National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000

Lourdes G. Birckelbaw, Walter E. McNeill,and Douglas A. Wardwell

NASA Technical Memorandum 110347

Aerodynamics Model for aGeneric ASTOVL Lift-FanAircraft

April 1995

National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000

NASA Technical Memorandum 110347

April 1995

Aerodynamics Model for aGeneric ASTOVL Lift-FanAircraft

Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell, Ames Research Center,Moffett Field, California

iii

Nomenclature

Aj individual jet exit area, ft2

AJ,total total jet exit area, ft2

b wing span, ft

c mean aerodynamic chord, ft

CD drag coefficient

CGFS fuselage station center of gravity, in.

CGWL waterline center of gravity, in.

Cl rolling moment (RM) coefficient

Clβ rolling moment due to sideslipderivative, 1/rad

Clp rolling moment due to roll ratederivative, 1/rad

Clr rolling moment due to yaw ratederivative, 1/rad

Clδrud rolling moment due to rudder deflectionderivative, 1/rad

CL lift coefficient

CLq lift coefficient due to pitch ratederivative, 1/rad

CLα̇ lift coefficient due to angle-of-attackrate derivative, 1/rad

Cm pitching moment (PM) coefficient

Cmq pitching moment due to pitch ratederivative, 1/rad

Cmα̇ pitching moment due to angle-of-attackrate derivative, 1/rad

Cn yawing moment (YM) coefficient

Cnβ yawing moment due to sideslipderivative, 1/rad

Cnp yawing moment due to roll ratederivative, 1/rad

Cnr yawing moment due to yaw ratederivative, 1/rad

Cnδrud yawing moment due to rudderdeflection derivative, 1/rad

Cy side force (FY) coefficient

Cyβ side force due to sideslip derivative,1/rad

Cyp side force due to roll rate derivative,1/rad

Cyδrud side force due to rudder deflectionderivative, 1/rad

de total equivalent circular jet diameter, ft:

d Ae j total= 2 , / π

D drag, lb

FY side force, lb

GE ground effect

h aircraft height from the bottom of thefuselage, ft

h/de nondimensional aircraft height

IGE in-ground effect

KGE ground effect washout factor

L lift, lb

LF lift fan

LN lift nozzle

MRC moment reference center

PM pitching moment, ft-lb

q pitch rate, rad/sec

q dynamic pressure, lb/ft2

RM rolling moment, ft-lb

RN rear nozzle, same as lift nozzle

S wing area, ft2

T total thrust, lb: T = TLF + TLN

TLF thrust of the lift fan, lb

TLN thrust of the lift nozzles, lb

Ve equivalent jet velocity ratio:

V q q A q Te j j j j, / /= =∞ ∞2

XMRC X-axis moment arm for varying CGFS,in.

YM yawing moment, ft-lb

ZMRC Z-axis moment arm for varyingCGWL, in.

α angle of attack, deg

β sideslip angle, rad

iv

δail aileron deflection angle, deg

δcanard canard deflection angle, deg

δflap flap deflection angle, deg

δrud rudder deflection angle, rad

δEQ equivalent jet angle, deg:

δ λ δ λ δEQ LF LN= ( ) + −( )1

δLF lift-fan nozzle deflection angle, deg

δLN lift nozzle deflection angle, deg

∆CDIGE unpowered in-ground effect dragincrement

∆CLIGE unpowered in-ground effect liftincrement

∆CmIGE unpowered in-ground effect pitchingmoment increment

∆L/T nondimensionalized jet-induced liftincrement

λ thrust split: λ = TLF/T

Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft

LOURDES G. BIRCKELBAW, WALTER E. MCNEILL, AND DOUGLAS A. WARDWELL

Ames Research Center

Summary

This report describes the aerodynamics model used in asimulation model of an advanced short takeoff andvertical landing lift-fan fighter aircraft. The simulationmodel was developed for use in piloted evaluations oftransition and hover flight regimes, so that only low speed(M ~ 0.2) aerodynamics are included in the mathematicalmodel. The aerodynamics model includes both the power-off aerodynamic forces and moments and the propulsionsystem induced aerodynamic effects.

