NASA-CR-165768-1981 Design and analysis of a fuel-efficient single-engine turboprop-powered business airplane《节能单发动机涡轮螺桨发动机驱动的商务机设计和分析》.pdf

《NASA-CR-165768-1981 Design and analysis of a fuel-efficient single-engine turboprop-powered business airplane《节能单发动机涡轮螺桨发动机驱动的商务机设计和分析》.pdf》由会员分享，可在线阅读，更多相关《NASA-CR-165768-1981 Design and analysis of a fuel-efficient single-engine turboprop-powered business airplane《节能单发动机涡轮螺桨发动机驱动的商务机设计和分析》.pdf（42页珍藏版）》请在麦多课文档分享上搜索。

1、a117600_666214j ,v“ I _11,_NASA Contractor Report 165768NASA-CR- 165768I _100 _24, q,2.,. DESIGNANDANALYSISOF A FUEL-EFFICIENTSINGLE-ENGINE,TURBOPROP-POWERED,BUSINESSAIRPLANEG. L. Martin, D. E. Everest, Jr., W. A. Lovell, J. E. Price,K. B. Walkley, and G. F. WashburnKENTRONINTERNATIONAL,Inc.HamptonT

2、echnical Center _ _. ,.,-,-.,_._,_-;_,_an LTV company L :,. _ ,-Hampton, Virginia 23666SEP! 5 1981_JM_L,:., ,_.:,.,-:.,:R:.; CEHTE tt.; B;,-“,R;CONTRACTNASl- 16000“ August 1981L-A_N/ ANationalAeronauticsandSpaceAdministrationLangleyResearchCenterHampton,Virginia23665Provided by IHSNot for ResaleNo r

3、eproduction or networking permitted without license from IHS-,-,-!Ipm lb8_J_Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARYA studywas conductedto determinewhethera generalaviationairplanepoweredby one turbopropenginecouldbe configuredto have

4、a speed, range,andpayloadcomparableto currenttwin-engineturbopropaircraftand also achieveo“ a significantincreasein fuelefficiency. An airplane.configurationwasdevelopedwhich can carry six peoplefor a no-reserverangeof 2 408 km(l 300 n.mi.)at acruise speedabove154 m/s (300 kt) and a cruise altitudea

5、bove9 144m (30000 ft). This cruisespeed is comparableto that of the fast-est of the currenttwin turboprop-poweredairplanes. The airplanehas a cruisespecificrangegreaterthan all twin turboprop-engineairplanesflyingin itsspeed rangeand most twin piston-engineairplanesflyingat considerablyslowercruise

6、airspeeds. The high thrust-weightratioand maximumuse of high-liftdevicesproducetakeoffand landingdistancesof less than 762m (2 500 ft) atmaximum grossweight for airportpressurealtitudesup to 1 830m (6 000 ft).INTRODUCTIONA large segment of the business community depends upon the use of twin-engine t

7、urboprop general aviation aircraft to satisfy their transportationrequirements. These aircraft have been developed to the point where they nowprovide reliable, all-weather capability and block times comparable to turbojetaircraft for short and medium ranges. This, combined with the ability to operat

8、eout of smaller airports, provides the business community with a very versatile airtransportation system.A continuing effort is being made by the NASAand industry toward improvingthe fuel efficiency of aircraft. This study was conducted to determine whethera single-engine turboprop business aircraft

9、 could be configured to have a speed,“ range, and payload comparable to current twin-turboprop business aircraft andalso have a substantial increase in fuel efficiency. Such a configuration wouldprovide a considerable savings in initial cost, maintenance costs_ and fuelcosts over current twin-engine

10、 business aircraft.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-In order to evaluate the performance characteristics of a single-engineturboprop-powered airplane and the fuel conservation improvements which couldbe gained, a six place configuratio

11、n was developed which would have a designcruise speed of 154 m/s (300 kt) and cruise at altitudes above 9 144 m(30 000 ft). The design range is 2 408 km (I 300 n.mi.) without reserves. The ostructure is of conventional aluminum construction with cabin pressurization suchthat a cabin altitude of 2 43

12、8 m (8 000 ft) could be maintained to the serviceceiling. A relatively small wing of 11.15 m2 (120 ft 2) and an aspect ratioof 8 was specified in order to achieve better cruise performance. The higherwing loading, resulting from the use of a small wing, required extensive use ofhigh-lift devices in

13、order to provide acceptable low-speed characteristics.SYMBOLSValues are given in this report in both International System Units (SI) andU.S. Customary Units. All calculations were made in U.S. Customary Units.A aspect ratioCD drag coefficient, D/qSCD. induced drag coefficient1CD parasite drag coeffi

14、cientPCDtri m trim drag coefficientCL lift coefficient, L/qSCL propeller integrated lift coefficientI Cp power coefficient .D drag, N (Ibf)J propeller advance ratioProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-L lift, N (Ibf)q dynamic pressure, Pa

