On the aerodynamic optimization of mini-RPV ... - CAFE Foundation

AI AA-84-2 163On the Aerodynamic Optimizationof Mini-RPV and Small GA AircraftF. R. Goldschmied, Monroeville, PAAlAA 2nd Applied Aerodynamics ConferenceAugust 21 -23, 1984/Seattle, WashingtonFor permission to copy or republish, contact the American Institute of Aeronautics and Astronautics1633 Broadway, New York, NY 10019

ON THE AERODYNAMIC OPTIMIZATION OF MINI-RPV AND SMALL GA AIRCRAFlFoblo R. Goldschmied"Monruevi 1 le, PA 15146vAbstractFWWing drag, lbA brief study has been carried out on the adap- H,Jet total-head @ Sta. 5, lblft'tation of an optimized system comprising an axisymmetricbody, suction boundary-layer control and AH,, Total-head rise of fan between Sta.stern jet-propul sion, which was developed origi- 2 and 5, lb/ft2nally for lighter-than-air application, to mini-RPVWing lift, 7band small GA aircraft by the addition of dynamicwina lift. For mini-RPV. consideration has been n Fan soeed, , RPMgiven to fuselage diameters of 20 and 34" with aFree-stream static pressure, lb/ft2gross weight range from 125 to 300 lb at the speedsof 100 and 150 Kn. The predicted powers rangedStatic base pressure @ Sta. 5, lblft'from 2.35 to 16.20 HP.P5For the GA aircraft,consideration has been givento fuselage diameters of 45 and 60" with a grossweight range from 1400 to 3400 lb at the speed of200 MPH. The predicted powers ranged from 60.6to 132.5 HP.AARBCAH, 5CH,, = -4"H, - PCHT,o= -q0LCL = qNomencl at ureWing area, ft2Aspect ratioWing span, ftWing chord, ftWing drag coefficientFan pressure-rise coefficientJet total-head coefficientWing lift coefficientP, - PCP5 = - Static base pressure coefficientq0 @ Sta. 5IWCTw = -qovo'66TOCTo = ___d,Dqovo.66Fan suction flow coefficientThrust coefficient for wingTotal thrust coefficientDiameter of stern jet and offan, ftDiameter of fuselage, ftqo = w,24Tu =tnd,nUOVNOEl- wouoHP550'IFVP$ = A n5 diutFree-stream dynamic pressure, lb/ft2Fan suction flow, ft'lsecThrust of fusel agelboundary-l ayercontrol/jet system, lbFan tip speed, ft/secFree-stream velocity, ft/secJet velocity @Fuselage volume, ft3Gross weight, lbSta. 5, ft/secAerodynamic efficiency indexFan total efficiencyKinematic viscosity of air, ft2/secMass density of air, lb sec2/ft4Fan flow parameterIntroductionThe concept of the optimum aerodynamic integrationof body pressure-distribution (with concomitantshape), slot-suction boundary-layer control andstern jet-propulsion was presented in 1967'; a windtunnelverification with a self-propelled test modelwas presented in 1982,' showing 50% power reductionas compared to the best streamlined body with sternwake-propeller. An optimized LTA system was derivedfrom the above data' and it was also shown that jetpropulsionof a subsonic body with free transitionwas achieved with jet total-head equal to freestream's.'It can be noted that, for the same massflow, a conventional free-stream jet propulsor wouldhave a jet total-head coefficient CHT, = 4 or, forthe same diameter, a coefficient CHT, = 2, as shownin Ref. 5, on the basis of the best conventionalstreamlined body of equal volume.The optimization of streamlined bodies by shape*Consulting Engineer; Associate Fellow AIAA. only was presented in 1974;6 for all-turbulent.-/ boundary-layers, little drag gain could be obtained@Copyright 1984 by Fabio R. Goldschmied, P.E. by optimizing the shape. This extensive programReleased Io AlAA Io publish in all lorms.