Introduction

NASA Ames Research Center is participating intechnology development for advanced short takeoff andvertical landing (ASTOVL) fighter aircraft as a memberof the Joint Advanced Strike Technology (JAST) andformerly the Advanced Research Projects Agency(ARPA) ASTOVL program. Integration of flight andpropulsion controls is one of the critical technologiesbeing pursued in that program. NASA’s role in thistechnical area is to participate in developing designguidelines for integrated flight/propulsion controls,support technology development for ASTOVL demon-strator aircraft, and provide consultation on integratedcontrol design to the program contractors. Specifically,NASA will carry out design guideline analyses for thecontrol system and conduct piloted simulations on theAmes Research Center Vertical Motion Simulator (VMS)to evaluate design guidelines and to assess the merits ofcontending design approaches.

The initial effort in this program was to develop amathematical model for simulation of a representativeASTOVL aircraft concept. This simulation model wasused in an experiment on the VMS to gain initialexperience with control system behavior and flyingqualities for this aircraft concept. A description of therepresentative ASTOVL aircraft’s integrated flight/propulsion control system, head-up display and thepropulsion system performance and dynamic response isprovided in reference 1. This report describes the repre-sentative aircraft’s subsonic, power-off aerodynamics andjet-induced aerodynamics in hover and forward flight,including ground effects.

Description of the ASTOVL Lift-FanAircraft

The representative ASTOVL lift-fan aircraft is a single-place, single-engine fighter/attack aircraft, featuring awing-canard arrangement with twin vertical tails, asshown in figure 1. Geometric characteristics of theconfiguration are summarized in table 1; mass propertiesare specified in table 2.

The propulsion system concept is presented in figure 2.It consists of a remote lift fan coupled to a lift cruiseturbofan engine to permit continuous transfer of energyfrom the lift cruise engine to the lift fan. The lift cruiseengine exhaust is either ducted aft to a thrust deflectingcruise nozzle in conventional flight, or diverted to twodeflecting lift nozzles in vertical flight. Throughouttransition flow can be continuously transferred betweenthe cruise and lift nozzles. Lift-fan and lift-nozzle thrustcan be deflected from 45 to 100 deg below the aircraftwaterline. The cruise nozzle can be deflected ±20 degvertically.

The basic flight control system consists of the canard,ailerons, and twin rudders for aerodynamic effectorsduring forward flight. For powered-lift operation, controlis provided by differential thrust transfer between the liftfan and lift nozzles, deflection of lift-fan and lift-nozzlethrust, and deflection of cruise-nozzle thrust. Pitch controlis achieved by a combination of canard deflection, thrusttransfer between the lift fan and lift nozzles, and deflec-tion of the cruise nozzle. Roll control is produced by theailerons and differential thrust transfer between the liftnozzles. Yaw control is derived from the combination ofrudder deflection, differential lift-nozzle deflection, andlateral lift-fan thrust deflection. As an option, reactioncontrol, powered by the engine compressor bleed air, canprovide additional control moments through nozzleslocated in the wing extremities and in the tail. Longi-tudinal acceleration is achieved through thrust transferbetween the lift fan, lift nozzles, and cruise nozzles andby deflection of the lift-fan and lift-nozzle thrust.

2

Aerodynamics Model

The aerodynamics model includes both the power-offaerodynamic forces and moments and the propulsionsystem induced aerodynamic effects. The simulationexperiment focused on transition and hover flightregimes, so that only low-speed (M ~ 0.2) aerodynamicsare included in the mathematical model.

The power-off aerodynamics data were generated usingthe U.S. Air Force Stability and Control DigitalDATCOM program (ref. 2) and a NASA Ames in-housegraphics program called VORVIEW (no referenceavailable) which allows the user to easily analyzearbitrary conceptual aircraft configurations using theVORLAX program (which is based on the vortex latticemethod of ref. 3). All the power-off coefficients andderivatives were calculated in the stability axes. The jet-induced data were generated using the prediction methodsof references 4–8. For the data shown in this report, themoment reference for Digital DATCOM was 30.889 ftaft of the nose, the moment reference for VORVIEW/VORLAX was 31.204 ft aft of the nose (–10 percent ofthe mean aerodynamic chord), and the moment referencefor the jet-induced effects was 31.11 ft aft of the nose. Inthe final simulation model, these data were all transferredto a moment reference center of 31.11 ft.

Due to certain Digital DATCOM limitations, somederivatives required special treatment because of thecanard configuration. For the α̇ derivatives, CLα̇ andCmα̇ , DATCOM methods do not exist for a ratio offorward-surface span to aft-surface span less than 1.5. Tosatisfy this requirement, the aft surface was truncated to aspan just less than two-thirds that of the canard. This wasconsidered a better choice than assuming the derivativeswere zero.

