aerodynamics model for a generic astovl lift-fan aircraft · nasa technical memorandum 110347...
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NASA Technical Memorandum 110347
Aerodynamics Model for aGeneric ASTOVL Lift-FanAircraft
Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell, Ames Research Center,
Moffett Field, California
April 1995
National Aeronautics andSpace Administration
Ames Research CenterMoffett Field, California 94035-1000
https://ntrs.nasa.gov/search.jsp?R=19950019882 2018-06-30T14:36:35+00:00Z
Nomenclature
Aj,total
b
CD
CGFS
CGWL
CI
CI[_
CIp
C1r
Cl&ud
CL
Crq
CL a
Cm
Cmq
Cm&
Cn
Cn[_
Cnp
Cn r
Cnsnd
Cy
Cy[_
individual jet exit area, ft 2
total jet exit area, ft 2
wing span, ft
mean aerodynamic chord, ft
drag coefficient
fuselage station center of gravity, in.
waterline center of gravity, in.
rolling moment (RM) coefficient
rolling moment due to sideslipderivative, l/rad
rolling moment due to roll ratederivative, 1/rad
rolling moment due to yaw ratederivative, 1/rad
rolling moment due to rudder deflectionderivative, 1/rad
lift coefficient
lift coefficient due to pitch ratederivative, l/rad
lift coefficient due to angle-of-attackrate derivative, l/rad
pitching moment (PM) coefficient
pitching moment due to pitch ratederivative, I/rad
pitching moment due to angle-of-attackrate derivative, I/rad
yawing moment (YM) coefficient
yawing moment due to sideslip
derivative, l/rad
yawing moment due to roll ratederivative, I/rad
yawing moment due to yaw ratederivative, l/rad
yawing moment due to rudderdeflection derivative, I/rad
side force (FY) coefficient
side force due to sideslip derivative,1/rad
Cyp
Cy&ud
de
D
FY
GE
h
h/d e
IGE
L
LF
LN
MRC
PM
q
RM
RN
S
T
TLF
TIN
Ve
XMRC
YM
ZMRC
ot
side force due to roll rate derivative,
I/rad
side force due to rudder deflection
derivative, 1/rad
total equivalent circular jet diameter, ft:
de = 2._Aj,tota I / n
drag, lb
side force, Ib
ground effect
aircraft height from the bottom of thefuselage, ft
nondimensional aircraft height
in-ground effect
ground effect washout factor
lift, Ib
lift fan
lift nozzle
moment reference center
pitching moment, ft-lb
pitch rate, rad/sec
dynamic pressure, lb/ft 2
rolling moment, ft-lb
rear nozzle, same as lift nozzle
wing area, ft 2
total thrust, Ib: T = TLF + TIN
thrust of the lift fan, lb
thrust of the lift nozzles, Ib
equivalent jet velocity ratio:
Ve.j= q_/qj =_/2Ajq_/Tj
X-axis moment arm for varying CGFS,in.
yawing moment, ft-lb
Z-axis moment arm for varyingCGWL, in.
angle of attack, deg
sideslip angle, rad
PRECEDING PAGE BLANK NOT FILMED
iii
5all
_canard
8flap
_rud
6_
aileron deflection angle, deg
canard deflection angle, deg
flap deflection angle, deg
rudder deflection angle, rad
equivalent jet angle, deg:
_EQ = _'(_LF) + (I - _')_LN
lift-fan nozzle deflection angle, deg
lift nozzle deflection angle, deg
_C DIGE
ACL_--,E
ACmlG E
AL/T
unpowered in-ground effect dragincrement
unpowered in-ground effect liftincrement
unpowered in-ground effect pitchingmoment increment
nondimensionalized jet-induced liftincrement
thrust split: X,= TLFfl"
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
This report describes the aerodynamics model used in asimulation model of an advanced short takeoff and
vertical landing lift-fan fighter aircraft. The simulation
model was developed for use in piloted evaluations of
transition and hover flight regimes, so that only low speed(M - 0.2) aerodynamics are included in the mathematical
model. The aerodynamics model includes both the power-
off aerodynamic forces and moments and the propulsion
system induced aerodynamic effects.
