Preliminary sizing and performance of aircraft

APPENDIX B DEFINITIONS OF INPUT The program is coded in FORTRAN Extended Version 4 (ref. 15). Provisions are unadefor two input files. One file contains propeller design data and is required only if propeller output data is to be computed. The data on this file must be • input in the fo ll owing fo rmat.

Col uLi.._ Type o f Data g -l O number c f propeller blades, a right justified integer @ 11-20 Mach number 21-30 propelle r advance ratio 31-40 propeller power coef f icient 41-50 propeller thrust coefficient i) s tais listed by increasing power coe f ficient at constant advance ratio, then by increasing advance ratio at constant Mach number, then by increasing Mach number.

During program execution, the file containing this data must be named TAP_3.

The other file contains the input for the baseline aircraft and sizing options. It is always required and during execution it must be called TAPE5. The input segments on TAPE5 are listed below in their order o f occurrence.

Input segments for sized aircraft " Aircraft Identification SPRNTCON $DESGNVB _TWINFUS ,_ , _ $141SSVB _ SRESRV _ STEC H CHG ,:".

Input segments for baseline aircraft Wave-Drag Geometry Deck or $SACGEM . $AIN SWTIN _EN I N Baseline Engine Characteristics I Segments prefixed by a $ sign are inputs in the NAMELISI format (ref. 15).

Oefinitions of input variables follow, with default values, if they exist, enclosed ,, parenthesis behind each definition. Variables with parenthesis are arrays and the included numbe r s indicate thei r maximum size.

. . ..... 14E

: r :i INPUT SEGMENTS FOR SIZED AIRCRAFT The first input is the aircraft identification. It is one free-field line of 8U characters.

_, $PRNTCON , __ Inputs in this segment control program output. All inputs are intege r s with a +_ value of eithe r I o r U. Unless noted, a value of 1 activates the option; (J suppresses it.

"_ _;_ IPPRNT = I prints all o utput _I : 0 prints only output for sized aircraft (default) , I I IPAERU prints aerodynamic data, (O) IPRPR U P prints engine characteristics (U) IPRWTS H prints weight statement, (0) ;] IPRMPRF pr+nts mission pr u file data, (O) _'i iSCOPE prints output showing progress of execution during interactive ._ terminal sessions, (0) _! $DESGNV_ This input segment contains aircraft design and sizing information. For clarity, the inputs have been separated into typical groups.

Miscellaneous Design Input ISSAC = C) for supersonic aircraft (default) = I for subsonic aircraft ISGEUM = {] if aircraft geometry is input in wave-drag format (default) = i if aircrdft geometry is input in $SACGEM IJP = I for aircraft usieg JP fuel, (0) IH2 = i for aircraft using hydrogen fuel, (0) IMTH = I for aircraft using methane fuel, (0) , NOFIW = I fo r airc r aft with hu fuel in wing, (U) i IBLAC = 0 for sizing of baseline aircraft (default) I = I for no sizing of baseline _ircraft I RHUJP density of JP fuel, lb / gal, (6.75)

2+ i

RHOH2 density of liquid hydrogen fuel, Ib / gal, (.592) } RHOMTH density of liquid methane fuel, Ib / gal, (3.54) I ; RHOCAR density of cargo, Ib / ft3 (I0) !

EIPEN redu c tion in thrust to a cc ount f or interferen c e effects between propeller and wing, percent, (0) !

i Th e n ex t t hr e e variables app l y to aircraft using hydrogen or methane fuel. i • KBFN = I for constant tuselage fineness ratio, (0) i KBL = I f o r con s tant fuselage length, (I) !

b KBR = I for constant fuselage radius, (0) I P a sse nger Cabin Design Input DGLW width of door-galley lavatory, ft, (7.3) NPPDGL number of passengers per door-galley-lavatory,integer, (75) LEVELS number of passenger levels, O, I, or 2; 0 provides the baseline value, (0) TP wall thickness of passenger cabin, inches, (4) STSAB number of seats abreast; 0 provides the baseline value STSABR, (0) TF wall thickness of fuel tanks for liquid hydrogen or methane fuel, inches (9) Specific Design Inputs F NDES _esign fu s elage fin e n es s ratio I _ IPASDES number of passengers, integer I- NENDE$ number of engines, integer NPDES number of engine nacelles, integer c RMAKDES design rad i u s of fuselage, ft ZLBDES design length of fuselage, ft NCARDES design cargo load, lb TUWDES design th. ' ust-to-_eight ratio TPEDES design single engine thrust, Ib WOSDES design wing loading, psf l ,j ' o; REFADES design wing referen c e area, f t 2 i WGDES design gr o ss weight, l b RNGDES design range, n.mi.

ZMCRDES design cruise Mach number The next six variab l es app l y to subsonic aircraft.

ASPRDES design wing aspect ratio EI]ES design endurance, h POWDES design shaft horsepower-to-weightratio, hp / I b L SHPDES design sha f t ho rsep o wer, hp SQCDES design sweep of win_ quarter chord line, deg _ WTOCDES design thickness ratio of wing, percent J I , If RNGDES is not input, the program solves for the range at input gross weight and passenger load. If RNGUES is input, the program so:yes foe the gro_s weight that provides this range at input passenger load.

To find the passenger load for a given r ange and gross _eight, use the following integer variable.

IGRAWFP = I (default is O) along with inputs for WGDES and RNGDES.

Each of the above variables, which defaults to its baseline value, is used to _, . _ hold that design parameter constant through one or more sizing passes during program I _ execution. If design parameters are to be changed in eacll sizing pass during execution, the following design variable arrays must be used. All arrays are one- dimensionaI.