proved conclusively that boundary-layer controland propulsion had to be integrated with the bodyshape if substantial power gain had to be achieved.The objective of this paper is to investigateheavier-than-air applications of this optimizedsystem, i.e. to investigate the installation ofa lifting wing onto the fuselagefboundary-layercontrolfjet propulsion system. The evaluation willbe carried out on the basis of an aerodynamic efficiencyindex 6 = W,U,/HP for two classes of aircraft,i.e. mini-RPV @ 100 and 150 Kn and smallGA (General Aviation) 2-seat and 4-seat aircraft0 200 MPH.Qtimizeo Body/Boundary-Layer Control/Jet-Propblsion System: *Tunnel Te-ss_tThe wind-tunnel test program of the 20" diameterself-propelled model was carried out in 1981 inthe 8x10 low-speed wind tunnel of the David TaylorNaval Ship R&D Center; the test program was quiteextensive as it comprised over 800 test pointsorganized in 86 test runs. The test results arepresented in Refs. 2, 3 and 4. The basic testmodel with open jet (Conf. 00) is shown in Fi1 (starboard photo) and in Fig. 2 (stern photoy:In Fig. 1 the 12" chord strut can be plainly seenas large as a wing would be; its interferenceeffect is already accounted for in all the testresults. Fioure 2 shows clearly the three radialrakes in the jet nozzle to measure the flow andthe jet total-head... .:. ./jFig. 2:.:.:.. . . 8Stern photo of wind-tunnel test model(Conf. 00)Fig. 1 Starboard photo of wind-tunnel test model(Conf. 00)The axial force (drag or thrust) measurementswere based both on the wake's momentum balance,as indicated by a wake rake, and on the forceexerted on the strut, as indicated by the windtunnelbalance which was supporting the strut.At the higher thrust coefficients, above CT =0.010, there was in all cases excellent agreementbetween the two types of axial force measurements,as shown in Figs. 7, 8 and 9; therefore the thrustdata cannot be disputed in any way.The model was also tested with a tailboom inthe jet nozzle (Conf. 01) as shown in Fig. 3.Fig. 3 Port photo of wind-tunnel test model withtailboom (Conf. 01)The layout of the test model's aftbody with thesuction slot, fan installation and jet nozzle isshown in Fig. 4. While the fan should have beenat Station 5, the arrangement shown had to be acceptedfor practical reasons. The fan mass-flowweighted mean pressure rise is computed betweenStations 2 and 5, with the flow being measured atStation 5; the fan air power is determined by theproduct of flow and pressure rise.Figure 5 presents a photo of the axial faninstallation in the forebody of the wind-tunnel testmodel: the fan discharae is in the left foresround,while the fan intake i s indicated by the Foundededge inside the forebody.._.2

L&&-R =ssx106 I... iMSz1.00 0.7 0.4 0.6 0.8 1.0 12 1.4 1.6 -... 1.8 L.-L 2.0 2.2 L...U. 2.4 2.612.8 1.0Dimanrianler~ hidl 0iildnce. XI0Fig. 6System AnalysisExperimental static-pressure distributionon test model @ 6" angle of attackFig. 4 Aftbody layout with suction-slot, faninstallation and jet nozzle (Conf. 00)A procedure has been developed for the additionof wings to the body and for the computation of theadditional thrust acquired from the jet to counteractthe wino's draa. The wino's liftldraa ratioshave been coGputed irom classical NACA data;' betterresults could be obtained with modern airfoils suchas the Liebeck, the NASA GA(W)-l and -2, etc. Thefan performance has been based on the tested NASAaxial rotor/stator stage 516; the stage's designand experimental performance is given in Ref. 8......:...Fig. 5 Photo of fan installation in wind-tunneltest modelThe typical stepwise static pressure distributionon the body i s shown in Fig. 6 at 0' and 6' angleof attack. The stepwise distribution is used (withsuitable boundary-layer control) to avoid the largegrowth of the boundary-layer momentum-thicknesswhich nonnallv occurs in the adverse Dressuregradient area "and to achieve very low external wakedrags, as shown in Refs. 1 and 2.While in Refs. 2, 3 and 4 the wind-tunnel testdata were plotted only up to CT = 0.020, since themain focus was on the equilibrium point (CT = O),in this paper all the available test data are plotted,for both free transition on the body and fortransition tripped at 10% length.The experimental wind-tunnel test data of Ref.2 have been replotted in the complete thrust rangeand are shown in Figs. 7, 8, 9, 11, 12 and 13. Thefan selection plot is given in Fig. 10, with the516 performance curves relating pressure, flow andefficiency.The computational procedure comprising 15 stepsis given below:1. Select the max. cruise speed U, and thecorresponding dynamic pressure qo; determine orestimate the gross weight Wo.2. Select fuselage diameter to yield adequatecabin space and/or equipment volume.3. Assume lift coefficient CL of wing @ max.cruising speed. Compute the wing loading qoCL andthe wing area A = Wo/qoCL. A lift coefficient CL= 0.40 is selected for the mini-RPV and CL : 0.30is selected for the GA aircraft.4. Compute the wing chord and span, assuminga wing aspect-ratio AR = 10 and constant chord:c =QB = 1ocActually the wing may have a taper ratio and thecomputed chord value is the niean chord.5. Compute the wing lift/drag ratio at theselected CL point, using one of the following twoexperimental equations from Ref. 7 (Fig. 18, p. 21)for wings of aspect ratio AR = 10:a. Cg=O.0068+0.0343 Cf (NACA 23018 Airfoil)b. Cg=0.0045+0.0383 Cf (NACA 653-418 Airfoil)A parametric wing design study should be made, inthe manner of Koegler,' but it is beyond the scopeof this brief paper.3

Compute6. Compute the drag of the wing:WOFw = cL/cDsince the wing lift must equal the gross weightW,. Compute the thrust coefficient:required to generate the thrust to counterbalancethe wing drag. Add 10% to the wing thrust coefficientfor wing/fuselage interference drag, althoughthe model was tested in the wind-tunnel with astrut large enough to be a wing; also add anotherthrust coefficient increment of 0.003 for theempennage and tailboms.CT, = CTw f 0.10 CTw f 0.003LIn the wind-tunnel tests of Ref. 2, an empennageadequate to yield neutral static stability to thebody up to 8' had an incremental power coefficientof only 0.0020.7. Obtain the value of the fan air power coeffi cient :aL._UcY rn0.02kM I 01*from the experimental plot CHP~J vs CT as presentedin Fig. 7, corresponding to the above CT, for thefree transition or for the tripped transition case,as warranted by the Reynolds number and by operationalconsiderations.8. Obtain the value of the fan flow coefficient:Fig. 70 0.01 0.02 0.03 0.MThrust Coefficient CT =Tk,Vo'66Fan air power coefficient CHP25 vs Thrustcoefficient CTfrom the experimental plot CQs vs CT as presentedin Fig. 8, corresponding to the above CT, for thefree transition or for the tripped transition case,as warranted by the Reynolds number and by operationalconsiderations. Compute the fan flow 4 =cq, x U0VQ.66.I ..* I I I9. Obtain the value of the fan pressure-risecoefficient :0.03-TransitionAnCHli = 2qofrom the experimental plot C H ~ S vs CT as presentedin Fig. 9, corresponding to the above CT, for thefree transition or for the tripped transition case,as warranted bv the Revnolds number and bv ooerationalconsid&ations. ~ the fan i r e b erise AH,,= CH, x q,.10. The selection of the best axial rotor/statorstage for the job requires the computation of thefan system resistance coefficient O'/$ in the O Ndomain :15--2LY 5a"@ = 3557Fig. 8Fan flow coefficient CQ5 vs Thrustcoefficient CT44

p 1.00 ,* i I I I INASA Axial1Transition0.4 -Thrust Coefficient CT = T/qoVa6Fig. 9The denomination ofFan pressure-rise coefficient CH2s vsThrust coefficient CT$ is the fan flow parameter:0’3 -Fan System Resistance0. 2 -@=$-7 d:utYThe denomination of + is the fan total-pressureparameter:The fan diameter corresponds to the jet diameterd, ; ut = ndSn is the fan tip speed and P is theair mass density. Figure 10 presents the $4 plotwith two fan system resistance curves, representingthe max. and min. encountered in this study, andwith the experimental performance curves $4 and$-n~ of the selected NASA axial rotor/stator stage518;’ it can be seen that a fairly good match isachieved, i.e. the fan will operate between 88.3and 90.6% efficiency. Fan aerodynamic selectionprocedures are discussed thoroughly in Section 6of Ref. 10.11. The fan speed is computed:n =-- cq5 ‘a 130.13 RPS$ VO.33The value of the fan flow parameter + is determinedin Fig. 10 from the intersection of the fan systemresistance curve with the NASA 518 performancecurve. The range of $ is from 0.485 to 0.508.Similarly, the fan efficiency n~ is determined inFig. 