Introduction
NASA Ames Research Center is participating in
technology development for advanced short takeoff andvertical landing (ASTOVL) fighter aircraft as a member
of the Joint Advanced Strike Technology (JAST) and
formerly the Advanced Research Projects Agency
(ARPA) ASTOVL program. Integration of flight and
propulsion controls is one of the critical technologies
being pursued in that program. NASA's role in thistechnical area is to participate in developing design
guidelines for integrated flight/propulsion controls,
support technology development for ASTOVL demon-
strator aircraft, and provide consultation on integrated
control design to the program contractors. Specifically,
NASA will carry out design guideline analyses for the
control system and conduct piloted simulations on theAmes Research Center Vertical Motion Simulator (VMS)
to evaluate design guidelines and to assess the merits of
contending design approaches.
The initial effort in this program was to develop a
mathematical model for simulation of a representative
ASTOVL aircraft concept. This simulation model was
used in an experiment on the VMS to gain initial
experience with control system behavior and flying
qualities for this aircraft concept. A description of the
representative ASTOVL aircraft's integrated flight/
propulsion control system, head-up display and the
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.
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, as
shown in figure I. Geometric characteristics of the
configuration are summarized in table 1; mass properties
are specified in table 2.
The propulsion system concept is presented in figure 2.
It consists of a remote lift fan coupled to a lift cruise
turbofan engine to permit continuous transfer of energy
from the lift cruise engine to the lift fan. The lift cruise
engine exhaust is either ducted aft to a thrust deflecting
cruise nozzle in conventional flight, or diverted to two
deflecting lift nozzles in vertical flight. Throughouttransition flow can be continuously transferred betweenthe cruise and lift nozzles. Lift-fan and lift-nozzle thrust
can be deflected from 45 to 100 deg below the aircraftwaterline. The cruise nozzle can be deflected +20 deg
vertically.
The basic flight control system consists of the canard,ailerons, and twin rudders for aerodynamic effectors
during forward flight. For powered-lift operation, controlis provided by differential thrust transfer between the liftfan and lift nozzles, deflection of lift-fan and lift-nozzle
thrust, and deflection of cruise-nozzle thrust. Pitch control
is 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 lift
nozzles. Yaw control is derived from the combination of
rudder deflection, differential lift-nozzle deflection, and
lateral lift-fan thrust deflection. As an option, reaction
control, powered by the engine compressor bleed air, can
provide additional control moments through nozzleslocated in the wing extremities and in the tail. Longi-
tudinal acceleration is achieved through thrust transfer
between the lift fan, lift nozzles, and cruise nozzles and
by deflection of the lift-fan and lift-nozzle thrust.
Aerodynamics Model
The aerodynamics model includes both the power-off
aerodynamic forces and moments and the propulsion
system induced aerodynamic effects. The simulation
experiment focused .on transition and hover flight
regimes, 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 Digital
DATCOM program (ref. 2) and a NASA Ames in-house
graphics program called VORVIEW (no reference
available) which allows the user to easily analyze
arbitrary conceptual aircraft configurations using the
VORLAX program (which is based on the vortex lattice
method of ref. 3). All the power-off coefficients and
derivatives were calculated in the stability axes. The jet-
induced data were generated using the prediction methods
of references 4-8. For the data shown in this report, the
moment 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 of
the mean aerodynamic chord), and the moment reference
for the jet-induced effects was 31. I I ft aft of the nose. Inthe final simulation model, these data were all transferred
to a moment reference center of 31.11 ft.
Due to certain Digital DATCOM limitations, some
derivatives required special treatment because of the
canard configuration. For the 6t derivatives, CL_ t and
Cmc t , DATCOM methods do not exist for a ratio offorward-surface span to aft-surface span less than 1.5. To
satisfy this requirement, the aft surface was truncated to a
span just less than two-thirds that of the canard. This was
considered a better choice than assuming the derivativeswere zero.
Also, the digital DATCOM program had no provision for
directly calculating the effects of deflected rudders. The
rudder effectiveness derivatives, Cyril, C Is_, andCn8 _ , were calculated by replacing the wing and canardwith an aft horizontal surface with exposed geometryidentical to that of the vertical tails and attached to a
radically slimmed body. At zero angle of attack, the
trailing-edge surfaces were deflected differentially, as
ailerons would be, and the change in rolling momentcoefficient was calculated. The same surfaces were
deflected symmetrically to generate changes in the lift
coefficient and the pitching moment coefficient, which
were converted to side force and yawing momentcoefficients, respectively. All coefficients were calculated
using the normal wing (aft lifting surface) reference
geometry.