15、(Ibf/ft 2)S reference area, m2 (ft 2)Vcruise cruise speed, m/s (kt)Vstal I stall speed, m/s (kt)ACD lift-dependent parasite drag coefficientPACDpower power-dependent drag coefficient6f trailing-edge flap deflection angle, degrees_s leading-edge slat deflection angle, degreesn propeller efficiencySub

16、scripts:min minimumAbbreviations:AF propeller activity factorFAR Federal Aviation RegulationMAC meanaerodynamic chord, m (ft)CONFIGURATIONDESCRIPTION “ The airplane is of conventional design with a low wing, a front mounted_. engine, and an aft located horizontal tail. This configuration was selecte

17、dinstead of a more unconventional configuration, such as those with pusher propel-lers or canards, in order to provide a more accurate comparison of the relativemerits of using one engine or two engines.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,

18、-The wing selected has an area of 11.15 m2 (120 ft2), a span of 9.39 m(30.8 ft), an aspect ratio of 8 and a taper ratio of 0.33. A high wing loading,1.9 Pa (39.6 Ib/ft2), was desired in order to provide a better match for the wingat the cruise condition. A 15 percent thick airfoil section, the NACA6

19、52-415,was selected for use on the unswept wing to reduce the wing structural weight. .In order to simplify the leading-edge slat and trailing-edge flap mechanism,straight leading and trailing edges were used on the wing. The taper ratio of.33 produces a nearly elliptical spanwise load distribution

20、for a straight-tapered wing, and also results in a lighter wing structure.The high-lift devices used on the wing consist of a single-slotted 30 per-cent chord trailing-edge fowler flap and a 15-percent chord leading-edge slat.The flaps have a 20-percent chord extension, 40 degrees of deflection and

21、extendfrom the fuselage to 90 percent of the semispan. The remaining I0 percent ofthe trailing edge is used for the aileron. Roll control is provided primarilyby spoilers located ahead of the flaps on the upper surface of the wing. Thesmall ailerons are used to provide linearity in the roll control

22、system. Fullspan leading-edge slats, with a deflection of 26 degrees, are used on thisconfiguration to increase the obtainable lift. The extensive use of high-liftdevices are required on this configuration in order to provide a stall speedwhich meets the requirement of FAR 23.49 (ref. I).The wing de

23、sign is of conventional riveted aluminum, skin-stringer construc-tion with spars located at 15 and 65 percent of the chord. These spars passthrough the fuselage under the cabin floor. Integral fuel tanks in the wingshave sufficient capacity to contain the fuel required for the design mission.The fus

24、elage cabin was designed to seat six 97.5-percentile men, includingthe pilot. The length of the cabin is 3.70 m (12.125 ft), the height is 1.42 m(56 in) at the center, and the width is 1.28 m (51 in) at elbow level. Thesecond and third rows of seats have pitches of .86 m (34 in) and 1.12 m (44 in) “

25、respectively and are separated by a .23 m (9 in) wide aisle. Entrance to the tcabin is through a door located at the left rear. An escape hatch is providedon the right side of the cabin. The baggage compartment is located aft of thepressure bulkhead. The layout and dimensions of the cabin are presen

26、ted infigures 1 and 2. This cabin has more room than most of the current six-placetwin-engine aircraft.4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The fuselage is designed for a pressure differential of 56.9 kPa (8.25 psi)which allows a 2 438 m

27、(8 000 ft) cabin pressure altitude to be maintained to a. cruise altitude of 12 190 m (40 000 ft). The cabin pressurization system isconventional, utilizing engine bleed air to maintain the required level ofpressurization.The configuration is equipped with a tricycle landing gear arrangement.The nos

28、e gear rotates and retracts aft into a compartment under the cabin floor.The wing-mounted main gear retracts inward into the wing root and fuselage.iA conventional empennage arrangement is used on the study aircraft. Thevertical tail has an area of 1.79 m2 (19.25 ft 2) with a 30 percent full spanrud

29、der. The horizontal tail has an area of 2.69 m2 (29 ft 2) and a 30 percentchord elevator. The horizontal tail is mounted so that the elevator is aft ofthe fuselage and extends the full span of the horizontal tail.A general arrangement drawing of the configuration is presented as figure3. A sunTnary

30、of the geometric characteristics is contained in table I.PROPULSIONANALYSISThe engine used for this study was the Pratt and Whitney Aircraft of CanadaPT6A-45Awhich is a lightweigh_free-power turbine, turboprop engine. Installedperformance for this engine was generated with the aid of an engine perfo

31、rmancecomputer program provided by Pratt and Whitney Aircraft of Canada. The engineperformance thus generated is based on the following installation effects andconstraints at all altitudes, airspeeds, and throttle settings:Inlet ram pressure recovery .98Service airbleed .113 kg/s (.25 Ibm/s)Accessor

