EGADS: A microcomputer program for estimating the aerodynamic performance of general aviation aircraft

ear relationship between ideal propeller efficiency and

Appendix A

velocity, equations (9) and (i I) must be solved in order to investigate the manner in which the available power and

Propulsion Computations

thrust vary with velocity and altitude (ref. 4). Above the critical altitude, engine horsepower decreases according to Equations (7)-(14) (appendix C) are used to estimate pro- equation (7).

peller efficiency, and engine thrust and power available throughout various flight regimes. Because of the nonlin-

PIIICqDoNG PA;,E ._ Ar4g Nql' Ft'. I_,_:"

II

rpA _ _ tNTEN'I'ION,A, LLY BLA_K

is first used to compute approximate values of wing and Appendix B horizontal tail lift curve slopes, then equation (31) is used to estimate the change in downwash angle at the tail with Aerodynamic Center Computations changes in angle of attack. Accurate estimation and input of _CMo /¢3CLw and the tail dynamic pressure ratio is A precise prediction of the location of the aerodynamic left up to_m_euser. Equation (33) is then used to estimate center requires the estimation and input of some aerody- the aerodynamic center of the aircraft.

namic parameters that are currently beyond the calculation capabilities of EGADS. Equation (30) (appendix C) I_IICEMNG PAGE BLANK NOT FU.MED PAGE _ INI'ENTIONALL_ BLAN_

Appendix C 9. Pshaft = 2pAV3 (I - rli) hp Pshafto qi 3 r Actuator disk theory assumptions (ref. 4) Partial List of Equations Used in EGADS cos */ |. n--_ IO. Pay = nni Pshart cos _) Available power reduced by non-ideal and Constant ram turns and climbs actuator diskpropeller efficiencies . q = lpv2 II. Ti : 2pAV2_ Definition of dynamic pressure Actuator disk theory assumptions (ref. 4) .

L = qSC L = nW = W cosy 12. T=nT i cos_ Available thrust reduced by non-ideal propeller Constant rate turns and climbs efficiency CL 2 D2/3 . 13. Tstatic =(2pAll/3 "shaft CD = CDo + _ARe Actuator disk theory assumptions Low speed, AR > 2, no flow separation, uncambered aircraft, constant rate turns and climbs 14. Ti = Tstatic(I - qi )|/3 Actuator disk theory assumptions .

D=qSC D = T- Wsin3f 15. V = Mona Thrust opposite drag, steady flight Definition of free-stream Mach number L 2 + (T- D) 2 = W 2 i6. S=S w Thrust opposite drag, steady flight (In EGADS, the reference area is the wing planform area) 17. Swe t =2 S/I+_-_- ) % "w / !

normally aspirated engine Wetted-area estimate for wing and .all surfaces (ref. I ) _ P h > hcritical, supercharged P= Pm_ h < hcritical, supercharged Linearly tapered, simple wing Po Approximate variation of reciprocating engine (I +).+ _. 2 ) power with altitude (tel'. 4) CMA C = 2Cmi d 3(I + _.)

Linearly tapered, simple wing Preq =DV Required power, steady level flight GTVhz =(xac_ - xac.) Shz CMACS Definition of geometric tail volume for horizontal stabilizer PAGE / " INTENTIONALLY BL_iK 30.

21.

_Cw/ bS Definition of geometric udl volume for vertical stabilizer 22.

f = CfSwe t Equivalent tim-plate area 31. _-_-=l.7 .AJRw(3L,r)ll4(l+ln_ ) Zf 23.

CDo = r=2 thz m=2 Zhz -zw Parasite drag coefficient b b V2 cosy _¢hz = XhZ -4"Xlghz - X w - Xac w 24. Turn radius = gumO Approximate downwash gradient (ref. I) Constant rate turns and climbs 32.

hn. = Xw +Xacv Vcosy 25. Turn rate = CMAC Turn radius Wing neutral point Constant rate turns and climbs BSFC 26. Fuel flow = 33.

hn = hnw +

P,h.e

I+F(CuAc/t )

Reciprocating engine 27.

Range = TTni C L tn / w / CLo" I_ BSFC C D _, w - Wfuel Approximate aircraft neumd point as a fraction Cruise without winds at constant altitude, of MAC (ref. 5) BSFC, and C L 34.

Xac(A/C _ = hnCMA C C_ M2 28.

Endurance = q11i (PS) I/2 Longitudinal location of neutral point BSFC C D hffiXcG 35.

CMAC [ ' ,

x ).2

(w - wfu¢l CG location as a fraction of MAC Cruise without winds at constant altitude, BSFC, and CL 36.

SM=hn-h Static margin as a fraction of MAC 29.

Biot-Savart estimate for horseshoe vorle4 fief. I ) (actuator disk theory) propeller efficiency qi- EGADS

AppendixD

attempts to determine the exact value of 11 i whenever it is required. The total available power (Pay) is then deter-

Assumptionsbehindthe Actuator Disk mined by equation (10) (appendix C). in which the total

propeller efficiency is assumed to be the product of the

Theory Used in EGADS

non-ideal and ideal propeller efficiencies. Because of the A complete discussion of the developmen', and assump- dependence of Tli (and Pay) on the velocity, the best rate tions um_lying actuato¢ disk theory can be found in speeds and ceilings are computed within EGADS using many introductory texts on aircraft aerodynamics and simple ite_tive techniques. The takeoff and landing roll flight mechanics. The assumptions underlying the theory distance computations use a simple explicit Euler-Cauchy are given by McCormick as the following: (I) velocity is time-integration technique for approximating the thrust constant over the propeller disk; (2) pressure is uniform and airspeed as a function of time during the ground roll, over the propeller disk; (3) rotation imparted to the flow and therefore require the computation of the variation in as it passes through the propeller is neglected; (4) flow Tli (and Tar ) with airspeed throughout the duration of the passing through the propeller can be separated from the ground rolls.

rest of the flow by a well-defined stream tube; and (5) the Some of the limitations of actuator disk theory can be flow is incompressible (ref. 6, p. 343). An additional seen in figure 20(b). As the airspeed increases, the actual assumption is that the pressure in the slipstream returns to propeller efficiency (and the available power) should the free-stream value far downstream of the pmpeUer reach a maximum and then decrease as a result of two disk.

effects: the angle of attack seen by the blades (and hence These assumptions lead to inaccuracies when using actua- the thrust) is continually reduced as the forward speed of tor disk theory to predict actual propeller efficiency. In the aircraft increases, and as the blades reach transonic order to make actuator di_k theory more realistic for use speeds, the formation of shock waves on the blades will in performance estimations, the efficiency factor I] (input further increase the blade drag while reducing the thrust.

by the user. and always less than I) is use:l to represent More advanced propeller estimation techniques based on additional non-ideal propeller losses neglected by the the- blade-element or vortex theory can compute improved ory. These include hub and tip losses as well as blade estimates of propeller efficiency, but they lack the sim- profile drag losses. From equations (9) and ( I 1) plicity of the actuator disk theory.

(appendix C), it is clear that a cubic or quadratic equation must be solved to determine the value of the ideal

along with the total wetted area, total equivalent fiat-plate Appendix E area, overall skin-friction coefficient, and aircraft zero-lift drag coefficient CDo (eq. (23)). Note that the wetted-area Drag Buildup Computations computations for the landing gear neglect the wetted areas of the struts and assume that there are two main gear The wetted area of the fuselage is computed using numer- wheels and one nose/tail wheel. The drag increment due ical integration. The wetted areas of the wing and tail s_w- to the landing gear is computed by multiplying the tire faces are estimated using equation (17) (appendix C) upon frontal areas by the value of CDg. Although the aircraft entry to the Drag Buildup worksheet. When the Recal- culate the Drag Buildup button or the keyboard carriage CDo and the Oswald efficiency factor appear in many of the performance worksheets, they can be permanently return key is pressed, the equivalent flat-plate areas for the changed only on this worksheet.

individual components are computed (eq. (22), appendix C) and displayed in the right-most column, PAGE., ; INTENTIONALLY BLANK

5. Use equation (! 4) to determine M.

Appendix F 6. Use equation (2) to determine q Climb and Turn Computations ?. Use equation (5) to determine T The equation numbers cited below refer to those in 8. Use equation (12) to determine Ti appendix C.

