Aerodynamics model for a generic ASTOVL lift-fan aircraft

NASA Technical Memorandum 110347

Aerodynamics Model for a

Generic ASTOVL Lift-Fan

Aircraft

Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell, Ames Research Center, Moffett Field, California April 1995 National Aeronautics and Space Administration Ames Research Center Moffett Field, California 94035-1000 side force due to roll rate derivative, Nomenclature Cyp I/rad individual jet exit area, ft 2 side force due to rudder deflection Cy&ud total jet exit area, ft 2 Aj,total derivative, 1/rad b wing span, ft total equivalent circular jet diameter, ft: de mean aerodynamic chord, ft d e = 2._Aj,tota I / n CD drag coefficient D drag, lb CGFS fuselage station center of gravity, in.

FY side force, Ib CGWL waterline center of gravity, in.

GE ground effect rolling moment (RM) coefficient CI h aircraft height from the bottom of the fuselage, ft rolling moment due to sideslip CI[_ derivative, l/rad nondimensional aircraft height h/d e rolling moment due to roll rate CIp IGE in-ground effect derivative, 1/rad ground effect washout factor rolling moment due to yaw rate C1r derivative, 1/rad L lift, Ib lift fan LF rolling moment due to rudder deflection Cl&ud derivative, 1/rad LN lift nozzle lift coefficient CL MRC moment reference center lift coefficient due to pitch rate

Crq

PM pitching moment, ft-lb derivative, l/rad pitch rate, rad/sec q lift coefficient due to angle-of-attack CL a rate derivative, l/rad dynamic pressure, lb/ft 2 RM pitching moment (PM) coefficient rolling moment, ft-lb Cm RN rear nozzle, same as lift nozzle pitching moment due to pitch rate Cmq derivative, I/rad S wing area, ft 2 pitching moment due to angle-of-attack Cm& T total thrust, Ib: T = TLF + TIN rate derivative, I/rad thrust of the lift fan, lb TLF yawing moment (YM) coefficient Cn thrust of the lift nozzles, Ib TIN yawing moment due to sideslip Cn[_ derivative, l/rad equivalent jet velocity ratio:

Ve

Ve.j= q_/qj =_/2Ajq_/Tj yawing moment due to roll rate Cnp derivative, I/rad X-axis moment arm for varying CGFS, XMRC yawing moment due to yaw rate in.

Cn r derivative, l/rad YM yawing moment, ft-lb yawing moment due to rudder Cnsnd Z-axis moment arm for varying ZMRC deflection derivative, I/rad CGWL, in.

side force (FY) coefficient Cy ot angle of attack, deg side force due to sideslip derivative, Cy[_ sideslip angle, rad 1/rad iii PRECEDING PAGE BLANK NOT FILMED 5all unpowered in-ground effect drag aileron deflection angle, deg _C DIGE increment _canard canard deflection angle, deg unpowered in-ground effect lift ACL_--, E 8flap flap deflection angle, deg increment _rud rudder deflection angle, rad unpowered in-ground effect pitching ACmlG E moment increment equivalent jet angle, deg: 6_ AL/T _EQ = _'(_LF) + (I - _')_LN nondimensionalized jet-induced lift increment lift-fan nozzle deflection angle, deg thrust split: X,= TLFfl" lift nozzle deflection angle, deg iv Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft LOURDES G. B IRCKELBAW, WALTER E. MCNEILL, AND DOUGLAS A. WARDWELL Ames Research Center Summary Description of the ASTOVL Lift-Fan Aircraft This report describes the aerodynamics model used in a simulation model of an advanced short takeoff and The representative ASTOVL lift-fan aircraft is a single- vertical landing lift-fan fighter aircraft. The simulation place, single-engine fighter/attack aircraft, featuring a model was developed for use in piloted evaluations of wing-canard arrangement with twin vertical tails, as transition and hover flight regimes, so that only low speed shown in figure I. Geometric characteristics of the (M - 0.2) aerodynamics are included in the mathematical configuration are summarized in table 1; mass properties model. The aerodynamics model includes both the power- are specified in table 2.

off aerodynamic forces and moments and the propulsion The propulsion system concept is presented in figure 2.

system induced aerodynamic effects.