32、y power extraction 7.46 kN (I0 HP)“ Propeller speed 1700 rpmConvergent nozzle exhaust area .058 m2 (90 in 2)- Nozzle discharge angle 0 (parallel to engine centerline)As a result of constraints and limits built into the PT6A-45A computer pro-gramjit was not possible to generate performance data above

33、 an altitude ofProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-9 144 m (30 000 ft). In order to provide performance data adequate to encompassthe desired airplane flight envelope, it was necessary to extrapolate engineperformance data to an altitude

34、of I0 668 m (35 000 ft).o-Propeller selection and performance estimation were based on the HamiltonStandard methods presented in reference 2. These methods are based on a seriesQof performance maps which provide systematic variations of the basic propellershape and aerodynamic parameters. The perfor

35、mance of a given propeller isaccurately defined by the map over the complete range of potential operatingconditions.Preliminary design considerations resulted in the definition of a four-bladed propeller with a 2.13 m (7 ft) diameter. A parametric analysis was thenperformed to determine the optimum

36、(i.e., highest efficiency) combination ofactivity factor and integrated lift coefficient for several combinations ofcruise power and speed at altitudes of 9 144 m (30 000 ft) and I0 668 m(35 000 ft). Based on this analysis, a final propeller was selected which wasbest matched to the required cruise

37、conditions and which also provided acceptableperformance throughout the flight envelope. Table II summarizes the propellerdesign point and characteristics.Figures 4 and 5 summarize the estimated propeller performance for altitudesfrom sea level to I0 668 m (35 000 ft). The engine power setting is th

38、at formaximumclimb/cruise power for each altitude. This performance summary indicatesincreasing efficiencies and decreasing thrusts with speed for a given altitude.Propeller efficiencies near 0.85 are indicated for the cruise conditions.Additional thrust is produced by the engine exhausts and varies

39、 accordingto both power setting and airspeed. This thrust ranges from 533.8 N (120 Ibf)at takeoff power and sea level static conditions to 13.3 N (3 Ibf) at cruiseconditions at I0 669 m (35 000 ft). At low power settings and high airspeed,such as may be experienced in descents, the thrust produced b

40、y the exhausts -becomes slightly negative.6Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WEIGHTSANALYSISThe empty weight was estimated to be 11.20 kN (2 520 Ibf) at atake-off gross weight of 21.1 kN (4 750 Ibf). This gross weight includes4.58 kN (I

41、 030 Ibf) for the fuel required to meet the no reserve range of 2 408 km (I 300 n.mi.) and 5.34 kN (I 200 Ibf) for the passengers and baggagewhich comprise the payload. The avionics included in the weight analysis arecomparable to those currently in use on general aviation aircraft operating inthe h

42、igh-altitude IFR environment. The design ultimate load factor used forthis study was 6.66, which places the study aircraft in the utility category.A detailed weight breakdown is presented in table III. For the purpose ofmaintaining conventional weights engineering terminology, .89 kN (200 Ibf) ofthe

43、 payload (pilot) is included in the operating weight empty of table III.The requirement of maintaining a 2 438 m (8 000 ft) cabin pressure altitudeto the service ceiling resulted in a cabin pressurization differential of56.88 kPa (8.25 psi). Additional structural and systems weight penalities forpre

44、ssurization to this level were estimated to be .89 kN (200 Ibf).The weight data for the engine and its accessories were obtained from ref-erence 3. Avionics and propeller weights were obtained from manufacturers data.The center of gravity travel for this airplane ranges from 3 percent MACat the oper

45、ating weight empty condition to 30 percent MACat the maximumgrossweight condition.AERODYNAMICANALYSlSThe airplane was designed primarily for high-speed cruise conditions:and_therefore, has a wing loading much higher than other single-engine general-. aviation aircraft. Because of the high wing loadi

46、ng, an extensive use of high-lift devices was required to produce sufficient lift to allow the configuration_“ to meet the FAR23.49 (ref. I) stall speed limit of 31.4 m/s (61 kt). The systemdesigned to meet this requirement incorporates the maximumpractical application ofhigh-lift devices with full-

47、span leading-edge slats and 90 percent span single-slotted7Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-fowler flaps. Someadditional lift could be obtained by using multiple-slottedflaps, but the complexity of such a mechanism does not lend itself

48、 to use on ageneral aviation aircraft.The lift characteristics were determined using the methods presented inreference 4. This method assumes that all slots in the flap system have beenoptimized. The lift curves are presented in figure 6 for flap deflections of0 degrees, 5 degrees, and 40 degrees. These flap settings are used for cruise,takeoff, and landing, respectively. Using the maximumtrimmed lift coefficientof 2.97, a minimum stall speed of 32.26 m/s (62.7 kts) can be obtained, which is.9 m/s (1.7 kt) above the FAR 23 requirement. A slight increase in wing areawould allow the configurat