10 from the NASA 51B efficiency curve at theabove 0 location.The fan diameter is assumed to correspond to thestern jet diameter: ds = 0.1625D.12. The fan shaft power can be computed now fromthe CHP25 value of Step 7 and the above efficiencydetermination for an available fan design:0.1 - I I I I0 0.2 0.3 0.4 0.5 0.6Fan Flw Parameter 0Fig. 10 Fan pressure parameter + and totalefficiency QF vs fan flow parameter $13. The jet total-head coefficientCHT,Hs- Poqo= __is determined from the plot of Fig. 11 and the jetvelocity ratio Us/U, is determined from the plotof Fig. 12 for the free transition or the trippedtransition cases, as warranted by Reynolds numberand opera t i onal con si dera ti on s.14. The jet static base pressure coefficient CP,is determined from the plot of Fig. 13. It can beseen that the static base pressure coefficient rangeis very high; CPs will be over 0.8 for the thrustcoefficient range of this study. As a reference,Ref. 11 shows that the’ base pressure coefficientof conventional boattail/jet afterbodies is in the0.10 to 0.15 range.15. The final step is the computation of theaerodynamic efficiency index:The index expresses the power required to fly theaircraft’s gross weight at the max. cruise speed;therefore, it i s a good indication of the efficiencyof the aerodynamic design when comparingtwo aircraft at the same speed.5

~Mini -RPVL& 9,- '\FreeTransiiionfiConi. Wake1 andTransitionTripped B 58% 0I I I I I0.01 0.02 0.03 0.04 0.050.66Thrust Ccefficient CT = T/q V0Fig. 11 Jet total-head coefficient CHT, vSThrust coefficient CTMini-RPV aerodynamic design has not achieved yetan adequate degree of efficiency for the missionspeed and endurance requirements. Considering typicalcurrent vehicles such as the Air Force/BoeingPave Tiger, the Atmy/Lockheed Aquila, the Israel WAircraft industries Scout, the Tadiran Mastiff MK3and the Developmental Sciences Sky Eye, it is foundthat the gross weight ranges from 220 to 380 lb,the maximum cruising speed ranges from 85 to 100Kn and the engine powers range from 22 to 30 HP.The aerodynamic efficiency index ranges from 2.5to 3.5.The wind-tunnel test model,2"'4 with its diameterD = 20.0 in. and 100 Kn speed, may be classifiedas a full-scale mini-RPV; Table 1 presents itsperformance data for 125, 150 and 175 lb grossweights with a suitable wing (CL = 0.40) and empennage.Table 1 20" Diameter (V = 6.2 ft3) Mini-RPVP 100 Kn (qo = 34.1 PSF)(Free Transition)1.8 I I I I1Gross Weight W, lbWing loading qoCL @ CL= 0.4PSF125 It13.6150 lb13.6175 lb13.60. 2 c01001 lo-II I I I J0.01 0.02 0. m 0. M 0.05Thrust Ccefficient CT = TPoVo'66Fig. 12 Jet velocity-ratio U,/U,coefficient CT1.4, I I I Ivs ThrustI I I I I0 0.01 0.02 0.03 0.M 0.05Thrust Ccefliclent Cl =T/qoVo'MFig. 13 Jet static base pressure coefficient CP,Thrust coefficient CTIvsWing area A = Wo/qoCL, ft2Wing chord, C ftWing span, B ftLift/drag ratio of wing,CL/CD @ CL = 0.40Drag of wing, Fw = WCLo / ~Thrust coeff. for wingCTwTotal thrust coeff. CTaFan air power coeff. CHP,Fan flow coeff. CQ,Fan flow 9CFSFan pressure-risecoeff.CHzsFan pressure-riseAH2 iPSFFan system-resi stancecoeff.@2/*Fan speed, nRPMFan diameter, d, in.Fan efficiency, '?F %Fan shaft powerHPJet total-head coeff. CHT5Jet velocity ratio U,/UoJet static basepressure coeff.cp5Aerodynamic efficiency F.index9.190.9589.5837.63.320.02850.034E0.0600.023713.502.5486.360.78533,9643.2589.752.352.661.351.0416.3411.031.05010.5037.63.990.03480.04120.06850.025214.352.8095.200.80535,7563.2590.502.782.901.421.1017.20__12.861.13411.3437.64.651.04061.04761.07551.026415.033.05103.700.81137,3113.2590.753.123.161.511.1018.22V6