The unpowered in-ground effects, ACLtGE, ACDIGE, and
ACmlGE, were calculated by Digital DATCOM asfunctions of angle of attack for a height of 6 ft at the wing
25 percent mean aerodynamic chord. For this purpose, the
configuration 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 in
equation !. The first term in this equation represents thepower-off lift, and the second term represents the lift
increment due to jet-induced effects.
aLTL = CL_S + (i)T
The equation for CL is shown in equation 2. Lift curves
for CL(0t, 8flap) and C L(Ct, 8canard) are shown in
figures 3 and 4, respectively. The curves shown in
figures 3 and 4 were generated using the vortex-lattice
program previously mentioned. Digital DATCOM was
used to predict the pitch rate derivative, CLq= 0.746/rad,and the CL¢i (0t) curve, shown in figure 5. DigitalDATCOM was also used to predict the lift coefficient
increment due to the influence of the ground plane,
ACLIGE(O0, shown in figure 6, as well as the groundeffect washout factor, KGE, shown in figure 7.
CL = CL(IX'Sflap) + ACLScanard + CLq 21_n
(2)
+C L. (Or) O_C + (Or)ct 2U B KGEACLIGE
where
ACLficanard = CL(t_,Scanard)
- C L (ft., 5canard = 0 °)
The expression for the jet-induced lift increment, Alfr,
is presented in equation 3. Note that the lift fan and liftnozzle terms use their respective nozzle angles, 8, and
velocity ratios, V e. However, the fountain term uses the
aircraft's equivalent nozzle angle and velocity ratio.
(2a)
2
-- = ,_LF, Ve,LFT :ALF
Figures 8-11 show the jet-induced lift increment due to
the lift fan for nozzle angles of 90, 75, 60, and 45 deg,
respectively. Figures 12-15 show the jet-induced lift
increment 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, 8EQ) angles of 90, 75, 60, and
45 deg, respectively.
Drag- The drag equation for the lift-fan model is shown
in equation 4. This equation accounts only for the power-
off drag.
D = CD _ S (4)
The equation for CD is shown in equation 5. Drag curves
for C D (_, 8flap) and C D (_, 8canard) are shown in
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
plane, ACDIGE(OQ, shown in figure 22.
C D = CD(a, Sflap) + ACD_canar d
where
+ KGEACDIGE (or)
(5)
ACD&zanard = CD(Ot,8canard)
(5a)
- CD(Ot, 8canard = 0 °)
Pitching moment- The pitching moment equation for the
lift-fan model is shown in equation 6. The first term in the
equation represents the power-off pitching moment, the
second term represents the jet-induced pitching moment
increment, and the remaining terms account for center-of-
gravity (e.g.) travel.
PM = Cm_SE +APMTdeTd e
+ (Lcosot+ Dsino0XMR C (6)
+ (Lsin c_- D coso0ZMR C
The equation for C m is shown in equation 7. Pitching
moment curves for Cm (or, 8flap) and Cm (or, 8canard) are
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.
qE
C m = Cm(Ot, Sflap)+ ACm_canar d +Cmq 2"_--B
where
6t_
+ C ma (°Q 2-'_B + KGEACmlGE (Oc)
(7)
ACm&zanard = C m (_, 8canard)(7a)
- Cm(O_, 8canard = 0 °)
The expression for the jet-induced pitching moment
increment, AF'M/Tde, is presented in equation 8.
APM VAPM( h 1]• VeI . F
LTde _-de :JLN
FAPM( h )]+/_/--,_EQ, VEQL TOe k,de Fount
Figures 27-30 show the jet-induced pitching moment
increment due to the lift fan for nozzle angles of 90, 75,
60, and 45 deg, respectively. Figures 31-34 show the
jet-induced pitching moment increment due to the lift
nozzles for angles of 90, 75, 60, and 45 deg, respectively.