Design Variable Arra_s.- TUWTB(7) an array containing design thrust-to-weight ratios Q TPETB(7) an array containing desig:lthrusts for a single engine, Ib NTOW the number of values in either the TUWTB or TPETB arrays, (O) WOSTB( 7 ) an array containing design wing loadings, psf REFArB( 7 ) an array containing a esign wing reference areas, ft2 NWOS the number of points in either the WOSTB or REFATB arrays, (0) AR T ,( I t) ) an array c o ntaining design wing aspe c t ratios ,!

i 30 N ASPR the nu mber o f poin t s in AR T B, (0 ) POWTB( / ) an array containing shaft horsepow_--to-weightratios N PU W t he n umber of v a lue s i n P OWTB, ( 0 ) Restrictions on Use of Variables. - Not a ll of the design variab l es are independent parameters. The fo ll owing restrictions, there f ore, are p l aced on the use o f severa l o f t hese variab l e s .

" I. Since P U W L )ES,TOW D ES, TPEDES, P L )WTB,T( J WTB,and TPETB a 1 1 control engine sizing, use o nly one of these variab l es.

• 2. Simi l arly, WOSDES, REFADES, WOSTB, and REFATB al l affect wing sizing, therefore, use on l y one.

3. The ARTB array must not be used with any other design va r iable array.

4. The integer variab l e IGRAWFP may not be used with any design variab l e array.

$T WINFUS This input segment contai n s input for the geometry of twin - fuse l age aircra f t.

The on l y component of this aircraft that can be sized is the propulsion system; therefore, on l y the design variab l es WGDES, TOWDES, TPEDES, and RNGDES are allowed. If changes in any other design variab l e, l ike wing size, p_.ssenger l oad, etc., are required, a new base l ine aircraft (with its aerodynamics weights a n d propu l sion) must be developed and input.

The on l y geometry input necessary for twin-fuselage aircraft are the r eference l engths and wetted areas of the components. These are required for ski r ,-friction ca l cu l ations during the mission analysis. To prevent program aborts, however, a comp l ete geometry input segment in either the wave-drag format or $SACG[M i_ a l so requi r ed but any available input segment will suffice because it will mot affect _._ resu l ts . : j ITWINF = U f o r c onventiona l aircraft, (de f au l t) ._. = = I for twin-fuse l age aircraft FURLiN fuse l age reference l ength, ft F USW T I,I fu s elag e wetted area, total of both fuselage, ft2 T WFREFA wing reference area, ftz • IWSEGIN number of wing segments WRLIN( Z U) a o ne-dimensiona l a r ray containing the reference l ength of each wing segment_ ft WSWETIN(ZO) a o ne-dimensiona l array c o ntaining the wetted area o f eac h wi r _g seg_nt , ft= C RLIN c anard r e feren c e leng t h, ft :'!i , , CSWETIN canard wetted area, ft2 ; , j ", HRLIN horizontal t a il r eference length, ft HSWETIN horizontal tail wetted area, ft2 IFINS number of vertical fins _', FRLIN(6) a one-dimensionalarray continuing the reference lengths of t h e vertical . fins ft d

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FSWETIN(6) a one-d_mensional array containing the wetted areas of the vertical v fins, ft_ PRLIN the reference length of one engine pod, ft PSWETIN the wetted area of one engine pod, ft2 ,; $m41 SSV9 _i This s egment provides the input s requir"d for the flight profile analysis. The ,_- inputs, again, have been separated into typical groups.

:._ Overall h,ssion Input IENDUR = I mission profile for long enduranr , :_ir_.aft, integer, (0) IRADIUS = I provides range to payload drop, integer (0) OEW operating empty weight, lb, (if greater than U, this value will replace the calculated L)EW) PAYLOAD total payload weight, Ib, (if greater than 0., this value will replace the calculated payload) P O LF payload factor, percent (I(]0% gives full payload), (10 U .)

NPASOL number of passengers in off-loaded aircraft, integer (-I, the default, gives no off-loading) WCAROL cargo weight in off-loaded aircraft, Ib (-10, the default, gives no off-loading) i ICALPRC = 1 provides data for the range-payload curve, integer, (0) OLRTB(IO) a one-dimensional array containing the ranges for fuel off-loading, ' I --| C ,I

_; 3 2

i RRTB(IO) a one-dimensional array containing reduced ranges with no fuel off-loading, n.mi.

NRR the number of values i n RRT_, integer, (0) IRFSOC = I start refue l ing at start of cruise, integer, (0) IRFEOC = I start refueling at end of cruise, integer, (0) RNGSRF range to start of refuel, n.mi.

- (refuelingmust occur sometime during cruise. If RNGSRF is set too low, refuel wi l l begin at the start of cruise; if set too high, refuel will begin at the end of cruise. To activate refueling, the value of RNGSRF must be greater than 10.), (0) $ RFRIGPM re f uel fl o w rate, gal / min, ( 600. ) WGARF the aircraft gross weight after refueling, Ib (-10, the default value, gives the design gross weight) DELTCG standard day temperature increment during main mission, deg., C (0.)

EM surface emissivity, (.8) TDFLM thrust deflection during main mission, deg., (0) Takeoff Input NOTO = I no takeof f calculations, (0) FWDTO fuel weight used during takeoff, Ib (may also be input as a fraction of gross weight), (0.)

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ALFDUT aircraft rotational speed, deg / sec, (3.) _I _ : t)ELTCTO standard day temperature increment during takeoff, deg, C, (0.)

% OTGRUP time for landlng gear retraction, sec, (10.)

