n t h a t less than 7,000 pounds of f u e l are required t o extend t h e operational range from 50 s t a t u t e miles t o 1,000 s t a t u t e miles.

Figure 7 presents t h e e f f e c t s of takeoff d i s t a n c e on t h e operational range f o r t h e fan-in-wing V/STOL airplane. This f i g u r e shows that t h e VTOL c a p a b i l i t y of t h i s airplane w i l l permit it t o f l y a 50-mile s t a g e length; but if t h e stage length exceeds 50 miles, t h e a i r p l a n e must use a s h o r t takeoff run. This f i g u r e a l s o shows t h a t approximately 16,000 pounds of f u e l are required t o extend t h e operational range from 50 s t a t u t e miles t o 1,000 s t a t u t e miles. It can be found from t h i s f i g u r e t h a t t h i s a i r p l a n e can f l y a 500-mile stage length using i t s V T O L c a p a b i l i t y if t h e passenger load is reduced from t h e design value of 60 t o a l e v e l of 22.

Figure 8 presents t h e e f f e c t of takeoff distance on t h e operational range f o r t h e propulsive wing 1,000-foot STOL airplane. This f i g u r e shows t h a t increasing t h e operational range from 50 s t a t u t e miles t o 1,000 s t a t u t e

miles increases t h e f u e l required by approximately 10,000 pounds - not

q u i t e as e f f i c i e n t as the turboprop propulsion system but considerably more e f f i c i e n t than t h e fan-in-wing propulsion system. A comparison of t h e data presented i n Figure 8a with t h e comparable d a t a presented i n Figures 5a through 7a shows t h a t t h e v a r i a t i o n of takeoff d i s t a n c e with range is not nearly so l i n e a r f o r t h e propulsive wing a i r p l a n e as f o r t h e turboprop o r fan-in-wing airplanes.

S e n s i t i v i t y of Airplane Designs t o A l t e r n a t e Design C r i t e r i a S e n s i t i v i t y of a i r p l a n e s design t o design s t a g e length. - I n order t o determine t h e s e n s i t i v i t y of t h e airplanes designed under Reference 2 t o the design stage length, a study has been made on t h e t i l t - w i n g VTOL air- plane and t h e fan-and-wing V/STOL a i r p l a n e s . For t h i s study t h e design range w a s reduced t o 300 s t a t u t e m i l e s , and t h e f u e l reserves were reduced t o simply t h a t f u e l required f o r e n t e r i n g t h e t r a f f i c p a t t e r n and making a landing on t h e first pass. It is considered t h a t t h e r e s u l t i n g a i r p l a n e s represent t h e minimum p r a c t i c a l s i z e s . One o t h e r change i n design c r i t e r i a 1 made f o r these a i r p l a n e s w a s t h a t t h e VTOL design c r i t e r i a were applied i only a t t h e landing condition a f t e r a 50-mile mission.

Table 2 presents a comparison of some of t h e more important character- i s t i c s of t h e a i r p l a n e s which have been optimized f o r t h e 300- and 500-mile stage length. A c l o s e a n a l y s i s of t h e data presented i n t h i s table w i l l show t h a t t h e weight of t h e turboprop V T O L airplane designed f o r 300 miles is approximately 90% of t h a t f o r the a i r p l a n e designed f o r 500 miles.

By c o n t r a s t , t h e fan-in-wing V/STOL airplane designed f o r 300 miles weighs approximately 80$ as much as t h e a i r p l a n e which was designed f o r 500 miles.

The reason f o r t h i s difference i n gross weight r a t i o comes about as a r e s u l t of t h e reduction i n t h e quantity of f u e l required.

The turboprop VTOL a i r p l a n e optimized f o r a stage length of 300 miles w i l l have an optimum c r u i s e a l t i t u d e of 25,000 feet.

A projection of t h e data presented i n t h i s t a b l e w i l l show t h a t t h e weight of t h e fan-and-wing V/STOL a i r p l a n e would equal t h e weight of t h e turboprop V T O L airplane a t a design stage length of approximately 175 s t a t u t e miles.

Propulsive w i n g V/STOL airplane. - During t h e study reported i n Refer-

ence 1, only STOL propulsive wing a i r p l a n e designs were developed. A s a r e s u l t of t h e promise of these STOL designs, it was considered appropriate t o develap a V/STOL propulsive wing airplane t o t h e same design c r i t e r i a used f o r t h e designs of Reference 1. A three-view drawing of t h e r e s u l t i n g propulsive wing V/STOL a i r p l a n e is presented i n Figure 9. This a i r p l a n e is f i t t e d with f o u r gas generators d r i v i n g f m r wing fans. The gas generators are connected t o t h e turbines which d r i v e these w i n g f a n s with an i n t e r - connecting hot-gas duct system. The design gross weight of t h e a i r p l a n e i s 73,300 pounds, and t h e a i r p l a n e has a design c r u i s e Wch number of 0.9 a t its design c r u i s e a l t i t u d e of 40,000 feet. This a i r p l a n e uses 59.5-inch diameter fans. The f o u r main gas generators produce 6,380 pounds of t h r u s t each.

The a i r p l a n e a l s o has two l i f t - t y p e gas generators located i n t h e nose of t h e fuselage t o provide hover and slow speed p i t c h t r i m and control.