Also, the digital DATCOM program had no provision fordirectly calculating the effects of deflected rudders. Therudder effectiveness derivatives, Cyδrud, Clδrud, andCnδrud, were calculated by replacing the wing and canardwith an aft horizontal surface with exposed geometryidentical to that of the vertical tails and attached to aradically slimmed body. At zero angle of attack, thetrailing-edge surfaces were deflected differentially, asailerons would be, and the change in rolling momentcoefficient was calculated. The same surfaces weredeflected symmetrically to generate changes in the liftcoefficient and the pitching moment coefficient, whichwere converted to side force and yawing momentcoefficients, respectively. All coefficients were calculatedusing the normal wing (aft lifting surface) referencegeometry.

The unpowered in-ground effects, ∆CLIGE, ∆CDIGE, and∆CmIGE, were calculated by Digital DATCOM asfunctions of angle of attack for a height of 6 ft at the wing25 percent mean aerodynamic chord. For this purpose, theconfiguration consisted of only the wing and regular(unslimmed) body.

The longitudinal aerodynamics terms are discussed nextand are followed by the lateral directional terms.

Longitudinal Aerodynamics

Lift– The lift equation for the lift-fan model is shown inequation 1. The first term in this equation represents thepower-off lift, and the second term represents the liftincrement due to jet-induced effects.

L C q SL

TTL = ∆

+ (1)

The equation for CL is shown in equation 2. Lift curvesfor CL (α, δflap) and CL (α, δcanard) are shown infigures 3 and 4, respectively. The curves shown infigures 3 and 4 were generated using the vortex-latticeprogram previously mentioned. Digital DATCOM wasused to predict the pitch rate derivative, CLq = 0.746/rad,and the CL ˙ ( )α α curve, shown in figure 5. DigitalDATCOM was also used to predict the lift coefficientincrement due to the influence of the ground plane,∆CLIGE(α), shown in figure 6, as well as the groundeffect washout factor, KGE, shown in figure 7.

C C flap C Cqc

UB

Cc

UBK C

L L L canard Lq

L GE LIGE

= ( ) + +

+ ( ) + ( )

α δ

α α α

δ

α

,

˙˙

∆

∆

2

2

(2)

where

∆C C canard

C canard

L canard L

L

δ α δ

α δ

= ( )

− = °( )

,

, 0(2a)

The expression for the jet-induced lift increment, ∆L/T,

is presented in equation 3. Note that the lift fan and liftnozzle terms use their respective nozzle angles, δ, andvelocity ratios, Ve. However, the fountain term uses theaircraft’s equivalent nozzle angle and velocity ratio.

3

∆ = ∆

+ ∆

+ ∆

L

T

L

T

h

dV

L

T

h

dV

L

T

h

dV

ee

ee

e

LF LFLF

LN LNLN

EQ EQFount

, ,

, ,

, ,

,

,

δ

δ

δ

(3)

Figures 8–11 show the jet-induced lift increment due tothe lift fan for nozzle angles of 90, 75, 60, and 45 deg,respectively. Figures 12–15 show the jet-induced liftincrement due to the lift nozzles for angles of 90, 75, 60,and 45 deg, respectively. Figures 16–19 show the jet-induced lift increment due to the fountain for equivalent(lift fan and lift nozzle, δEQ) angles of 90, 75, 60, and45 deg, respectively.

Drag– The drag equation for the lift-fan model is shownin equation 4. This equation accounts only for the power-off drag.

D C SD q = (4)

The equation for CD is shown in equation 5. Drag curvesfor CD (α, δflap) and CD (α, δcanard) are shown infigures 20 and 21, respectively. The curves shown infigures 20 and 21 were generated using the vortex-latticeprogram. Digital DATCOM was used to predict the dragcoefficient increment due to the influence of the groundplane, ∆CDIGE(α), shown in figure 22.

C C flap C

K C

D D D canard

GE DIGE

= ( ) +

+ ( )

α δ

α

δ, ∆

∆ (5)

where

∆C C canard

C canard

D canard D

D

δ α δ

α δ

= ( )

− = °( )

,

, 0 (5a)

Pitching moment– The pitching moment equation for thelift-fan model is shown in equation 6. The first term in theequation represents the power-off pitching moment, thesecond term represents the jet-induced pitching momentincrement, and the remaining terms account for center-of-gravity (c.g.) travel.