~~4VWith free transition the fan shaft powers are3 HP and less at the 100 Kn speed; the aerodynamicefficiency index is over 16. The engine may bean available Fox Twin, yielding 3 HP @ 14,500 RPMand weighing 3 lb with mount and muffler. Table2 presents the performance data of the same 20 in.fuselage at 150 Kn speed with free transition,while Table 2a presents the corresponding data withtransition tripped @ 10% length on the fuselage,for gross weights of 150, 175 and 200 lb. It isfound that the fan shaft powers are less than 7HP with free transition and less than 8 HP withtripped transition; the corresponding aerodynamicefficiency index values are over 13 and 10, respectively.As a direct comparison with a current mini-RPVdesign, the fuselage diameter was increased to 34in. to match both the lateral dimension and thelength of the Air Force/Boeing Pave Tiger's fuselage;Table 3 presents the performance data @ 100Kn with free transition while Table 3a presentsthe corresponding data with transition tripped @10% length, for gross weights of 225, 250 and 2751 b.7Table 2 20" Diameter (V = 6.2 ft') Mini-RPV@ 150 Kn (qo = 77.3 PSF)(Free Transition)~ __175 lb 203 lbWing loading qoCL @ C~=o.430.8 30.8Wing area A = Wo/qoCL,Wing span, BLift/drag ratio of w'CL/CO @ CL = 0.40ft2Wing chord, C =tCLDrag of wing, Fw = W0/%Thrust coeff. for wing CTwTotal thrust coeff.Fan air power coeff. CHP,Fan flow coeff. CQ,Fan flow QCT,CFSFan pressure-risecoeff. CH2 5Fan pressure-riseAH,,PSFFan system-resi stancecoeff. '$2/ $Fan speed, nRPMFan diameter, dSin.Fan efficiency, 'iF %Fan shaft powerHPJet total-head coeff. CHTsJet velocity ratioJet static basepressure coeff.Aerodynamic efficiency4 indexU,/Uocp5E4.870.6976.9737.63.990.015:0.01910.039!0.020;17.201.94150.00.74644,0713.2589.05.292.061.160.8613.005.680.7537.5337.64.650.01790.02260.04320.021017.892.05158.80.76345,4123.2589.65.742.201.200.8914.006.490.8058.0537.65.320.02050.02550.04710.021618.402.16166.90.76646,6173.2589.76.262.301.250.9314.69With free transition the fan shaft powers areless than 6 HP while with tripped transition thepowers are less than 20 HP; this may be comparedwith the 28 HP enaine of the 250 lb Pave Tiaer.The aerodynamic efficiency index values are over14 and 11, respectively.Finally, Table 4 presents the performance dataof the 34 in. diameter fuselage at 150 Kn withgross weights of 250, 275 and 300 lb and withtripped transition. It can be noted that this casewith 300 lb represents a substantial performanceimprovement over the Pave Tiger, in both speed (50%gain) and weight (20% gain); the fan shaft poweri s 16 HP and the aerodynamic efficiency index is8.75. It can be noted that, with tripped transition,there is no laminar flow risk and that theturbulent power coefficients should actually belower because the Reynolds number is higher by thefactor 1.7 x 1.5 = 2.55.A schematic layout of the proposed mini-RPV configurationi s shown in Fig. 14; the pylon/wingarrangement was proposed by Larrabee" for glidersso as to maximize the wing's lift. It can be notedthat the pylon/fuselage interference effect wasalready simulated in the wind-tunnel tests by thestrut. The wing span is 162 in. and the usefulfuselage length is 83 in., while the overall fuselagelength is 127 in. The empennage is supportedby a single boom.Small General Aviation AircraftSmall GA aircraft cornwise another cateoorv.~ ..to which this system analysis may be applied withinteresting results.Table 2a20" Diameter (V = 6.2 ft3) Mini-RPV@ 150 Kn (qo = 77.3 PSF)(Transition Tripped @ 10% Length)Gross Weight Wo lbTotal thrust coeff. CT,Fan air power coeff. CHPZSFan flow coeff. CQ,Fan flow QCFSFan pressure-risecoeff.CH,Fan pressure-riseAH* 5PSFFan system-resi stancecoeff.'$2/$Fan speed, nRPMFan diameter, d5 in.Fan efficiency, 'i~ %Fan shaft powerHPJet total-head coeff. CHTSJet velocity ratio U,/UoJet static basepressure coeff.cp5Aerodynamic efficiency Eindex__150 110.019c0.0500.022;18.92.24173.10.78149,0343.2588.36.752.281.280.9810.40~175 I t0.0220.054:0.022719.32.35181.60.77849,8353.2589.97.192.401.321.0211.19-200 lb0.02550.05870.023419.92.48191.70.78350,1403.2590.27.752.501.361.0611.84__7