The p i t c h engines a r e s i z e d s o that each is capable of providing t h e maximum longitudinal t r i m f o r t h e hover mode, plus 20 percent of t h e hover p i t c h c o n t r o l requirements, and t h e r e s u l t i n g engines a r e capable of developing 15,250 pounds of t h r u s t each. The exhaust system f o r t h e s e engines is arranged s o that they are run at f u l l t h r u s t when i n use. The gas exhaust from these engines is varied between t h e f r o n t and aft o u t l e t s i n order t o vary t h e p i t c h i n g moment. A weight breakdown of t h e propulsive wing V/STOL a i r p l a n e is presented i n Table 3.

Direct operating c o s t comparisons between t h e propulsive wing 1,000- f o o t STOL a i r p l a n e and t h e propulsive wing V/STOL a i r p l a n e have been made using parametric-type costing equations rather than t h e modified ATA costing methodology used i n Reference 1. The parametric costing equations show t h a t d i r e c t operating c o s t s f o r t h e V/STOL a i r p l a n e were j u s t s l i g h t l y higher than those of a 1,000-foot STOL airplane. Since t h e V/STOL a i r p l a n e i s approximately 1 4 heavier than t h e 1,000-foot STOL a i r p l a n e , t h e deprecia- t i o n c o s t s should be approximately 1 % greater than t h e depreciation c o s t s of t h e propulsive wing 1,000-foot STOL airplane. The f u e l r e q u i r e d i s approxi- mately 18% g r e a t e r f o r t h e propulsive wing V/STOL a i r p l a n e than f o r t h e propulsive wing 1,000-foot STOL airplane; theref ore, t h e f l y i n g operations c o s t s w i l l be higher ( t o a lesser percentage). Wintenance c o s t s would approximately equal t h e maintenance c o s t s t h a t were determined f o r t h e propulsive wing 1,000-foot STOL airplane. As a r e s u l t of these considera- t i o n s , it is projected t h a t a d e t a i l e d costing a n a l y s t s of t h e propulsive wing V/STOL a i r p l a n e would show d i r e c t operating c o s t s were between 10 and 1 5 percent g r e a t e r f o r t h e propulsive wing V/STOL a i r p l a n e t h s n f o r propulsive S1Y)L airplane.

wing 1,000-f oot P r o p e l l e r RPM-Engine R P M M t c h 1, t h e p r o p e l l e r s of a l l t h e turboprop I n t h e study of Reference a i r c r a f t were designed f o r maximum s t a t i c t h r u s t . I&iximum s t a t i c t h r u s t w a s obtained with a p r o p e l l e r t i p speed of 1,000 f e e t p e r second ( f p s ) .

It w a s found during t h e course of t h e study t h a t c r u i s e performance, r a t h e r than takeoff performnce, was c r i t i c a l f o r s i z i n g t h e propulsion system of t h e turboprop STOL a i r c r a f t . The best c r u i s e speed occurred f o r an N R P s e t t i n g and at a p r o p e l l e r R P M t h a t w a s between 70 and 80 percent of t h e RPM needed t o give a 1,000 f p s p r o p e l l e r t i p speed a t takeoff. "he use of t h i s low percentage of t h e design engine f r e e - t u r b i n e RPM caused t h e engine performance t o be penalized; consequently, a study was made 100 percent engine of d i f f e r e n t takeoff p r o p e l l e r t i p speeds coupled with free-turbine R P M (i.e., d i f f e r e n t engine free-turbine t o p r o p e l l e r g e a r r a t i o s ) with d i f f e r e n t p r o p e l l e r a c t i v i t y f a c t o r s and i n t e g r a t e d design l i f t coefficients. By matching t h e 100 percent engine f r e e - t u r b i n e R P M with an 800 f p s p r o p e l l e r t i p speed instead of t h e o r i g i n a l 1000 f p s propeller t i p speed, t h e c r u i s e speed was increased from 340 knots t o 370 knots with a n e g l i g i b l e change i n takeoff performance f o r both t h e turboprop 1,000-foot STOL and 2,000-foot STOL a i r p l a n e s (Reference 1).

This reduction i n p r o p e l l e r takeoff t i p speed would a l s o provide a l a r g e and t h e s e e f f e c t s w i l l be reduction i n p r o p e l l e r noise during takeoff, discussed later.

I n l i g h t of t h e s e performance improvements f o r t h e turboprop STOL airplanes, an a d d i t i o n a l study was conducted t o determine if similar improvements could be obtained f o r t h e turboprop VTOL 60-passenger a i r p l a n e by rematching t h e p r o p e l l e r takeoff R P M with t h e engine free-turbine RPM.

Figures 10 through 14 summarize t h e r e s u l t s of varying t h e p r o p e l l e r take- off t i p speed, t h e engine free-turbine BPM during takeoff ( t h e engine free- t u r b i n e can be operated at 125 percent of t h e design RPM without adversely affecting t h e s t r u c t u r a l i n t e g r i t y of t h e engine), and t h e engine shaft horsepower ( S H P ) level. The e f f e c t s of t h e s e v a r i a b l e s on payload are presented i n Figure 10, on takeoff weight i n Figure 11, and on c r u i s e speed i n Figure 12. The r e s u l t i n g change i n operating c o s t s is given i n Figures 13 and 14. Reducing t h e p r o p e l l e r takeoff t i p speed from 1 0 0 0 f p s t o 900 f p s f o r the engine free-turbine operating a t 100 percent R P M reduces t h e VTOL takeoff weight (because of t h e reduction i n s t a t i c t h r u s t ) and payload by 3,200 pounds and increases t h e c r u i s e speed from 339 knots t o 362 knots (because of a b e t t e r p r o p e l l e r RPM-engine f r e e - t u r b i n e RPM match a t c r u i s e ) .