9. Use equation (1 i) to determine Tli Calculations based on Mach number:.

10. Use equations (7) and (9) to determine hPr and Pshaft 1. Use atmosphere calculations to determine a and p I 1. Use equation (24) to determine turn radius 2. Use equation (I) to determine n 12. Use equation (25) to determine turn rate 3. Use equation (15) to determine V Calculations based on airspeed: 4. Use equation (2) to determine q i. Use atmosphere calculations to determine a and p 5. Use equation (3) to determine CL 2. Use equation (I) to determine n 6. Use equation (4) to determine CD 3. Use equation (15) to determine M.

7. Use equation (5) to determine T 4. Use equation (2) to determine q 8. Use equation (I 2) to determine T i 5. Use equation (3) to determine C L 9. Use equation (11) to determine rli 6. Use equation (4) to determine C D 10. Use equations (7) and (9) to determine hPr and Pshaft 7. Use equation (5) to determine T 12. Use equation (24) to determine turn radius 8. Use equation (I 2) to determine Ti 13. Use equation (25) to determine turn rate 9. Use equation (! !) to determine Qi Calculations based on CL: 10.

Use equations (7) and (9) to determine hPr and Pshaft 1. Use atmosphere calculations to determine a and p il.

Use equation (24) to determine turn radius 2. Use equation (!) to determine n 12.

Use equation (25) to determine turn rate 3. Use equation (4) m determine C D 4. Use equation (3) to determine V PAGF. _ iI,,,ITF.,,NTl()l_sl,,.i,,, i ... ,,,,_

4. U_¢ equation (4) to determine C D Appendix G 5. Use equation (5) to determine T Range Computations 6. Use equation (! 2) to determine T i The equation numbers cited below refer to those in 7. Use equation (11) to determine vii appendix C.

8. Use equations (7) and (9) to determine hPr and Pshaft 1. Use atmosphere calculations to determine a and p 9. Use equation (26) to determine fuel flow 2. Use equation (3) to determine q when half of trip fuel is consumed i0. Use equation (27) to determine range 3. Use equation (2) to determine V ! !. Use equation (28) to determine endurance IllqUiClDIIde PAGE BLANK NOT FILMED PAGE _ INTENTIONALL_I ' {J_'_;tK

Calculations for Vy (best rate of climb speed): Appendix H 1. Use atmosphere calculations to determine a and p Speed Computations 2. Since vii is a function of V, V_ is found by using a bisection search assuming Pshaft = hPr Pshafto..

The equation numbem cited below refer to those in appendix C. Calculations for Vx (best angle of climb speed): Calculations for Vmin: 1. Use atmosphere calculations to determine a and p I. Use atmosphere calculations to determine a and p 2. Since vii is a function of V, V x is found by using a bisection search resuming Pshsfl = hPrPslmflo., 2. Use equation (3) to determine q at CLmax Calculations for best power-off glide speed assume 3. Use equation (2) to determine V C L = (x • AR CDo)I/2 for maximum glide ratio.

4. Use equation (15) to determine Moo Calculations for best power-off glide endurance use a 5. Use equation (4) to determine C D bisection search to determine the C L for maximum endurance.

6. Use equation (5) to determine T 7. Use equation (12) to determine Ti 8. Use equation (II) to determine vii 9. Use equations (7) and (9) to determine hPr and Pshafl Calculations for Vrnax: I. Use atmosphere calculations to determine a and p 2. Since vli is a function of V, Vmax is found by using a bisection search assuming Pshafl = hPrPshaflo=, Pltli4Gm_lG PAGE BI_A!'tK NOT FILMED PAGE .-_--.-.-- INTENTION,",LL'_ .....

absolute ceiling calculation also uses two bisection Appendix I searches: an outer search is performed for the altitude at which Vy = 0 fi/min, and an inner bisection search CelUngs Computations determines the maximum Vy at the trial altitude. In both cases, equations (7) and (9) (appendix C) are used to The service ceiling calculation uses two bisection determine the shaft power available as a function of alti- searches: an outer search is performed for the altitude at tude, and equation (I I) is used to determine the variation which Vy = i00 ft/min, and an inner bisection search in ideal thrust with airspeed.

determines the maximum Vy at the trial altitude. The BI,ANK NOT FILMI[O INTENTIOPiALLY [_L_

Appendix J imity on the induced drag is calculated by usi,_, equa- tion (29). The integration time-step is 0.02 sec. For the takeoff roll, the integration is halted upon reaching the Takeoff and Landing Roll Computations rotation airspeed--no estimations of rotation and first- segment climb distances are made. The value of the thrust The equations of motion for the aircraft during the takeoff at the beginning of the landing roll is computed using the and landing rolls are integrated forward in time using a approach airslx_ed and descent angle, then the thrust is simple explicit Euler-Cauchy technique. The ideal static linearly decreased to zero during the Time to Zero Power.

thrust is calculated using equation (13), and then equa- The landing-roll integration is halted as soon as the air- tion (9) (appendix C) is solved to determine qi as a func- craft comes to a complete stop. No estimations of the flare don of airspeed. Equations (12) and (14) are then used to and landing rotation distances are made.

compute the actual thrust. The effect of the ground's prox- Nicolai, Leland M.: Fundamentals of Aircraft Design.

References METS, Inc., San Jose, Calif,, 1984.

1. Torenbeek, E.: Synthesis of Subsonic Airplane Perkins, Counland D., and Hage, Rohen E.: Airplane Design. Delft University Press and Martinus Performance, Stability, and Control. John Wiley Nijhoff Publishers, Delft, Holland, 1982.

& Sons, New York, N.Y., 1949.

2. U.S. Standard Atmosphere. U.S. Government Raymer, Daniel P.: Aircraft Design" A Conceptual Printing Office, Washington, D.C., 1962.

Approach. American Institute of Aeronautics and 3. Hoak, D.; Finck, R.; et al.: USAF Stability and Astronautics, Inc., Washington, D.C., 1989.

Control DATCOM. Air Force Flight Dynamics Roncz, John G.: Designing Your Homebuilt. Sport Laboratory, Wright-Patterson AFB, Ohio, 1978.

Aviation, Feb. 1990, pp. 37--40.

4. Crawford, D. R.: A Practical Guide to Airplane Roncz, John G.: Sizing Your Wings. Sport Aviation, Performance and Design. Crawford Aviation, March 1990, pp. 34-39.

1981.

Roncz, John G.: Wing Incidence & Tail Size. Sport 5. Sivier, K.: AAE 319 Course Notes. University of Aviation, April 1990, pp. 23-28.

Illinois, Urbana, II!., 1984.

Roncz, John G.: Forward Sweep & The Great Tire Crisis.

6. McCormick, B. W.: Aerodynamics, Aeronautics, and Sport Aviation, May 1990, pp. 43--47.

Flight Mechanics. John Wiley & Sons, New York, 1979.

Roncz John G." Questions and Answers. Sport Aviation, June 1990, pp. 41-45.

Bibliography Roncz John G." Tail Incidence, Part I. Sport Aviation.

Aug. 1990, pp. 36--40.

Abbot, Ira H.; and yon Doenhoff, Albert E.: Theory of Roncz, John G.: Tail Incidence, Part 2. Sport Aviation, Wing Sections. Dover Publications, Inc., New York, N.Y., 1959. Sept. 1990, pp. 35-40.