IIEOUT = I one en g ine out during takeoff, integer, (0) • RFF ro l ling friction factor, (.02) TXTME taxi time, rain,(10.)

• GRALFA angle of attack before start of rotation, deg (-4) TOALFA angle of attack after rotation, deg, (5.5) TOOELF flap deflection during takeoff, deg, ( 30.)

TOFLTO thrust deflection during takeoff, deg, (0.)

I TFACT f a c tor for in c reasing or de c rea s ing net t hrust during takeo f f, (I.)

WFFACT corresponding f actor for fuel flow, (I.)

HUBS obstacle height, ft, (35.)

HNOGE a l titude for disappearance of aerodynamic in-ground effect, ft, i (wing span) d TOCLGR takeoff climb gradient, percent, (6.8) : Climb Input HASNT(50) a one-dimensionalarray containing the altitudes in the climb profile ZMASNT(50) a one-dimensionalarray containing the correspondingMach numbers in the climb profile ICLPR the number of points in the climb profile I Zr_STTH f_achnumber at which engine throttling may begin to conserve climb fuel, i !

(.61 I

ZMMPCL Mach number above which maximum engine thrust is used, (I.)

Cruise Input NCRP the number of points at which cruise calculations are made, (4) ZI_CR cruise Mach number ROC_IIN the minimum rate of climb during cruise, ft / min CRALTI altitude for start of cruise, ft (-10) CRALT2 altitude for second leg of two-step cruise (-10)

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I_FCR = I _ives cruise at best available !_reguetfactor, integer, (U) _ ICACR = I gives cruise at constant altitude, integer, (0) ICLCR = I gives cruise at constant lift coeffficient, integer, (U) ALFCR angle of attack for cruise, deg (Note: this input may be used only if CLAR or ALPHAT are input in NAMELIST segment $AIN) CRT U T_ cruise t l lrottlingfactor (0, ) (if CRTOTM = I., tull power is applied; if defaulted, CRTOTM is not used) 3 4 c T he fo ll o wing input is fo r subs o nic cruise l egs o n a supers o nic mission. Range or time f or cruise may be input, but not both at the same time. Cruise on both legs may occur.

For subsonic cruise on the outbound l eg: ZMS I _C O B cruise Hach number HSBCUB cruise a l titude, ft • SSRNGOB cruise range, n.mi.

TSBCOB cruise time, min ' For subsonic cruise on the inbound leg: ZMSBCIB cruise _lachnumbe r H SBCIB cruise altitude, ft SSRNGIF_ cruise range, n.mi.

TSBCIB cruise time, rain Input f or Long-EnduranceCruise CLMAX maximum l ift coe ff icient, (2.2) NENCR number of engines operating during cruise (-I) DCFACT increment in drag coefficient for engines not operating, (0.)

VKCRI velocity at start of cruise, k IC(_NVCR = I for constant cruise velocity, (U) _ Descent Input ICALDS = I descent calcu l ati o ns are ma d e = 0 descent calculations are not made . NUSP the num b er of po ints at which descent calcu l ati o ns are ma_e, ( 4 ) DSFF the f raction of fuel flow at maximum power that is use d during descent,

(.o67)

WFDRCF an initial estimate of the fuel used during descent normalized by aircraft gross weight, (.005)

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. -#_ i , .!,I v . A p p ro a c h Input _, APALFA a n g l e of att a ck du r ing app r oach, deg, (-40) _ , APDELF fl ap deflection du r ing approach, deg_ (30.)

•' [FXAFLP : I use s input f l ap deflection _ ; : O finds flap def l e c tion for minimum d r ag, (default) ] "_ (used on l y i f VAPIN is input)

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VAPIN approach speed, k, (-10) _. (if defaulted, approach speed is calculated internally) ' No t e: VAPIN is used only i f aerodynamic data f or more than one f lap def l ection is input in SAIN.

J SRESRV i The i n put for the reser v e fue l legs are placed in this segment.

i! RFWT reserve fue l weight, in either pounds or as fraction of gross weight (use t h is input on l y if reserve fue l calculations are to the bypassed) : i IRCRO = 0 ca l culates c l imb, cruise, and descent segments of reserve fue l l egs = I ca l cu l a tes on l y cruise segments of reserve fuel l egs, (1_ Trip Fuel A11owance : TPFATME _s e at fue, flow at start of cruise, h or rain for subsonic aircraft), (0.)

TPFA percent of trip f_Jel (use for supersonic aircraft), (U.)

Missed Approach ICALMA = ( J uses R I ' IAFA to calculate fuel requirement } = I c a lc u l ates mi s sed approach to find fuel requirement, (i) RMAFA t ime a t ta k eo ff fue l fl ow, min, (2. 0 ) .

I A l ternate Airport l i AAU distance to a l terna i ceairport n.mi., (0.)

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F o r this l eg, the program normal l y iterates throug h Mach number, a l titude, and engine th o tt l ing during c l imb to find the conditions for l ea s t fue l . If any of t h e fo ll owing variab l e s are input, iteration on that condition wi ll not be performed.

HALT a l titu d e for cruise to a l ternate airport 3 6 ; ' T U T MA L T f a c t o r f o r engine thr o ttli n g d uring c l imb to c ruise, v a l ue between O _.

and I.

_, .

. , ZMA LT Macilnumber for cruise to alternate airport !

I : H ol d _ ; . TH L D time for h old, rain I i , H H L D altitude for hold, ft _ , ZM HL D Mach number for h ol d i Again, optimum conditions for least fuel are normally determined but inputs in I I HLD i and / or Z t I HLDwill e l iminate the iteration on t h at condition.

_m $TEC HC HG !