~~~~Table 334" Diakter (V = 30.5 ft3) Mini-RPV@ 100 Kn (so = 34.1 PSF)(Free Transition)Gross Weight Wo lbWing loading qoCL @ C~=0.4PSFWing area A = Wo/qoCL, ftZWing chord, C ftWing span, B ftLift/drag ratio of wing,CL/CO @ CL = 0.40Drag of wing, Fw =CLThrust coeff. for wing CTwTotal thrust coeff. CT,Fan air power coeff. CHP,,Fan flow coeff. CQ5Fan flow QCFSFan pressure-risecoeff.CHssFan pressure-riseAH, _ _PSFFan system-resi stancecoeff. $2/*Fan speed, nRPMFan diameter, d in.Fan efficiency, n~ %Fan shaft powerHPJet total-head coeff. CHT5Jet velocity ratio U,/UoJet static basepressure coeff.CPSAerodynamic efficiency Eindex-225 lb13.616.51.28412.8437.65.981.01801.02281.04361.021034.952.0569.90.763:7,9165.52589.104.982.201.200.9014.0__!50 I t13.618.41.35613.5637.66.651.02011.02511.046f1.021f35.942.1472.9D.7738,3545.52589.405.292.301.240.9214.6__175 lb13.620.21.42114.2137.67.31'LO2211.02738.0495'.022036.602.2376.03.7708.6195.52539.605.652.381.260.9515.1Table 434" Diameter (V = 30.5 ft') Mini-RPV@ 150 Kn (so = 77.3 PSF)(Transition Tripped @ 10% Length)Gross Weight Wo lbWing loading qoCL @ C~'0.4PSFWing area A = Wo/qoCL, ftZWing chord, C ftWing span, B ftLift/drag ratio of wing,CL/CD @ CL = 0.40CLDrag of wing, Fw = W o / r lb0Thrust coeff. for wing CTwTotal thrust coeff. CT,Fan air power coeff. CHP,,Fan flow coeff. CQsFan flow QCFSFan pressure-risecoeff.CH25Fan pressure-riseAH*,PSFFan system-resi stancecoeff.m2/*Fan speed, nRPMFan diameter, d in.Fan efficiency, QF %Fan shaft powerHPJet total-head coeff. CHTsJet velocity ratio U&Jet static base CP 5pressure coeff.Aerodynamic efficiency Eindex250 lb 275 lb30.8 30.86.65 7.310.0089 0.00980.0128 0.01380.0390 0.04050.0198 0.020248.8 49.81.92 1.98148.4 153.00.724 0.73125,862 ' 26,2765.525 5.52588.4 88.615.00 15.701.90 1.961.15 1.160.87 0.897.728.19-I00 lb__30.89.740.9869.8637.67.981.01061.01461.04161.020550.52.02156.30.7385,5575.52588.816.202.001.180.908.75 WWTable 3a 34" Diameter (V = 30.5 ft') Mini-RPVP 100 Kn (qo = 34.1 PSF)(Transition Tripped @ 10% Length)Gross Weight Wo lbTotal thrust coeff. CTOFan air power coeff. CHP,Fan flow coeff. CQ,Fan flow QCFSFan pressure-risecoeff.CH,,Fan pressure-riseAH25PSFFan system-resistancecoeff.m2/+Fan speed, nRPMFan diameter, d5 in.Fan efficiency, n~ %Fan shaft powerHPJet total-head coeff. CHT,Jet velocity ratio U,/UoJet static basepressure coeff.CP 5Aerodynamic efficiency Eindex~275 lb0.02730.06120.023739.432.5486.60.78520,0185.52589.86.902.561.381.0812.26Table 5 lists the pertinent parameters of threecurrent 2-seat personal GA aircraft by Beech,Cessna and Piper; the gross weight is nearly thesame (1670 to 1675 lb) and so is the speed (121to 127 MPH) and the engine power (108 to 115 HP).The aerodynamic efficiency index is in the vicinityof 5.0.The fuselage diameter is selected at 45 in. soas to accommodate two tandem seats; the cabinarrangement layout is shown in Fig. 15. It hasbeen observed that minimum cabin height is 42 in.and the minimum seat spacing dimension is 36 in.-The speed has been selected to be 200 MPHinstead of 125 MPH because it represents presentGA speed for small personal aircraft. Table 6 presentsthe performance data for gross weights of1400, 1675 and 1800 lb with tripped transition.It can be seen that the fan shaft power i s lessthan 72 HP and the aerodynamic efficiency indexis over 12.0. This performance yields a 60% speedimprovement, a 33% power gain and a 7.5% grossweight enhancement. lhe schematic layout of this2-seat GA configuration is shown in Fig. 16, witha mid-wing arrangement with 0.45 taper ratio and277 in. overall span. The useful fuselage lengthis 110 in. while the overall fuselage length is130 in. The empennage is supported by a twin boomand it comorises twin rudders.8