Roncz John G.: Tail Incidence, Part 3. Sport Aviation, Anderson, John D., Jr.: Introduction to Flight. McGraw- Hill, Inc., New York, N.Y., 1989. Oct. 1990, pp. 45-49.

Roncz, J_Jln G.: Tail Incidence, Part 4. Sport Aviation, Dommasch, Daniel O.; Sherby, Sydney S.; and Connolly, Thomas: Airplane Aerodynamics. Pittman Nov. 1990, pp. 41-44.

Publishing Co., N.Y., 1967.

Roncz, John G.: Ground Effect. Sport Aviation, Dec.

Etkin, Bernard: Dynamics of Atmospheric Flight. John 1990, pp. 37-40, 98.

Wiley & Sons, New York, N.Y., 1972.

Roncz, John G.: Canards and Other Unsolved Mysteries.

Hoerner, Sighard F.: Fluid-Dynamic Drag. Hoerner Fluid Sport Aviation, Jan. 1991, pp. 57--61.

Dynamics, Midland Park, N.J., 1965.

Roncz, John G.: Evolution of a Homebuilt Design. Sport Kuethe, Arnold M.; and Chow, Chuen-Yen: Foundations Aviation, Feb. 1991, pp. 29-36.

of Aer_/namics: Bases of Aerodynamic Shevell, Richard S.: Fundamentals of Flight. Prentice Design. John Wil,,y & Sons, New York, N.Y., Hall, Englewood Cliffs, N.J., 1989.

1986.

Smith, Hubert: The Illustrated Guide to Aerodynamics.

[.an, C. Edward; and Roskam, Jan: Airplane TAB Books, Blue Ridge Summit, Pa., 1985.

Aerodynamics and Performance. Roskam Aviation and Engineering Co., Ottawa, Kans., Talay, Theodore A.: Introduction to the Aerodynamics of 1988. Flight. NASA SP-367, 1975.

White, Frank M.: Fluid Mechanics. McGraw-Hill, Inc., Loftin, Laurence K., Jr.: Quest for performance. NASA New York, N.Y., 1986.

SP-468, 1985.

Milne-Tbomson, L. M.: Theoretical Aerodynamics. Dover publications, Inc. New York, N.Y., 1978.

PRIGiOII_ P_,._E td_.AI',IK NOT F._Lli4.1_ .,.1 _ .... _:. _ Figure 7. EGADS s{artup screen.

{START EGADS )

/

(QUITEGADS_, MaiiMenu ,_> No Hasthe geometry changed since the lastdragbuildup?

Redodragbuildup Figure Z. Simplifiedflowchart of EGADS.

.t3 _Ldk I | IIIllllJJ AI[ Files Layout Propulslon IV & R Performance J Sense info about (6AOS i Now to get started Ivhat (GAOl ¢on't do .., get I Dlsclalnsors and responsibilities J Print the curranl workshoot It P QWI (with Save options) log I_ Eontrol-.iim.-ILayoul Propulsion IV & B Performance ] Save the currant aircraft file It S J

J Reed e stored aircraft file II el

Sea parameters of some funsoas aircraft J (b) I 1_ Control A/C Files iinuuid Propulsion IV & B Performance I Fuselage N f IvIng ItIV Tall liT Landing gear IlL Plan vluw Front view Sldo vlow Blg wlng S-view N$ (c) [ _ Control A/( Files Layout =,,=,,w,,.1,,..- W & B Performance } [ Piston & Propeller N(I (d} I _ Control A/C Flies Layout Propulsion iilmmll Performance I C6 location N C I Ilorodynonslc canter X RJ (e) iliDlOllllmllla I _ Control A/( Flies Lauoul Propulsion W & B Drag buildup II O Clinsb end turn Range Speeds X U Ceilings Takeoff roll Landing roll Standard atnsosphera Perensetrlc plots Lift distributions (f} Figure 3. Menu Selections. (a) Control menu, (b) A/C Files menu, (c) Layout menu, (d) Propulsion menu, (e) Weight and Balance (W&B) menu, (f) Performance menu.

EGADS : (E)asj (6)eneral (A)viation (D)esign (S)ystem Author: JohnMelton Applied Aerodynamics Branch, Aerodynamics Division (RAA) M/S 227-6, NASA Ames Research Center Moffett Field, CA 94035 ATTN : EGADS Environment : MICROSOFT® QuIckBASIC on the Applee Macintosh'" Documentation: NASATM 104013 EGADS is a collection of "computer worksheets" intended to help the designers of home-built and general aviation aircraft. Because aircraft design Is an iterativeprocess, EGADS provides the user with : I : Design variables that are quick to input and easy to change 2 : Fast recalculationof the results 3 : Lots of graphics The organization of EGADS allows the designer to run through a multitude of design alterations in a minimum amount of time, and have fun in the process.

I Figure 4. Controlmenu: $ome-lnfo-about-EGAD$worksheet.

Getting Started with EGAOS EGADS :Easy General Aviation Design System The easiest way to learn EGADS is to begin by modifying an existing aircraft. Simply choose one of the deslgns already stored on this disk (see the "Read a stored aircraft file" option under the "AIC Files" menu), then proceed to alter the wing, tail, and fuselage geometry (under the "Layout" menu). After the geometry has been altered, calculate the new performance characteristics (under the "'Performance" menu) All of the inputs are "standard Macintosh" : program controls are mouse-activated by pulling down a menu or clicking on a button. Use Command-Stuff-3 to create a screen dump file of a current worksheet (on many Macs, the screen must be .set to two-color mode for this to work), or use the "Print the current worksheet" command from the "Control" menu to send the worksheet straight to the printer. Please refer to the EGADS user's guTde (NASA TM 104013) for additional information, detailed operating instruc- tions, and a discussion of the capabilities and limitations of EGADS Figure 5. Control menu: How-to-Get-Struttedworksheet.

Limitations and the Future of EGADS EGADS • Easy General Aviation Design System EGADS currently has no provisions for Transonic/supersonic calculations Jet er=gines Power effects on stability Cockpit layouts Weight estimation Rotation and flare distances Lateral stability derivatives Structures calculations Lots of other good things ...

But someday ... !

Figure 6. Control menu: What-EGADS-Can't-Do . . . Yet worksheet.

Boring Legal Stuff about EGADS EGADS ' Easy General Aviation Design System The author of EGADS does not warrant, guarantee, or make any representations regarding the use of, or the results of the use of, the EGADS program in terms of correctness, accuracy, reliability, currentness, or otherwise; the user of the program relies on it and its results solely at his/her own risk.

I II I II Figure 7. Control menu: Disclaimers-and-Responsibdities worksheet.

pen the folder containing the d_sign file

"__

l3 A6M (ZERO) "_ _ I Icx HD I"I Beech V-Tail Bonanza _..

[ ,,,.,,: ]

[ Open _'_] ID Piper Cherokee _ [ Cancel ]

- Point and click on the name of the design file

3 - Press this button to open and read the design file-

Figure 8. A/C Files menu: Read-a-Stored-Aircraft-File worksheet.

_X_- Choose a folder for storing the design file

i

lea[snOSl

C:3 Aircraft _ _-_ I Icx HD !IA$1[ [ ompilPr

] EjPcl

C:3 B._SIC overlays | ['._ llO(llhl(I [ illll,e ] r-_Correspondence Fo...

r_ Documentation Enter the A/C fllename [ Sape J

111 I ["co.,.,]

/

E2 - Type the name of the design file in this rectangle

3 - Press this button af[er typing the name and choosing a folder --

Figure 9. A/C Files menu; Save-the-Current-Axcraft-File worksheet.