By using the gear r a t i o which gives a p r o p e l l e r t i p speed of 900 f p s a t 100 percent engine free-turbine RPM and overspeeding t h e engine free-turbine a t takeoff t o 111 percent ( i n order t o g e t a takeoff p r o p e l l e r t i p speed of 1,000 fps), t h e takeoff weight and payload are reduced by only 450 pounds and t h e c r u i s e speed is increased from 339 knots t o 357 knots.

Further overspeeding of t h e engine free-turbine f o r takeoff while main- t a i n i n g a 1,000 f p s p r o p e l l e r t i p speed would cause a more rapid drop i n payload.

Increasing t h e i n s t a l l e d engine s h a f t horsepower makes possible t h e use of lower p r o p e l l e r takeoff t i p speeds and/or f u r t h e r overspeeding of t h e engine free-turbine during takeoff i n order t o provide a b e t t e r match between t h e hover and c r u i s e t h r u s t requirements while s t i l l maintaining a constant passenger load.

Figure 13 presents t h e r e l a t i v e d i r e c t operating c o s t s on a cost-per- airplane-mile basis associated with rematching t h e p r o p e l l e r takeoff t i p speed, t h e engine free-turbine EZPM during takeoff, and t h e percentage increase i n shaft horsepower over t h a t used f o r t h e b a s i c design. This f i g u r e shows t h a t overspeeding t h e engine free-turbine f o r takeoff and reducing t h e takeoff p r o p e l l e r t i p speed s i g n i f i c a n t l y reduces t h e d i r e c t a per-airplane-mile basis; but increasing t h e engine operating c o s t s on shaft horsepower does not mike an appreciable (less than one percent) e f f e c t .

If t h e VTOL ground rules a r e retained and accounting f o r t h e change i n payload i s made by varying t h e passenger load (assuming space is a v a i l - a b l e f o r a d d i t i o n a l passengers and/or f u e l , as appropriate), t h e e f f e c t s on t h e r e l a t i v e d i r e c t operating c o s t s on a cost-per-seat-mile b a s i s are This curve has been developed assuming t h e number of shown i n Figure 14.

passengers c a r r i e d equals t h e payload (Figure 10) divided by 220 ( t h e weight including baggage and revenue cargo).

allowance p e r passenger, a better match between engine and p r o p e l l e r These curves show t h a t R P M can be made f o r turboprop V/STOL short-haul t r a n s p o r t a i r c r a f t than w a s used f o r t h e turboprop point design a i r c r a f t of Reference 1. A5 an reducing t h e takeoff p r o p e l l e r t i p speed t o 950 f p s , increasing example, t h e engine takeoff free-turbine speed t o 118 percent of its design value, and increasing t h e i n s t a l l e d shaft horsepower by 10% over t h e value used i n Reference 1 would reduce the d i r e c t operating c o s t s per-seat-mile by approximately seven percent compared t o those c o s t s determined i n Reference 1.

Drag Polars I n o r d e r to provide a more basic understanding of some of t h e funda- mental aerodynamic c h a r a c t e r i s t i c s used i n configuring t h e a i r p l a n e s developed i n response t o Reference 2, landing drag p o l a r s have been developed f o r f o u r of these a i r p l a n e s and are presented i n Figures 1 5 through 18.

These landing polars are f o r operating on sea level, 86°F day ambient atmospheric conditions.

Figure 15 presents t h e landing drag polar f o r t h e turboprop V/STOL This p o l a r is f o r a condition where t h e wing is t i l t e d up 20 airplane.

degrees and t h e 48 percent chord, f u l l span, double-slotted f l a p s are The angles of a t t a c k a r e varied from zero degree deflected 60 degrees.

t o a positive 12 degrees, and t h e t h r u s t c o e f f i c i e n t , based on slipstream dynamic pressure, is varied from 0.5 t o 0.8. The symbol i n t h i s f i g u r e , located a t a lift c o e f f i c i e n t of approximately 10 and a dr'ag c o e f f i c i e n t of approximately 1.5, represents t h e condition f o r an 800-foot-per-minute rate of descent at a 54-knot f l i g h t speed. This condition represents t h e c r i t i c a l STOL landing conditions as s p e c i f i e d by Reference 2. It can be seen from t h i s f i g u r e t h a t a t t h i s landing condition, and with t h i s wing incidence and f l a p configuration, t h e a i r p l a n e is operating c l o s e t o t h e buffet onset boundary. F l i g h t experience with t h e XC-142A a i r p l a n e shows that the i n i t i a l b u f f e t is m i l d . This curve shows t h a t increasing t h e t h r u s t coefficient from .65 t o .75 ( t h e equivalent t o increasing the engine power from approximately 30% t o 40%) w i l l give a normal a c c e l e r a t i o n increase of 0.30 g's. I f a p i l o t should encounter an undesirable f l i g h t condition while f l y i n g s o close t o t h e b u f f e t onset boundary, a l i g h t a p p l i c a t i o n of power w i l l c o r r e c t it; therefore, it is expected that t h e a i r p l a n e would be s a f e f o r such operations.