_, This segment contains the inputs for finding the effects of ch a nges in the t ech n o l og y of t h e various air c raft systems, i TIML the l o west _lachnumber at w h ich changes apply, (-1) TIMH t h e i,lghestF1ac h number at which cilangesapply, (10.)

l TDCDT(15) a one-dimensiona] array of percentage changes in total drag coef f icient, a function of ' 4achnumber; a positive value gives dr a g reductions TMDCUT(15) t h e Mach n umbers corresponding to t i levalues in T DCDT i NDCDT t h e numoer of values in the T DCDT a rray, (U) j]!, The following input is al l in pe r centages. A positive value gives a technology , improvement. Default values for all in p uts are O. L.;_ DSFC e ngine spe c i f ic f u el c on s ump t ion !

DCDW wave drag DCDRUF roughness drag DCDAC air-c o nditioning drag m DCDBL bleed drag I DCDO z ero-lift d rag DCDL d r ag due to lift i, DCDT total drag 3 7 .

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, <++' ., i- Friction Drag Changes DCDFWNG wing . r _: DCDFCS canards, ho r izontal tails, and vertical fins ¢ _. DCDFBDY fuselage '_+_J UCDFPDS na c e l ies "+ t)CD F total friction draQ _"_ Weight Changes ' t ' DWWING wing +:, DWCHV canard, horizontal tail, vertical fin +J ! DWFUS fuselage ++_ DWLG landing gear '_ UWNAC naceIles DWST total structure weight " DWENG engines DWTR thrust reversers \ UWTPR propellers DWTGB gear box 'JWMIS miscellaneous propulsion system welghts DWPLP _, plumbing system DWFT fuel tank for hyd r ogen or methane fue l ., DWINS fue l tank insulation f or hydrogen or methane fue l DWPPS total wei i. propulsion system ght DWSC surface controls J ' j I )WAP auxiIiary power DW[ N ST in s truments I DWHYD hydrauIics DWELE electrical +++m _ IIIF $ I + • , . . 'W_',-+'+,. _,_m._+ DWAV(JN avionics DWFE{_ furnishings and equipment DWACONLI air conditioning DWANTICE anti-icing DWSEQ total systems and equipment weight DWEMP empty weight DWOEW operating empty weight Laminar Flow The following inputs express the percentage of laminar flow on the surface. A value of O. gi_es all turbulent flow; a value of 10U. gives a11 laminar flow. All default values are O.

PCLAMW wing PCLA_ I F fuselage PCLAMC canard PCLAMH h orizontal tail ?

PCLAMBV body vertical fins PCLAMWV wing vertical fins PCLAMP nace11es INPUT FOR BASELINE AIRCRAFT GEOMETRY I : __ The geof._etry for the baseline airc r aft can be input by one of t a o m_thods. The first method which applies o n ly to supersonic aircraft use_ the zero-lift wave-drag program format described in reference 7 . This input is very detailed and unless it is already available from the analysis of wave d r ag, the effort required to assemble " the i nfo rmati u n i s no t ju s tified f o r this program. The alternate method available is the use of $SACGEM.

$ SACGEr4 Wing RE F A wing re f erence area, ft2 WXO distance from fu s elage apex to wing apex at wing centerline, f t 3 9 ' _ _ WRC ro o t ch o rd at wing c en t er li ne, ft '.j •I WTC tip chord, ft i WSPAH wing span, ft _ ! WDXT longitudin a l distance from wing apex to leading e d ge of tip chord, ft _I (positive rearward) YWCREX spanwise distance from wing centerline to exposed root cho r d, ft NWBP nul_Iber o f breakpoints in wing l eading and trai l ing edge s , an integer.

Wing apex and t i p are blotinc l uded. If l eading edge and trai l ing edge break at the same spanwise s tation, count as one breakpoint.

WDXBP(12) a one-dimensiona l array of t h e l ongitudina l distances from wing apex to l eading edge at breakpoints, ft (positive re a rward_ WYBP( I 2) a o n e -dinension a l array o f t h e s panwi s e di st ances to the br_akpoints f t ' - WBPC(12) a one-dimensional array of wing c h o r ds at the breakpoints ft _.

WTOC average maximum wing thickness-chord ratio in percent Fuselage FLGTH l engt h , ft FDPTH H X f n aximumdepth, ft FWDTHMX maximum width, ft FRAD maximum radius, ft ( Note: If FRA , . ) is input, Ft)PTHI_X and FWDTHMX inputs are not required) FWETA total fusel a ge surface area, ft2 FV (J LT U T t o t al fuse l age v o lume, ft3 Ho r izonta l Su r faces T h e di_:,s i o r _s t or hori z ontal surface( . a re input in the t olloiwng one- dime n siona l arrays.

NH the number of horizontal surfaces, (Z maximu=_1) HDXT(2) the longitudin a l distance from the apex of the surface to its leading edge at the tip, ,t TI H RC(2) root chord, ft HTC(2) tip ch o rd, ft

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: I I HSPAN(2) span, ft HTOC(2) average thickness-to-chord ratio, percent HXO(2} longitudinal distance from fuselage apex to surface apex, ft Vertical Surfaces The di l aensionsfor the vertical surfaces are inp,Jt in the following one- dimensional ar r ays. A spanwise mounted surface is assumed to be located symmetrically a bout the fuselage centerline and is counted as one surface.