Figures 35-38 show the jet-induced pitching moment
increment due to the fountain for equivalent (lift fan
and lift nozzle, 5130) angles of 90, 75, 60, and 45 deg,
respectively.
Lateral Directional Aerodynamics
The Digital DATCOM program was used to predict most
of the lateral directional stability derivatives. The static
derivatives, Cyfl, Cii _, CnB, were obtained for the complete
aircraft configu'ratioh by a'dding the individual airframe
components: body, wing, canard, and vertical tails, a
procedure which assumed the absence of interference.
Side force-- The side force equation is shown in
equation 9, and the expansion of the power-off side force
coefficient is presented in equation 10.
FY = Cy _ S (9)
Cy = Cyl3 (_)_ + Cyp (0 0 2-_B(10)
+ CySrud 8rud + Cy (a, 8all)
Digital DATCOM was used to predict the side force
coefficients for C y13(c0 and C yp(or); these curves areshown in figures 39 and 40, respectively. Digital
DATCOM was used to predict the rudder derivative:
Cy_t_l = 0.2063/rad. The side force coefficient due to
aileron deflection, Cy (_ Sail), is shown in figure 41and was generated using the vortex-lattice program.
Rolling moment- The rolling moment equation is shown
in equation 1 t. The first term accounts for the power-off
rolling moment, the second term represents the jet-
induced rolling moment increment, and the third termaccounts for c.g. travel.
RM = CI_Sb +ARMTde+FYZMRC (11)Td e
The equation for CI is presented in equation 12. DigitalDATCOM was used to predict the rolling moment
coefficients for CI[_(_ ), C 1 (o0, C Ir (_), and Clsmd (00;these curves are shown in _gures 42-45, respectively.
The rolling moment coefficient due to aileron deflection,
CI (_ Sail), is shown in figure 46 and was generated
using the vortex-lattice program.
CI = Ci_ (°0_ + CIp (°0 2_B + Clr (°0 2"_B (12)
+ C 18rud (ct)Srud + C I(¢X,8ail)
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 rollingmoment assumes that the effects of i_ are linear and
should therefore be limited to 13< 10 deg. Predictions for
jet-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 sideslip
were calculated and were assumed height independent.
Td e = _ _-'_, LF, e,LF [3L e_ \ e ] JLF
+(13)
L Tde kde J JFo-nt
Figures 4"/-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 nozzles
for angles of 90, 75, 60, and 45 deg, respectively. Sinceonly out-of-ground effects were accounted for, and since
the fountain is only felt in-ground effect, the fountaincontribution was zero.
Yawing moment- The yawing moment equation is
shown in equation 14. The first term accounts for the
power-off yawing moment and the second term accounts
for e.g. travel. The jet-induced yawing moment incrementcould not be predicted very well, but it was assumed to be
small, and therefore neglected.
YM = C n_Sb +FYXMR C (14)
The equation for Cn is presented in equation 15. DigitalDATCOM was used to predict the yawing moment
coefficients for C nl_(°0' C np (_), C n r (°0' and C n6axt(o0;these curves are shown in figures 55-58, respectively.
The yawing moment coefficient due to aileron deflection,
Cn (or, Sail), is shown in figure 59 and was generated
using the vortex-lattice program.
Cn = Cnl3(o0_+Cnp(Ot ) 2uBPb +Cnr(002._ B (15)
+ Cn&ud (c08rud + C n (a, Sail)
Conclusions
This report describes the aerodynamics model used in asimulation model of an advanced short takeoff and
vertical landing lift-fan fighter aircraft. The simulation
model was developed for use in piloted evaluations of
transition and hover flight regimes, so that only low speed(M - 0.2) aerodynamics are included in the mathematical
model. The aerodynamics model includes the power-off
aerodynamic forces and moments and the propulsion
system induced aerodynamic effects, including groundeffects.
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 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.
References
1. Chung, W. W. Y.; Borchers, P. F.; and Franklin,
J. A.: Simulation Model of the Integrated
Flight/Propulsion Control System, Displays, and
Propulsion System for an ASTOVL Lift Fan
Aircraft. NASA TM-108866, Apr. 1995.