It consists of a remote lift fan coupled to a lift cruise turbofan engine to permit continuous transfer of energy Introduction from the lift cruise engine to the lift fan. The lift cruise engine exhaust is either ducted aft to a thrust deflecting NASA Ames Research Center is participating in cruise nozzle in conventional flight, or diverted to two technology development for advanced short takeoff and deflecting lift nozzles in vertical flight. Throughout vertical landing (ASTOVL) fighter aircraft as a member transition flow can be continuously transferred between of the Joint Advanced Strike Technology (JAST) and the cruise and lift nozzles. Lift-fan and lift-nozzle thrust formerly the Advanced Research Projects Agency can be deflected from 45 to 100 deg below the aircraft (ARPA) ASTOVL program. Integration of flight and waterline. The cruise nozzle can be deflected +20 deg propulsion controls is one of the critical technologies vertically.

being pursued in that program. NASA's role in this technical area is to participate in developing design The basic flight control system consists of the canard, ailerons, and twin rudders for aerodynamic effectors guidelines for integrated flight/propulsion controls, support technology development for ASTOVL demon- during forward flight. For powered-lift operation, control is provided by differential thrust transfer between the lift strator aircraft, and provide consultation on integrated fan and lift nozzles, deflection of lift-fan and lift-nozzle control design to the program contractors. Specifically, thrust, and deflection of cruise-nozzle thrust. Pitch control NASA will carry out design guideline analyses for the is achieved by a combination of canard deflection, thrust control system and conduct piloted simulations on the transfer between the lift fan and lift nozzles, and deflec- Ames Research Center Vertical Motion Simulator (VMS) tion of the cruise nozzle. Roll control is produced by the to evaluate design guidelines and to assess the merits of ailerons and differential thrust transfer between the lift contending design approaches.

nozzles. Yaw control is derived from the combination of The initial effort in this program was to develop a rudder deflection, differential lift-nozzle deflection, and mathematical model for simulation of a representative lateral lift-fan thrust deflection. As an option, reaction ASTOVL aircraft concept. This simulation model was control, powered by the engine compressor bleed air, can used in an experiment on the VMS to gain initial provide additional control moments through nozzles experience with control system behavior and flying located in the wing extremities and in the tail. Longi- qualities for this aircraft concept. A description of the tudinal acceleration is achieved through thrust transfer representative ASTOVL aircraft's integrated flight/ between the lift fan, lift nozzles, and cruise nozzles and propulsion control system, head-up display and the by deflection of the lift-fan and lift-nozzle thrust.

propulsion system performance and dynamic response is provided in reference I. This report describes the repre- sentative aircraft's subsonic, power-off aerodynamics and jet-induced aerodynamics in hover and forward flight, including ground effects.

Aerodynamics Model The unpowered in-ground effects, ACLtGE, ACDIGE, and ACmlGE, were calculated by Digital DATCOM as The aerodynamics model includes both the power-off functions of angle of attack for a height of 6 ft at the wing aerodynamic forces and moments and the propulsion 25 percent mean aerodynamic chord. For this purpose, the system induced aerodynamic effects. The simulation configuration consisted of only the wing and regular experiment focused .on transition and hover flight (unslimmed) body.

regimes, so that only low-speed (M - 0.2) aerodynamics are included in the mathematical model.

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

The power-off aerodynamics data were generated using the U.S. Air Force Stability and Control Digital Longitudinal Aerodynamics DATCOM program (ref. 2) and a NASA Ames in-house graphics program called VORVIEW (no reference Lift- The lift equation for the lift-fan model is shown in available) which allows the user to easily analyze equation !. The first term in this equation represents the arbitrary conceptual aircraft configurations using the power-off lift, and the second term represents the lift VORLAX program (which is based on the vortex lattice increment due to jet-induced effects.

method of ref. 3). All the power-off coefficients and derivatives were calculated in the stability axes. The jet-

aL T

L = CL_S + (i) T induced data were generated using the prediction methods of references 4-8. For the data shown in this report, the The equation for CL is shown in equation 2. Lift curves moment reference for Digital DATCOM was 30.889 ft for CL(0t, 8flap) and C L(Ct, 8canard) are shown in aft of the nose, the moment reference for VORVIEW/ figures 3 and 4, respectively. The curves shown in VORLAX was 31.204 ft aft of the nose (-10 percent of figures 3 and 4 were generated using the vortex-lattice the mean aerodynamic chord), and the moment reference program previously mentioned. Digital DATCOM was for the jet-induced effects was 31. I I ft aft of the nose. In used to predict the pitch rate derivative, CLq= 0.746/rad, the final simulation model, these data were all transferred and the CL¢i (0t) curve, shown in figure 5. Digital to a moment reference center of 31.11 ft.