L >I48" I4-Seat Cabin Arrangement - 60" Oia. Fuselage- 110" LFig. 146.7'0.2725.04nUSchematic layout of mini-RPY configurationTable 5Empty weight W lbGross weight Wo IbWing span B ftWing area A ft'Wing loading W/A2-Seat E4 Aircraft @ 120 MPH(qo = 37.0 PSF)___IBeech Cessna Piper77 152 PA-38-112,kipper ierobat Tomahawk 2PSFLength 2. ftEngine powerMax. cruisespeedWing liftcoeff.AerodynamicefficiencyindexUoCLHPMPHE__1,1031,67530.013012.8824.01151210.3414.70124.724.1 23.11 122 1270.3225.05Note: The above data were derived from AviationWeek & Space Technology, March 12, 1984, p. 144.Fig. 152-Seat Cabin Arrangement - 45" Dia. FuselageCabin arrangement layout for GA aircraftTable 6 45" Diameter (V = 70.6 ft3) E4 Aircraft@ 200 MPH (qo = 104 PSF)(Transition Tripped @ 10% Length)--Gross Weight Wo lb 1400 1 b 675 1 I 800 1 b__ -Wing loading qoCL @ C~=0.3 31.2 31.2 31.2PSFWing area A = Wo/qoCL, ft2 44.8 53.7 57.7Mean wing chord, C ft 2.11 2.31 2.40Wing span, B ft 21.1 23.1 24.0Lift/drag ratio of wing,CL/CO @ CL i 0.3038.3 38.3 38.3Drag of wing, Fw = Wo/rlbCL0Thrust coeff. for wing CTw36.50.020643.71.024;47.01.0265Total thrust coeff. CTo 0.0256 1.030: 1.0321Fan air power coeff. CHP, 0.0588 1.0651 1.0685Fan flow coeff. CQ,0.0234 1.0241 1.0248Fan flow QCFS 117.1 122.1 124.1Fan pressure-rise2.45 2.66 2.74coef f . C"* 5Fan pressure-rise254.8AH,,PSF276.6 284.9Fan system-resistance0.793 0.794 0.796coeff. m'nFan speed, nRPM 25,777 !6,82! 17,211Fan diameter, d5 in. 7.31 7.31 7.31Fan efficiency, '1~ % 90.1 90.2 90.3Fan shaft powerHP 60.6 68.6 71.6Jet total-head coeff. CHTs 2.5 2.66 2.74Jet velocity ratio U,/Uo 1.37 1.41 1.44Jet static base cp5 0.92 1.10 1.11pressure coeff.Aerodynamic efficiency E 12.1 13.0 13.3index--9

~~~Table 7 lists th'e pertinent parameters of fivecurrent 4-seat GA aircraft by Maule, Mooney, Piper,Cessna and Beech, with speed from 196 to 201 MPH,gross weights from 2500 to 3400 lb and enginepowers from 200 to 285 HP. The aerodynamic efficiencyindex ranges from 6.5 to 7.9.Table 7 4-Seat 6R Aircraft @ 200 MPH(so = 104 PSF)The cabin dimensions for four seats are typically43 in. width, 48 in. height and 92 in.length; a fuselage diameter of 60 in. has beenselected and the cabin layout arrangement i s shownin Fig. 15.Table 8 presents the performance data for grossweights of 2500, 2900 and 3400 lb at 200 MPH withtripped transition. It is seen that for the lowestweight the fan shaft power is 110, for a gain of47%; for the middle weight the fan power is 120HP, for a gain of 40%, while for the top weightthe fan power is 132 HP, for a gain of 53%.r - r f l; Schematic 2-Seat GA Aircraft Layout1 Confiouration1675 6 Gross Weight68 HP (Tripped Trans. ILengthEngine powerftHPMax' cruise U, MPHspeedWing liftcoeff.AerodynamicefficiencyindexCLE__2125340033.5181.018.824.4 24.7 27.3 28.6 26.7210.0 200.0 200.0 235.0 285.0196.0 201.0 198.0 199.0 198.00.157 0.148 0.167 0.172 0.1846.49 7.47 7.90 7.22 6.50Fig. 16Schematic layout of 2-seat GA aircraftconfigurationGross Weight W, lbWing loading qoCL @ C~=0.3PSFWing area A = Wo/qoCL, ft2Mean wing chord, C ftWing span, 8 ftLift/drag ratio of wing,CL/CO @ CL = 0.30Drag of wing, Fw = Wo/-lbCLCDThrust coeff. for wing CTwTotal thrust coeff. CTOFan air power coeff. CHP,,Fan flow coeff. CQ,Fan flow QCFSFan pressure-ri secoeff.CH..L DFan pressure-riseAH"./

The aerodynamic efficiency index is over 12.0.The schematic layout of the 4-seat GA configurationis shown in Fig. 17; a high-wing arrangement wasselected with a span of 360 in. and a 0.45 taperratio. The useful fuselage length i s 147 in. whilethe overall fuselaoe lenoth i s 174 in. The emDennageis supported -by twin booms and it comprisestwin rudders.ConclusionsIt has been shown that the simole addition ofa conventional NACA wing to the tested optimizedsystem comprising axisymmetric body, boundary-layercontrol and stern jet propulsion yields a predictedaircraft perfonnance which is superior to currentmini-RPV and small GA aircraft levels.Schematic Layout4-Seat GA AircraftConfiguration2903 Ib Gross Weight2% MPH Speed120 HP (Tripped Trans. )A 250 lb gross weight mini-RPV could be flownat 100 Kn with 7.0 HP; a 2-seat 1675 lb GA aircraftcould be flown at 200 MPH with 69 HP, and a 4-seat2900 lb GA aircraft could be flown at 200 MPH with120 HP.The general trend of the aerodynamic efficiencyindex E against the thrust coefficient CT is shownin Fig. 18 for both free transition and transitiontripped at 10% length; it can be seen that thethrust coefficient