2. Williams, J. E.; and Vukelich, S. R.: The USAF
Stability 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 for
Subsonic 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 Induced
on Jet STOVL Aircraft Hovering In Ground
Effect. WL-TR-93-3046, Flight Dynamics
Directorate, Wright Patterson Air Force Base,
Ohio, Aug. 1993.
5. Kuhn, R. E.: An Engineering Method for Estimatingthe Lateral/Directional Characteristics of
V/STOL Configurations in Transition. NADC
81031-60, Naval Air Development Center,Warminster, Pa., Feb. 1981.
6. Stewart, V. R.; and Kuhn, R. E.: A Method for
Prediction of the Aerodynamic Stability andControl Parameters of STOL Aircraft Config-
urations; Volume II: STOL Aerodynamic
Stability 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 for
Prediction of the Aerodynamic Stability andControl Parameters of STOL Aircraft Con-
figurations; Volume III: General Backup
Information, 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/STOL
Aerodynamics, Stability & Control Manual
(Supplement 1). NADC 80017-60, NAVAL Air
Systems Command, Department of the Navy,
Washington, D.C., Jan. 1983.
Wing
Canard
Verticaltail(each)
TableI. Aircraftgeometry
OveralllengthOverallheight
Area
Span
Mean aerodynamic chord
Aspect ratio
Leading-edge sweep
Trailing-edge sweep
Airfoil
Area
Span
Mean aerodynamic chord
Aspect ratio
Leading-edge sweep
Trailing-edge sweep
Airfoil
Area
Span
Mean aerodynamic chord
Aspect ratio
Leading-edge sweep
Trailing-edge sweep
Airfoil
55.4 ft
14.16 ft
523.3 ft 2
36.17 ft
18.42 ft
2.50
40.0 deg
30.0 deg
NACA 64A005
243.1 ft2
24.65 ft
12.55 ft
2.50
40.0 deg
30.0 deg
NACA 64A004.5
39.0 f12
6.98 ft
7.11 ft
1.25
40.0 deg
30.0 deg
NACA 64A004.5
Table 2. Mass properties
Weight
x c.g. location
y c.g. location
z c.g. location
Pitch moment of inertia
Roll moment of inertia
Yaw moment of inertia
Product of inertia
30,000 lb
373.3 in.
0.0 in.
96.0 in.
91,200 slug-ft 2
14,300 slug-ft 2
101,000 slug-ft 2
0 slug-ft 2
Figure 1. ASTOVL lift-fan aircraft
Lift Fan
I I
I II I
Lift FanNozzle
Lift-Cruise Engine2D-CD Nozzle
• a I _"- i i • r
Lift Nozzles _
Figure 2. Propulsion system configuration
O
00
%_M
1
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
9
,I,,,#
It)° ,,,,,I
¢9
ID
,.d
1.!
it
-0.3, j •
¢"-0.6 7"
-10 -5 0 5 10 15
Alpha, deg
A
canard = 10 °
canard = 0°
canard = -10 °
canard = -20 ° -
canard = -30 °
20 25 30
Figure 4. Lift coefficient for various canard deflections, M = 0.2
10
CL ° , 1rad
O.&
-0.4'-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 5. Lift coefficient due to angle-of-attack rate
A C L IGE
0.10.
0.05'
0.00 /
-0.05 / r
-0.10
-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 6. Lift coefficient increment due to ground plane influence
11
K GE
1.0t
0._'t
0.6' _
0.4_ _._ _ •
, \0.2_
0.0
o 10 15 20 25 30 35 40 45
Gear height, ft
Figure 7. Power-off ground effect washout factor
12
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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
25
o°1-1
°,-i
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m
0.3
0.2
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I
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w
-0.05,
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Alpha, dog
Figure 21. Drag coefficient for various canard deflections, M = 0.2
26
A CDIGE
0.01,
0.00-
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-0.03.....
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-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 22. Drag coefficient increment due to ground plane influence
27
ID¢dl
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0.11
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Flap = -15 °
Nap=if'
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Flap = 30 °
Flap = 45 °
-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 23. Pitching moment coefficient for various flap deflections, M = 0.2
28
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° _...q
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canard = 0 °
canard = 10 °
----o--- canard = 20 °
+ canard = 30 °
-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
29
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Alpha, deg
Figure 25. Pitching moment coefficient due to angle-of-attack rate
0.02-
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0.00
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Alpha, deg
Figure 26. Pitching moment coefficient increment due to ground plane influence
30
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30
Figure 39. Side force coefficient due to sideslip
1
Cyp , rad
0.20-
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J
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-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 40. Side force coefficient due to roll rate
43
Cy
D
h.