NV the number of vertical surfaces, (4 maximum) o VDXT(4) the longitudinal distance from the apex of the surface to its leading edge at the tip, ft VRC(4) root chord, ft VTC(4) tip chord, ft VSPAN(4) span, ft VTOC(4) average thickness-to-chord ratio, percent VXO(4) longitudinal distance from fuselage apex to surface apex, ft VYO(4) spanwise distance from fuselage centerline to surface, ft J NacelIes XIJP(9) a one-dimensional array containing the longitudinal distances from the fuselage apex to the nacelle, ft _ YOP(9) a one-dimensional array containing the spanwise distances from the fuselage centerline to the nacelle centerline, ft '_ NPI the number of nacelles in the above array ; Spanwise mounted nacelles are counted in tile sa , ile manner as the vertical s urfaces.

$AIN The baseli,leaerodynamic input is located in this segment. Trimmed angle of attack, lift, and drag data are preferred.

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Takeo ff Aer o dynami c Input With Ground E f fect i

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j T ALP TO G ( 15,4) a t wo-limen_i o nal array of angle of att.ack for various flap ' "i deflec_i o ns, deg 1 "I : _ TrLTOG(15,_) a similar array for the l i ft coLm q =ficients CL 1 I ' i ! TCDTUG(15,4) a similar array for the drag coefficients,CD i

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] NTOG the number of v , _luesin each of the above arrays at each flap , deflection, (15 maximum) (If NTL)G = O, an internal r outine is used to calculate values in the i 1 i ab ov e arrays w h e n the follo wing takeoff input is provided.) , i Takeoff Aerodynamic Input Without Ground Effect TALPHT(J (15,4) ang l e-of-attackarray TCLTO(15,4) l ift-coefficient array TCDT(J(15,4) drag-coef f icientarray I NT U P the number of values in the above arrays at each flap deflection, i I (15 ma x imum ) , ( U ) i NFD the number of fl ap def l ections, (4 maximum) ( U ) I i ' TFSET(4) a one-dimensionalarray containing the flap deflections, deg CDGRT(15) a one-dimensiona l array o f increments in drag coef f icient of the l anding gear as a function of li ft coefficient CLGRT(I5) the corresponding array of l ift coef f icient s NLGU the number points in the above arrays, (0) DC I JLG increment in drag coefficient of landing gear (independentof li ft coe ff icient, use i f IILGD = U r, l ainMission Ae r odynamic Input CLT(15,15) a two-dimensiona l array of l ift coe f ficients for various Mach n u mbers E ALPHAT(15,15) the r-rresponding array for angles of attack, deg CDPT(15,15 ) the corresponding array for the drag coefficients NAI the number of values in each of the above arrays at each M_ch number (the va l ue of the first dimension; NAI points must bL,input at each 14achnumber ) IPBOL = 0 aerodynamic input in CLT, ALPHAT, and CDPT will be used for i aerodynamic data (the default value) = 1 the parabolic polar inputs (CDOR, CLAR, DLFR, and CLOR de f ined be low will be used for aerodynamic data CDOR(15) a one-dimensional array containing the minimum drag coefficients at various Mach numbers CLAR(15) a similar array for the lift-curve slope, per deg " DLFR(15) the array for drag-due-to-lift facto r s CLf}R(15) the array for lift coefficients at minimum drag coefficient NAJ the number of Mach numbers for the lift and drag data (15 maximum) MAEROT(15) a one-dimensional a,ray containing the Mach numbers for the lift and drag data THARg(15) a one-dimensional _rray containing the al t . itudes at each Mach number at which the skin friction contributions to drag were calculated DELCD(15) a one-dimensional array containing increments in zero-lift drag coefficient for the Mach numbers in MAEROT CDW(Iq) a one-dimensional array containing the wave-drag coefficients at each Mach number in MAEROT CDRUF(15) a similar array of roughness drag coefficients J CDAC(15) a simila r array of air-conditioningdrag coefficients I .

CDBL(15) a similar array of engine-bleed drag coefficients I Note: Input for CDAC and CDBL is required only if values are included in CDPT !_ or CDOR.

SWTIN This segment contains the weights and all other input for the baseline aircraft but not the engine data. Except where noted default values are 0 and input weights are in pounds.

$ Structural Weight Input WWING wlng WHT horizontal tall i WVT vertical tails (sum of all) WCAN canard WFUS fuselage WLG landing gear Weight Input for Cryogenic Tanks • (Hydrogen on Methane Fuel) WFT fuel tank, total WINS insulation, total WFTINS combined weight of fuel tank and insulation (required only if WFT and WINS cannot be separated) ,!

TSA tota l surface area of fue l tank, ft2 i System and Equipment Weight Input WSC surface controls WAP auxiliary power W I NST instrw,_ e nts J WHYD hydrau I ics g ELE electrical WAVON avi o nics -_ WFEQ furnishings and equipmen t ___. .

WACOND air conditioning WANTICE anti-icing Operating Weight Inputs WFCR weight of f l ight crew IFCR number of flight crew _embers, integer WCCR weight of cabin crew ICCR number of cabin crew members, integer W U FUEL unusab l e f u e l 44 _ " I WENO engine oi1 WPSER passenger service WCCONT cargo containets WOPIN sum of a| ] operating weights (use only if a ll individual weights cannot be supp l ied) • Payload Weight Input WPAS weight of passengers " I PA_R number of passengers, integer WPB passenger baggage WCAR weight of c a rgo Mi sc ellan e ou s Input WGREF gross _eig i ltof base l ine aircraft ULF design load factor, (3.5) ZMDESBL desion Hach number f or baseline aircraft STSA6R numbe r of se a ts abreast LVLR number of passenger levels, (2 maximum) XPASC l ength ot passenger compartment, ft VCARMX maximum volume for cargo, ft3 ' BFCILBS maximum fue l capacity in fuselage, Ib WFCILBS m aximum fue l capacity in wing, Ib BFCIGAL f ue l capacity in fuseIa£e, ga l WFCIE_AL fuel capacity in wing, gal (Note: input for rFCIGAL and WFCIGAL is not required if BFCILBS and WFCILBS are used) SENIN Data fo r t h e engine and nacelle are input in this s egment.