._?

i Use these parameters as design gu0delines: Aircraft CL CO Wing Aspect Gross Oswald HP max nmn Area Ratao Weight Factor Fokker D-7 1.28 .0404 11S 6.$8 1238 0.72 185 Ryam NYP 1.2S .0379 319 6.63 $135 0.74 220 Cessna 152 1.70 .037Z 160 6.97 1670 0.77 160 Cessna 172 2.10 .0319 175 7.3Z 2300 0.77 160 Piper Cherokee 1.75 .0358 170 6.02 2400 O.7S 180 0eech V-3S 1.8S .0192 181 6.20 3400 0.7S 285 P-S10 1.70 .0163 233 S.86 10100 O.7S 1490 B-17G 1.90 .030Z 1420 7.$8 SSOOO 0.82 4800 Thorp T-18 2.10 .0353 8S.4 S.08 1500 0.74 1S0 Glasair III 2.20 .0180 81.3 6.68 2500 O.75 300 Figure 70. A/C Files menu: See-Parameters-of-Some-Famous-Aircraft worksheet.

m Edit fields for current cross-section dimensions

;"Close box_ view of the current cross section

\ Highlighted current cross section

in quasi-isometric view

Section: g 12 Fue |th 22.5 (ft) Wing upon 33.3 (ft) [ Save this section ) ( Oelete this section J ( Replace section J Next section ] ( Orew isometric )

Help button

Operation buttons _-

Buttons for drawing front, side, and plan views

Fuselage Information Input the dimensions and location of each fuselage cross section X Is the longitudinal location of the cross section, Z _s the height of the section above the ground, and Radius ts the section corner radius All dimensions ere in decimal flat. "lso ingle" refers to the angle (in degrees) at which the sections are drawn _n the quasi-isometric view when the 'Draw isometric" button Is pushed The fuselage wetted area is celculeted ln the "Drag Buildup) worksheet under the "Performance" menu, and the fuselage volume and fineness ratio are displayed when the "EstFusAaro" button Is pressed in the "Lift Distributions" worksheet IMPORTANT - Cross sections are save¢l onlg wf_en the buttons IoDe!ed "Save this section" or Replace section ere pressed Tne number of fuselage cross sections _scurrently ltmtteOtoamaxlmumof 30 After pressing the "'Update" button, use the "F",'S .and 'P buttons tovlew the front, sloe. end plan (top) viewsor the aircraft Figure 1 I. Layout menu. (a) Fuselaqe worksheet, (b) Help information.

e,

Location of mean aerodynamic chord (MAC)

Edit fields for entering wing dimensions Flap--7

Aileron

-- Win| X-lecetleit Wmg Z-lacetlen S_a (¢t) Root chord (ft) Tip Lkertl (ft) DtltedreI (de I) Twist (dog) Incidence (dog) LEdge sweep AIIorln chord if t) Aileron spelt (It) All tnboord (ft) Flop chord (ft) Flop split (ft) MAC 4.04 Flip Inboard (ft) Aspect ratio 6.94 114 chord sweep -0•4 (do I ) Up4_ Tepmr ratio 0.714 AlteroelWIng (qwoos) 0.00 Arls 159.00 (rt'2) Flop/Wing (areas) 0.09

Buttons for drawing front, side, and plan views

Help button

(a)

Wing Information Input the wing. flap, and aileron dimensions in feet end degrees IMPORTANT - This is the only worksheet where PERMANENr changes to the wing planform con be made Although the wing area and aspect ratio show up in other worksheets, they ere pray,deal m those worksheets only for plremetnc anelysi_, aria "whet if" studies "Twist" is also kf_own es "washout". I_J "lfl¢l(hlflce" refers to the Inclination of the wing root with respect to the Z:O fuselage datum plane The location of the MAC shows up as aM extended chord line After pressing the •'Update" button, use the F".

"S", and "P" buttons to view the front, side. and plan (top) views Of the aircraft Figure 1Z. Layout menu. (a) Wing worksheet, (b) Help information.

4O

Editfieldsforentering taildimensions Vertical tail

GrossweightCGmarker

Hz X-location 22.5 (ft) Wing 33.3 (ft) Hz Z-location Hz Sped Hz root titre1 Hz tip chord Hz dlhe4rol HI ll¢ldeNre Hz LE sweep Elevator ¢hord Elevetor inn Ver K-Iocetiu Var height Vet root chord Vsr lip chord Ver LE swoop Rudder chord Area (rt'2) AR Tell volume Hz: 29.50 3.39 0.485 Rudder span Rudder inboard Ver: 12.01 !.42 0.031

Buttons for drawing front, side, and plan views

Help button

(a)

Toll information Input the tad surface dimensions end locations in feet and degrees IMPORTANT - This is the on|y worksheet where PERMANENT changes to the tail planforms can be made. Many of the inputs are similar to those used to describe the wing geometry The tall volume coefficients ere based only on the geometries of the wlng end ta_i, end are NOT a function of the CG location Note that t_evertlcal talllsdrawnlylngonltssldenext terse horizontal stabilizer in the plan view After pressing the "Update" button, usa the 'F', 'S', and "P" buttons to view the front, side, and plan (top) views of the aircraft Figure 13. Layout menu. (a) Tail worksheet, (b) Help information.

..... •.qlt I •

Edit fields for entering gear information --7

F Edit fields for entering weight information /

I Main Hell/Till / I I_ Gross Weight (Ib) _ I-IocIUIO (ft) r'7_"--_ 3.1-"_5_ --1 I

I I Gross Weight CG I--_ Strut length (ft) I--_ i_6_ LJ

i I Elaptg Weight (Ib) I-_ Tire diameter If t) I__ /

"7 Emptg Weight CG I-_ Tiri width If t) I_ .4_ --J l Itln Weight (Ih) I_ Gear laid (Ib) +563.1 *545.6 L Itlo Weight CG _ Wing height above ground *6.0 If t) FvsologO-GrNmd Incidence *2.26 (dug) Wlag-Gro_end incldenHce *4.26 (dog) 0 IlutllCtSlelS gear loads using:

M,.,.. I

® .

Gross weight CG

butt°n _RHeel___tton

Ground line

Landing Gear Information Input the landing gear dimensions and locations in feet, then use the buttons to calculate the static toads on the struts for the different welght'- andCGlocat;ons EGADS assumes that the main gear has two struts and the nose (or tall) gear has one strut Strut length refers to the dlstance between the OOttom of the fuselage and the center of the wheel, and Is used to calculate the angle of incidence between tne ground and the wing and fuselage The real aircraft will have an increasing tenclency to tile over as the CG moves out from between the landing gear The user will De warned of this situation by a bead tone anO e gear IoaO that is less than or equal to zero If changes ere made to the Ioadings contained in the weight lnforl_natlon edit fields, i, he user can reset this weight mforrnat_on to the original values bg pressing the "'R'"(reset) button (permanen, changes to the weight and )oadtng (:late can only be made in the "W & B : CG location" worksheet) Oe sure to lnOtcate whether the gear is ftxedorrotractaLleby p_lsslng the 'Retractable' button if the gear can I)e retracted so that the landlng gear drag w_ll be correcUg accounted for _n the "'Performance" worksheets

(b)

Figure 14. Layout menu. (a) LandingGear worksheet, (b) Helpinformation.

Span: 33.30 (ft)

24.08 (ft)

Overall length :

Aircraft:Cessna 152

Figure 15. Layout menu: Plan View worksheet:.

FRONT VIEW

33.30 (ft)

Span :

24.08 (ft)

Overall length:

Aircraft:Cessna 152

Figure 16. Layou¢ menu: Front View wo.kshee_.

24.08 (ft)

Overoll length'

7.44 (ft)

Overall height'

Aircraft:Cessna 152

I I Figure 17. Layout menu: Side View work_heet.

Span : 33.30 (ft)

Area : 15g.80 (ft'2)

AR : 6.94

Aircraft:Cessna 152

Figure 18. Layout menu: Big Wing worksheet.