Figure 16 presents t h e landing drag p o l a r f o r t h e turboprop 2,000-foot STOL airplane. For t h i s curve, t h e angles of a t t a c k a r e varied f r o n zero degree t o a p o s i t i v e 12 degrees, and t h e t h r u s t c o e f f i c i e n t s are varied from 0 . 1 t o 0.7. The symbol shown a t a lift c o e f f i c i e n t of approximately 3.7 and a drag c o e f f i c i e n t of approximately 0.4 represents t h e aerodynamic conditions t h a t are required f o r descending a t 800 feet p e r minute while f l y i n g a t 86 knots, t h e c r i t i c a l landing condition s p e c i f i e d by Reference 2 f o r t h i s airplane. From t h i s f i g u r e it can be determined t h a t increasing t h e angle of a t t a c k from approximately s i x degrees t o approximately 8.5 degrees w i l l provide an 0.lg normal a c c e l e r a t i o n as required by Reference 2 f o r t h i s s i t u a t i o n where one engine has failed. It can be a l s o seen from this figure t h a t increasing t h e t h r u s t c o e f f i c i e n t from approximately .25 t o approximately .29 w i l l a l s o give an 0 . l g normal a c c e l e r a t i o n c a p a b i l i t y t o t h e a i r p l a n e , another a l t e r n a t e design condition s p e c i f i e d by Reference For t h e same f l i g h t condition, increasing t h e angle of a.ttack from 6 2.

degrees t o approximately t e n degrees and increasing t h e t h r u s t c o e f f i c i e n t from approximately .25 t o approximately .35, o r simply increasing t h e t h r u s t c o e f f i c i e n t t o .45 with no angle of a t t a c k change, w i l l give an increase i n t h e normal f o r c e c o e f f i c i e n t of 0.3, another of t h e requirements of Reference 2. I n summary then, it can be seen t h a t t h i s a i r p l a n e has adequate margin i n all of t h e c r i t i c a l conditions of t h e landing mode of operation.

Since t h e wing geometry f o r t h e turboprop V/STOL a i r p l a n e and t h e turboprop 2,000-foot STOL a i r p l a n e are similar, t h e p o l a r s f o r t h e s e air- planes w i l l be similar f o r comparable wing incidences and f l a p d e f l e c t i o n conditions. A comparison of Figures 15 and 16 gives an i n d i c a t i o n of t h e e f f e c t s of w i n g tilt on these polars. As an example, Figure 16, a zero wing tilt condition, shows t h a t at a t h r u s t c o e f f i c i e n t of 0.7 and an angle of attack of 8 O , t h i s a i r p l a n e w i l l have a lift c o e f f i c i e n t of approximately 7.5 and a drag c o e f f i c i e n t of approximately -1.2. Figure 15, f o r a wing tilt wing condition of 20 degrees, shows t h a t a t t h e same t h r u s t c o e f f i - c i e n t and angle of a t t a c k , t h e a i r p l a n e develops a l i f t c o e f f i c i e n t of approximately 10.7 and a drag c o e f f i c i e n t of a p o s i t i v e 1.4; therefore, adding 20 degrees of wing incidence has increased t h e trimmed l i f t c o e f f i - c i e n t by over 3.2, and t h e drag coefficient has increased by a p p r o x i m t e l y 2.6. Thus, t h e s e two f i g u r e s i l l u s t r a t e the operational f l e x i b i l i t y a v a i l - able t o the p i l o t of a tilt wing V/STOL airplane.

The p i l o t of such an a i r p l a n e has t h e a b i l i t y t o a d j u s t h i s wing tilt t o provide a wide l a t i t u d e of safe f l i g h t conditions i n t h e slow speed f l i g h t modes.

Figure 17 presents t h e landing drag polar f o r t h e fan-in-wing V/STOL a i r p l a n e developed i n response t o Reference 2. This drag p o l a r is specifi- c a l l y f o r a condition of f l y i n g a t 54 knots a t sea l e v e l on an 86°F day.

The symbol located at a l i f t c o e f f i c i e n t Of approximately 7.0 and a drag c o e f f i c i e n t of approximately 1.25 indicates t h e f l i g h t conditions f o r making an 800-foot-per-minute rate of descent at a 54-knot f l i g h t condition. It should be kept i n mind, while r e f e r r i n g t o t h i s figure, t h a t t h i s p o l a r assumes t h e nose f a n is not operative, and t h e nose f a n makes a l a r g e contribution t o t h e normal f o r c e on t h i s airplane. (The nose f a n l i f t w i l l provide a l i f t c o e f f i c i e n t change of approximately 1.5 a t t h i s f l i g h t condition.) This f i g u r e shows t h a t increasing t h e wing f a n t h r u s t from a p p r o x i m t e l y 60$ t o approximately 75% f o r t h e condition where t h e w i n g f a n louvers are deflected aft by 1 0 ' w i i i provide G . l g normal acceier%tion required by Reference 2 f o r t h e engine-out f l i g h t s i t u a t i o n . It can a l s o be seen from t h i s f i g u r e that increasing the parer t o 90 percent a t a constant angle of a t t a c k w i l l increase t h e l i f t c o e f f i c i e n t t o approximately 9.5, a value needed t o provide a .3g normal a c c e l e r a t i o n with a l l engines operating, another of t h e conditions specified by Reference 2. It does not appear from t h i s f i g u r e t h a t increasing the angle of a t t a c k , alone, w i l l provide t h e c a p a b i l i t y of increasing t h e normal f o r c e c o e f f i c i e n t by 0.1, one of the a l t e r n a t i v e s s p e c i f i e d by Reference 2.