NENR the number of engines, integer NEW the number of engines mounted on the wing, integer NEF the number of engines mounted on the f uselage, integer NPODSR the number of engine nacelles, integer PDAVG the average diameter of the nace!le, ft } PLGTH the length of the nacelle, ft i l ESE X P na c e ll e length s izing e x p o nent, ( .438 ) EWEXP engine weight sizing exponent, (I.085) SLST H R sea leve l static thrust f or the engine having the fo l lowing weight, I b WENG weight of one engine, lh WTR weight of one thrust reverser, l b WNAC weight of one engine nace ll e, Ib ,j WMIS total miscellaneous propu l sion system weight, I b i i i WPLM total propulsion plumbing system weight, ]b i The following t h ree variables are provided for engine bleed and air-c o ndition drag. These variab l es should be used only if the drag it-.msare not inc l uded in gross t h rust values or by entries in CDAC and CDBL in SAIN.

DOQINT(15) a one-dinmnsional array containing the sum of engine bleed and air- condition drags divided by dynamic pressure for a range of Mach numbers, f t_ _' DL}QMT(15) the corresponding array of Mach numbers NBDP the number of Mach nu_bers in the above arrays • I 1 Note: These variables c ann ot b e used f or engines with two-dimensional inlets.

J (nput for Propeller-EngineCombinations

t

Propeller-drivenaircraft require no additional input if: (I) propeller output i data is n o t re q uired, (2) the weight of the propeller and gear box are included in the engine weight, WENG, and (3) the design inputs TL)WUES,TPEDES, TOWTB, or TPETB, rather than PUWDES, SHPDES, or POWTB are used to s i ze the propulsion system. This analysis treats the a i rcraft as a jet-driven aircraft.

4 6 • -_- , -_ _ . _ .

Ir j. Th e prop e ller will be treated in more d e tail, however, if T APE3, described previously, is pro v ided and th e following data is input.

AUVRDES design a d vanced rat i o ", . _ CPDES design power coefficient _: C T OES d e s ign t h r u st c oef f i c ient I EPRDES prope l ler e ff iciency at design conditioning - EOVUES overall e ff iciency o f t he pr o pel l er-engine combination at design conditions P O L120ES shaft_horsepower-propellerdiameter squared ratio at design conditions, ' hp / ft z PTIPS propeller tips speea at design conditons, fps DIAMPBL propeller diameter, ft (-10.)

SI4PREF shaft horsepower of the engine at sea level static conditions and having - _ its weight entered in WENG RPMEBL engine RPM RPr_PBL propeller RPFI PSEXP propeller weight sizing exponent (Z.488) WTGB weight of one uearbox, lb (-10.)

WTPR weight of one propeller, Ib (-10.)

The propeller weight, WIPR, is used to test if a jet analysis the (default case) or a propeller analysis is . .tobe made. When WTPR is input, all of the above propeller inputs and the P I L} X" values, described below, a r e required. I Input for Engine with Two-Dimensional Inlets INL2b --I engine nacelles are two-dimensional = 0 engine nacelles are circular, (default) £ :. PWIL)2D width of the nacelle, ft P H GT21) height of the nacelle, ft PLSTH length of the nacelle, ft PSWETZD wetted area of one nacelle, ftz I I BASELINE ENGINE CHARACTERISTICS This final input segment contains four groups of data for the installed engine. These include the identification of the baseline engine, full-power data as a function of altitude and _lachnumber, part-power data, and for engines with two- dimensional inlets, full-power data as a function of Mach number, angle of attack, and altitude. The engine identification, the first input of this segment, is one l i n e o f free - fie l d inp u t l imited t o 80 c haracter s .

T he ne x t i n p u t pr o vi d e s the full-power characteri s tics of the engine as a function of Mach number and altitude. The d ata format i s as f o llow s : Colu mn T_p_ of Datd , 1-5 Mach nu mber 6- 1 5 a l titude , ft 21-30 gro s s thru s t, I b 31-40 ram drag, Ib 41-50 ,uel f lo w, I b / h 51-60 P / D2, hp / ft2 Where P / D2 is the s haft horsep o wer-propel l er diameter s qu_red ratio. P / D2 data i s required on l y if the prope l ler data described above is input.

The data is arranged by increa s ing a l titude at c o n s tant Mach number, then by increa s ing Math number. Up to 15 separate values of altitude and 15 s eparate values o f Mach num b er are al lo wed. On the la s t line of this data segment, the characters 9. located in columns 2 and 3 are required.

The ne x t g r o up of data pertains o nly to engines with two- d imensional inlets and i s required only if the effect s of angle of attack are to be included in the full- p o wer data. To indicate the pre s ence of thi s data group, which is restricted to s u per so nic and higher Mach number s , the fir s t line of inp u t mu s t contain the three characters -9. in columns 2, 3, and 4. O n s ucceeding lines thereafter, the data i s input in the following format.