&
Aii = 300R, -30°L
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Ail = 10°R, -10°L
Ail = 0°
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Alpha, deg
Figure 41. Side force coefficient for various aileron deflections, M = 0.2
44
1
C113 md
0.05
0.00
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Alpha, deg
Figure 42. Rolling moment coefficient due to sideslip
C1 , 1p rad
0.05"
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Alpha, deg
Figure 43. Rolling moment coefficient due to roll rate
45
C1 , 1r rad
0.04"
0.03
0.02 _ J
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/
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5 I0 15
Alpha, deg
20 25 30
Figure 44. Rolling moment coefficient due to yaw rate
1
C 1_irud ' rad
0.03"
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-0.02-
-10 -5 0 5 I0 15 20 25 30
Alpha, deg
Figure 45. Rolling moment coefficient due to rudder deflection
46
+ All = 30°R, -30°L
All =20°R, -20°L
---O-- All = 10°R, -10°L
+ All = 0 °
- All -- -10°R, 10°L
+ All =-20°R, 20°L
All = -30°R, 30°I..
0.020
0.010
C1 o.ooo
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Alpha, deg
Figure 46. Rolling moment coefficient for various aileron deflections, M = 0.2
47
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0.015'
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Alpha, deg
Figure 55. Yawing moment coefficient due to sideslip
C1
n _ m
p rad
0.01-
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-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 56. Yawing moment coefficient due to roll rate
56
-0.070-
-0.075-
-0.080"
-0.085"
\\
-0.090.
-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 57. Yawing moment coefficient due to yaw rate
-0.079.
-0.080.
-0.081 r
-0.082 _
-0.083
-0.084. ....
-0.085 ...... _ _ _t_
\ . .,e¢,
-0.086-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 58. Yawing moment coefficient due to rudder deflection
57
Cn
0.006'
n All = 30°R, -30°L
All =20°R, -20°L
Ail= 10°R,-10°L
All = 0°
-- All= dO°R, 10°L
All =-20°R, 20°L
t All = -30°R, 30%
0.004
0.000
-0.002
"0.00_ '
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-10 -5 0 5 10 15 20 25 30
Alpha, deg
Figure 59. Yawing moment coefficient for various aileron deflections, M = 0.2
58
Form Approved
REPORT DOCUMENTATION PAGE OMBNO.0704-0188
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED
April 1995 Technical Memorandum4. TITLE AND SUBTITLE 5. FUNDING NUMBERS
Aerodynamics Model for a Generic ASTOVL Lift-Fan Aircraft
6. AUTHOR(S)
Lourdes G. Birckelbaw, Walter E. McNeill, and Douglas A. Wardwell
7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
Ames Research Center
Moffett Field, CA 94035-1000
9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)
National Aeronautics and Space Administration
Washington, DC 20546-0001
505-68-32
8. PERFORMING ORGANIZATIONREPORT NUMBER
A-950051
10. SPONSORING/MONITORINGAGENCY REPORT NUMBER
NASA TM-110347
11. SUPPLEMENTARY NOTES
Point of Contact: Lourdes G. Birckelbaw, Ames Research Center, MS 237-2, Moffett Field, CA 94035-1000
(415) 604-5592
12a. DISTRIBUTION/AVAILABILITY STATEMENT
Unclassified -- Unlimited
Subject Category 02
12b. DISTRIBUTION CODE
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.
14. SUBJECT TERMS
ASTOVL, Lift fan, Aerodynamics model
17. SECURITY CLASSIFICATIONOF REPORT
Unclassified
NSN 7540-O1-280-5500
18. SECURITY CLASSIFICATION
OF THIS PAGE
Unclassified
19. SECURITY CLASSIFICATIONOF ABSTRACT
15. NUMBER OF PAGES
6316. PRICE CODE
A04
20. LIMITATION OF ABSTRAC1
Standard Form 298 (Rev. 2-89)Prescribed by ANSI Std. Z39-1e