P C SPM

TdTd

L D X

L D Z

m qe

e

MRC

MRC

M c = + ∆

+ +( )

+ −( )

cos sin

sin cos

α α

α α

(6)

The equation for Cm is shown in equation 7. Pitchingmoment curves for Cm (α, δflap) and Cm (α, δcanard) areshown in figures 23 and 24, respectively. The curves offigure 23 were generated using the vortex-lattice program.The curves shown in figure 24 were generated usingDigital DATCOM. DATCOM was also used to predictthe pitch rate derivative, Cmq = –1.589/rad, the curve forCm ˙ ( )α α , shown in figure 25, and the pitching momentcoefficient increment due to the influence of the groundplane, ∆CmIGE(α), shown in figure 26.

C C flap C Cqc

UB

Cc

UBK C

m m m canard mq

m GE mIGE

= ( ) + +

+ ( ) + ( )

α δ

α α α

δ

α

,

˙˙

∆

∆

2

2

(7)

where

∆C C canard

C canard

m canard m

m

δ α δ

α δ

= ( )

− = °( )

,

, 0(7a)

The expression for the jet-induced pitching momentincrement, ∆PM/Tde, is presented in equation 8.

∆ ∆

∆

∆

=

+

+

PM

Tde

PM

Tde

h

deVe

PM

Tde

h

deVe

PM

Tde

h

deV

LF LFLF

LN LNLN

EQ EQFount

, ,

, ,

, ,

,

,

δ

δ

δ

(8)

Figures 27–30 show the jet-induced pitching momentincrement due to the lift fan for nozzle angles of 90, 75,60, and 45 deg, respectively. Figures 31–34 show thejet-induced pitching moment increment due to the liftnozzles for angles of 90, 75, 60, and 45 deg, respectively.Figures 35–38 show the jet-induced pitching momentincrement due to the fountain for equivalent (lift fanand lift nozzle, δEQ) angles of 90, 75, 60, and 45 deg,respectively.

Lateral Directional Aerodynamics

The Digital DATCOM program was used to predict mostof the lateral directional stability derivatives. The staticderivatives, Cyβ, Clβ, Cnβ, were obtained for the completeaircraft configuration by adding the individual airframecomponents: body, wing, canard, and vertical tails, aprocedure which assumed the absence of interference.

4

Side force– The side force equation is shown inequation 9, and the expansion of the power-off side forcecoefficient is presented in equation 10.

FY C Sy= q (9)

C C Cpb

U

C rud C ail

y y ypB

y rud y

= ( ) + ( )

+ + ( )

β

δ δ

α β α

α δ

2

,

(10)

Digital DATCOM was used to predict the side forcecoefficients for Cyβ (α) and Cyp (α); these curves areshown in figures 39 and 40, respectively. DigitalDATCOM was used to predict the rudder derivative:Cyδrud = 0.2063/rad. The side force coefficient due toaileron deflection, Cy (α, δail), is shown in figure 41and was generated using the vortex-lattice program.

Rolling moment– The rolling moment equation is shownin equation 11. The first term accounts for the power-offrolling moment, the second term represents the jet-induced rolling moment increment, and the third termaccounts for c.g. travel.

R C SRM

TdTd FY Zl q

ee MRCM b = + ∆ + (11)

The equation for Cl is presented in equation 12. DigitalDATCOM was used to predict the rolling momentcoefficients for Clβ (α), Clp (α), Clr (α), and Clδrud (α);these curves are shown in figures 42–45, respectively.The rolling moment coefficient due to aileron deflection,Cl (α, δail), is shown in figure 46 and was generatedusing the vortex-lattice program.

C C Cpb

UC

rb

U

C rud C ail

l l lpB

lrB

l rud l

= ( ) + ( ) + ( )

+ + ( )( )

β

δ

α β α α

α δ α δ

2 2

,

(12)

The jet-induced rolling moment increment, ∆RM/Tde,was predicted using the methods of reference 5, and ispresented in equation 13. The prediction for rollingmoment assumes that the effects of β are linear andshould therefore be limited to β < 10 deg. Predictions forjet-induced rolling moment per degrees of sideslip in-ground effect could not be predicted; however, out-of-ground effect numbers were better defined. Therefore,only out-of-ground effect rolling moments due to sideslipwere calculated and were assumed height independent.