Figure 18 p r e s e n t s t h e landing drag polar f o r t h e propulsive w i n g s p e c i f i c a l l y f o r t h e 2,000-foot STOL airplane. This landing drag p o l a r is operational conditions on a sea l e v e l , 86°F day, and it is f o r t h e nose f a n The symbol s h a m at a l i f t c o e f f i c i e n t of approximately inoperative case.

3.4 a t a drag c o e f f i c i e n t of approximately 0.4 i n d i c a t e s t h e operational condition for an 800-foot-per-minute r a t e of s i n k at a f l i g h t condition of 86 knots. (The nose f a n l i f t w i l l provide a l i f t c o e f f i c i e n t increase of approximately 0.6 a t t h i s f l i g h t condition.) From t h i s curve, it can be its angle of a t t a c k at a constant power seen t h a t t h e a i r p l a n e can increase setting t o g i v e a change i n normal acceleration of 0.1 with a f l a p d e f l e c t i o n

of 90" - one of the engine-out requirements s p e c i f i e d by Reference 2. The

propulsion system can maintain SO$ t h r u s t with one engine f a i l e d by operating The airplane can increase power and angle t h e engines a t emergency power.

of a t t a c k t o g e t t h e increase i n normal acceleration of 0.3 t o s a t i s f y t h e margin requirements f o r a l l engines operating as s p e c i f i e d by Reference 2.

Noise

Effects of a i r c r a f t s i z e . - Under Reference 2, 60-, gO-, and 120-

passenger a i r p l a n e s were developed f o r s e l e c t e d turboprop, fan-in-wing, propulsive wing V/STOL designs. Figures 1 9 through 2 1 present perceived noise l e v e l contours during t h e takeoff mode of f l i g h t f o r 60- and 120- passenger a i r c r a f t designed around each of these t h r e e V/STOL concepts.

These contours describe t h e noise l e v e l s f o r ground-based observers with an assumed climbout angle of 20°. Figure 1 9 shows t h e e f f e c t of a i r c r a f t s i z e on perceived noise l e v e l f o r t h e turboprop VTOL airplane. This curve shows t h a t f o r t h e turboprop concept, the noise l e v e l a t most distances f r o m t h e source f o r t h e 120-passenger a i r p l a n e is from 5 t o 7 PNdb higher t h a n f o r t h e 60-passenger a i r c r a f t .

Figure 20 presents t h e e f f e c t of s i z e on t h e perceived noise l e v e l This f o r the fan-in-wing V/STOL a i r p l a n e during t h e takeoff f l i g h t mode.

f i g u r e shows t h a t t h e perceived noise l e v e l is approximately 10 decibels it is f o r t h e 60-passenger higher f o r t h e 120-passenger a i r p l a n e than airplane.

Figure 2 1 presents t h e e f f e c t s of s i z e on perceived noise l e v e l f o r This curve t h e propulsive wing 2,000-foot STOL a i r p l a n e during takeoff.

shows d i f f e r e n t r e s u l t s than have t h e two previous curves i n t h a t t h e perceived noise l e v e l f o r t h e l a r g e r a i r p l a n e is lower than it is f o r t h e smaller a i r p l a n e . This unusual change i n trend occurs because t h e j e t engine RPM increases as t h e a i r p l a n e s i z e increases from the 60-passenger s i z e t o a 120-passenger s i z e . This increase i n engine RPM s h i f t s t h e spectrum peak beyond t h e last octave band; thus, t h e perceived noise l e v e l e f f e c t s from t h e higher octave bands a r e lowered.

Effect of reduced p r o p e l l e r t i p speed. - It has been mentioned pre-

viously t h a t f o r t h e turboprop 2,000-foot STOL a i r p l a n e , t h e p r o p e l l e r t i p speed can be reduced and provide a more e f f i c i e n t match between t h e desired propeller performance c h a r a c t e r i s t i c s f o r takeoff and c r u i s e f l i g h t con- ditions. Another b e n e f i t t h a t can be derived from reducing t h e takeoff propeller t i p speed is a reduction i n t h e p r o p e l l e r noise i n t h e takeoff mode of f l i g h t . Figure 22 presents a d e s c r i p t i o n of t h e e f f e c t s of t h e propeller t i p speed on t h e perceived noise l e v e l contours f o r t h e turboprop 2,000-foot STOL a i r p l a n e during a takeoff. This curve shows perceived noise l e v e l contours f o r both 1,000-f oot-per-second p r o p e l l e r t i p speeds and 800-f oot-per-second p r o p e l l e r t i p speeds. This curve shows t h a t f o r t h e airplane f i t t e d with p r o p e l l e r s having a n 800-f oot-per-second t i p speed, t h e perceived noise l e v e l is nearly 10 decibels lower thzn f o r t h e a i r p l a n e f i t t e d with propellers using a 1,000-foot-per-second t i p speed.

Figure 23 a l s o shows t h e e f f e c t s of t h e p r o p e l l e r t i p speed on noise This curve presents t h e maximum r a d i a l distance during t h e takeoff mode.

from the a i r p l a n e a t which a given perceived noise l e v e l is detected.

Curves a r e presented f o r t h e turboprop V/STOL a i r p l a n e f i t t e d with p r o p e l l e r s r o t a t i n g a t a 1,000-foot-per-second t i p speed and f o r t h e turboprop 2,000- f o o t STOL a i r p l a n e f i t t e d with propellers rotating with p r o p e l l e r t i p speeds of 1,000-f oot-per-second and 800-f oot-per-second. The primary difference between noise l e v e l f o r t h e turboprop V/STOL a i r p l a n e and t h e turboprop a propeller r o t a t i n g a t 1,000-foot-per- 2,000-foot STOL a i r p l a n e f i t t e d with second t i p speeds are t h e power differences between these two airplanes.