DATCOM was also used to predict the lift coefficient Due to certain Digital DATCOM limitations, some increment due to the influence of the ground plane, derivatives required special treatment because of the ACLIGE(O0, shown in figure 6, as well as the ground effect washout factor, KGE, shown in figure 7.

canard configuration. For the 6t derivatives, CL_ t and Cmc t , DATCOM methods do not exist for a ratio of forward-surface span to aft-surface span less than 1.5. To CL = CL(IX'Sflap) + ACLScanard + CLq 21_n satisfy this requirement, the aft surface was truncated to a (2) span just less than two-thirds that of the canard. This was +C L. (Or) O_C + (Or) considered a better choice than assuming the derivatives ct 2U B KGEACLIGE were zero.

where Also, the digital DATCOM program had no provision for directly calculating the effects of deflected rudders. The ACLficanard = CL(t_,Scanard) rudder effectiveness derivatives, Cyril, C Is_, and (2a) Cn8 _ , were calculated by replacing the wing and canard - C L (ft., 5canard = 0 °) with an aft horizontal surface with exposed geometry identical to that of the vertical tails and attached to a The expression for the jet-induced lift increment, Alfr, radically slimmed body. At zero angle of attack, the is presented in equation 3. Note that the lift fan and lift trailing-edge surfaces were deflected differentially, as nozzle terms use their respective nozzle angles, 8, and ailerons would be, and the change in rolling moment velocity ratios, V e. However, the fountain term uses the coefficient was calculated. The same surfaces were aircraft's equivalent nozzle angle and velocity ratio.

deflected symmetrically to generate changes in the lift coefficient and the pitching moment coefficient, which were converted to side force and yawing moment coefficients, respectively. All coefficients were calculated using the normal wing (aft lifting surface) reference geometry.

The equation for C m is shown in equation 7. Pitching -- = ,_LF, Ve,LF moment curves for Cm (or, 8flap) and Cm (or, 8canard) are T :ALF shown in figures 23 and 24, respectively. The curves of figure 23 were generated using the vortex-lattice program.

The curves shown in figure 24 were generated using Digital DATCOM. DATCOM was also used to predict the pitch rate derivative, Cmq = -l.589/rad, the curve for Cm6 c (_), shown in figure 25, and the pitching moment coefficient increment due to the influence of the ground plane, ACmlGE(ff.), shown in figure 26.

Figures 8-11 show the jet-induced lift increment due to qE the lift fan for nozzle angles of 90, 75, 60, and 45 deg, C m = Cm(Ot, Sflap)+ ACm_canar d +Cmq 2"_-- B respectively. Figures 12-15 show the jet-induced lift (7) 6t_ increment due to the lift nozzles for angles of 90, 75, 60, and 45 deg, respectively. Figures 16--19 show the jet- + C ma (°Q 2-'_B + KGEACmlGE (Oc) induced lift increment due to the fountain for equivalent where (lift fan and lift nozzle, 8EQ) angles of 90, 75, 60, and 45 deg, respectively.

ACm&zanard = C m (_, 8canard) (7a) Drag- The drag equation for the lift-fan model is shown - Cm(O_, 8canard = 0 °) in equation 4. This equation accounts only for the power- off drag.

The expression for the jet-induced pitching moment D = CD _ S (4) increment, AF'M/Tde, is presented in equation 8.

The equation for CD is shown in equation 5. Drag curves APM VAPM( h 1] for C D (_, 8flap) and C D (_, 8canard) are shown in

• VeI . F

figures 20 and 21, respectively. The curves shown in figures 20 and 21 were generated using the vortex-lattice program. Digital DATCOM was used to predict the drag coefficient increment due to the influence of the ground

L Tde _-de :JLN

plane, ACDIGE(OQ, shown in figure 22.