Aircraft:Cessna 152

I II III Figure 19. Layout menu: 3-View worksheet.

Supercharged button

Minimum Maximum Step Airspeed (KTAS) 30 I ! 00 115 I Altitude (ft) 0 II0000 i Prop diameter (ft) Prop If fitter.ell .05 SII level In01nI hp 10_ .---I -OSuparcbirged Critical altitude (ft) _ ] PriaI button to plot:

[,rap .m.*..., ,,..*.,paed) (Thro.t ,,..b'.pead I [Sbait po.,e. ,,,.,,iludel

[ Available Power US Airlpaed )

Help button /

Figure 20. Propulsion menu. (a) Piston & Propeller worksheet.

80 0

as]

75 2000 70 4000 65 6000 60 10000 .=0 3S _"_"_'_ I 60 70 80 '0 lO0 _i Airspeed (KTAS) __

/----N---"

Cursor Print button _

Text output button _

(b)

Piston & Propeller I'_formatlon Plots of the effects of airspeed and altituOe on engine anti prop311or performance can be easily made with this worksheet. This Is the only worksheet where permanent changes to _,he shaft horsepower and propeller can be made tf the engine is supercimr'gad, the button should be pressed and the critical altitude er_tered-below the critical altitude, no loss _n or, pine shaft horsepower occurs w_th increased altitude The propeller- efficiency edit field refers to the non-ideal losses associated with the propeller drag and slipstream swirl. This efficiency lactates always mult=plied by the ideal eff|cl_ncy in order to de;.ermme the total propeller effic:ency Theldeelpropellerefficlencylsefunctionof the speed, diameter, altitude, and shaft power of the engine, and Is calculated w,.arever necessan; In EGADS using simple actuator disk theory Use the mouse to point to any place within the graph, end the x- and y-values automatically appear above toe chart Pressing the "TexL" button produces a data file containing the information for each curve.

(c)

Figure ZO. Concluded. (b) Sample plot, (c) Help infc,, marion.

lllt'ld II'_ ,llirl i,rl b_ fll IJl<llrlll I : _'<,_,11,) 't, _ Component Weight (Ib) X-location Eft) Wing O.I!

Horizontal 10.46 (Ih/ft'2) 20-36 Wing loading Vertical Fuselage 6_5 1672.0 (Ib) Grass Weight

1042.0 (Ih)

Engine 1.6 Emptg Weight Main Gear 7.9 I 152.0 (Ib) Minimum Weight Nose/Tell Gear Oetterl l_5 6.36 (ft) Grass Weight CG 6.13 (ft) Propeller 0 Erupt g Weight C6 Avionics 5 6.09 (ft) Mln Weight CG Fuel g Max Pilot 5.7 MIn Pllot 5.7 Pus *' I 5.7 Pass °2 5.7 Recalculate the (:G limits 1 Pose d'3 i 5.7 Baggage 7 i (l:entroids}/i _

Help button --/

(_

Weights Information Input the major components of the aircraft and their longitudinal location from the X:O datum line, Press the "Centrolds" button to dlsploy the area centroids of the wing and horizontal ano vertical stal)lhzers Although temporary changes to the wing and tall locations can be made in this worksheet, permanent changes can only Be made in the "Layout" worksheets. This workshaet is absolutely NOT adequate for the detailed computations essential for an actual design, where EVERY component must be accurately wetghe(I and its location and moment precisely determined

(b)

Figure 21. W & 8 menu. (a) CG Location worksheet, (b) Help information.

4?

_J_ b_d-_g_ 7"_ Wi_WgF T or zero to negreCl _O Ognomic pressure roUo ot tail Downweeh gradient _(eps) I _(AoA) .375 Wing Lift Curve Slope (per degree) .075 Tell Lift Curve Slope (par degree) .057 0.31 Sweep of wing eeredgnamlc centers (dog) Min Weight Gross Weight Acre Center I CG CG Neutral Point h: 1.257 i.312 1.636 x (ft): .09 .36 7,.92 CG iS always ahead of Acre Canter - nir_reft Is static stable Static margin range : +0.323 +0.379 ( Recalculate the tic )j_

_/

Help button

{at

Aerodynamic Center Information This worksheet provldes a quick analytic estimate of the aerodynamic center (neutral point) for the configuration, and computes the stick-fixed static marglns for the varlou$ CGtocattons Remember that the aircraft must be both stable and trlmmable throughout the range of expected lift coefficients Note that the neutral Delhi can also be estimated with the "Lift Distribution'" wort(shlNIt I_j adlustmg the CG local:on (t_e mint about which the moments ere computed) until there is no change on the aircraft pitching moment wlth angle of attack

(b)

Figure 22. W & B menu. (a) Aerodynamic Center worksheet, (b) Help information.

Overlap Wetted Flot-plqte Thlck area %Overlap area area Component CF ratio (ft'2) wetted (ft'2) (ft'2) ACDo Wing __ 18.2 _ 294.04 2.4990.0156 Horlz. toll 2.4 55.71 0.479 0.0030 Vert. tall _1.08 ] 25.62 0.225 0.0014 Fuselage _ COl/ 202.92 1.806 0.0113 Main Gear _ I 1.74 0.467 0.0029 Nose/Tall Gear _ 3.20 0.176 0.0011 Miscellaneous _ 0.0019 Gear retracted Gear extended Total wetted area 570.3 (ft°2) 593.3 (ft'2) Minimum COo 0.0332 0.0372 Equlvalent flat-plate area 5.31 (ft°2) 5.95 (ft°2) Overall CF 92 100 Oswald factor (<I) I .77 !

CF is in drqg counts.

CD_ Is based on gear frontal area. IRecalculate drag buildup_j__._.__

/

"- Help button

Drag Buildup Information This worksheet can be used to calculate the wetted and equivalent flat- plate areas for each of the aircraft components Allowance c_t be the wing or tells which may he mslcle the fuselage by entering he percentage of the overlapping area under the column labeled "'%Overlap wetted" If none of the overlapping plenform area of a component Is to be Included In a component's wetted area, then a zero should be entered In the appropriate I Overlap" field. An estlmetton of the absolute minimum CDo (parasite drag coefficient) due to skin friction alone is then computed.

USa the miscellaneous Input to tncluda the effects of form, cooling, base, lmd other drags Note thatCF(skinfrictloncoefftclent) tsexpressedln drag counts, w_th a CF of O000l equal to i drag count. Note also that the COo found ul_on entry to the "'Performance" worksneets )s chosen according to the type (fixed or retractable) of the lending gear. anO tt_e Oswald factor for the "'Performance worxsheets ts taken from this worksheet For most aircraft, the Oswald factor lies In the rings between 0 7 to 0 aS, and CD_I for most fixed landing gear ranges from 0 17 (full g faired) to 085 (circular stl'lJt with no wheel fatrlngs)

A,.r.c r.o f ! T yp,¢.o.I .o__.tr .e ! ,_ C.F. _(:,_ _d_r.tg" to_.n_ t.=_)

Cessna 150 io0 Piper Cherokee 9¢_ Composite HomebuH t _0 Beech Starshlp _4 P-51 Mustang 30

_)

Figure 23. Performance menu. (a) Drag Buildup worksheet, (b) Help information,.

I1eChn_ Climb angle _d_ 3 Aircraft COo Airspeed (KTAS) Oswald factor (<l) .77 Aircraft CL -.407 I Altitude (It) Prep diameter (ft) Wing urou (ft'2) Prop efflc|encg Aspect retie 6.04 Ideul prop afflcloncg O.OOg Lied fuctor = L/W 0.999 Gross Weight (Ib) .LAZ;._.