The engines of t h e turboprop V/STOL airplane develop approximately 60% more power than do t h e engines of the turboprop 2,000-foot STOL airplane.

It is important t o note t h a t while the source noise l e v e l between using 1,000-f oot-per-second and 800-f oot-per-second t i p speed i s not g r e a t at distances very close t o the airplane, s h a r p reductions i n noise do occur as t h e distance from t h e a i r p l a n e i s increased. These reductions occur p r i - marily because t h e low frequency band noise l e v e l s have been reduced f o r t h e p r o p e l l e r having an 800 f p s t i p speed. The higher frequency noise levels, which have not been appreciably reduced, a t t e n u a t e much more r a p i d l y than do t h e lower frequency noises.

A.ccuracy of noise predictions methods. - I n order t o g e t an assessment

of t h e accuracy of t h e noise prediction methods t h a t have been u t i l i z e d i n t h i s study a n d t h e study reported i n Reference 1, a comparison has been made of measured and calculated perceived n o i s e l e v e l s f o r t h e X C - 1 4 a a i r p l a n e and the Breguet 941 airplane. Figure 24 presents a comparison of t h e measured and calculated perceived noise l e v e l s f o r t h e X C - 1 4 2 ~ a i r p l a n e i n hover. The calculated Curves ceEe QQ% as pure c i r c l e s about t h e hover point, whereas t h e measured d a t a have lobes located 45 degrees t o l e f t o r r i g h t i n f r o n t and a f t around t h e airplane.

Figure 24 shows t h a t these lobes i n the quadrants a f t of the a i r p l a n e f o r t h e 80 PNdb noise l e v e l go beyond t h e calculated l i n e s s l i g h t l y . The lobes i n t h e forward quadrants of t h e airplane do not extend t o t h e c a l - culated l i n e s . For t h e 90 PNdb level, t h e measured lobes ext,end t o t h e calculated l i n e s i n t h e a f t quadrant and again do not extend t o t h e calcu- l a t e d l e v e l s i n t h e forward positions. When t h e measured l i n e s extend beyond t h e c a l c u l a t e d l i n e s , t h e noise is g r e a t e r than would be calculated.

These curves show t h a t the calculations can be as much as 7 decibels i n It should be e r r o r f o r t h i s p a r t i c u l a r f l i g h t condition and t h i s airplane.

noted that f o r t h e 100 PNdb level, t h e calculations very c l o s e l y agree with t h e measured values.

Figure 25 presents a comparison of measured and calculated noise l e v e l s f o r t h e Breguet 941 as measured from a s i d e - l i n e p o s i t i o n during a takeoff One was 70 f e e t t o t h e s i d e of t h e ground r o l l . Two microphones were used.

c e n t e r l i n e of t h e runway and t h e other 370 f e e t t o t h e s i d e of t h e runway c e n t e r l i n e as shown on Figure 25. The calculated values are compared with measured values t h a t were m d e during f o u r d i f f e r e n t takeoff runs. I n general, t h e c a l c u l a t i o n s f o r microphone number 1 p o s i t i o n a r e higher than

t h e measured values - by as much as 5 decibels f o r one frequency range. For

t h e microphone l o c a t i o n number 2, t h e calculations are much more accurate; but i n t h e higher frequency bands, one position was found t o be c a l c u l a t i n g excessive noise by nearly 9 decibels.

Figures 24 and 25 show t h a t t h e e x i s t i n g prediction methods can make reasonably close e s t i m t e s of noise i n general; but these f i g u r e s a l s o i l l u s t r a t e that t h e e x i s t i n g c a l c u l a t i o n methods are t o t a l l y inadequate f o r making accurate estimates of noise f o r a wide v a r i e t y of conditions and a t a l l octave bands. It should be kept i n mind that an e r r o r of f i v e t o t e n decibels out of 115 seems l i k e a very small percentage, but an increase of s i x decibels a t any l e v e l means t h a t t h e noise f o r t h e higher d e c i b e l l e v e l is twice as loud as f o r t h e lower l e v e l . Additional improvement is needed on noise estimating methods f o r V/STOL a i r c r a f t t h a t u t i l i z e p r o p e l l e r s .

It i s also expected t h a t improvements w i l l be required on noise estimating methods f o r j e t powered V/STOL a i r c r a f t .

S U M M A R Y As a r e s u l t of t h e a d d i t i o n a l examinations and perturbations m a d e on t h e designs developed i n response t o Reference 2 and reported i n Reference 1, t h e following conclusions a r e draim: 1. A V/STOL short-haul. t r a n s p o r t a i r p l a n e should have s e r i o u s consid- e r a t i o n given i n t h e s e l e c t i o n of i t s design c h a r a c t e r i s t i c s t o t h e possi- b i l i t y t h a t t h i s a i r p l a n e may have t o operate at nonoptimm c r u i s e conditions.

Such considerations would probably r e s u l t i n redesigning t h e a i r c r a f t of Reference 1 which were optimized for a 500-rni1.e s t a g e length. This redesign would permit t h e a i r c r a f t t o operate a t higher equivalent a i r speeds than would be required if t h e a i r p l a n e were a t optimum c r u i s e conditions.