C D = CD(a, Sflap) + ACD_canar d +/_/--,_EQ, VEQ FAPM( h )] L TOe k, de Fount (5) + KGEACDIGE (or) Figures 27-30 show the jet-induced pitching moment increment due to the lift fan for nozzle angles of 90, 75, where 60, and 45 deg, respectively. Figures 31-34 show the ACD&zanard = CD(Ot,8canard) jet-induced pitching moment increment due to the lift (5a) nozzles for angles of 90, 75, 60, and 45 deg, respectively.

- CD(Ot, 8canard = 0 °) Figures 35-38 show the jet-induced pitching moment increment due to the fountain for equivalent (lift fan Pitching moment- The pitching moment equation for the and lift nozzle, 5130) angles of 90, 75, 60, and 45 deg, lift-fan model is shown in equation 6. The first term in the respectively.

equation represents the power-off pitching moment, the second term represents the jet-induced pitching moment Lateral Directional Aerodynamics increment, and the remaining terms account for center-of- gravity (e.g.) travel.

The Digital DATCOM program was used to predict most of the lateral directional stability derivatives. The static PM = Cm_SE +APMTde derivatives, Cyfl, Cii _, CnB, were obtained for the complete Td e aircraft configu'ratioh by a'dding the individual airframe components: body, wing, canard, and vertical tails, a + (Lcosot+ Dsino0XMR C (6) procedure which assumed the absence of interference.

+ (Lsin c_- D coso0ZMR C Side force-- The side force equation is shown in equation 9, and the expansion of the power-off side force Td e = _ _-'_, LF, e,LF [3 L e_ \ e ] JLF coefficient is presented in equation 10.

FY = Cy _ S (9) +

(13)

Cy = Cyl3 (_)_ + Cyp (0 0 2-_B (10) + CySrud 8rud + Cy (a, 8all) L Tde k de J JFo-nt Digital DATCOM was used to predict the side force Figures 4"/-50 show the jet-induced rolling moment coefficients for C y13 (c0 and C y p (or); these curves are increment due to the lift fan for nozzle angles of 90, 75, shown in figures 39 and 40, respectively. Digital 60, and 45 deg, respectively. Figures 51-54 show the jet- DATCOM was used to predict the rudder derivative: induced rolling moment increment due to the lift nozzles Cy_t_l = 0.2063/rad. The side force coefficient due to for angles of 90, 75, 60, and 45 deg, respectively. Since aileron deflection, Cy (_ Sail), is shown in figure 41 only out-of-ground effects were accounted for, and since and was generated using the vortex-lattice program.

the fountain is only felt in-ground effect, the fountain contribution was zero.

Rolling moment- The rolling moment equation is shown in equation 1 t. The first term accounts for the power-off Yawing moment- The yawing moment equation is rolling moment, the second term represents the jet- shown in equation 14. The first term accounts for the induced rolling moment increment, and the third term power-off yawing moment and the second term accounts accounts for c.g. travel.

for e.g. travel. The jet-induced yawing moment increment could not be predicted very well, but it was assumed to be RM = CI_Sb +ARMTde+FYZMRC (11) small, and therefore neglected.

Td e YM = C n_Sb +FYXMR C (14) The equation for CI is presented in equation 12. Digital DATCOM was used to predict the rolling moment The equation for Cn is presented in equation 15. Digital DATCOM was used to predict the yawing moment coefficients for CI[_(_ ), C 1 (o0, C Ir (_), and Clsmd (00; these curves are shown in _gures 42-45, respectively.

coefficients for C nl_(°0' C n p (_), C n r (°0' and C n 6axt (o0; The rolling moment coefficient due to aileron deflection, these curves are shown in figures 55-58, respectively.

CI (_ Sail), is shown in figure 46 and was generated The yawing moment coefficient due to aileron deflection, using the vortex-lattice program.

Cn (or, Sail), is shown in figure 59 and was generated using the vortex-lattice program.