∆ = ∆

+ ∆

+ ∆

RM

Td

RM

Td

h

dV

RM

Td

h

dV

RM

Td

h

dV

e e ee

e ee

e e

LF LFLF

LN LNLN

EQ EQFount

βδ β

δ β

δ β

, ,

, ,

, ,

,

, (13)

Figures 47–50 show the jet-induced rolling momentincrement due to the lift fan for nozzle angles of 90, 75,60, and 45 deg, respectively. Figures 51–54 show the jet-induced rolling moment increment due to the lift nozzlesfor angles of 90, 75, 60, and 45 deg, respectively. Sinceonly out-of-ground effects were accounted for, and sincethe fountain is only felt in-ground effect, the fountaincontribution was zero.

Yawing moment– The yawing moment equation isshown in equation 14. The first term accounts for thepower-off yawing moment and the second term accountsfor c.g. travel. The jet-induced yawing moment incrementcould not be predicted very well, but it was assumed to besmall, and therefore neglected.

Y C S FY Xn q MRCM b = + (14)

The equation for Cn is presented in equation 15. DigitalDATCOM was used to predict the yawing momentcoefficients for Cnβ (α), Cnp (α), Cnr (α), and Cnδrud(α);these curves are shown in figures 55–58, respectively.The yawing moment coefficient due to aileron deflection,Cn (α, δail), is shown in figure 59 and was generatedusing the vortex-lattice program.

C C Cpb

UC

rb

U

C rud C ail

n n npB

nrB

n rud n

= ( ) + ( ) + ( )

+ + ( )( )

β

δ

α β α α

α δ α δ

2 2

,

(15)

Conclusions

This report describes the aerodynamics model used in asimulation model of an advanced short takeoff andvertical landing lift-fan fighter aircraft. The simulationmodel was developed for use in piloted evaluations oftransition and hover flight regimes, so that only low speed(M ~ 0.2) aerodynamics are included in the mathematicalmodel. The aerodynamics model includes the power-offaerodynamic forces and moments and the propulsionsystem induced aerodynamic effects, including groundeffects.

5

The power-off aerodynamics data were generatedusing the U.S. Air Force Stability and Control DigitalDATCOM program and a NASA Ames in-house graphicsprogram called VORVIEW which allows the user toeasily analyze arbitrary conceptual aircraft configurationsusing the VORLAX program. The jet-induced data weregenerated using the prediction methods of R. E. Kuhnet al., as referenced in this report.

References

1. Chung, W. W. Y.; Borchers, P. F.; and Franklin,J. A.: Simulation Model of the IntegratedFlight/Propulsion Control System, Displays, andPropulsion System for an ASTOVL Lift FanAircraft. NASA TM-108866, Apr. 1995.

2. Williams, J. E.; and Vukelich, S. R.: The USAFStability and Control Digital DATCOM;Volumes I, II, and III. AFFDL-TR-79-3032,Apr. 1979.

3. Miranda, L. R.; Elliot, R. D.; and Baker, W. M.:A Generalized Vortex Lattice Method forSubsonic and Supersonic Flow Applications.NASA CR-2865, Dec. 1977.

4. Kuhn, R. E.; Stewart, V. R.; and Wardwell, D. A.:Estimation of Lift and Pitching Moment Inducedon Jet STOVL Aircraft Hovering In GroundEffect. WL-TR-93-3046, Flight DynamicsDirectorate, Wright Patterson Air Force Base,Ohio, Aug. 1993.

5. Kuhn, R. E.: An Engineering Method for Estimatingthe Lateral/Directional Characteristics ofV/STOL Configurations in Transition. NADC81031-60, Naval Air Development Center,Warminster, Pa., Feb. 1981.

6. Stewart, V. R.; and Kuhn, R. E.: A Method forPrediction of the Aerodynamic Stability andControl Parameters of STOL Aircraft Config-urations; Volume II: STOL AerodynamicStability and Control Estimation Methods.AFWAL-TR-87-3019, vol. II, secs. 4 and 14,Flight Dynamics Laboratory, Wright PattersonAir Force Base, Ohio, June 1987.

7. Stewart, V. R.; and Kuhn, R. E.: A Method forPrediction of the Aerodynamic Stability andControl Parameters of STOL Aircraft Con-figurations; Volume III: General BackupInformation, Derivation, and Verification.AFWAL-TR-87-3019, vol. III, secs. E, H,and K, Flight Dynamics Laboratory, WrightPatterson Air Force Base, Ohio, June 1987.

8. Henderson, C.; Clark, J.; and Walters, M.: V/STOLAerodynamics, Stability & Control Manual(Supplement 1). NADC 80017-60, NAVAL AirSystems Command, Department of the Navy,Washington, D.C., Jan. 1983.