W/S (Ib/ft'2) 10.46 Bank anglo (dog) 0 Climb late (ftlmln) -430.0 Airspeed (mph) 95.5 Throet real (Ib) t261.9 Airspeed (KEAS) 79.0 -0.16 Re number 5.93E-06 Thrust/Weight COl (induced) 0.0147 Altitude abaft hp -80,3 COilCD (total) 0.203 Weight/Power (Iblhp) * 10.93 C0 (total) 0.0519 See level shaft hp -99.5 CL/CD 9.57 Turn radius (ft) O.OOE400 Lift (Ib) 1659.7 Turn rata (dee/sac) 0.00 hi (Ib) 174.4 Recalculate using : Oeualtg (slugalft'3) .0021-;3 _f lU'rsPoed |_'CL-_

_Tz

Restore button

Help button

(s]

Climbing and Turning Flight Information This worksheet Is very useful for calculating the lift, drag, and power required during level, climblng, end steady turning flight The user should enter the desired flight tlech number, airspeed, or lift coefficient (CL) tn the appropriate edit field, then click on the corresponding button or press , ";.,, kejboord carriage return Key tn order to perform the calculations The two corresponding values will then be automatically displayed in the neighboring edit fields, along with a full display of other flight parameters in the lower huh of the worksheet For example, iI flight at a specific flilch n_be¢ It destrlNl, tMn the Mach number sMuld first Pa entered tats the "MilCh numOer" edit field, arid the "Mech" button (or keyOosrd carriage return ke U) shOuld be proseed. The carr_apondlng values of airspeed and lift coefficient will t.hen De dr|played In their respective edit fields, along with s large variety of additional Inrc_rmatlon for this flight condition.

Use the "R" I_tton to restart the values taken from the "Lo_ou_" ohd "'CG Location" wOrkShlltl Figure Z4. Performance menu. (a) Climb and Turn worksheet, (b) Help reformation.

5O Aircraft CDo Crb.oe CL J_5 ] OaWL'd factor (< i ) Fraction of fuel used (<I) Cruise altitude if t)

3500 1.0 I

Wing area (ft'2) Ideal prop efflclon¢ g 0.930 Aspect ratio AvO fuel flow (gph) 6.44 OSFC (Iblhp-IIr) True airspeed (KTAS) 97.2 .5 Prop diameter if t) True airspeed (mph) I I 1.0 5.75 .05 EQOIY airspeed (KEAS) 92.5 Prop offlcluncg Moch O. 149 Fuel (gel) 25 Re number 4.600E_06 Gross Weight (Ib) 1672 0.0097 Fuel/Grass Range (SM) 347.4 0.0445 CD Range (NM) 301.9 L/O 7.066 Tlw.a In hours 3. I I 0.7903 Roq hp at altitude 77.3 Optimum CL 64.4 Opt oirupaod (KTAS) Roq hp at see level 07.0 560.2 Successful...

Optimum range (SM) 493.7 Optimum range (NM)

( Recalculate the rangL(____ j

Restore button -/Z'-Help button

Range Information This wart(sheet calculates the aircraft's range and endurance using the classic Bregeut equations The optimum CL for maximum range is displayed, along with the range and flight conditions for the input CL Flights of different lengths con be simulated by adjusting the fraction of fuel used, wlth a fraction of t equal to the maximum tank-empt_ range BSFC ,s the ratio of fuel flow (Ib/hr) to horsepower produced, end I_es e value of approximately 0 5 for most air-cooled reciprocet_n 9 engines The user _s cautioned that the Bregaut. equation includes no a_iowances for taxiing, climb, descent, or reserve Use the R button to restore the values taken from the"La_out' and 'CGLocatlon' worksheets

(b)

Figure 25. Performance menu. (a) Range worksheet, (b) Help information.

.... • _Vl • Aircraft COo .0372 Available power at sea level (hp) Oswald factor (<I) .77 [ I05 I Altitude (ft) 3500 Prop diameter (ft) i 5.75 i Wing area (ft'2) 159.0 Prop afflcloncg I .05

Aspect raUa i

6.94 CL max ! 1.7 W at altitude (Ib) Shaft hp at altitude 93.3 Flight speeds Y rain V max VU Vx Mach number 0.069 O. 160 0.097 0.077 Ideal prop afflcloncg 0.773 0.'.)32 0.005 0.72 I Shaft hp required 43;.2 93.2 93;.3 93.3 True alrspaad (KTAS) 44.9 104.7 63.5 50.6 Re number 2.13E_06 4.96E+06 3;.01E_06 2.39E*05 CL 1.700 0.313 0.845 1.333 CD 0.2093 0.0430 0.0797 O. 1430 CLICD O. 12 7.27 10.60 9.3;2 Rate af climb (ftlmln) 0.0 0.2 655.5 502.9 Angle of climb (dog) +0.00 -0.00 +5.05 +6.54 IRecalculate the speeds] [ Shaw glide spa __eds__

Restore button Help button

(a)

Spaaas Information Thls workshae', detarrntnes the stall speed, maximum speed, bast rate of chmD speeO (Vy). bast angle of climb speeO (Vx). and the speeds for best glide dJstanca and endurance (power off) Note thesansttlv_ttjtoavatlabla horsepower and propeller efficiency - accurate engine data are absolutely essent_alfor accurate performancepra(:l_ct_ons Use the 'R" but Lon to restore the values taken from the "'La_jout'" and "CG Location" workshaats

(b)

Figure 26. Performance menu. (a) Speeds workshee¢, (b) Help information.

'- _,Jl ,I Available power at see level (hp)

I' los I

Wing erie (it'2) I 15g.O Prop diameter (it) Aspect ratio Prop afficlenc U et altitude 1.05 I Weight at altitude (Ib) CeiIings Service lbiolutl Altitude (it) 10065.5 20607.9 Much number O. I 12 O. I 16 Shaft hp required 53.6 48.0 Ideal prop afflclancu 0.044 0.053 True airspeed (KTAS) 69.5 71.2 True eiropied (mph) 00.0 O!.9 Re number 2.26E+06 2.16E*06 CL 1.074 I. I 15 CO O, 1059 O. I I 12 CL/CO I0.14 10.02 Rats of climb (ft/mln) + I00.0 +0.0 Recalculate Ihe ceilings )___

Restore button J L

Help button

¢0

Ceilings Information This wart(sheet estimates the service end absolute ceilings end the flight parameters at these altitudes. One half of the fuel we;grit is subtracted from the gross weight upon Intr_l to this worksheet Note also the slnsltivity of the ceilings to the aveiloi;le horsepower and propeller characteristics AS in the "Sl_lldS" worksheet, accurate an(leas data are ebso;utelyessentlal foreccureteperformanceestlmates use the R bUtton tO restore the values token from the "Layout end C5 Location" workshiets SERVICE CEILING altitude at which maximum rate of climb is lO0 ft/mln ABSOLUTE CEILING :highest possible eltituOe for steady level flight

(b)

Figure ZT. Performance menu. (a) Ceilings worksheet, (b) Help information.

,, _'qlI Inl! _._t 1'_.'

_tT_dl[ gear _0332_ Oowold factor (¢1) .77 Prop diemtor (ft)[5.75 ] Airport altitude (ft) 0 I_edwlnd (knots) Ground friction coefficient Takeoff roll CL .0 Pavement : 0.02 Sea level takeoff hp 105 Gross field : O.lO Takeoff weight (Ib) 1672_ Gear flat-plate (ft'2) .54 I' .02 I Y(rotutu)IV(ateII) 1.2 Wind height ubovo greund if t)

[5 ]

Prop offlcloncg .05 CLmex of takeoff configuretlen Wing urea (ft'2) 159.0 Aspect ratio 5.94 I J.3 I Static thrust (Ib) 532.3 COo wlth Door 0.0372 C01 at llftoff 0.0340 Grouad effect factor 0.093 CD = COo • COl 0.0712 Stall speed (KTAS) 48.0 Rotate airspeed (KTAS) 50.5 Ground speed (mph) 57.3 (Recalculate the takeoff roll I Ground roll (ft) 707.5 Takeoff hp at altitude 105.0 b.,o.