Co l u mn Ty pe o t ;_at d 1-5 Mach number 6-15 altitude, f t ' 16-20 angle of attack, deg 2 1- 3 n gr o ss thr us t, Ib 31-40 ram d rag, Ib 4 1-50 fuel f l ow, Ib / h The d ata is arranged by increa s ing a l tit u de at c o n s tant angle of attack and c o n s tant Mach n u m b er, then by increasing a n g l e of attack at c o nstant Mac h n u mber, then by i n crea s ing Mach number. Input i s restricted t o 15 va lu es o f Mach number, 3 va lu e s o f angle o f attack at each Mach num b er, and 3 value s o f a l titude at each t h e same number of angles of attack at each Mach _umber and the same number of ii_ 1 angle of attack--Mach number combination. Sufficient data must be input to provide , a l titudes at ea ch ang l e of attack--Mach number combination. If this requirement is 1 not met, the progra n lwi l l abort. To end this data group, t h e two characters 9. must _. be l ocated in columns 2 and 3 on the last line of input.

- The last input provides the p a rt-power characteristics of the engine . T h e format for this data is: Column T_pe of Data : _j I-5 I1achnumber WJ 21-30 gross t h rust, Ib

t

31-40 r am drag, Ib " 41-50 fue l f l ow, ]b / h Data is arranged by increasing thrust at constant Mach number, then by increasing Mach number. The l ast input of this group must contain t he two characters 9.

l ocat_d in columns Z and 3.

SAMPLE INPUT LISTING An input l isting tilati ll ustrates t h e content of the previous section is given fuel ) shown in figure 22 and the resu l ting output is discussed in Appendix C.

While the sizing "thumbprint" (fig. 22) provides fairly accurate va l ues of W / S and T / W for t l _edesign point, the gross weight value is very approximate. To more I. _ il in Tab l e B1. This input is f or th e design point of the supersonic transport (JP accurately determine this value, the computer must be directed to find the gross I weight that provi d es L he design range at the design point conditions. This is done I_ in input segment SDES(_NV_by inputs f or the required range (RNGDES = 4000.), wing l oa d ing (WOSLiES = 8 2 .), rindthrust-weight ratio (TOWDES = . 3 2). These and the other inpuLs for the sized airc r a f t are shown in Table BI.

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ENDIX C o UTP U T The o utput from the pr o gram i s c o ntained on three mass st o rage files. U nle s s c ' _angedby the user at execution time, these files are labeled TAPE6, TAPE11, and _APE14. TAPE 6 contains a listing of all the results provided by the program, ; T APE11 co ntains s ele c ted data in a format suitable for irput to a plotting program, and TAPE14 contains the basis for preparing sizing "thumbprints°" T APE6 On this file, several types of output are available. These include the preliminary output, sized aircraft output, a short l isting of selected output quantities, weight statement, and a mission profile output.

Preliminar_ output.- This printout , which is optional and obtained with IPPRNI"=I in S P RNTCON, contains a listing of the input values for all input segments. Because of its length, an example of this output is not given here.

SILeu Aircraft uutput.- This printout, which cannot be suppressed, contains the results for each sizing variation requested. A listing of the output that results from the input for the design-point aircraft given in Table BI is given in Table CI. A short summary of this listing is also output and this is shown in Table C2. A li s ting of the weight statement, if selected by IPRWTSM=I, is given in Table C3. With IPRMPRFn_I, mission profile results are output and these are listed in Table C 4 .

TAPE11 This file, produced for every compute r cycle, contains the same parameters as the short output list (Table C2) located on TAPE6. The file is formatted to allow the sizing and performance results to be plotted during interactive plotting se s sions where the abscissa and ordinate of the plots are identified by the indices of the chosen variables. The indices are the numbers of the variables on the short output list in Table C2. An i11_stration of the listing this file contains is shown in Ta b le C5 and the different sections of data correspond to different computer " , cycle s in whi c h th e si z ing pa r ameters are varied. The first section contains the same re s ults as the short ou c put list (lable C2) for the design point aircraft; the ; remaining s e c ti o n s c ontain the results for a systematic variation in thrust-to- " weight ratio at constant wing loading. The integer at the beginning of each section has a constant value for all sections of data that belong to the same curve (in this ca s e, a con s tant value of wing loading).

T APE1 4 Thi s file contains the basis f o r preparing sizing "thumbprint s "like that s h o wn in figure 22. Unlike TAPE6 and TAPE11, thls file is pr o duced o nly when a matrix o f wing and engine si z e variati o ns are t o be c o mputed. To pr o duce the "thumbprint" f or the s upersonic transport, (fig. 22), inputs were required f o r the d esign range (RNGDES : 4000.), an array of values for wing loading (WOSTB = 11( } .,I00., 90., 80., 70., 60., 50.), an array o_ t_rust-to-weight ratios (TOWTB = .25, .30, .35, .40, .45, .50, .55), along with NWOS _ 7 and NTOW = 7. The values that resulted from one ; computer run are shown in Tab|e C6. The listing contains the abscissa values of wing loading, the ordinate values of thrust-to-weight ratio, and the values of the parameters to be contour_d that are functions of wing and engine size. The rows provide the change in dat_ ;_ithW / S at constant T / W, whereas the columns give the data change with T / W at constant W / S. The large negative values (-1000000) are default values used where no s , )lutions were obtained.

Tab|e C6 identifies all the parameters that are available for contouring. To _W obtain the coordinates nY the actual contours, the data on this file can be either cross-plotted by hand o r TAPE14 can be used to prepare an input file for a contouring program.

P i _ ..... -- . L -._m_ _ _,-_ _m_ _ ' '_ I

i J APPENDIX D n I ,i JOl_ CONTROL CARDS l The program executes on the Network Operating System (NOS) Version 1.4 currently _i in use at the Langley Research Center. The control cards required for execution • are: ] GET, ASP / UN = 273347N I ASP (F5, F6, FII, F14) where F5 is the file containing the baseline aircraft input and sizing options (TAPE5) Fb i s the ou t put f ile (TAPEO) F I I is the file c ontaining plot data (TAPE11) F14 is the file contailiqg the data for contour plots (TAPE14) ,.| Note: Although the file containing the propeller design data (TAPE3) does not appear in the exeoation card ASP, it must be a local file du,ing execution.