CI = Ci_ (°0_ + CIp (°0 2_B + Clr (°0 2"_B (12) Cn = Cnl3(o0_+Cnp(Ot ) 2uBPb +Cnr(002._ B (15) + C 18rud (ct)Srud + C I(¢X, 8ail) + Cn&ud (c08rud + C n (a, Sail) The jet-induced rolling moment increment, ARM/Tde, was predicted using the methods of reference 5, and is presented in equation 13. The prediction for rolling Conclusions moment assumes that the effects of i_ are linear and This report describes the aerodynamics model used in a should therefore be limited to 13< 10 deg. Predictions for simulation model of an advanced short takeoff and jet-induced rolling moment per degrees of sideslip in- vertical landing lift-fan fighter aircraft. The simulation ground effect could not be predicted; however, out-of- model was developed for use in piloted evaluations of ground effect numbers were better defined. Therefore, transition and hover flight regimes, so that only low speed only out-of-ground effect rolling moments due to sideslip (M - 0.2) aerodynamics are included in the mathematical were calculated and were assumed height independent.

model. The aerodynamics model includes the power-off aerodynamic forces and moments and the propulsion system induced aerodynamic effects, including ground effects.

5. Kuhn, R. E.: An Engineering Method for Estimating The power-off aerodynamics data were generated the Lateral/Directional Characteristics of using the U.S. Air Force Stability and Control Digital V/STOL Configurations in Transition. NADC DATCOM program and a NASA Ames in-house graphics 81031-60, Naval Air Development Center, program called VORVIEW which allows the user to Warminster, Pa., Feb. 1981.

easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induced data were 6. Stewart, V. R.; and Kuhn, R. E.: A Method for generated using the prediction methods of R. E. Kuhn Prediction of the Aerodynamic Stability and et al., as referenced in this report.

Control Parameters of STOL Aircraft Config- urations; Volume II: STOL Aerodynamic Stability and Control Estimation Methods.

References AFWAL-TR-87-3019, vol. II, secs. 4 and 14, 1. Chung, W. W. Y.; Borchers, P. F.; and Franklin, Flight Dynamics Laboratory, Wright Patterson J. A.: Simulation Model of the Integrated Air Force Base, Ohio, June 1987.

Flight/Propulsion Control System, Displays, and 7. Stewart, V. R.; and Kuhn, R. E.: A Method for Propulsion System for an ASTOVL Lift Fan Prediction of the Aerodynamic Stability and Aircraft. NASA TM-108866, Apr. 1995.

Control Parameters of STOL Aircraft Con- 2. Williams, J. E.; and Vukelich, S. R.: The USAF figurations; Volume III: General Backup Stability and Control Digital DATCOM; Information, Derivation, and Verification.

Volumes I, II, and III. AFFDL-TR-79-3032, AFWAL-TR-87-3019, vol. III, secs. E, H, Apr. 1979.

and K, Flight Dynamics Laboratory, Wright Patterson Air Force Base, Ohio, June 1987.

3. Miranda, L. R.; Elliot, R. D.; and Baker, W. M.: A Generalized Vortex Lattice Method for 8. Henderson, C.; Clark, J.; and Walters, M.: V/STOL Subsonic and Supersonic Flow Applications.

Aerodynamics, Stability & Control Manual NASA CR-2865, Dec. 1977.

(Supplement 1). NADC 80017-60, NAVAL Air Systems Command, Department of the Navy, 4. Kuhn, R. E.; Stewart, V. R.; and Wardwell, D. A.: Washington, D.C., Jan. 1983.

Estimation of Lift and Pitching Moment Induced on Jet STOVL Aircraft Hovering In Ground Effect. WL-TR-93-3046, Flight Dynamics Directorate, Wright Patterson Air Force Base, Ohio, Aug. 1993.

Table I. Aircraft geometry

55.4 ft

Overall length

14.16 ft

Overall height

Area

Wing 523.3 ft 2

36.17 ft Span 18.42 ft Mean aerodynamic chord 2.50 Aspect ratio Leading-edge sweep 40.0 deg Trailing-edge sweep 30.0 deg Airfoil NACA 64A005

Canard Area 243.1 ft2

24.65 ft Span 12.55 ft Mean aerodynamic chord 2.50 Aspect ratio Leading-edge sweep 40.0 deg Trailing-edge sweep 30.0 deg Airfoil NACA 64A004.5 Area

Vertical tail(each) 39.0 f12

6.98 ft Span 7.11 ft Mean aerodynamic chord 1.25 Aspect ratio Leading-edge sweep 40.0 deg Trailing-edge sweep 30.0 deg Airfoil NACA 64A004.5 Table 2. Mass properties 30,000 lb Weight 373.3 in.

x c.g. location 0.0 in.

y c.g. location 96.0 in.

z c.g. location Pitch moment of inertia 91,200 slug-ft 2 Roll moment of inertia 14,300 slug-ft 2 Yaw moment of inertia 101,000 slug-ft 2 Product of inertia 0 slug-ft 2 Figure 1. ASTOVL lift-fan aircraft Lift Fan Lift-Cruise Engine 2D-CD Nozzle I I

I I

I I • a I _"- i i • r Lift Fan Nozzle Lift Nozzles _ Figure 2. Propulsion system configuration O

%_

M Flap = 45 °

-0.8. 1' T

-10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 3. Lift coefficient for various flap deflections, M = 0.2 1.!