Restore button

Tikeoff information This wortsl_est ten be used to make takeoff roll predictions "Takeoff roll CL'"is the lift coefficient of the affcreft is it roils down the runw_, end is assumed to remain constant until rot.alien. "CLmax of takeoff ¢onfigu- retlun" refers to the maximum lift coefficient of the aircraft tn the takeoff conftgurehon The "Ground friction coefficient' Is used to characterize the r relative rougl_ness of different runwaL, js "Aircraft COo' refers to the COo of the aircraft wathout landing gear; the gear flee-plate erie ts now IncluOed es on Input, end the CDo used In the takeoff roll computations combines tnenlrcr_ft end lapdlng gear drag Note that the takeoff roll calculated bore Is onlV the ground ro?l portion of the riquired takeoff distance u_)to the start of rotation, end dolt not Include the additional distance consumed by the rotation er_dInitial climb to minimum obstacle iclearance height Use the R button to restore the values tek'en from the "Leuout ' duo "CG Location' worksheltS

(b)

Figure Z8. Performance menu. Ca) Takeoff Roll workshee¢, (b) Help information.

6ear flit-plato Eft'2) I .64 : Grossflald : 0.2

V(approach) I V(atell) I 1.2

1.5 I

Wing area Eft'2) Wing height above ground Eft)

Aspect ratio

15 I

Prop diameter Eft) CLmax of lending configuration Approach anglo (dog) i I 1.7 i See level approach hp 17.00 CDo with gear 0.0372 Ground effect factor 0.093 CD app : CDo + CDi O. 1200 Stall speed (KTAS) 42.6 Descent rate (It/rain) -271 Approach airopd (KTAS) 51.2 Ground speed (mph) 50.R [ Recalculate the lending roll ) Lending roll Eft) 435.1 CL appronch i. I0 b

Re sto re bu tton H_L_Ip button

(a)

Lending Information This workshset Is very similar to the "Takeoff" workshest, hut It should be remembered that only the ground roll portion of the lansing distance Is cal- culated, and no allowance is made for flare The 'Seconds to zero power" variable Is the time in secondsafter touchdown that It takes to reduce the throttle from the approach setting to Idle Note that the computed approach airspeed Is the TRUE(not INDICATED)alrsCssd Use the "R" button to restore the values taken from the "Layout" end "CG Location" worksheets

(b)

Figure 29. Performance menu. (a) Landing Roll worksheet, (b) Help information.

Altitude (ft) [5000 Mech number I. I 5 Chord (ft) I 4.04 .0020482 Density (elugelft*3) 1760.06 Pressure (Iblft'2) 41.17 Temperature ('F) 5.10 Temperature ('C) 3.636E-07 Ylscesltg (Ib-eeclft'2) 1.775E-04 Kin vlscosltg (ft'21suc) 1097.09 Speed of sound (ftlsec) 164.56 True airspeed (ftlaec) 97.5 True airspeed (knots) 112.2 True airspeed (mph) 90.8 Equivalent airspeed (knots) 4.487E+06 Re number [ Recalculate quentlUes J_ /-

Help button J

¢Q

Atmosphere Information After entering the altitude. Plach number, end chord length combination of Interest, pressing the "Recalculate quantities" button or keyboard carriage return kay causes thQ corresponding atmospheric conditions and Reynolds rumber to be computed end displayed in the tower half of the werksfleet The atrnospherlc quantities are calculated in accordance with the U S. Extension to the IRA(} Standard Atmosphere (1950) These formulas atmospheric quantities up to I,OOO,OOO feet (190 mllas).

service ceiling of the high*flying U-2 sPyplone Is 70.000 ft. and most pace shuttle ortJtts era Oalow 130 miles (680.000 ft)) The equivalent airspeed ts calculated using the pitot pressure ratio e_d see-level density For most general aviation aircraft, the equivalent airspeed ts the same as the :ndlcatad airspeed (assuming no instrument or position errors for flight at low I'lach numbers) Performance menu. (a) Standard Atmosphere worksheet, (b) Help information.

Figure 30.

Parametric variable

Vertical axis variable Outlined edit field Horizontal ax__ Mlnlmm Ilexlmxm Fixed 1672 1672 weipt lib) _ [_ ( IIJO I0 P-slop wins erda (tt_)_ [.| [ 33.3 33.3 spu (ft) _E| [ ] ( Fixed 6.94 It.U4 ,ep,ct r, uo _l ( ] ( Fixed 0 0 Altitude (it) _|[] [ I00 5 X-step ,Irspoed (KTAS) Ill() [ 0 0 nech number I-I_[ ] [ 0 0 CL I It) ( 0 D CO [][)( Fixed .0372 .1)372 CUe [] I ] [ Fixed 77 ..-.- . (Oeweid) [] !) [ CLICD [I []1)[ 0 , - 0 , - Yhrue, (,,) []_il Power (hp) [._ II I05 1105 Now edit outlined veluoo Eqlletlon _req=DV=VqS(CDo*(MI(qS))21(_IARo)) _1 ( S.l.¢,uori.bi.. ] ( M.k. ihepio,_ /- J -/--"-r Help button L_ Equation to be plotted

_Q

Curve for wing area = 190 sq ft --_ Curve labels _arametric variable "7 " _L_ ' w'u_' _'° v'''c'''_ ,1_

"_1 .," \ *ino ore (rrt°') I

O0 190

0o] \ I/

70 J 160

\ / ii_V

$0 4o 6o 8o Airspeed (KTAS) -- _a

!

--Cursor position in graph _ Curso_ Horizontal axis variable J Curve for wing area ,, 130 ft 2.1

Co)

Figure 31. Performance menu. (a) Parametric Plots worksheet, (b) Sample plot.

m ,.

Parametric Plots Information This workshlot provides en easy way to produce many of the standard plots used in aircraft design end to develop a go_d understanding of the relationship between many of the basic design parameters After selecting the values for the x- anti y-axes of the plot, choose the parametric raft1 then edit the outlined values as required. Click on the "'Make the plot" button to draw the graph. Each individual curve represents o different valu of the paramstrlc variable (nots the labels on the right side of the plot) Use the mouse to point to any place within the graph, and the x- and y-values automatically appear above the chart. Press the "Text" button to writs a data file containing the information for each curve. (Nots - if the step increments are too smell, the labels may overwrite each other and become practically illegible ) The table below summarizes the combi- nations of vat'loblos end ecluatlons that can De plotted wlth this worksheet Effect On For different values Using of (X) (V) of parameter (P) equation .............................................................

S AR Span AR : Span 2 / S Span AR S AR = Span'2 / S AR CO CL,CDo,e CD : CDo * CL'2 / (flARe) V CD W,S,AR,h.CDo.e CD : CDo * (W/(qS))'2 / (flARe)

(c)

For different values Using equation ....................................

Mach CD W,S,AR,h,CDO.o CD = (:Do ° (W/(QS))'2 / (flARe) CL C_ AR, CDO, o CO = CDo + CL'2 / (fARe) hach CL W, S, h CL = W / (qS) CD CL AR. COo, a CD -- CDo * CL'2 / OIARe) V CL W, S, h CL : W / (ClS) Mach CL/CD W.S,AR,h,CDo.e CL : W / (qS) : CD : COo * CL'2 / (flARe) V CL/CD W,S,AR.h,CDo,e CL : W / (qS) CO : CDo * CL'2 / (flARe) h Mech V Math = V / Speed Of sound V Moth h Math = V / Spied of sound V PreQ W,S.AR,h,CDo,s Preq : DV : VqS (CDo ° (W/(qS))'2 / (fARe)) Span S AR AR : Span'2 / Wing area S Span AR AR : Span'2 / Wing area Mech Treq W,S,AR,h,CDo,e Trod : D -- qS (CDo * (Wit(IS))'2 / (flARe)) V Treq W,S,AR,h,CDo,e Treq : D = qS (COo * (W/(qS))'2 / (flARe)) W V S,h,CL L • W = I/2 rho V'2 S CL where h : Altitude, q = 112 rho V'2, S : wing area, V : Velocily, W : Weight

(d)

Figure 3 I. Concluded. (c) Help information, (d) Help information.