2. If space is a v a i l a b l e f o r f u e l , V/STOL a i r c r a f t can use s l i g h t l y increased takeoff distance and obtain a large increase i n t h e maxisum operational stage length.

The design of V/STOL a i r c r a f t is very s e n s i t i v e t o t h e design s t a g e 3 .

length, and t h e choice of t h e b e s t V/STOL arrangement may vary as t h e design s t a g e length is varied.

Proper matching of t h e p r o p e l l e r takeoff RPM and t h e engine takeoff 4.

RPM f o r turboprop V/STOL a i r c r a f t designs can provide DOC b e n e f i t s and reductions i n t h e far f i e l d noise c h a r a c t e r i s t i c s of t h e s e airplanes.

These changes d i d not reduce t h e takeoff performance of t h e turboprop STOL airplanes, but they did give increased c r u i s e speed. For t h e turboprop VTOL airplanes, t h e reduced p r o p e l l e r takeoff t i p speed and t h e increased engine takeoff R P M reduced t h e hover performance, and, hence, it w a s neces- s a r y t o increase t h e engine s i z e .

5 . I n general, as t h e a i r c r a f t s i z e increases, t h e perceived noise l e v e l c h a r a c t e r i s t i c s i n takeoff of t h e V/STOL a i r p l a n e s increase.

6 . The e x i s t i n g noise p r e d i c t i o n methods are inadequate t o make accurate predictions of t h e noise of propeller-driven a i r c r a f t .

1 . Narsh, K . R., “Study on t h e F e a s i b i l i t y of V/STOL Concepts f o r Short -Haul Transport A i r c r a f t , I’ NASA CR-670, January 1967 2. Contract NM2-3036, “Study on t h e F e a s i b i l i t y of V/STOL Concepts f o r Short-Haul Transport A i r c r a f t ” TABLE I V/STOL SHORT-HAUL TRANSPORTS DESIGN GROUND RULFS Passenger plus baggage weight is 200 pounds per passenger Revenue cargo is 10% of t h e design passenger weight The perceived noise l e v e l i n t h e cargo compartment s h a l l not exceed 75 decibels i n takeoff o r 70 decibels i n c r u i s e The landing gear is designed f o r a 12 fps r a t e of s i n k The a i r p l a n e s t r u c t u r a l design c r i t e r i a is t h a t defined by Federal Aviation Regulations, P a r t 25, Airworthiness Standard: Transport Category Airplanes Takeoff and landing performance is based on sea level, 86°F day Special VTOL design c r i t e r i a : T/W = 1.15, a l l engines operating, no c o n t r o l input T/W = 1.05, a l l engines operating, 50% of t h e maximum c o n t r o l about t h e c r i t i c a l a x i s p l u s 20% about t h e o t h e r two axes T/W = 1.05, t h e c r i t i c a l engine inoperative, no c o n t r o l input T/W = 1.0, t h e c r i t i c a l engine inoperative, 5O$ of t h e maximum c o n t r o l about t h e c r i t i c a l a x i s plus 20% about t h e other two axes Special STOL design c r i t e r i a : Takeoff f i e l d length is c a l c u l a t e d assuming a c r i t i c a l engine is failed Landing f i e l d length required i s t h e c a l c u l a t e d required landing distance divided by 0.60 The rate of descent s h a l l not exceed 800 fpm during t h e landing approach The maximum d e c e l e r a t i o n roll during t h e landing ground r o l l s h a l l not exceed 0.5 g ' s TABLE 2 COMPARISON OF AIRPLANES DESIGNED FOR 300- AND 500-MILE STATUTE MILE STAGE LENGTHS Item Turboprop VTOL Fan-in-Wing V/STOL Design S t w e Length, S.Mi. 500 300 500 300 Gross Weight, lb. 62,300

55,950 79,587 63 , 300

Design VTOL Weight, lb. 62,300

52 , 320 72,827 56,555

Fuel Load, lb. 6,407 3,835 3-7 , 190 7,210

SHF' o r Thrust per Engine 5,960 5,080 6,400 5,160 Propeller o r Wing Fan Diameter 18.3 Ft. 16.1 Ft. 87 In. 79 In.

Optimum Cruise Altitude, Ft. 35,000 25,000 35,000 35,000 I Optimum Cruise Speed, Knots 350 395 460 460 TABLE 3 ESTIMATED WEIGHT BREAKDOWN 60-PASSENGER PROPULSIVE W I N G V/STOL AIRPLANE Weight.

Component Pounds W i n g Group . . . . . . . . . . . . . . . .

. . . . . .

4. 966 TailGroup . . . . . . . . . . . . . . . .

. . . . . .

1 . 559 B o d y Group . . . . . . . . . . . . . . . .

. . . . . .

7. 445 Alighting Gear . . . . . . . . . . . . . .

. . . . . .

2. 743 Flight Controls Group . . . . . . . . . .

. . . . . .

3. 596 Nacelle Group . . . . . . . . . . . . . .

. . . . . .

2. 238 Engines . . . . . . . . . . . . . . . . .

. . . . . .

4. 760 Exhaust System . . . . . . . . . . . . . .

. . . . . .

Lubricating System . . . . . . . . . . . .

. . . . . . 140

Fuel System . . . . . . . . . . . . . . .

. . . . . .

Engine Controls . . . . . . . . . . . . .

. . . . . .

Starting System . . . . . . . . . . . . .

. . . . . . 200

Fan System . . . . . . . . . . . . . . . .

. . . . . .

6. 127 Hot G a s Ducting System (including d i v e r t e r values) . .