The storage requirement for program execution is about 21U,UO0 (octal) _ords.

i ° ' , 'I o • .-, 8_"i H (Jbm_,_ = . _e,. ..

I REFERENCES 1. Fetterman, David E., Jr.: Preliminary Sizing and Performance Evaluation of Supersonic Cruise Aircraft. NASA TM X-73936, 1976.

2. Staff of the Langley Research Center: Noise and Performance Calibration Study of a Mach 2.2 Supersonic Cruise Aircraft. NASA TM-80043, 1979.

3. Staff of the Langley Research Center: Jet Noise a n d Performance Comparison Study of a Mach 2.55 Supersonic C r uise Aircraft. NASA ° TM-80094, 1979.

4. Sears, W. R.: On Projectiles of Minimum Drag. Quart. Appl. Math, • vol. 4, no. 4, 1947, pp. 361-366.

5. Morrison, W. D., Jr.: Advanced Airfoil Design Empirically Based Transonic Aircraft-Drag Buildup Technique. NASA CR-137928, 1976.

6. Ashley, Holt; and Tandahl, Marten: Aerodynamics of Wings and Bodies.

Addison-Wesley Publishing Company, Inc., 1965, pp. 1 7 8-181.

7. Harris, Roy V., J r .: An Analysis and Correlation of Aircraft Wave Drag. NASA TM X-947, 1964.

8. Rubesln, ti.W.; and Johnson, H. A.: A Critical Review of Skin Friction and Heat Transfer Solutions of the Laminar Boundary Layer of a Flat Plate. Trans. ASME, vol. 71, no. a, May 1 9 49, pp. 383-388.

9. Monoghan, R. J.: On the Behavior of Boundary Layers at Supersonic Speeds. Fifth InternationalAeronautical Conference, Los Angeles, California, June 20-23, 1955, Institute Aeronautical Sciences, Inc., pp. 277-315.

10. Schlichting, H.: Boundary Layer Theory. Pergamon Press, 1955. _ 11. F o ss, Willard E., Jr.: A C o mp u ter Program for Detailed Analysis of the J_ T ake o ff and Approa c h Performance Capabilities of Transport Category _ Aircraft. NASA TM-80120, 1979.

12. Jackson, Charlie M., Jr.: Estimation of Flight Performance With Closed- Form Approximations to the Equations of Motion. NASA TR R-228, 1966.

13. Brewer, G. D.; and Morris, R. E.: Advanced Supersonic Technology Concept Study - Hydrogen Fueled Configuration, Summary Report. NASA CR-114 7 17, 1974.

14. Johnson, Vicki S.: Compar i son of Advanced Turboprop and Turbofan Airplanes. NASA TM-85692, 1983.

15. FORTRAN Extended Version 4 Reference Manual. Control Data Corp., Ig80.

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{ 0 0 _ t I 0 0 gO0 OOQ 00 QO_ 000 " _ * _Q_ - 0_ - G_ . 00_ " 00_ • _ .... _ " _ _ _ __ " _ _ - - _ o? .... 9 _ e • • - II II II II I' II _ . -- ..... _ ...... _ ...... _ ..... _ ...... _ ......

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o . oo ° . .............. ° ...... . ° ......... o °° ....... _;,; d g;_ d2 _g g _ dgggg ; ....... d2 £ d £ dd .... .. ° i ° * *.o.. , o .o°.o .. ..°. °...o ° ° . o _ . oo o I 4 1111111 i | I 113 I NASA .TM- _6 3 57 "4. T,tl_ a n d SuM, t i e § . RelIx) r t Oete PRELIMINARY SIZING AND PERFORMANCE OF AIRCRAFT July 1985 6. P _f / Or mL r _ I Orpn,z_tlml 505-43-43-01 I A ut h o r(s) 8 . Pe r fo r min g Org _ rt;llt l o n R ep o r t No.

David E. Fetterman, Jr .

10. Wor k Un i t No.

9. P e rfo* ' mmg O rgan, z at = on Nl m e _ mdAddrm I NASA L angley Research Center 'It c_t , __ G _l N o Hampton, VA 23665 !

13 . T y p m of Re p o r t m (I Pl e t_ l Co _ I "_ s _ .w ,_ _ v ,_ - ,.._ Ad o , ,, , Technical Memorandum National Aeronautics an d Space Administration 14 so_._,_ A w .-vc o d, Washington, DC 205 4 6 I S Supt _ e m entary No t cm

i

16 A bstract t The basic pr o c ess e s o f a program that performs sizing operations on a ba s eline aircraft and determines their subsequent effects on aerodynamics, propu ls ion, weights, and mission performance are described. Input req u iremen L s are defined an d o utput listings explained . Results obtained by applying the method to severa | type s of aircraft are discusse d .

i ; i r ; 17 . K e y Wo r ds ( _ ggmt K I by A u thor(s)) 18. O,str,but l on S tltlmem ;

t De s i gn

Mission analysis Unclassified - U n limited 1 Hydroge, Meth_,_e Jet Subject Category 05 P , ' c_e ', ", er Un c l a ss ifle d Unc l a s sifie_ 114 A06 , . J o _ For s a l e bythe N a t i onal Tec _ ¢,llInf 0_ l b on S erv,ce. S p mlfield. V =f i _ ,n _ Z 2 1 61 - n