Wing Horizontal Fueoluge Incidence (dog) 2 i-2 Cmo Twist (dog) 2 0 e of Vortices 15 I0 (per degree) Airfoil Cmn¢ AoA start AoA finish i-.015 9 Airfoil ZLA (dug) l0 0 _--_ (dug) [_L_(dng) C6 X-locution (ft) J6.35 [EstFusnuro}_(CLhorz][COwing][COhorz]{" Ruculculete CL CD CM } [ Uortlces ][BMwing][BMhorz}[ c*CLwlng][ c'CLhorz][ Text output ] [CL us non][ co us CL ][ CM us CL ][ Save twist end incidence ) [](_ ---I Fuselage Volume : 140.2 ft3 Fineness Ratio : 6.i0 Results for AoA = + I0.0 "--Ib_ 0 Rr,toscale Wing Hertz Fuselage induced Alrcruft Plot limits CL +0.014 *0.027 +0.000 -0.041 +0.041 A0_-2 -I !0 II CD +0.0407 *0.0059 +0.0113 *0.0359 CM -0.107 -0.074 +0.029 -0.101 +0.0731 CL -0.152 C0 • 0.9212 0.1275 0.9039

o.go3g c. -I:.z_.L2__)J

-- Fuselage information

-- Autoscale button

(,i)

Gross weight CG marker

Cursor position in graph

Y: 790 X: 12.99 Control point

Cursor

L Horshoe vortex trailing filament

Print button

(b)

Figure 32.. Performance menu. (a) Lift Distribution worksheet, (b) Horseshoe vortices and control points.

+ • .. ..... A _ -- &_lS_IV . J_ _.2 ] AoA (tie|) ! 6.0 ?.6 .8

t o

.4 /_ a.e _ 1.6 ,2 .4

o _/ "°

- .2 5 .4 3 .2 I 0 t .9 .O /.// ,6

Print button -.----J /

Cursor ----/

,f

Text output button -/

(c)

1.2 I ,8 ou,.o,

(d)

Figure 32. Continued. (c) Wing load distributions, (d) C L vs. Angle of Attack.

6O

I..,,.,LI1,¥

.I .09 .08 .OS .04 _., .03 .0!

o I _ ' ._, i

.oz Cursor

(e)

Cursor position in graph

C!1 x : 0.4059 Y : - 0.0474 .15 I -05 -I -15 -2

Cursor ---_ Te_tobuU_°u? button

(f)

Figure 32. Continued. (e) Co vs. CL, (f) CM vs. CL.

Lift l,_formatioh This wor1(sheet uses s,mpla lifting-llne thaor_ to astlmate the load distributions off the wing _ toil (or wing Mid ¢irlerd) Total etr'¢reft lift, dro_ moment, and Oswald efficiency ore dleplayed, along wlth component coefficients The component lifts, drags, _ moments are referenced to the wln_ planform area. (md include their pQros|te drag contributions.

The data under the "induced" column combine the Jnvi$clP wing an(_ hz results The tredit;onal CL vs AoA. CD vS CL, (rod CM ¥s CL plots are made using the total e)rcroft coefficients The vortices, lift, load, end moment distributions can also be vieweO by pressln 9 the appropriate _uttons RamemDar that the total CM for the all-craft must _)e zero for the aircraft to be In trim Cmac values are the sectional pltchtr_ moment coefficients for the wing and tail airfoils about their aerodynamic centers, and ZLA are the zero lift angles of attack EGADS displays approximations for the fuaeloga zero-lift pitching moment (Cmo), the fuselage pitching-moment- curve slope (Cma), and the fuselage volume and fineness ratio whenever the "';:stFusAero" button Ts prassaO Pressing the "Text output" t)utton produces a data file containing a complete tabulation of worl<sheet results Use the "'R" b_tton to restore the values taken from tha "Layout" onO CG Location' worksheats. See the f'GADS documentation for further information about Ithls workshaet When the number of vortices (currently I im_teO to a total of 40) gets large, be patientl

(o)

Figure 3Z. Concluded. (g) Help information.

(; Control AICFllos Layout Propulsion W&B Performance Figure 33. Control menu: Quir-(with-Save-Oprions) final screen.

REPORT DOCUMENTATION PAGE Fo,.,_

OMB No. 0704.0188 I I II Ill PulSe miX.iRe Ixmlen Iw _ co_cton otIn_cmuNkm _ _lmmm 0oan,mqe t hour INrreq_nm. _ckdnll me tree Im r_ m_l ._ mlra_q ex_ d_N _, Om md m_kt_nkql me ,_m need_l, and _m_eenlD md re_ Ohe colece_n Be mtoemi(m. Oend mnmwxe m_ role Ixmlen mar rely olw mpe_ Be role cmlltcil0et ot MirmmmtlGm.MC:_ _ fw m4ude_ Ih_ bwdem, toW_ _t $eneiae, _i___ _ 1216 Jeilomon oe,_ HlOt,,,mV. tkJto '0a04. _btte_wn. VAm.4._, endWe,eO_ oe ttemoemmt w,,'J Su_l_e_ eW_m_k _ _ _tW. W_. _ m.

1. AGENCY USE ONLY (Leave blenk) I S. REPORT DATE f$1 REPORT TYPE AND DATES COVERED '

I Januar 1994 I Technical Mem°randum

I 4. TITLE AND EUETITLE S. FUNDING NUMBERS

EGADS: A Microcomputer Program for Estimating the Aerodynamic

Performance of General Aviation Aircraft

I •

505-59-53

S. AUTHOR(B)

John E. Melton

II. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESE(EB) REPORT NUMBER

Ames Research Center

A-93066

Moffett Field, CA 94035-1000

I

9. SPONSORING/MONITORING AGENCY NAME(S) AND AOONESS(ES) _o. EPDNEORIN_/MONITORI'NG

AGENCY REPORT NUMBER

National Aeronautics and Space Administration NASA TM- 104013

Washington, DC 20546-0001

I 11. SUPPLEMENTARY NOTES Point of Contact: John E. Melton, Ames Research Center, MS 227-6, Moffett Field, CA 94035-1000

(415) 604-1461

II i 1;le. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified-Unlimited

Subject Category - 05

I I I I I I I 13. ABSTRACT (Mlxlmum 200 words) EGADS is a comprehensive preliminary design tool for estimating the performance of light, single- engine general aviation aircraft. The software runs on the Apple® Macintosh" series of personal com- puters and assists amateur designers and aeronautical engineering students in performing the many repetitive calculations required in the aircraft design process. The program makes full use of the mouse and standard Macintosh interface techniques to simplify the input of various design parameters. Exten- sive graphics, plotting, and text output capabilities are also included.

Ill I 11. NUMBER OF pAGIII 14. IUIJECTTIAMI

7O

General aviation, Microcomputer software, Aircraft design

lB. FA0Cll CODe

A04

IIII St. ||CU_ITY CLA|SI_ICATION |0,' LIMIT'ATMN OF ABSTRACT 17. llCUflITV (:IJliIIIiIICATION 1t. I_CUAITV CI_IISIFICATIOR OiI ABITRACT OF TNIB PAGE OII REPORT

Unclassified Unclassified

im _._,._-_"'

8tamcl_d Vo_m _ (lqtv. _-N) II1._01