should be kept above a valueof 0.025 in order to achieve aerodynamic efficiencyindex values above 12.0. This can be done byadjusting the gross weight Wo against the fuselagevolume V and the speed Uo.Better results can be expected if the liftsystem should be integrated, i.e. if the fuselagesuction air mass flow should be directed laterallyinto the wings to supply jet flaps or circulationcontrolwall-jets. The limitation then would bethe allowable wing loading and the wing's induceddrag, as the lift coefficient is substantiallyincreased by the jet-flap or by the circulationcontrol.The air mass flow would serve three purposes:(a) fuselage boundary-layer control, (b) wing liftenhancement, and (c) propulsion. Also the suction'smomentum drag is minimized since only theinner boundary-layer flow i s drawn into the slotat Sta. 1 (Fig. 5); it is estimated that a 33%drag saving is achieved, as compared to free-streamflow intake, at the point of equilibrium for thebody alone. The next step of this program is expectedto be the wind-tunnel test of the 20 in.diameter model in the arrangement of Fig. 14; thewing pylon would be installed exactly as thepresent strut. A second step would be the investigationof the optimum fineness ratio; the 2.72fineness ratio was selected originally for LTAapplication. A preliminary theoretical parametricstudy indicates that the optimum should be near6 and that further power reduction can be expected.174" I181 Free Transitionand16 - Transition TrippedT aQ 58%\,**--o-C 14 - # .ox-Ea-.$< 12E=-lO-"35u II 8.-gi0- ,2'6--I'2-SeatGAB 123 MPH- I// Mini-RPV2 -,'@1WKn 701I I I I0 0.01 0.02 0.03 0.04 0.05Thrust Coefficient CT = T/qoVo'66-_/--/'# s t Transition,*' Tripped B 10%;:LMPH 7Fig. 17Schematic layout of 4-seat GA aircraftconfigurationFig. 18 Aerodynamic efficiency index E vsThrust coefficient CT11

~~ --~ ~~,References1. F. R. Goldschmied, "Integrated Hull Design,Boundary-Layer Control and Propulsion of SubmergedBodies," AIAA Journal of Hydronautics,Vol. 1, No. 1, pp. 1-11 (July 1967).2. F. R. Goldschmied. "Inteorated Hull Oesion.~~~Boundary-Layer Control and Propulsion of SubmergedBodies: Wind-Tunnel Verification,"AIAA Paper 82-1204, AIAA/SAE/ASME 18th JointPropulsion Conference, Clevela nd, OH (June1982).3. F. R. Goldschmied, "Wind-Tunnel Demonstrationof an Optimized LTA System with 65% PowerReduction and Neutral Static Stability," AIAAPaper 83-1981, AIAA Lighter-Than-Air SystemsConference, Anaheim, CA (July 1983).4. F. R. Goldschmied, "Jet-Propulsion of SubsonicBodies with Jet Total-Head Equal to Free-Stream's,'' AIAA Paper 83-1790, AIAA AppliedAerodynamics Conference, Danvers, MA (July1983).5. F. R. Goldschmied, "Aerodynamic Integrationof Ax i symme t ri c Body Pressure- Di s t r i but i on ,Slot-Suction Boundary-Layer Control and SternJet-Propulsion, I1 - Propulsion Evaluation,"AIAA Journal of Aircraft (to be published).6. J. S. Parsons, R. E. Goodson and F. R. Goldschmied,"Shaping of Axismetric Bodies forMinimum Drag in Incompressible Flow," AIAAJournal of Hydronautics, Vol. 8, No. 3, pp.100-107 (July 1974).7. I. H. Abbott, A. E. von Doenhoff and L. S.Stivers, Jr., "Sumnary of Airfoil Data," NACAReport 824 (1945).8. G. Kovich and R. J. Steinke, "Performance ofLow-Pressure-Ratio Low-Tip-Speed Fan Stagewith Blade Tip Solidity of 0.65," NASA TMX-3341 (February 1976).9. J. A. Koegler, Jr., "A Parametric Wing DesignStudy for a Modern Laminar Flow Wing," NASATM 80154 (December 1979).10. C. 0. Wood and F. R. Goldschmied, "Design andSpecification Guidelines for Large Draft Fan:.and Systems," Electric Power Research instituteReport EPRI CS-3431 (January 1984).11. D. E. Reubush and J. F. Runckel, "Effect 0:Fineness Ratio on Boattail Drag of Circular-Arc Afterbodies Having Closure Ratios of 0.50with t Exhaust at Mach Numbers up to 1.30,"I3SA ThJ-7192 (May 1973).12. E. E. Larrabee, "Preservation of Wing Leading-Edge Suction at the Plane of !Fetry as aFactor in Wing-Fuselage Design, Proceedingsof the NASA-Industry-University General AviationDrag Reductic? Workshop, J. Roskam, Ed.,University of Kansas (July 1975), pp. 107-119.12