6

Table 1. Aircraft geometry

Overall length 55.4 ft

Overall height 14.16 ft

Wing Area 523.3 ft2

Span 36.17 ft

Mean aerodynamic chord 18.42 ft

Aspect ratio 2.50

Leading-edge sweep 40.0 deg

Trailing-edge sweep 30.0 deg

Airfoil NACA 64A005

Canard Area 243.1 ft2

Span 24.65 ft

Mean aerodynamic chord 12.55 ft

Aspect ratio 2.50

Leading-edge sweep 40.0 deg

Trailing-edge sweep 30.0 deg

Airfoil NACA 64A004.5

Vertical tail (each) Area 39.0 ft2

Span 6.98 ft

Mean aerodynamic chord 7.11 ft

Aspect ratio 1.25

Leading-edge sweep 40.0 deg

Trailing-edge sweep 30.0 deg

Airfoil NACA 64A004.5

Table 2. Mass properties

Weight 30,000 lb

x c.g. location 373.3 in.

y c.g. location 0.0 in.

z c.g. location 96.0 in.

Pitch moment of inertia 91,200 slug-ft2

Roll moment of inertia 14,300 slug-ft2

Yaw moment of inertia 101,000 slug-ft2

Product of inertia 0 slug-ft2

7

Figure 1. ASTOVL lift-fan aircraft

8

Figure 2. Propulsion system configuration

9

Figure 3. Lift coefficient for various flap deflections, M = 0.2

10

Figure 4. Lift coefficient for various canard deflections, M = 0.2

11

Figure 5. Lift coefficient due to angle-of-attack rate

Figure 6. Lift coefficient increment due to ground plane influence

12

Figure 7. Power-off ground effect washout factor

13

14

15

16

17

18

19

20

21

22

23

24

25

Figure 20. Drag coefficient for various flap deflections (includes CDmin), M = 0.2

26

Figure 21. Drag coefficient for various canard deflections, M = 0.2

27

Figure 22. Drag coefficient increment due to ground plane influence

28

Figure 23. Pitching moment coefficient for various flap deflections, M = 0.2

29

Figure 24. Pitching moment coefficient for various canard deflections, M = 0.2

30

Figure 25. Pitching moment coefficient due to angle-of-attack rate

Figure 26. Pitching moment coefficient increment due to ground plane influence

31

32

33

34

35

36

37

38

39

40

41

42

43

Figure 39. Side force coefficient due to sideslip

Figure 40. Side force coefficient due to roll rate

44

Figure 41. Side force coefficient for various aileron deflections, M = 0.2

45

Figure 42. Rolling moment coefficient due to sideslip

Figure 43. Rolling moment coefficient due to roll rate

46

Figure 44. Rolling moment coefficient due to yaw rate

Figure 45. Rolling moment coefficient due to rudder deflection

47

Figure 46. Rolling moment coefficient for various aileron deflections, M = 0.2

48

49

50

51

52

53

54

55

56

Figure 55. Yawing moment coefficient due to sideslip

Figure 56. Yawing moment coefficient due to roll rate

57

Figure 57. Yawing moment coefficient due to yaw rate

Figure 58. Yawing moment coefficient due to rudder deflection

58

Figure 59. Yawing moment coefficient for various aileron deflections, M = 0.2

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11. SUPPLEMENTARY NOTES

Unclassified Unclassified

Unclassified — UnlimitedSubject Category 02

A-950051

NASA TM-110347

April 1995

Ames Research CenterMoffett Field, CA 94035-1000

National Aeronautics and Space AdministrationWashington, DC 20546-0001

505-68-32

63

A04

Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft

Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell

This report describes the aerodynamics model used in a simulation model of an advanced short takeoffand vertical landing lift-fan fighter aircraft. The simulation model was developed for use in piloted evalua-tions of transition and hover flight regimes, so that only low speed (M ~ 0.2) aerodynamics are included inthe mathematical model. The aerodynamic model includes the power-off aerodynamic forces and momentsand the propulsion system induced aerodynamic effects, including ground effects.

The power-off aerodynamics data were generated using the U.S. Air Force Stability and Control DigitalDATCOM program and a NASA Ames in-house graphics program called VORVIEW which allows the userto easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induceddata were generated using the prediction methods of R. E. Kuhn et al., as referenced in this report.

ASTOVL, Lift fan, Aerodynamics model

Technical Memorandum

Point of Contact: Lourdes G. Birckelbaw, Ames Research Center, MS 237-2, Moffett Field, CA 94035-1000 (415) 604-5592