1 . 052 Instrument Group . . . . . . . . . . . . .

. . e . . .

Hydraulic and Pneumatic Group . . . . . .

. . . . . .

E l e c t r i c a l Group . . . . . . . . . . . . . . . . . . .

1 . 336 Electronics Group . . . . . . . . . . . . . . . . . .

Furnishing Group . . . . . . . . . . . . .

. . . . . .

5. 391 Air-conditioning Group and Anti-Icing . . . .

. e . . 1 . 423 A u x i l i a r y Gear Group . . . . . . . . . . .

. . . . . . 40

TOTAL EMPTY WEIGHT . . . . . . . . . . . . . . . . . .

45. 475

Water. Food. Beverage. e t c . . . . . . . . . . . . . . 633

C r e w Plus Baggage . . . . . . . . . . . . . . . . . . 520

Passengers Plus Baggage . . . . . . . . . . . . . . . 12. 000

Cargo . . . . . . . . . . . . . . . . . . . . . . . . 1 . 200

Fuel (including unusable f u e l ) . . . . . . . . . . . .

13. ...

O i l . . . . . . . . . . . . . . . . . . . . . . . . . 250

T O T A L USEFUL LOAD . . . . . . . . . . . . . . . . . . 27. 825

TAKEOFF GROSS WEIGHT . . . . . . . . . . . . . . . . .

73. 300 a J m 60-PASSENCER TURBOPROP VTOL m I 9 30 X

E

I

i!

! 2o

DIRECT OPERATING COSTS - $/SEAT M I L E

Figure 2 . D . O . C . Versus C r u i s e A l t i t u d e 60-PASSENGER TURBOPROP 2000-FT STOL CRUISE ALTITUDE- 1OOC FT

DIRECT OPERATING COSTS - $/SEAT MILE

Figure 3. D . O . C . Versus Cruise A l t i t u d e 60-PASSENGER PROPULSIVE WING 2000-FT STOL I Figure 4 . D.O.C. Versus Cruise Altitude W t a l Distance t o Clear a 50-Ft Obstacle S E A LEVEL 86°F ONE ENGINE FAILED FIGURE 5a TAKEOFF

DISTANCE - FT

GROSS WEIGHT -

1000 LB 0 200 400 600 1000 RANGE - S. M I .

Figure 5. Effect of Operational Range on Takeoff Distance .21 TURBOPROP 1000-FT STOL Total Distance to Clear a 50-Ft Obstacle SEA LEVEL 86°F ONE ENGINE FAILED FIGURE 6a TAKEOFF

DISTANCE - FT

GROSS WEIGHT- 1000 LB 0 200 600 800 1000 RANGE - S.MI.

Figure 6. Effect of Operational Range on Takeoff Distance F A N I N WING V/STOL Total Distance t o Clear a 5 O - F t , Obstacle SEA LEVEL 86°F ONE ENGINE FAILED TAKEOFF

DISTANCE -

FT 800 U GROSS

WEIGHT -

1000 LB 0 200 400 600 800 1000 RANGE - S . M I .

Figure 7. Effect of Operational Range on Wkeoff Distance PROPULSIVE WING 1000-FT STOL Total Distance to Clear a 50-Ft Obstacle SEA LEVEL 06°F ONE ENGINE FAILED TAKEOFF 80C

DISTANCE - FT

FIGURE 8b GROSS

WEIGHT -

1000 LB 0 200 400 600 800 1000 RANGE - S. MI.

Figure 8 . Effect of Operational Range on Takeoff Distance -86 ET 6 IN.

(1038 IN. ) I /i-27.4/ Figure 9. Propulsive Wing V/STOL Airplane 60-PASSENGER TURBOPROP VTOL DESIGNED BASE AIRPLANE: FOR A 500 STA M I STAGE LENGTH Figure 10. Effect of Takeoff Propeller Tip Speed, Engine Overspeeding, and SHP on Payload

SLX - a3ms 3smt13 ~ A V

c, m u a u) -3 c u 0;' 0;' TURBOPROP VTOL VARIABLE PASSENGER LOAD 500 MILE STAGE LENGTH

* Ratio of alternate

airplane D . O . C . to basic airplane D . O . C .

Figure 1 4 . Relative D . O . C . for Tip Speed, Engine Overspeed and Horsepower Variations cLmmI Figure 1 6 . Landing Drag %lar CL -2 -1 CD 0 1 2 3 4 -5 -4 -3 Figure 17. T a i l - O f f , Nose Fan InoDerative, Power-on Rlar f o r the 60--Passeng& Fan-in-Wing V/STOL CL -1 -2 -3 0 .4 .8 1 . 2

- 1 . 6 -1.2 - a8 -.4

cD Nose Fan Inoperative, Figure 18. Tail-Off, Power-on Polar for the 6O-passenger Propulsive 2000-Ft STOL Wing 0 0 cu

3 \o

Zd 0007: - 33NvJJsIa

I B ( u cr I I

8 20

L1: H

*

:: 12

I 3 60 70 00 90 100 110 120

PERCEIVED NOISE LEmL - DECIBELS

* Distance i s the m a x i m u m radial distance from the airplane a t which

t h e PNdb is at t h e l e v e l indicated.

Figure 23. Effects of Power and Propeller Tip Speed on Noise,Takeoff XC-142A AIRPLANE IN HOVER Figure 24. Comparison of Measured and Calculated Perceived Noise NASA-Langley, 1961 - 2 c ~ - 6 7 0 ( 01)