,I,,,# It) ° ,,,,,I ¢9 ID ,.d canard = 10 ° A canard = 0 ° it canard = -10 ° canard = -20 ° - -0.3, j • canard = -30 °

¢"

-0.6 7" -10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 4. Lift coefficient for various canard deflections, M = 0.2 O.&

CL ° , 1

rad -0.4' -10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 5. Lift coefficient due to angle-of-attack rate 0.10.

0.05' A C L IGE 0.00 / -0.05 / r -0.10 0 5 10 15 20 25 30 -10 -5 Alpha, deg Figure 6. Lift coefficient increment due to ground plane influence 1.0t 0._'t 0.6' _ K GE 0.4_ _. _ _ •

, \

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-10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 20. Drag coefficient for various flap deflections (includes CDmin), M = 0.2 0.3 I D canard = I0° I ":- canard = 0° - canard = -lO ° canard = -20 ° 0.2 - cana_l = -30 ° o °1-1 °,-i o 0.1 m 0.0 w -0.05, -10 -5 0 5 10 15 20 2S 30 Alpha, dog Figure 21. Drag coefficient for various canard deflections, M = 0.2 0.01,

0.00-

-0.01' _

A C DIGE

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-0.03 .....

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canard = -30 ° 0.20 canard = -20 ° + canard =-10 ° 0.15- canard = 0 ° canard = 10 ° e- ° _...q 0.10- ----o--- canard = 20 ° .o + canard = 30 ° o O o 0o00 o O E lat) ° F,..0 e- -0.10 -0.30 -10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 24. Pitching moment coefficient for various canard deflections, M = 0.2 0.10" 0.05 •

/

Cm. , m o_ rad O.O(Y. _/ -0.05-- L_ d

• _i

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• \

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0.010' -10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 55. Yawing moment coefficient due to sideslip 0.01- C n _ m p rad -0.04.

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\

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\

-0.085" -0.090.

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-0.080.

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\ . .,e

-0.083 ¢, -0.084. ....

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Cn

0.000 -0.002 "0.00_ ' _.006 -10 -5 0 5 10 15 20 25 30 Alpha, deg Figure 59. Yawing moment coefficient for various aileron deflections, M = 0.2 Form Approved

REPORT DOCUMENTATION PAGE OMB NO. 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED April 1995 Technical Memorandum 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft 505-68-32 6. AUTHOR(S) Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell 8. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER Ames Research Center A-950051 Moffett Field, CA 94035-1000 10. SPONSORING/MONITORING 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TM-110347 Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES Point of Contact: Lourdes G. Birckelbaw, Ames Research Center, MS 237-2, Moffett Field, CA 94035-1000 (415) 604-5592 12b. DISTRIBUTION CODE 12a. DISTRIBUTION/AVAILABILITY STATEMENT Unclassified -- Unlimited Subject Category 02 13. ABSTRACT (Maximum 200 words) This report describes the aerodynamics model used in a simulation model of an advanced short takeoff and 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 in the mathematical model. The aerodynamic model includes the power-off aerodynamic forces and moments and 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 Digital DATCOM program and a NASA Ames in-house graphics program called VORVIEW which allows the user to easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induced data were generated using the prediction methods of R. E. Kuhn et al., as referenced in this report.

15. NUMBER OF PAGES 14. SUBJECT TERMS ASTOVL, Lift fan, Aerodynamics model 16. PRICE CODE A04 20. LIMITATION OF ABSTRAC1 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 17. SECURITY CLASSIFICATION OF ABSTRACT OF THIS PAGE OF REPORT Unclassified Unclassified Standard Form 298 (Rev. 2-89) NSN 7540-O1-280-5500 Prescribed by ANSI Std. Z39-1e